FMT (Flight Software Memory Tracker) For Cassini Spacecraft-Software Engineering Using JAVA

Science.gov (United States)

Kan, Edwin P.; Uffelman, Hal; Wax, Allan H.

1997-01-01

The software engineering design of the Flight Software Memory Tracker (FMT) Tool is discussed in this paper. FMT is a ground analysis software set, consisting of utilities and procedures, designed to track the flight software, i.e., images of memory load and updatable parameters of the computers on-board Cassini spacecraft. FMT is implemented in Java.

Risk Assessment of Cassini Sun Sensor Integrity Due to Hypervelocity Impact of Saturn Dust Particles

Science.gov (United States)

Lee, Allan Y.

2016-01-01

A sophisticated interplanetary spacecraft, Cassini is one of the heaviest and most sophisticated interplanetary spacecraft humans have ever built and launched. Since achieving orbit at Saturn in 2004, Cassini has collected science data throughout its four-year prime mission (2004-08), and has since been approved for first and second extended missions through September 2017. In late 2016, the Cassini spacecraft will begin a daring set of ballistic orbits that will hop the rings and dive between the upper atmosphere of Saturn and its innermost D-ring twenty-two times. The "dusty" environment of the inner D-ring region the spacecraft must fly through is hazardous because of the possible damage that dust particles, travelling at speeds as high as 31.4 km/s, can do to spacecraft hardware. During hazardous proximal ring-plane crossings, the Cassini mission operation team plans to point the high-gain antenna to the RAM vector in order to protect most of spacecraft instruments from the incoming energetic ring dust particles. However, this particular spacecraft attitude will expose two Sun sensors (that are mounted on the antenna dish) to the incoming dust particles. High-velocity impacts on the Sun sensor cover glass might penetrate the 2.54-mm glass cover of the Sun sensor. Even without penetration damage, craters created by these impacts on the surface of the cover glass will degrade the transmissibility of light through it. Apart from being directly impacted by the dust particles, the Sun sensors are also threatened by some fraction of ricochet ejecta that are produced by dust particle impacts on the large antenna dish (made of graphite fiber epoxy composite material). Finally, the spacecraft attitude control system must cope with disturbances due to both the translational and angular impulses imparted on the large antenna dish and the long magnetometer boom by the incoming high-velocity projectiles. Analyses performed to quantify the risks the Sun sensors must contend

Cassini Solstice Mission Maneuver Experience: Year Two

Science.gov (United States)

Arrieta, Juan; Ballard, Christopher G.; Hahn, Yungsun

2012-01-01

The Cassini Spacecraft was launched in October 1997 on a mission to observe Saturn and its moons; it entered orbit around Saturn in July 2004 for a nominal four-year Prime Mission, later augmented by two extensions: the Equinox Mission, from July 2008 through September 2010, and the Solstice Mission, from October 2010 through September 2017. This paper provides an overview of the maneuver activities from August 2011 through June 2012 which include the design of 38 Orbit Trim Maneuvers--OTM-288 through OTM-326-- for attaining 14 natural satellite encounters: seven with Titan, six with Enceladus, and one with Dione.

The Cassini-Huygens visit to Saturn an historic mission to the ringed planet

CERN Document Server

Meltzer, Michael

2015-01-01

Cassini-Huygens was the most ambitious and successful space journey ever launched to the outer Solar System. This book examines all aspects of the journey: its conception and planning; the lengthy political processes needed to make it a reality; the engineering and development required to build the spacecraft; its 2.2-billion mile journey from Earth to the Ringed Planet; and the amazing discoveries from the mission. The author traces how the visions of a few brilliant scientists matured, gained popularity, and eventually became a reality. Innovative technical leaps were necessary to assemble such a multifaceted spacecraft and reliably operate it while it orbited a planet so far from our own. The Cassini-Huygens spacecraft design evolved from other deep space efforts, most notably the Galileo mission to Jupiter, enabling the voluminous, paradigm-shifting scientific data collected by the spacecraft.  Some of these discoveries are absolute gems. A small satellite that scientists once thought of as a dead pi...

Radioisotopic heater units warm an interplanetary spacecraft

International Nuclear Information System (INIS)

Franco-Ferreira, E.A.

1998-01-01

The Cassini orbiter and Huygens probe, which were successfully launched on October 15, 1997, constitute NASA's last grand-scale interplanetary mission of this century. The mission, which consists of a four-year, close-up study of Saturn and its moons, begins in July 2004 with Cassini's 60 orbits of Saturn and about 33 fly-bys of the large moon Titan. The Huygens probe will descend and land on Titan. Investigations will include Saturn's atmosphere, its rings and its magnetosphere. The atmosphere and surface of Titan and other icy moons also will be characterized. Because of the great distance of Saturn from the sun, some of the instruments and equipment on both the orbiter and the probe require external heaters to maintain their temperature within normal operating ranges. These requirements are met by Light Weight Radioisotope Heater Units (LWRHUs) designed, fabricated and safety tested at Los Alamos National Laboratory, New Mexico. An improved gas tungsten arc welding procedure lowered costs and decreased processing time for heat units for the Cassini spacecraft

Cassini-Huygens Science Highlights: Surprises in the Saturn System

Science.gov (United States)

Spilker, Linda; Altobelli, Nicolas; Edgington, Scott

2014-05-01

The Cassini-Huygens mission has greatly enhanced our understanding of the Saturn system. Fundamental discoveries have altered our views of Saturn, its retinue of icy moons including Titan, the dynamic rings, and the system's complex magnetosphere. Launched in 1997, the Cassini-Huygens spacecraft spent seven years traveling to Saturn, arriving in July 2004, roughly two years after the northern winter solstice. Cassini has orbited Saturn for 9.5 years, delivering the Huygens probe to its Titan landing in 2005, crossing northern equinox in August 2009, and completing its Prime and Equinox Missions. It is now three years into its 7-year Solstice mission, returning science in a previously unobserved seasonal phase between equinox and solstice. As it watches the approach of northern summer, long-dark regions throughout the system become sunlit, allowing Cassini's science instruments to probe as-yet unsolved mysteries. Key Cassini-Huygens discoveries include icy jets of material streaming from tiny Enceladus' south pole, lakes of liquid hydrocarbons and methane rain on giant Titan, three-dimensional structures in Saturn's rings, and curtain-like aurorae flickering over Saturn's poles. The Huygens probe sent back amazing images of Titan's surface, and made detailed measurements of the atmospheric composition, structure and winds. Key Cassini-Huygens science highlights will be presented. The Solstice Mission continues to provide new science. First, the Cassini spacecraft observes seasonally and temporally dependent processes on Saturn, Titan, Enceladus and other icy satellites, and within the rings and magnetosphere. Second, it addresses new questions that have arisen during the mission thus far, for example providing qualitatively new measurements of Enceladus and Titan that could not be accommodated in the earlier mission phases. Third, it will conduct a close-in mission at Saturn yielding fundamental knowledge about the interior of Saturn. This grand finale of the

Cassini-Huygens makes first close approach to Titan

Science.gov (United States)

2004-10-01

Purple zaze hi-res Size hi-res: 88 kb Credits: NASA/JPL/Space Science Institute Purple haze around Titan This NASA/ESA/ASI Cassini-Huygens image of Titan was taken with the narrow-angle camera on 3 July 2004, from a distance of about 789 000 kilometres from Titan. The image scale is 4.7 kilometres per pixel. This image shows two thin haze layers. The outer haze layer is detached and appears to float high in the atmosphere. Because of its thinness, the high haze layer is best seen at the moon's limb. The image was taken using a spectral filter sensitive to wavelengths of ultraviolet light centred at 338 nanometres. The image has been falsely coloured, the globe of Titan retains the pale orange hue our eyes would usually see, but both the main atmospheric haze and the thin detached layer have been brightened and given a purple colour to enhance their visibility. At the time of the closest approach, which is scheduled for 18:44 CEST, the spacecraft will be travelling only 1200 kilometres above the surface of the moon, almost grazing the outer atmosphere, at a speed of six kilometres per second (21 800 kilometres per hour)! Confirmation that the fly-by was successful and that all the data were received will not take place until 03:30 CEST on 27 October. This fly-by not only allows important surface science to be performed, such as radar analysis at close quarters, but also it significantly changes the orbit of the spacecraft around Saturn. Currently Cassini-Huygens has an orbital period of four months, which will change to 48 days, thus setting the course for the next close Titan fly-by on 13 December 2004 and the Huygens probe release on 25 December. Several of the observations performed during this fly-by will provide important information for ESA’s Huygens team, who will be using the data gathered to double-check atmospheric models for entry and descent on 14 January 2005. The Huygens probe will need to perform reliably in some of the most challenging and remote

Titan after Cassini Huygens

Science.gov (United States)

Beauchamp, P. M.; Lunine, J.; Lebreton, J.; Coustenis, A.; Matson, D.; Reh, K.; Erd, C.

2008-12-01

In 2005, the Huygens Probe gave us a snapshot of a world tantalizingly like our own, yet frozen in its evolution on the threshold of life. The descent under parachute, like that of Huygens in 2005, is happening again, but this time in the Saturn-cast twilight of winter in Titan's northern reaches. With a pop, the parachute is released, and then a muffled splash signals the beginning of the first floating exploration of an extraterrestrial sea-this one not of water but of liquid hydrocarbons. Meanwhile, thousands of miles away, a hot air balloon, a "montgolfiere," cruises 6 miles above sunnier terrain, imaging vistas of dunes, river channels, mountains and valleys carved in water ice, and probing the subsurface for vast quantities of "missing" methane and ethane that might be hidden within a porous icy crust. Balloon and floater return their data to a Titan Orbiter equipped to strip away Titan's mysteries with imaging, radar profiling, and atmospheric sampling, much more powerful and more complete than Cassini was capable of. This spacecraft, preparing to enter a circular orbit around Saturn's cloud-shrouded giant moon, has just completed a series of flybys of Enceladus, a tiny but active world with plumes that blow water and organics from the interior into space. Specialized instruments on the orbiter were able to analyze these plumes directly during the flybys. Titan and Enceladus could hardly seem more different, and yet they are linked by their origin in the Saturn system, by a magnetosphere that sweeps up mass and delivers energy, and by the possibility that one or both worlds harbor life. It is the goal of the NASA/ESA Titan Saturn System Mission (TSSM) to explore and investigate these exotic and inviting worlds, to understand their natures and assess the possibilities of habitability in this system so distant from our home world. Orbiting, landing, and ballooning at Titan represent a new and exciting approach to planetary exploration. The TSSM mission

On the spacecraft attitude stabilization in the orbital frame

Directory of Open Access Journals (Sweden)

Antipov Kirill A.

2012-01-01

Full Text Available The paper deals with spacecraft in the circular near-Earth orbit. The spacecraft interacts with geomagnetic field by the moments of Lorentz and magnetic forces. The octupole approximation of the Earth’s magnetic field is accepted. The spacecraft electromagnetic parameters, namely the electrostatic charge moment of the first order and the eigen magnetic moment are the controlled quasiperiodic functions. The control algorithms for the spacecraft electromagnetic parameters, which allows to stabilize the spacecraft attitude position in the orbital frame are obtained. The stability of the spacecraft stabilized orientation is proved both analytically and by PC computations.

Effective stability around the Cassini state in the spin-orbit problem

Science.gov (United States)

Sansottera, Marco; Lhotka, Christoph; Lemaître, Anne

2014-05-01

We investigate the long-time stability in the neighborhood of the Cassini state in the conservative spin-orbit problem. Starting with an expansion of the Hamiltonian in the canonical Andoyer-Delaunay variables, we construct a high-order Birkhoff normal form and give an estimate of the effective stability time in the Nekhoroshev sense. By extensively using algebraic manipulations on a computer, we explicitly apply our method to the rotation of Titan. We obtain physical bounds of Titan's latitudinal and longitudinal librations, finding a stability time greatly exceeding the estimated age of the Universe. In addition, we study the dependence of the effective stability time on three relevant physical parameters: the orbital inclination, , the mean precession of the ascending node of Titan orbit, , and the polar moment of inertia,.

Flight Path Control Design for the Cassini Solstice Mission

Science.gov (United States)

Ballard, Christopher G.; Ionasescu, Rodica

2011-01-01

The Cassini spacecraft has been in orbit around Saturn for just over 7 years, with a planned 7-year extension, called the Solstice Mission, which started on September 27, 2010. The Solstice Mission includes 205 maneuvers and 70 flybys which consist of the moons Titan, Enceladus, Dione, and Rhea. This mission is designed to use all available propellant with a statistical margin averaging 0.6 m/s per encounter, and the work done to prove and ensure the viability of this margin is highlighted in this paper.

Prospects for Measuring Planetary Spin and Frame-Dragging in Spacecraft Timing Signals

Energy Technology Data Exchange (ETDEWEB)

Schärer, Andreas; Bondarescu, Ruxandra [Department of Physics, University of Zurich, Zurich (Switzerland); Saha, Prasenjit [Department of Physics, University of Zurich, Zurich (Switzerland); Institute for Computational Science, University of Zurich, Zurich (Switzerland); Angélil, Raymond [Institute for Computational Science, University of Zurich, Zurich (Switzerland); Helled, Ravit [Institute for Computational Science, University of Zurich, Zurich (Switzerland); Department of Geosciences, Tel Aviv University, Tel Aviv (Israel); Jetzer, Philippe, E-mail: andreas.schaerer@physik.uzh.ch [Department of Physics, University of Zurich, Zurich (Switzerland)

2017-09-05

Satellite tracking involves sending electromagnetic signals to Earth. Both the orbit of the spacecraft and the electromagnetic signals themselves are affected by the curvature of spacetime. The arrival time of the pulses is compared to the ticks of local clocks to reconstruct the orbital path of the satellite to high accuracy, and implicitly measure general relativistic effects. In particular, Schwarzschild space curvature (static) and frame-dragging (stationary) due to the planet's spin affect the satellite's orbit. The dominant relativistic effect on the path of the signal photons is Shapiro delays due to static space curvature. We compute these effects for some current and proposed space missions, using a Hamiltonian formulation in four dimensions. For highly eccentric orbits, such as in the Juno mission and in the Cassini Grand Finale, the relativistic effects have a kick-like nature, which could be advantageous for detecting them if their signatures are properly modeled as functions of time. Frame-dragging appears, in principle, measurable by Juno and Cassini, though not by Galileo 5 and 6. Practical measurement would require disentangling frame-dragging from the Newtonian “foreground” such as the gravitational quadrupole which has an impact on both the spacecraft's orbit and the signal propagation. The foreground problem remains to be solved.

Orbit Determination of Spacecraft in Earth-Moon L1 and L2 Libration Point Orbits

Science.gov (United States)

Woodard, Mark; Cosgrove, Daniel; Morinelli, Patrick; Marchese, Jeff; Owens, Brandon; Folta, David

2011-01-01

The ARTEMIS mission, part of the THEMIS extended mission, is the first to fly spacecraft in the Earth-Moon Lissajous regions. In 2009, two of the five THEMIS spacecraft were redeployed from Earth-centered orbits to arrive in Earth-Moon Lissajous orbits in late 2010. Starting in August 2010, the ARTEMIS P1 spacecraft executed numerous stationkeeping maneuvers, initially maintaining a lunar L2 Lissajous orbit before transitioning into a lunar L1 orbit. The ARTEMIS P2 spacecraft entered a L1 Lissajous orbit in October 2010. In April 2011, both ARTEMIS spacecraft will suspend Lissajous stationkeeping and will be maneuvered into lunar orbits. The success of the ARTEMIS mission has allowed the science team to gather unprecedented magnetospheric measurements in the lunar Lissajous regions. In order to effectively perform lunar Lissajous stationkeeping maneuvers, the ARTEMIS operations team has provided orbit determination solutions with typical accuracies on the order of 0.1 km in position and 0.1 cm/s in velocity. The ARTEMIS team utilizes the Goddard Trajectory Determination System (GTDS), using a batch least squares method, to process range and Doppler tracking measurements from the NASA Deep Space Network (DSN), Berkeley Ground Station (BGS), Merritt Island (MILA) station, and United Space Network (USN). The team has also investigated processing of the same tracking data measurements using the Orbit Determination Tool Kit (ODTK) software, which uses an extended Kalman filter and recursive smoother to estimate the orbit. The orbit determination results from each of these methods will be presented and we will discuss the advantages and disadvantages associated with using each method in the lunar Lissajous regions. Orbit determination accuracy is dependent on both the quality and quantity of tracking measurements, fidelity of the orbit force models, and the estimation techniques used. Prior to Lissajous operations, the team determined the appropriate quantity of tracking

The Cassini-Huygens mission

CERN Document Server

The joint NASA-ESA Cassini-Huygens mission promises to return four (and possibly more) years of unparalleled scientific data from the solar system’s most exotic planet, the ringed, gas giant, Saturn. Larger than Galileo with a much greater communication bandwidth, Cassini can accomplish in a single flyby what Galileo returned in a series of passes. Cassini explores the Saturn environment in three dimensions, using gravity assists to climb out of the equatorial plane to look down on the rings from above, to image the aurora and to study polar magnetospheric processes such as field-aligned currents. Since the radiation belt particle fluxes are much more benign than those at Jupiter, Cassini can more safely explore the inner regions of the magnetosphere. The spacecraft approaches the planet closer than Galileo could, and explores the inner moons and the rings much more thoroughly than was possible at Jupiter. This book is the second volume, in a three volume set, that describes the Cassini/Huygens mission. Thi...

Saturn's Internal Magnetic Field Revealed by Cassini Grand Finale

Science.gov (United States)

Cao, H.; Dougherty, M. K.; Khurana, K. K.; Hunt, G. J.; Provan, G.; Kellock, S.; Burton, M. E.; Burk, T. A.

2017-12-01

Saturn's internal magnetic field has been puzzling since the first in-situ measurements during the Pioneer 11 Saturn flyby. Cassini magnetometer measurements prior to the Grand Finale phase established 1) the highly axisymmetric nature of Saturn's internal magnetic field with a dipole tilt smaller than 0.06 degrees, 2) at least an order of magnitude slower secular variation rate compared to that of the current geomagnetic field, and 3) expulsion of magnetic fluxes from the equatorial region towards high latitude. The highly axisymmetric nature of Saturn's intrinsic magnetic field not only challenges dynamo theory but also makes an accurate determination of the interior rotation rate of Saturn extremely difficult. The Cassini spacecraft entered the Grand Finale phase in April 2017, during which time the spacecraft dived through the gap between Saturn's atmosphere and the inner edge of the D-ring 22 times before descending into the deep atmosphere of Saturn. The unprecedented proximity to Saturn (reaching 2500 km above the cloud deck) and the highly inclined nature of the Grand Finale orbits provided an ideal opportunity to decode Saturn's internal magnetic field. The fluxgate magnetometer onboard Cassini made precise vector measurements during the Grand Finale phase. Magnetic signals from the interior of the planet, the magnetospheric ring current, the high-latitude field-aligned current (FAC) modulated by the 10.7 hour planetary period oscillation, and low-latitude FACs were observed during the Grand Finale phase. Here we report the magnetometer measurements during the Cassini Grand Finale phase, new features of Saturn's internal magnetic field revealed by these measurements (e.g., the high degree magnetic moments of Saturn, the level of axisymmetry beyond dipole), and implications for the deep interior of Saturn.

An overview of the risk uncertainty assessment process for the Cassini space mission

International Nuclear Information System (INIS)

Wyss, G.D.

1996-01-01

The Cassini spacecraft is a deep space probe whose mission is to explore the planet Saturn and its moons. Since the spacecraft's electrical requirements will be supplied by radioisotope thermoelectric generators (RTGs), the spacecraft designers and mission planners must assure that potential accidents involving the spacecraft do not pose significant human risk. The Cassini risk analysis team is seeking to perform a quantitative uncertainty analysis as a part of the overall mission risk assessment program. This paper describes the uncertainty analysis methodology to be used for the Cassini mission and compares it to the methods that were originally developed for evaluation of commercial nuclear power reactors

Modeling Saturn's Inner Plasmasphere: Cassini's Closest Approach

Science.gov (United States)

Moore, L.; Mendillo, M.

2005-05-01

Ion densities from the three-dimensional Saturn-Thermosphere-Ionosphere-Model (STIM, Moore et al., 2004) are extended above the plasma exobase using the formalism of Pierrard and Lemaire (1996, 1998), which evaluates the balance of gravitational, centrifugal and electric forces on the plasma. The parameter space of low-energy ionospheric contributions to Saturn's plasmasphere is explored by comparing results that span the observed extremes of plasma temperature, 650 K to 1700 K, and a range of velocity distributions, Lorentzian (or Kappa) to Maxwellian. Calculations are made for plasma densities along the path of the Cassini spacecraft's orbital insertion on 1 July 2004. These calculations neglect any ring or satellite sources of plasma, which are most likely minor contributors at 1.3 Saturn radii. Modeled densities will be compared with Cassini measurements as they become available. Moore, L.E., M. Mendillo, I.C.F. Mueller-Wodarg, and D.L. Murr, Icarus, 172, 503-520, 2004. Pierrard, V. and J. Lemaire, J. Geophys. Res., 101, 7923-7934, 1996. Pierrard, V. and J. Lemaire, J. Geophys. Res., 103, 4117, 1998.

Observations and temperatures of Io's Pele Patera from Cassini and Galileo spacecraft images

Science.gov (United States)

Radebaugh, J.; McEwen, A.S.; Milazzo, M.P.; Keszthelyi, L.P.; Davies, A.G.; Turtle, E.P.; Dawson, D.D.

2004-01-01

Pele has been the most intense high-temperature hotspot on Io to be continuously active during the Galileo monitoring from 1996-2001. A suite of characteristics suggests that Pele is an active lava lake inside a volcanic depression. In 2000-2001, Pele was observed by two spacecraft, Cassini and Galileo. The Cassini observations revealed that Pele is variable in activity over timescales of minutes, typical of active lava lakes in Hawaii and Ethiopia. These observations also revealed that the short-wavelength thermal emission from Pele decreases with rotation of Io by a factor significantly greater than the cosine of the emission angle, and that the color temperature becomes more variable and hotter at high emission angles. This behavior suggests that a significant portion of the visible thermal emission from Pele comes from lava fountains within a topographically confined lava body. High spatial resolution, nightside images from a Galileo flyby in October 2001 revealed a large, relatively cool (Pele has lavas with ultramafic compositions. The long-lived, vigorous activity of what is most likely an actively overturning lava lake in Pele Patera indicates that there is a strong connection to a large, stable magma source region. ?? 2003 Elsevier Inc. All rights reserved.

Constraints on the Mass and Location of Planet 9 set by Range and VLBI Observations of Cassini

Science.gov (United States)

Jacobson, Robert Arthur; Folkner, William; Park, Ryan; Williams, James

2018-04-01

Batygin and Brown, 2016 AJ, found that Kuiper belt objects (KBOs) with well determined orbits having periods greater than 4000 years are apsidally aligned. They attribute this orbital clustering to the existence of a distant planet, Planet 9, well beyond Neptune, with a mass roughly ten times that of Earth. If such a planet exists, it would affect the motion of the known solar system planets, in particular Saturn, which is well observed with radiometric ranging from the Cassini spacecraft and VLBI observations of Cassini. The current planetary ephemerides do not account for the postulated Planet 9, yet their fit to the observational data shows no obvious effect that could be attributed to neglecting that planet. However, it is possible that the effect could be absorbed by the estimated parameters used to determine the ephemerides. Those parameters include the planetary orbital elements, mass of the Sun, and the masses of the asteroids that perturb the Martian orbit. We recently updated the Cassini data set and extended it through the end of the mssion in 2017 September. We analyze the sensitivity of these data to the tidal perturbations caused by the postulated Planet 9 for a range of positions on the sky and tidal parameters (the ratio of the mass of Planet 9 to the cube of its distance from Saturn). We determine an upper bound on the tidal parameter and the most probable directions consistent with the observational data.

Spacecraft on-orbit deployment anomalies - What can be done?

Science.gov (United States)

Freeman, Michael T.

1993-04-01

Modern communications satellites rely heavily upon deployable appendage (i.e. solar arrays, communications antennas, etc.) to perform vital functions that enable the spacecraft to effectively conduct mission objectives. Communications and telemetry antennas provide the radiofrequency link between the spacecraft and the earth ground station, permitting data to be transmitted and received from the satellite. Solar arrays serve as the principle source of electrical energy to the satellite, and recharge internal batteries during operation. However, since satellites cannot carry backup systems, if a solar array fails to deploy, the mission is lost. This article examines the subject of on-orbit anomalies related to the deployment of spacecraft appendage, and possible causes of such failures. Topics discussed shall include mechanical launch loading, on-orbit thermal and solar concerns, reliability of spacecraft pyrotechnics, and practical limitations of ground-based deployment testing. Of particular significance, the article will feature an in-depth look at the lessons learned from the successful recovery of the Telesat Canada Anik-E2 satellite in 1991.

Passive radiative cooling of a HTS coil for attitude orbit control in micro-spacecraft

Science.gov (United States)

Inamori, Takaya; Ozaki, Naoya; Saisutjarit, Phongsatorn; Ohsaki, Hiroyuki

2015-02-01

This paper proposes a novel radiative cooling system for a high temperature superconducting (HTS) coil for an attitude orbit control system in nano- and micro-spacecraft missions. These days, nano-spacecraft (1-10 kg) and micro-spacecraft (10-100 kg) provide space access to a broader range of spacecraft developers and attract interest as space development applications. In planetary and high earth orbits, most previous standard-size spacecraft used thrusters for their attitude and orbit control, which are not available for nano- and micro-spacecraft missions because of the strict power consumption, space, and weight constraints. This paper considers orbit and attitude control methods that use a superconducting coil, which interacts with on-orbit space plasmas and creates a propulsion force. Because these spacecraft cannot use an active cooling system for the superconducting coil because of their mass and power consumption constraints, this paper proposes the utilization of a passive radiative cooling system, in which the superconducting coil is thermally connected to the 3 K cosmic background radiation of deep space, insulated from the heat generation using magnetic holders, and shielded from the sun. With this proposed cooling system, the HTS coil is cooled to 60 K in interplanetary orbits. Because the system does not use refrigerators for its cooling system, the spacecraft can achieve an HTS coil with low power consumption, small mass, and low cost.

The architecture of the Cassini division

Science.gov (United States)

Hedman, M.M.; Nicholson, P.D.; Baines, K.H.; Buratti, B.J.; Sotin, Christophe; Clark, R.N.; Brown, R.H.; French, R.G.; Marouf, E.A.

2010-01-01

The Cassini Division in Saturn's rings contains a series of eight named gaps, three of which contain dense ringlets. Observations of stellar occultations by the Visual and Infrared Mapping Spectrometer onboard the Cassini spacecraft have yielded 40 accurate and precise measurements of the radial position of the edges of all of these gaps and ringlets. These data reveal suggestive patterns in the shapes of many of the gap edges: the outer edges of the five gaps without ringlets are circular to within 1 km, while the inner edges of six of the gaps are eccentric, with apsidal precession rates consistent with those expected for eccentric orbits near each edge. Intriguingly, the pattern speeds of these eccentric inner gap edges, together with that of the eccentric Huygens Ringlet, form a series with a characteristic spacing of 006 day-1. The two gaps with non-eccentric inner edges lie near first-order inner Lindblad resonances (ILRs) with moons. One such edge is close to the 5:4 ILR with Prometheus, and the radial excursions of this edge do appear to have an m = 5 component aligned with that moon. The other resonantly confined edge is the outer edge of the B ring, which lies near the 2:1 Mimas ILR. Detailed investigation of the B-ring-edge data confirm the presence of an m = 2 perturbation on the B-ring edge, but also show that during the course of the Cassini Mission, this pattern has drifted backward relative to Mimas. Comparisons with earlier occultation measurements going back to Voyager suggest the possibility that the m = 2 pattern is actually librating relative to Mimas with a libration frequency L 006 day-1 (or possibly 012 day -1). In addition to the m = 2 pattern, the B-ring edge also has an m = 1 component that rotates around the planet at a rate close to the expected apsidal precession rate (?? ?? ?? B ??? 5.??06 day -1). Thus, the pattern speeds of the eccentric edges in the Cassini Division can be generated from various combinations of the pattern speeds

Generating Animated Displays of Spacecraft Orbits

Science.gov (United States)

Candey, Robert M.; Chimiak, Reine A.; Harris, Bernard T.

2005-01-01

Tool for Interactive Plotting, Sonification, and 3D Orbit Display (TIPSOD) is a computer program for generating interactive, animated, four-dimensional (space and time) displays of spacecraft orbits. TIPSOD utilizes the programming interface of the Satellite Situation Center Web (SSCWeb) services to communicate with the SSC logic and database by use of the open protocols of the Internet. TIPSOD is implemented in Java 3D and effects an extension of the preexisting SSCWeb two-dimensional static graphical displays of orbits. Orbits can be displayed in any or all of the following seven reference systems: true-of-date (an inertial system), J2000 (another inertial system), geographic, geomagnetic, geocentric solar ecliptic, geocentric solar magnetospheric, and solar magnetic. In addition to orbits, TIPSOD computes and displays Sibeck's magnetopause and Fairfield's bow-shock surfaces. TIPSOD can be used by the scientific community as a means of projection or interpretation. It also has potential as an educational tool.

Global optimum spacecraft orbit control subject to bounded thrust in presence of nonlinear and random disturbances in a low earth orbit

Directory of Open Access Journals (Sweden)

Tamer Mekky Ahmed Habib

2012-06-01

Full Text Available The primary objective of this work is to develop an effective spacecraft orbit control algorithm suitable for spacecraft orbital maneuver and/or rendezvous. The actual governing equation of a spacecraft orbiting the earth is merely nonlinear. Disturbance forces resulting from aerodynamic drag, oblateness of the earth till the fourth order (i.e. J4, and random disturbances are modeled for the initial and target orbits. These disturbances increase the complexity of nonlinear governing equations. Global optimum solutions of the control algorithm parameters are determined throughout real coded genetic algorithms such that the steady state difference between the actual and desired trajectories is minimized. The resulting solutions are constrained to avoid spacecraft collision with the surface of the earth taking into account limited thrust budget.

Spacecraft design project: Low Earth orbit communications satellite

Science.gov (United States)

Moroney, Dave; Lashbrook, Dave; Mckibben, Barry; Gardener, Nigel; Rivers, Thane; Nottingham, Greg; Golden, Bill; Barfield, Bill; Bruening, Joe; Wood, Dave

1991-01-01

This is the final product of the spacecraft design project completed to fulfill the academic requirements of the Spacecraft Design and Integration 2 course (AE-4871) taught at the U.S. Naval Postgraduate School. The Spacecraft Design and Integration 2 course is intended to provide students detailed design experience in selection and design of both satellite system and subsystem components, and their location and integration into a final spacecraft configuration. The design team pursued a design to support a Low Earth Orbiting (LEO) communications system (GLOBALSTAR) currently under development by the Loral Cellular Systems Corporation. Each of the 14 team members was assigned both primary and secondary duties in program management or system design. Hardware selection, spacecraft component design, analysis, and integration were accomplished within the constraints imposed by the 11 week academic schedule and the available design facilities.

The Global Precipitation Measurement (GPM) Spacecraft Power System Design and Orbital Performance

Science.gov (United States)

Dakermanji, George; Burns, Michael; Lee, Leonine; Lyons, John; Kim, David; Spitzer, Thomas; Kercheval, Bradford

2016-01-01

The Global Precipitation Measurement (GPM) spacecraft was jointly developed by National Aeronautics and Space Administration (NASA) and Japan Aerospace Exploration Agency (JAXA). It is a Low Earth Orbit (LEO) spacecraft launched on February 27, 2014. The spacecraft is in a circular 400 Km altitude, 65 degrees inclination nadir pointing orbit with a three year basic mission life. The solar array consists of two sun tracking wings with cable wraps. The panels are populated with triple junction cells of nominal 29.5% efficiency. One axis is canted by 52 degrees to provide power to the spacecraft at high beta angles. The power system is a Direct Energy Transfer (DET) system designed to support 1950 Watts orbit average power. The batteries use SONY 18650HC cells and consist of three 8s x 84p batteries operated in parallel as a single battery. The paper describes the power system design details, its performance to date and the lithium ion battery model that was developed for use in the energy balance analysis and is being used to predict the on-orbit health of the battery.

Guidance and Navigation for Rendezvous and Proximity Operations with a Non-Cooperative Spacecraft at Geosynchronous Orbit

Science.gov (United States)

Barbee, Brent William; Carpenter, J. Russell; Heatwole, Scott; Markley, F. Landis; Moreau, Michael; Naasz, Bo J.; VanEepoel, John

2010-01-01

The feasibility and benefits of various spacecraft servicing concepts are currently being assessed, and all require that the servicer spacecraft perform rendezvous, proximity, and capture operations with the target spacecraft to be serviced. Many high-value spacecraft, which would be logical targets for servicing from an economic point of view, are located in geosynchronous orbit, a regime in which autonomous rendezvous and capture operations are not commonplace. Furthermore, existing GEO spacecraft were not designed to be serviced. Most do not have cooperative relative navigation sensors or docking features, and some servicing applications, such as de-orbiting of a non-functional spacecraft, entail rendezvous and capture with a spacecraft that may be non-functional or un-controlled. Several of these challenges have been explored via the design of a notional mission in which a nonfunctional satellite in geosynchronous orbit is captured by a servicer spacecraft and boosted into super-synchronous orbit for safe disposal. A strategy for autonomous rendezvous, proximity operations, and capture is developed, and the Orbit Determination Toolbox (ODTBX) is used to perform a relative navigation simulation to assess the feasibility of performing the rendezvous using a combination of angles-only and range measurements. Additionally, a method for designing efficient orbital rendezvous sequences for multiple target spacecraft is utilized to examine the capabilities of a servicer spacecraft to service multiple targets during the course of a single mission.

A simple method to design non-collision relative orbits for close spacecraft formation flying

Science.gov (United States)

Jiang, Wei; Li, JunFeng; Jiang, FangHua; Bernelli-Zazzera, Franco

2018-05-01

A set of linearized relative motion equations of spacecraft flying on unperturbed elliptical orbits are specialized for particular cases, where the leader orbit is circular or equatorial. Based on these extended equations, we are able to analyze the relative motion regulation between a pair of spacecraft flying on arbitrary unperturbed orbits with the same semi-major axis in close formation. Given the initial orbital elements of the leader, this paper presents a simple way to design initial relative orbital elements of close spacecraft with the same semi-major axis, thus preventing collision under non-perturbed conditions. Considering the mean influence of J 2 perturbation, namely secular J 2 perturbation, we derive the mean derivatives of orbital element differences, and then expand them to first order. Thus the first order expansion of orbital element differences can be added to the relative motion equations for further analysis. For a pair of spacecraft that will never collide under non-perturbed situations, we present a simple method to determine whether a collision will occur when J 2 perturbation is considered. Examples are given to prove the validity of the extended relative motion equations and to illustrate how the methods presented can be used. The simple method for designing initial relative orbital elements proposed here could be helpful to the preliminary design of the relative orbital elements between spacecraft in a close formation, when collision avoidance is necessary.

Modernization of the Cassini Ground System

Science.gov (United States)

Razo, Gus; Fujii, Tammy

2014-01-01

The Cassini Spacecraft and its ground system have been operational for over 16 years. Modernization presents several challenges due to the personnel, processes, and tools already invested and embedded into the current ground system structure. Every mission's ground system has its own unique complexities and challenges, involving various organizational units. As any mission from its inception to its execution, schedules are always tight. This forces GDS engineers to implement a working ground system that is not necessarily fully optimized. Ground system challenges increase as technology evolves and cyber threats become more sophisticated. Cassini's main challenges were due to its ground system existing before many security requirements were levied on the multi-mission tools and networks. This caused a domino effect on Cassini GDS tools that relied on outdated technological features. In the aerospace industry reliable and established technology is preferred over innovative yet less proven technology. Loss of data for a spacecraft mission can be catastrophic; therefore, there is a reluctance to make changes and updates to the ground system. Nevertheless, all missions and associated teams face the need to modernize their processes and tools. Systems development methods from well-known system analysis and design principles can be applied to many missions' ground systems. Modernization should always be considered, but should be done in such a way that it does not affect flexibility nor interfere with established practices. Cassini has accomplished a secure and efficient ground data system through periodic updates. The obstacles faced while performing the modernization of the Cassini ground system will be outlined, as well as the advantages and challenges that were encountered.

Examining the Combined Saturn and Ring Exosphere/Ionosphere using Cassini's Proximal orbits

Science.gov (United States)

Tucker, O. J.; Tseng, W. L.; Johnson, R. E.; Perry, M. E.

2017-12-01

Neutral molecules that are emitted from Saturn's exobase (i.e., H2) and the main rings (i.e., H2, O2, H) are a source of material for both the Saturn and ring ionospheres as well as Saturn's magnetosphere (Tseng et al., 2013 [PSS 85 164 - 167]). However, the density gradient of H2 produced from the main rings is very different than that produced by Saturn's exospheric flux due to its emission from the ring plane and distance from Saturn. Cassini measurements obtained during the proximal orbits can likely be used to identify contributions from Saturn and the rings. Here we present results obtained from Monte Carlo models of the Saturn and ring exosphere used to analyze INMS data of neutrals and ions measured along the trajectories of the Proximal orbits. Understanding the sources of neutrals and the concomitant ions can help provide insight about the dynamics occurring in the Saturn system.

Cassini radar: Instrument description and performance status

Science.gov (United States)

Johnson, W. T. K.; Im, E.; Borgarelli, L.; ZampoliniFaustini, E.

1995-01-01

The spacecraft of the Cassini mission is planned to be launched towards Saturn in October 1997. The mission is designed to study the physical structure and chemical composition of Titan. The results of the tests performed on the Cassini radar engineering qualification model (EQM) are summarized. The approach followed in the verification and evaluation of the performance of the radio frequency subsystem EQM is presented. The results show that the instrument satisfies the most relevant mission requirements.

Cassini: The Journey and the Legacy

KAUST Repository

Porco, Carolyn

2018-01-01

An international mission to explore, in depth, the Saturnian system ﾖthe planet Saturn and its magnetosphere, glorious rings, and many moons- begun over 27 years ago. After seven years of development, the Cassini spacecraft was launched in 1997

Searching sequences of resonant orbits between a spacecraft and Jupiter

International Nuclear Information System (INIS)

Formiga, J K S; Prado, A F B A

2013-01-01

This research shows a study of the dynamical behavior of a spacecraft that performs a series of close approaches with the planet Jupiter. The main idea is to find a sequence of resonant orbits that allows the spacecraft to stay in the region of the space near the orbit of Jupiter around the Sun gaining energy from each passage by the planet. The dynamical model considers the existence of only two massive bodies in the systems, which are the Sun and Jupiter. They are assumed to be in circular orbits around their center of mass. Analytical equations are used to obtain the values of the parameters required to get this sequence of close approaches. Those equations are useful, because they show which orbits are physically possible when taking into account that the periapsis distances have to be above the surface of the Sun and that the closest approach distances during the passage by Jupiter have to be above its surface

On-orbit supervisor for controlling spacecraft

Science.gov (United States)

Vandervoort, Richard J.

1992-07-01

Spacecraft systems of the 1990's and beyond will be substantially more complex than their predecessors. They will have demanding performance requirements and will be expected to operate more autonomously. This underscores the need for innovative approaches to Fault Detection, Isolation and Recovery (FDIR). A hierarchical expert system is presented that provides on-orbit supervision using intelligent FDIR techniques. Each expert system in the hierarchy supervises the operation of a local set of spacecraft functions. Spacecraft operational goals flow top down while responses flow bottom up. The expert system supervisors have a fairly high degree of autonomy. Bureaucratic responsibilities are minimized to conserve bandwidth and maximize response time. Data for FDIR can be acquired local to an expert and from other experts. By using a blackboard architecture for each supervisor, the system provides a great degree of flexibility in implementing the problem solvers for each problem domain. In addition, it provides for a clear separation between facts and knowledge, leading to an efficient system capable of real time response.

Proposed gravity-gradient dynamics experiments in lunar orbit using the RAE-B spacecraft

Science.gov (United States)

Blanchard, D. L.; Walden, H.

1973-01-01

A series of seven gravity-gradient dynamics experiments is proposed utilizing the Radio Astronomy Explorer (RAE-B) spacecraft in lunar orbit. It is believed that none of the experiments will impair the spacecraft structure or adversely affect the continuation of the scientific mission of the satellite. The first experiment is designed to investigate the spacecraft dynamical behavior in the absence of libration damper action and inertia. It requires stable gravity-gradient capture of the spacecraft in lunar orbit with small amplitude attitude librations as a prerequisite. Four subsequent experiments involve partial retraction, ultimately followed by full redeployment, of one or two of the 230-meter booms forming the lunar-directed Vee-antenna. These boom length change operations will induce moderate amplitude angular librations of the spacecraft.

Simulated Aging of Spacecraft External Materials on Orbit

Science.gov (United States)

Khatipov, S.

Moscow State Engineering Physics Institute (MIFI), in cooperation with Air Force Research Laboratory's Satellite Assessment Center (SatAC), the European Office of Aerospace Research and Development (EOARD), and the International Science and Technology Center (ISTC), has developed a database describing the changes in optical properties of materials used on the external surfaces of spacecraft due to space environmental factors. The database includes data acquired from tests completed under contract with the ISTC and EOARD, as well as from previous Russian materials studies conducted within the last 30 years. The space environmental factors studied are for those found in Low Earth Orbits (LEO) and Geosynchronous orbits (GEO), including electron irradiation at 50, 100, and 200 keV, proton irradiation at 50, 150, 300, and 500 keV, and ultraviolet irradiation equivalent to 1 sun-year. The material characteristics investigated were solar absorption (aS), spectral reflectance (rl), solar reflectance (rS), emissivity (e), spectral transmission coefficient (Tl), solar transmittance (TS), optical density (D), relative optical density (D/x), Bi-directional Reflectance Distribution Function (BRDF), and change of appearance and color in the visible wavelengths. The materials tested in the project were thermal control coatings (paints), multilayer insulation (films), and solar cells. The ability to predict changes in optical properties of spacecraft materials is important to increase the fidelity of space observation tools, better understand observation of space objects, and increase the longevity of spacecraft. The end goal of our project is to build semi-empirical mathematical models to predict the long-term effects of space aging as a function of time and orbit.

BepiColombo — The Next Step of Mercury Exploration with Two Orbiting Spacecraft

Science.gov (United States)

Benkhoff, J.

2018-05-01

BepiColombo is a joint project between ESA and JAXA. The mission consists of two orbiters — the Mercury Planetary Orbiter and the Mercury Magnetospheric Orbiter. From dedicated orbits, the spacecraft will be studying the planet and its environment.

Five Fabulous Flybys of the Small Inner Moons of Saturn by the Cassini Spacecraft

Science.gov (United States)

Buratti, B. J.; Momary, T.; Clark, R. N.; Brown, R. H.; Filacchione, G.; Mosher, J. A.; Baines, K. H.; Nicholson, P. D.

2017-12-01

The Saturn system possesses a number of small unique moons, including the coorbitals Janus and Epimetheus; the ring moons Pan and Daphnis; and Prometheus, Pandora, and Atlas, which orbit near the edge of the main ring system. During the last phases of the Cassini mission, when the spacecraft executed close passes to the F-ring of Saturn, five "best-ever" flybys of these moons occurred. Pan, Daphnis, Atlas, Pandora, and Epimetheus were approached at distances ranging from 6000-40,000 km. The Visual Infrared Mapping Spectrometer (VIMS) captured data from the spectral range spanning 0.35-5.1 microns, as well as capturing solar phase angles not observed before. When combined with spectra from different regions of the moons obtained throughout the mission, the VIMS observations reveal substantial changes in the depth of water-ice absorption bands and color over the moons' surfaces. These measurements show the accretion of main-ring material onto the moons, with leading sides exhibiting stronger water-ice signatures in general. Atlas and Pandora have red visible spectra similar to the A-ring and unlike other icy moons, which are blue, further revealing accretion of main ring material onto the small inner moons. In general the visible spectra of the moons gets bluer with distance from Saturn until the surface of the moons is dominated by contamination from the E-ring, which is composed of fresh ice. There is a weak correlation between color and albedo, with lower-albedo moons being redder, suggesting the existence of a dark reddish contaminant from the main ring system. The solar phase curves of the moons are similar to those of larger icy moons (unfortunately no opposition surge data was gathered). 2017 California Institute of Technology. Government sponsorship acknowledged.

On-Orbit 3-Dimensional Electrostatic Detumble for Generic Spacecraft Geometries

Science.gov (United States)

Bennett, Trevor J.

In recent years, there is a growing interest in active debris removal and on-orbit servicing of Earth orbiting assets. The growing need for such approaches is often exemplified by the Iridium-Kosmos collision in 2009 that generated thousands of debris fragments. There exists a variety of active debris removal and on-orbit servicing technologies in development. Conventional docking mechanisms and mechanical capture by actuated manipulators, exemplified by NASA's Restore-L mission, require slow target tumble rates or more aggressive circumnavigation rate matching. The tumble rate limitations can be overcome with flexible capture systems such nets, harpoons, or tethers yet these systems require complex deployment, towing, and/or interfacing strategies to avoid servicer and target damage. Alternatively, touchless methods overcome the tumble rate limitations by provide detumble control prior to a mechanical interface. This thesis explores electrostatic detumble technology to touchlessly reduce large target rotation rates of Geostationary satellites and debris. The technical challenges preceding flight implementation largely reside in the long-duration formation flying guidance, navigation, and control of a servicer spacecraft equipped with electrostatic charge transfer capability. Leveraging prior research into the electrostatic charging of spacecraft, electrostatic detumble control formulations are developed for both axisymmetric and generic target geometries. A novel relative position vector and associated relative orbit control approach is created to manage the long-duration proximity operations. Through detailed numerical simulations, the proposed detumble and relative motion control formulations demonstrate detumble of several thousand kilogram spacecraft tumbling at several degrees per second in only several days. The availability, either through modeling or sensing, of the relative attitude, relative position, and electrostatic potential are among key concerns

Titan's cold case files - Outstanding questions after Cassini-Huygens

Science.gov (United States)

Nixon, C. A.; Lorenz, R. D.; Achterberg, R. K.; Buch, A.; Coll, P.; Clark, R. N.; Courtin, R.; Hayes, A.; Iess, L.; Johnson, R. E.; Lopes, R. M. C.; Mastrogiuseppe, M.; Mandt, K.; Mitchell, D. G.; Raulin, F.; Rymer, A. M.; Todd Smith, H.; Solomonidou, A.; Sotin, C.; Strobel, D.; Turtle, E. P.; Vuitton, V.; West, R. A.; Yelle, R. V.

2018-06-01

The entry of the Cassini-Huygens spacecraft into orbit around Saturn in July 2004 marked the start of a golden era in the exploration of Titan, Saturn's giant moon. During the Prime Mission (2004-2008), ground-breaking discoveries were made by the Cassini orbiter including the equatorial dune fields (flyby T3, 2005), northern lakes and seas (T16, 2006), and the large positive and negative ions (T16 & T18, 2006), to name a few. In 2005 the Huygens probe descended through Titan's atmosphere, taking the first close-up pictures of the surface, including large networks of dendritic channels leading to a dried-up seabed, and also obtaining detailed profiles of temperature and gas composition during the atmospheric descent. The discoveries continued through the Equinox Mission (2008-2010) and Solstice Mission (2010-2017) totaling 127 targeted flybys of Titan in all. Now at the end of the mission, we are able to look back on the high-level scientific questions from the start of the mission, and assess the progress that has been made towards answering these. At the same time, new scientific questions regarding Titan have emerged from the discoveries that have been made. In this paper we review a cross-section of important scientific questions that remain partially or completely unanswered, ranging from Titan's deep interior to the exosphere. Our intention is to help formulate the science goals for the next generation of planetary missions to Titan, and to stimulate new experimental, observational and theoretical investigations in the interim.

Gravity-gradient dynamics experiments performed in orbit utilizing the Radio Astronomy Explorer (RAE-1) spacecraft

Science.gov (United States)

Walden, H.

1973-01-01

Six dynamic experiments were performed in earth orbit utilizing the RAE spacecraft in order to test the accuracy of the mathematical model of RAE dynamics. The spacecraft consisted of four flexible antenna booms, mounted on a rigid cylindrical spacecraft hub at center, for measuring radio emissions from extraterrestrial sources. Attitude control of the gravity stabilized spacecraft was tested by using damper clamping, single lower leading boom operations, and double lower boom operations. Results and conclusions of the in-orbit dynamic experiments proved the accuracy of the analytic techniques used to model RAE dynamical behavior.

High-angular-resolution stellar imaging with occultations from the Cassini spacecraft - III. Mira

Science.gov (United States)

Stewart, Paul N.; Tuthill, Peter G.; Nicholson, Philip D.; Hedman, Matthew M.

2016-04-01

We present an analysis of spectral and spatial data of Mira obtained by the Cassini spacecraft, which not only observed the star's spectra over a broad range of near-infrared wavelengths, but was also able to obtain high-resolution spatial information by watching the star pass behind Saturn's rings. The observed spectral range of 1-5 microns reveals the stellar atmosphere in the crucial water-bands which are unavailable to terrestrial observers, and the simultaneous spatial sampling allows the origin of spectral features to be located in the stellar environment. Models are fitted to the data, revealing the spectral and spatial structure of molecular layers surrounding the star. High-resolution imagery is recovered revealing the layered and asymmetric nature of the stellar atmosphere. The observational data set is also used to confront the state-of-the-art cool opacity-sampling dynamic extended atmosphere models of Mira variables through a detailed spectral and spatial comparison, revealing in general a good agreement with some specific departures corresponding to particular spectral features.

Constraints on the Mass and Location of Planet 9 set by Range and VLBI Observations of Spacecraft at Saturn

Science.gov (United States)

Jacobson, Robert A.; Folkner, William M.; Park, Ryan S.; Williams, James G.

2017-06-01

Batygin and Brown, 2016 AJ, found that all Kuiper belt objects (KBOs) with well determined orbits having periods greater than 4000 years share nearly the same orbital plane and are apsidally aligned. They attribute this orbital clustering to the existence of a distant planet, Planet 9, well beyond Neptune, with a mass roughly ten times that of Earth. If such a planet exists, it would affect the motion of the known solar system planets, in particular Saturn, which is well observed with radiometric ranging from the Voyager and Cassini spacecraft and VLBI observations of Cassini. The current planetary ephemerides do not account for the postulated Planet 9, yet their fit to the observational data shows no obvious effect that could be attributed to neglecting that planet. However, it is possible that the effect could be absorbed by the estimated parameters used to determine the ephemerides. Those parameters include the planetary orbital elements, mass of the Sun, and the masses of the asteroids that perturb the Martian orbit. We recently updated the Voyager and Cassini data sets and extended the latter through 2017 March. We analyze the sensitivity of these data to the tidal perturbations caused by Planet 9 for a range of positions on the sky and tidal parameters (the ratio of the mass of Planet 9 to the cube of its distance from Saturn). We determine an upper bound on the tidal parameter and the most probable directions consistent with the observational data.

RTG performance on Galileo and Ulysses and Cassini test results

International Nuclear Information System (INIS)

Kelly, C. Edward; Klee, Paul M.

1997-01-01

Power output from telemetry for the two Galileo RTGs are shown from the 1989 launch to the recent Jupiter encounter. Comparisons of predicted, measured and required performance are shown. Similar comparisons are made for the RTG on the Ulysses spacecraft which completed its planned mission in 1995. Also presented are test results from small scale thermoelectric modules and full scale converters performed for the Cassini program. The Cassini mission to Saturn is scheduled for an October 1997 launch. Small scale module test results on thermoelectric couples from the qualification and flight production runs are shown. These tests have exceeded 19,000 hours are continuing to provide increased confidence in the predicted long term performance of the Cassini RTGs. Test results are presented for full scale units both ETGs (E-6, E-7) and RTGs (F-2, F-5) along with mission power predictions. F-5, fueled in 1985, served as a spare for the Galileo and Ulysses missions and plays the same role in the Cassini program. It has successfully completed all acceptance testing. The ten years storage between thermal vacuum tests is the longest ever experienced by an RTG. The data from this test are unique in providing the effects of long term low temperature storage on power output. All ETG and RTG test results to date indicate that the power requirements of the Cassini spacecraft will be met. BOM and EOM power margins of at least five percent are predicted

The SPQR experiment: detecting damage to orbiting spacecraft with ground-based telescopes

Science.gov (United States)

Paolozzi, Antonio; Porfilio, Manfredi; Currie, Douglas G.; Dantowitz, Ronald F.

2007-09-01

The objective of the Specular Point-like Quick Reference (SPQR) experiment was to evaluate the possibility of improving the resolution of ground-based telescopic imaging of manned spacecraft in orbit. The concept was to reduce image distortions due to atmospheric turbulence by evaluating the Point Spread Function (PSF) of a point-like light reference and processing the spacecraft image accordingly. The target spacecraft was the International Space Station (ISS) and the point-like reference was provided by a laser beam emitted by the ground station and reflected back to the telescope by a Cube Corner Reflector (CCR) mounted on an ISS window. The ultimate objective of the experiment was to demonstrate that it is possible to image spacecraft in Low Earth Orbit (LEO) with a resolution of 20 cm, which would have probably been sufficient to detect the damage which caused the Columbia disaster. The experiment was successfully performed from March to May 2005. The paper provides an overview of the SPQR experiment.

Electrical interferences observed in the Cassini CIRS spectrometer

Science.gov (United States)

Chan, Cheong; Albright, Shane; Gorius, Nicolas; Brasunas, John; Jennings, Don; Flasar, F. Michael; Carlson, Ronald; Guandique, Ever; Nixon, Conor

2015-06-01

The Composite Infrared Spectrometer (CIRS) carried onboard the Cassini spacecraft has now operated successfully for 17 years, following launch in 1997. Following insertion into Saturnian orbit in July 2004, the instrument has taken data nearly continuously, returning over 100 million interferograms (spectra) to date. Although of generally high quality, and resulting in more than 100 peer-reviewed scientific articles, the spectra are afflicted with several types of instrumental electrical (non-random) noise artifacts. These noise artifacts require either mitigation strategies (prevention), removal from the observed data, or else awareness of the affected spectral areas which must be excluded from scientific analysis. The sources and nature of these varied noise types were not readily identified until after launch. The purpose of this article is to inform users of the noise in the CIRS dataset and to serve as a `lesson-learned' guide for designers of future instruments.

Continuing Improvement in the Planetary Ephemeris with VLBA Observations of Cassini

Science.gov (United States)

Jones, Dayton L.; Folkner, William M.; Jacobson, Robert A.; Jacobs, Christopher S.; Romney, Jonathan D.; Dhawan, Vivek; Fomalont, Edward B.

2016-06-01

During the past decade a continuing series of measurements of the barycentric position of the Saturn system in the inertial International Celestial Reference Frame (ICRF) has led to a significant improvement in our knowledge of Saturn's orbit. This in turn has improved the current accuracy and time range of the solar system ephemeris produced and maintained by the Jet Propulsion Laboratory. Our observing technique involves high-precision astrometry of the radio signal from Cassini with the NRAO Very Long Baseline Array, combined with solutions for the orbital motion of Cassini about the Saturn barycenter from Doppler tracking by the Deep Space Network. Our VLBA astrometry is done in a phase-referencing mode, providing nrad-level relative positions between Cassini and angularly nearby extragalactic radio sources. The positions of those reference radio sources are tied to the ICRF through dedicated VLBI observations by several groups around the world. We will present recent results from our astrometric observations of Cassini through early 2016. This program will continue until the end of the Cassini mission in 2017, although future improvement in Saturn's orbit will be more incremental because we have already covered more that a quarter of Saturn's orbital period. The Juno mission to Jupiter, which will orbit Jupiter for about 1.5 years starting in July 2016, will provide an excellent opportunity for us to apply the same VLBA astrometry technique to improve the orbit of Jupiter by a factor of several. The National Radio Astronomy Observatory is a facility of the National Science Foundation operated under cooperative agreement by Associated Universities, Inc. This work made use of the Swinburne University of Technology software correlator, developed as part of the Australian Major National Research Facilities Program and operated under license. Part of this research was carried out at the Jet Propulsion Laboratory, California Institute of Technology, under contract

SSRPT (SSR Pointer Trackeer) for Cassini Mission Operations - A Ground Data Analysis Tool

Science.gov (United States)

Kan, E.

1998-01-01

Tracking the resources of the two redundant Solid State Recorders (SSR) is a necessary routine for Cassini spacecraft mission operations. Instead of relying on a full-fledged spacecraft hardware/software simulator to track and predict the SSR recording and playback pointer positions, a stand-alone SSR Pointer Tracker tool was developed as part of JPL's Multimission Spacecraft Analysis system.

RTG performance on Galileo and Ulysses and Cassini test results

International Nuclear Information System (INIS)

Kelly, C.E.; Klee, P.M.

1997-01-01

Power output from telemetry for the two Galileo RTGs are shown from the 1989 launch to the recent Jupiter encounter. Comparisons of predicted, measured and required performance are shown. Similar comparisons are made for the RTG on the Ulysses spacecraft which completed its planned mission in 1995. Also presented are test results from small scale thermoelectric modules and full scale converters performed for the Cassini program. The Cassini mission to Saturn is scheduled for an October 1997 launch. Small scale module test results on thermoelectric couples from the qualification and flight production runs are shown. These tests have exceeded 19,000 hours are continuing to provide increased confidence in the predicted long term performance of the Cassini RTGs. Test results are presented for full scale units both ETGs (E-6, E-7) and RTGs (F-2, F-5) along with mission power predictions. F-5, fueled in 1985, served as a spare for the Galileo and Ulysses missions and plays the same role in the Cassini program. It has successfully completed all acceptance testing. The ten years storage between thermal vacuum tests is the longest ever experienced by an RTG. The data from this test are unique in providing the effects of long term low temperature storage on power output. All ETG and RTG test results to date indicate that the power requirements of the Cassini spacecraft will be met. BOM and EOM power margins of at least five percent are predicted. copyright 1997 American Institute of Physics

Particle-in-Cell Simulation Study on the Floating Potential of Spacecraft in the Low Earth Orbit

International Nuclear Information System (INIS)

Tang Daotan; Yang Shengsheng; Zheng Kuohai; Qin Xiaogang; Li Detian; Liu Qing; Zhao Chengxuan; Du Shanshan

2015-01-01

In order to further understand the characteristics of the floating potential of low earth orbit spacecraft, the effects of the electron current collection area, background electron temperature, photocurrent emission, spacecraft wake, and the shape of spacecraft on spacecraft floating potential were studied here by particle-in-cell simulation in the low earth orbit. The simulation results show that the electron current collection area and background electron temperature impact on the floating potential by changing the electron current collection of spacecraft. By increasing the electron current collection area or background electron temperature, the spacecraft will float at a lower electric potential with respect to the surrounding plasma. However, the spacecraft wake affects the floating potential by increasing the ion current collected by spacecraft. The emission of the photocurrent from the spacecraft surface, which compensates for the electrons collected from background plasma, causes the floating potential to increase. The shape of the spacecraft is also an important factor influencing the floating potential. (paper)

Combined spacecraft orbit and attitude control through extended Kalman filtering of magnetometer, gyro, and GPS measurements

Directory of Open Access Journals (Sweden)

Tamer Mekky Ahmed Habib

2014-06-01

Full Text Available The main goal of this research is to establish spacecraft orbit and attitude control algorithms based on extended Kalman filter which provides estimates of spacecraft orbital and attitude states. The control and estimation algorithms must be capable of dealing with the spacecraft conditions during the detumbling and attitude acquisition modes of operation. These conditions are characterized by nonlinearities represented by large initial attitude angles, large initial angular velocities, large initial attitude estimation error, and large initial position estimation error. All of the developed estimation and control algorithms are suitable for application to the next Egyptian scientific satellite, EGYPTSAT-2. The parameters of the case-study spacecraft are similar but not identical to the former Egyptian satellite EGYPTSAT-1. This is done because the parameters of EGYPTSAT-2 satellite have not been consolidated yet. The sensors utilized are gyro, magnetometer, and GPS. Gyro and magnetometer are utilized to provide measurements for the estimates of spacecraft attitude state vector where as magnetometer and GPS are utilized to provide measurements for the estimates of spacecraft orbital state vector.

Contingency Trajectory Design for a Lunar Orbit Insertion Maneuver Failure by the LADEE Spacecraft

Science.gov (United States)

Genova, A. L.

2014-01-01

This paper presents results from a contingency trajectory analysis performed for the Lunar Atmosphere & Dust Environment Explorer (LADEE) mission in the event of a missed lunar-orbit insertion (LOI) maneuver by the LADEE spacecraft. The effects of varying solar perturbations in the vicinity of the weak stability boundary (WSB) in the Sun-Earth system on the trajectory design are analyzed and discussed. It is shown that geocentric recovery trajectory options existed for the LADEE spacecraft, depending on the spacecraft's recovery time to perform an Earth escape-prevention maneuver after the hypothetical LOI maneuver failure and subsequent path traveled through the Sun-Earth WSB. If Earth-escape occurred, a heliocentric recovery option existed, but with reduced science capacapability for the spacecraft in an eccentric, not circular near-equatorial retrograde lunar orbit.

GPHS-RTGs in support of the Cassini Mission. Semi annual technical progress report, 1 April 1996--29 September 1996

International Nuclear Information System (INIS)

1996-01-01

This technical progress report discusses work on the Radioisotope Generators and Ancillary Activities for the Cassini spacecraft. The Cassini spacecraft is expected to launch in October 1997, and will explore Saturn and its moons. This progress report discusses issues in: spacecraft integration and liason, engineering support, safety, qualified unicouple fabrication, ETG fabrication and testing, ground support equipment, RTG shipping and launch support, designs, reviews and mission application. Safety analysis of the RTGs during reentry and launch accidents are covered. This report covers the period of April 1 to September 29, 1996

Design and analysis of RTGs for CRAF and Cassini missions

International Nuclear Information System (INIS)

Schock, A.; Noravian, H.; Or, C.; Sankarankandath, K.

1991-01-01

The paper describes the design and analysis of Radioisotope Thermoelectric Generators (RTGs) integrated with the Jet Proplusion Laboratory's CRAF (Comet Rendezvous and Asteroid Flyby) and Cassini Spacecraft. The principal purpose of the CRAF mission is the study of Asteroids and comets, and the principal purpose of the Cassini mission is the study of asteroids, Saturn, and its moons (particularly Titan). Both misions will employ the Mariner/Mark-2 spacecraft, and each will be powered by two GPHS-RTGs (General Purpose Heat Source-RTGs). JPL's spacecraft designers wish to locate the two RTGs in close proximity to each other, resulting in mutual and unsymmetrical obstruction of their heat rejection paths. To support JPL's design studies, the U.S. Department of Energy asked Fairchild to determine the effect of the RTGs' proximity on their power output. As described in the paper, this required the development of novel analysis methods and computer codes for the coupled thermal and electrical analysis of obstructed RTGs with axial and circumferential temperature, voltage, and current variations. The code was validated against measured data of unobstructed RTG tests, and was used for the detailed analysis of the obstructed CRAF and Cassini RTGs. Also described is a new method for predicting the combined effect of fuel decay and thermoelectric degradation on the output of obstructed RTGs, which amounts for the effect of diminishing temperatures on degradation rates. For the 24-degree separation angle of JPL's original baseline design, and for the 35-degree RTG separation of JPL's revised design, the computed results indicate that the mutually obstructed GPHS/RTGs with standard fuel loading and operating temperatures can comfortably meet the JPL-specified power requirements for the CRAF mission and almost meet the specified requirements for the Cassini mission

Characterizing Observed Limit Cycles in the Cassini Main Engine Guidance Control System

Science.gov (United States)

Rizvi, Farheen; Weitl, Raquel M.

2011-01-01

The Cassini spacecraft dynamics-related telemetry during long Main Engine (ME) burns has indicated the presence of stable limit cycles between 0.03-0.04 Hz frequencies. These stable limit cycles cause the spacecraft to possess non-zero oscillating rates for extended periods of time. This indicates that the linear ME guidance control system does not model the complete dynamics of the spacecraft. In this study, we propose that the observed limit cycles in the spacecraft dynamics telemetry appear from a stable interaction between the unmodeled nonlinear elements in the ME guidance control system. Many nonlinearities in the control system emerge from translating the linear engine gimbal actuator (EGA) motion into a spacecraft rotation. One such nonlinearity comes from the gear backlash in the EGA system, which is the focus of this paper. The limit cycle characteristics and behavior can be predicted by modeling this gear backlash nonlinear element via a describing function and studying the interaction of this describing function with the overall dynamics of the spacecraft. The linear ME guidance controller and gear backlash nonlinearity are modeled analytically. The frequency, magnitude, and nature of the limit cycle are obtained from the frequency response of the ME guidance controller and nonlinear element. In addition, the ME guidance controller along with the nonlinearity is simulated. The simulation response contains a limit cycle with similar characterstics as predicted analytically: 0.03-0.04 Hz frequency and stable, sustained oscillations. The analytical and simulated limit cycle responses are compared to the flight telemetry for long burns such as the Saturn Orbit Insertion and Main Engine Orbit Trim Maneuvers. The analytical and simulated limit cycle characteristics compare well with the actual observed limit cycles in the flight telemetry. Both have frequencies between 0.03-0.04 Hz and stable oscillations. This work shows that the stable limit cycles occur

METRIC: A Dedicated Earth-Orbiting Spacecraft for Investigating Gravitational Physics and the Space Environment

Directory of Open Access Journals (Sweden)

Roberto Peron

2017-07-01

Full Text Available A dedicated mission in low Earth orbit is proposed to test predictions of gravitational interaction theories and to directly measure the atmospheric density in a relevant altitude range, as well as to provide a metrological platform able to tie different space geodesy techniques. The concept foresees a small spacecraft to be placed in a dawn-dusk eccentric orbit between 450 and 1200 km of altitude. The spacecraft will be tracked from the ground with high precision, and a three-axis accelerometer package on-board will measure the non-gravitational accelerations acting on its surface. Estimates of parameters related to fundamental physics and geophysics should be obtained by a precise orbit determination, while the accelerometer data will be instrumental in constraining the atmospheric density. Along with the mission scientific objectives, a conceptual configuration is described together with an analysis of the dynamical environment experienced by the spacecraft and the accelerometer.

ADRC for spacecraft attitude and position synchronization in libration point orbits

Science.gov (United States)

Gao, Chen; Yuan, Jianping; Zhao, Yakun

2018-04-01

This paper addresses the problem of spacecraft attitude and position synchronization in libration point orbits between a leader and a follower. Using dual quaternion, the dimensionless relative coupled dynamical model is derived considering computation efficiency and accuracy. Then a model-independent dimensionless cascade pose-feedback active disturbance rejection controller is designed to spacecraft attitude and position tracking control problems considering parameter uncertainties and external disturbances. Numerical simulations for the final approach phase in spacecraft rendezvous and docking and formation flying are done, and the results show high-precision tracking errors and satisfactory convergent rates under bounded control torque and force which validate the proposed approach.

Model predictive control for spacecraft rendezvous in elliptical orbit

Science.gov (United States)

Li, Peng; Zhu, Zheng H.

2018-05-01

This paper studies the control of spacecraft rendezvous with attitude stable or spinning targets in an elliptical orbit. The linearized Tschauner-Hempel equation is used to describe the motion of spacecraft and the problem is formulated by model predictive control. The control objective is to maximize control accuracy and smoothness simultaneously to avoid unexpected change or overshoot of trajectory for safe rendezvous. It is achieved by minimizing the weighted summations of control errors and increments. The effects of two sets of horizons (control and predictive horizons) in the model predictive control are examined in terms of fuel consumption, rendezvous time and computational effort. The numerical results show the proposed control strategy is effective.

Gravimagnetic effect of the barycentric motion of the Sun and determination of the post-Newtonian parameter γ in the Cassini experiment

Science.gov (United States)

Kopeikin, S. M.; Polnarev, A. G.; Schäfer, G.; Vlasov, I. Yu.

2007-07-01

The most precise test of the post-Newtonian γ parameter in the solar system has been achieved in measurement of the frequency shift of radio waves to and from the Cassini spacecraft as they passed near the Sun. The test relies upon the JPL model of radiowave propagation that includes, but does not explicitly parametrize, the impact of the non-stationary component of the gravitational field of the Sun, generated by its barycentric orbital motion, on the Shapiro delay. This non-stationary gravitational field of the Sun is associated with the Lorentz transformation of the metric tensor and the affine connection from the heliocentric to the barycentric frame of the solar system and can be treated as gravimagnetic field. The gravimagnetic field perturbs the propagation of a radio wave and contributes to its frequency shift at the level up to 4×10-13 that may affect the precise measurement of the parameter γ in the Cassini experiment to about one part in 10 000. Our analysis suggests that the translational gravimagnetic field of the Sun can be extracted from the Cassini data, and its effect is separable from the space curvature characterized by the parameter γ.

Automation of Cassini Support Imaging Uplink Command Development

Science.gov (United States)

Ly-Hollins, Lisa; Breneman, Herbert H.; Brooks, Robert

2010-01-01

"Support imaging" is imagery requested by other Cassini science teams to aid in the interpretation of their data. The generation of the spacecraft command sequences for these images is performed by the Cassini Instrument Operations Team. The process initially established for doing this was very labor-intensive, tedious and prone to human error. Team management recognized this process as one that could easily benefit from automation. Team members were tasked to document the existing manual process, develop a plan and strategy to automate the process, implement the plan and strategy, test and validate the new automated process, and deliver the new software tools and documentation to Flight Operations for use during the Cassini extended mission. In addition to the goals of higher efficiency and lower risk in the processing of support imaging requests, an effort was made to maximize adaptability of the process to accommodate uplink procedure changes and the potential addition of new capabilities outside the scope of the initial effort.

Effects of the Eccentricity of a Perturbing Third Body on the Orbital Correction Maneuvers of a Spacecraft

Directory of Open Access Journals (Sweden)

R. C. Domingos

2014-01-01

Full Text Available The fuel consumption required by the orbital maneuvers when correcting perturbations on the orbit of a spacecraft due to a perturbing body was estimated. The main goals are the measurement of the influence of the eccentricity of the perturbing body on the fuel consumption required by the station keeping maneuvers and the validation of the averaged methods when applied to the problem of predicting orbital maneuvers. To study the evolution of the orbits, the restricted elliptic three-body problem and the single- and double-averaged models are used. Maneuvers are made by using impulsive and low thrust maneuvers. The results indicated that the averaged models are good to make predictions for the orbital maneuvers when the spacecraft is in a high inclined orbit. The eccentricity of the perturbing body plays an important role in increasing the effects of the perturbation and the fuel consumption required for the station keeping maneuvers. It is shown that the use of more frequent maneuvers decreases the annual cost of the station keeping to correct the orbit of a spacecraft. An example of an eccentric planetary system of importance to apply the present study is the dwarf planet Haumea and its moons, one of them in an eccentric orbit.

In-orbit evaluation of the control system/structural mode interactions of the OSO-8 spacecraft

Science.gov (United States)

Slafer, L. I.

1979-01-01

The Orbiting Solar Observatory-8 experienced severe structural mode/control loop interaction problems during the spacecraft development. Extensive analytical studies, using the hybrid coordinate modeling approach, and comprehensive ground testing were carried out in order to achieve the system's precision pointing performance requirements. A recent series of flight tests were conducted with the spacecraft in which a wide bandwidth, high resolution telemetry system was utilized to evaluate the on-orbit flexible dynamics characteristics of the vehicle along with the control system performance. The paper describes the results of these tests, reviewing the basic design problem, analytical approach taken, ground test philosophy, and on-orbit testing. Data from the tests was used to determine the primary mode frequency, damping, and servo coupling dynamics for the on-orbit condition. Additionally, the test results have verified analytically predicted differences between the on-orbit and ground test environments, and have led to a validation of both the analytical modeling and servo design techniques used during the development of the control system.

Chaos and its control in the pitch motion of an asymmetric magnetic spacecraft in polar elliptic orbit

Energy Technology Data Exchange (ETDEWEB)

Inarrea, Manuel [Universidad de La Rioja, Area de Fisica Aplicada, 26006 Logrono (Spain)], E-mail: manuel.inarrea@unirioja.es

2009-05-30

We study the pitch attitude dynamics of an asymmetric magnetic spacecraft in a polar almost circular orbit under the influence of a gravity gradient torque. The spacecraft is perturbed by the small eccentricity of the elliptic orbit and by a small magnetic torque generated by the interaction between the Earth's magnetic field and the magnetic moment of the spacecraft. Under both perturbations, we show that the pitch motion exhibits heteroclinic chaotic behavior by means of the Melnikov method. Numerical methods applied to simulations of the pitch motion also confirm the chaotic character of the spacecraft attitude dynamics. Finally, a linear time-delay feedback method for controlling chaos is applied to the governing equations of the spacecraft pitch motion in order to remove the chaotic character of initially irregular attitude motions and transform them into periodic ones.

Chaos and its control in the pitch motion of an asymmetric magnetic spacecraft in polar elliptic orbit

International Nuclear Information System (INIS)

Inarrea, Manuel

2009-01-01

We study the pitch attitude dynamics of an asymmetric magnetic spacecraft in a polar almost circular orbit under the influence of a gravity gradient torque. The spacecraft is perturbed by the small eccentricity of the elliptic orbit and by a small magnetic torque generated by the interaction between the Earth's magnetic field and the magnetic moment of the spacecraft. Under both perturbations, we show that the pitch motion exhibits heteroclinic chaotic behavior by means of the Melnikov method. Numerical methods applied to simulations of the pitch motion also confirm the chaotic character of the spacecraft attitude dynamics. Finally, a linear time-delay feedback method for controlling chaos is applied to the governing equations of the spacecraft pitch motion in order to remove the chaotic character of initially irregular attitude motions and transform them into periodic ones.

On-Orbit Assembly of a Universally Interlocking Modular Spacecraft (7224-110), Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Physical Sciences Inc. and Advanced Solutions, Inc. propose a novel approach for on-orbit assembly of a modular spacecraft using a unique universal, intelligent,...

Laboratory investigations: Low Earth orbit environment chemistry with spacecraft surfaces

Science.gov (United States)

Cross, Jon B.

1990-01-01

Long-term space operations that require exposure of material to the low earth orbit (LEO) environment must take into account the effects of this highly oxidative atmosphere on material properties and the possible contamination of the spacecraft surroundings. Ground-based laboratory experiments at Los Alamos using a newly developed hyperthermal atomic oxygen (AO) source have shown that not only are hydrocarbon based materials effected but that inorganic materials such as MoS2 are also oxidized and that thin protective coatings such as Al2O3 can be breached, producing oxidation of the underlying substrate material. Gas-phase reaction products, such as SO2 from oxidation of MoS2 and CO and CO2 from hydrocarbon materials, have been detected and have consequences in terms of spacecraft contamination. Energy loss through gas-surface collisions causing spacecraft drag has been measured for a few select surfaces and has been found to be highly dependent on the surface reactivity.

Searching Less Perturbed Circular Orbits for a Spacecraft Travelling around Europa

Directory of Open Access Journals (Sweden)

J. P. S. Carvalho

2014-01-01

Full Text Available Space missions to visit the natural satellite of Jupiter, Europa, constitute an important topic in space activities today, because missions to this moon are under study now. Several considerations have to be made for these missions. The present paper searches for less perturbed circular orbits around Europa. This search is made based on the total effects of the perturbing forces over the time, evaluated by the integral of those forces over the time. This value depends on the dynamical model and on the orbit of the spacecraft. The perturbing forces considered are the third-body perturbation that comes from Jupiter and the J2, J3, and C22 terms of the gravitational potential of Europa. Several numerical studies are performed and the results show the locations of the less perturbed orbits. Using those results, it is possible to find near-circular frozen orbits with smaller amplitudes of variations of the orbital elements.

Estimation of Gravitation Parameters of Saturnian Moons Using Cassini Attitude Control Flight Data

Science.gov (United States)

Krening, Samantha C.

2013-01-01

A major science objective of the Cassini mission is to study Saturnian satellites. The gravitational properties of each Saturnian moon is of interest not only to scientists but also to attitude control engineers. When the Cassini spacecraft flies close to a moon, a gravity gradient torque is exerted on the spacecraft due to the mass of the moon. The gravity gradient torque will alter the spin rates of the reaction wheels (RWA). The change of each reaction wheel's spin rate might lead to overspeed issues or operating the wheel bearings in an undesirable boundary lubrication condition. Hence, it is imperative to understand how the gravity gradient torque caused by a moon will affect the reaction wheels in order to protect the health of the hardware. The attitude control telemetry from low-altitude flybys of Saturn's moons can be used to estimate the gravitational parameter of the moon or the distance between the centers of mass of Cassini and the moon. Flight data from several low altitude flybys of three Saturnian moons, Dione, Rhea, and Enceladus, were used to estimate the gravitational parameters of these moons. Results are compared with values given in the literature.

Enhancing Cassini Operations & Science Planning Tools

Science.gov (United States)

Castello, Jonathan

2012-01-01

The Cassini team uses a variety of software utilities as they manage and coordinate their mission to Saturn. Most of these tools have been unchanged for many years, and although stability is a virtue for long-lived space missions, there are some less-fragile tools that could greatly benefit from modern improvements. This report shall describe three such upgrades, including their architectural differences and their overall impact. Emphasis is placed on the motivation and rationale behind architectural choices rather than the final product, so as to illuminate the lessons learned and discoveries made.These three enhancements included developing a strategy for migrating Science Planning utilities to a new execution model, rewriting the team's internal portal for ease of use and maintenance, and developing a web-based agenda application for tracking the sequence of files being transmitted to the Cassini spacecraft. Of this set, the first two have been fully completed, while the agenda application is currently in the early prototype stage.

LO2/LH2 propulsion for outer planet orbiter spacecraft

Science.gov (United States)

Garrison, P. W.; Sigurdson, K. B.

1983-01-01

Galileo class orbiter missions (750-1500 kg) to the outer planets require a large postinjection delta-V for improved propulsion performance. The present investigation shows that a pump-fed low thrust LO2/LH2 propulsion system can provide a significantly larger net on-orbit mass for a given delta-V than a state-of-the-art earth storable, N2O4/monomethylhydrazine pressure-fed propulsion system. A description is given of a conceptual design for a LO2/LH2 pump-fed propulsion system developed for a Galileo class mission to the outer planets. Attention is given to spacecraft configuration, details regarding the propulsion system, the thermal control of the cryogenic propellants, and aspects of mission performance.

A Small Spacecraft Swarm Deployment and Stationkeeping Strategy for Sun-Earth L1 Halo Orbits

Science.gov (United States)

Renea Conn, Tracie; Bookbinder, Jay

2018-01-01

Spacecraft orbits about the Sun-Earth librarian point L1 have been of interest since the 1950s. An L1 halo orbit was first achieved with the International Sun-Earth Explorer-3 (ISEE-3) mission, and similar orbits around Sun-Earth L1 were achieved in the Solar and Heliospheric Observatory (SOHO), Advanced Composition Explorer (ACE), Genesis, and Deep Space Climate Observatory (DSCOVR) missions. With recent advancements in CubeSat technology, we envision that it will soon be feasible to deploy CubeSats at L1. As opposed to these prior missions where one large satellite orbited alone, a swarm of CubeSats at L1 would enable novel science data return, providing a topology for intersatellite measurements of heliophysics phenomena both spatially and temporally, at varying spatial scales.The purpose of this iPoster is to present a flight dynamics strategy for a swarm of numerous CubeSats orbiting Sun-Earth L1. The presented method is a coupled, two-part solution. First, we present a deployment strategy for the CubeSats that is optimized to produce prescribed, time-varying intersatellite baselines for the purposes of collecting magnetometer data as well as radiometric measurements from cross-links. Second, we employ a loose control strategy that was successfully applied to SOHO and ACE for minimized stationkeeping propellant expenditure. We emphasize that the presented solution is practical within the current state-of-the-art and heritage CubeSat technology, citing capabilities of CubeSat designs that will launch on the upcoming Exploration Mission 1 (EM-1) to lunar orbits and beyond. Within this iPoster, we present animations of the simulated deployment strategy and resulting spacecraft trajectories. Mission design parameters such as total Δv required for long-term station keeping and minimum/maximum/mean spacecraft separation distances are also presented.

IMF dependence of Saturn's auroras: modelling study of HST and Cassini data from 12–15 February 2008

Directory of Open Access Journals (Sweden)

E. S. Belenkaya

2010-08-01

Full Text Available To gain better understanding of auroral processes in Saturn's magnetosphere, we compare ultraviolet (UV auroral images obtained by the Hubble Space Telescope (HST with the position of the open-closed field line boundary in the ionosphere calculated using a magnetic field model that employs Cassini measurements of the interplanetary magnetic field (IMF as input. Following earlier related studies of pre-orbit insertion data from January 2004 when Cassini was located ~ 1300 Saturn radii away from the planet, here we investigate the interval 12–15 February 2008, when UV images of Saturn's southern dayside aurora were obtained by the HST while the Cassini spacecraft measured the IMF in the solar wind just upstream of the dayside bow shock. This configuration thus provides an opportunity, unique to date, to determine the IMF impinging on Saturn's magnetosphere during imaging observations, without the need to take account of extended and uncertain interplanetary propagation delays. The paraboloid model of Saturn's magnetosphere is then employed to calculate the magnetospheric magnetic field structure and ionospheric open-closed field line boundary for averaged IMF vectors that correspond, with appropriate response delays, to four HST images. We show that the IMF-dependent open field region calculated from the model agrees reasonably well with the area lying poleward of the UV emissions, thus supporting the view that the poleward boundary of Saturn's auroral oval in the dayside ionosphere lies adjacent to the open-closed field line boundary.

Improved Atlases of Mimas and Enceladus derived from Cassini-ISS images

Science.gov (United States)

Roatsch, T.; Kersten, E.; Matz, K. D.; Bland, M. T.; Becker, T. L.; Patterson, G. W.

2017-12-01

The Cassini Imaging Science Subsystem (ISS) took a couple of high-resolution images of the Icy satellites Mimas and Enceladus during the last few years of the Cassini mission. Both satellites were captured over a period of non-targeted flybys: Mimas in 2016 and 2017 in orbits 230, 249, and 259 and Enceladus in 2015 and 2016 in orbits 224, 228, and 250. We used the new Mimas images to improve the existing semi-controlled mosaic of Mimas. A new controlled Enceladus mosaic was published recently [1] and was now updated using the latest Enceladus images. Both new mosaics are the baseline for improved atlases of Mimas in 3 tiles with a scale of 1:1,000,000 and Enceladus in 15 tiles with a scale of 1:500,000. The nomenclature for both satellites was proposed by the Cassini-ISS team and approved by the IAU and was not changed here. Examples of the improved atlases will be shown in this presentation. Reference: [1] Bland, M.T. et. al., A new Enceladus base map and global control network in support of geological mapping, 46th Lunar and Planetary Science Conference (2015) , abstract 2303.

Comparison of technologies for deorbiting spacecraft from low-earth-orbit at end of mission

Science.gov (United States)

Sánchez-Arriaga, G.; Sanmartín, J. R.; Lorenzini, E. C.

2017-09-01

An analytical comparison of four technologies for deorbiting spacecraft from Low-Earth-Orbit at end of mission is presented. Basic formulas based on simple physical models of key figures of merit for each device are found. Active devices - rockets and electrical thrusters - and passive technologies - drag augmentation devices and electrodynamic tethers - are considered. A basic figure of merit is the deorbit device-to-spacecraft mass ratio, which is, in general, a function of environmental variables, technology development parameters and deorbit time. For typical state-of-the-art values, equal deorbit time, middle inclination and initial altitude of 850 km, the analysis indicates that tethers are about one and two orders of magnitude lighter than active technologies and drag augmentation devices, respectively; a tether needs a few percent mass-ratio for a deorbit time of a couple of weeks. For high inclination, the performance drop of the tether system is moderate: mass ratio and deorbit time increase by factors of 2 and 4, respectively. Besides collision risk with other spacecraft and system mass considerations, such as main driving factors for deorbit space technologies, the analysis addresses other important constraints, like deorbit time, system scalability, manoeuver capability, reliability, simplicity, attitude control requirement, and re-entry and multi-mission capability (deorbit and re-boost) issues. The requirements and constraints are used to make a critical assessment of the four technologies as functions of spacecraft mass and initial orbit (altitude and inclination). Emphasis is placed on electrodynamic tethers, including the latest advances attained in the FP7/Space project BETs. The superiority of tape tethers as compared to round and multi-line tethers in terms of deorbit mission performance is highlighted, as well as the importance of an optimal geometry selection, i.e. tape length, width, and thickness, as function of spacecraft mass and initial

Towards a standardized grasping and refuelling on-orbit servicing for geo spacecraft

Science.gov (United States)

Medina, Alberto; Tomassini, Angelo; Suatoni, Matteo; Avilés, Marcos; Solway, Nick; Coxhill, Ian; Paraskevas, Iosif S.; Rekleitis, Georgios; Papadopoulos, Evangelos; Krenn, Rainer; Brito, André; Sabbatinelli, Beatrice; Wollenhaupt, Birk; Vidal, Christian; Aziz, Sarmad; Visentin, Gianfranco

2017-05-01

Exploitation of space must benefit from the latest advances in robotics. On-orbit servicing is a clear candidate for the application of autonomous rendezvous and docking mechanisms. However, during the last three decades most of the trials took place combining extravehicular activities (EVAs) with telemanipulated robotic arms. The European Space Agency (ESA) considers that grasping and refuelling are promising near-mid-term capabilities that could be performed by servicing spacecraft. Minimal add-ons on spacecraft to enhance their serviceability may protect them for a changing future in which satellite servicing may become mainstream. ESA aims to conceive and promote standard refuelling provisions that can be installed in present and future European commercial geostationary orbit (GEO) satellite platforms and scientific spacecraft. For this purpose ESA has started the ASSIST activity addressing the analysis, design and validation of internal provisions (such as modifications to fuel, gas, electrical and data architecture to allow servicing) and external provisions (such as integrated berthing fixtures with peripheral electrical, gas, liquid connectors, leak check systems and corresponding optical and radio markers for cooperative rendezvous and docking). This refuelling approach is being agreed with European industry (OHB, Thales Alenia Space) and expected to be consolidated with European commercial operators as a first step to become an international standard; this approach is also being considered for on-orbit servicing spacecraft, such as the SpaceTug, by Airbus DS. This paper describes in detail the operational means, structure, geometry and accommodation of the system. Internal and external provisions will be designed with the minimum possible impact on the current architecture of GEO satellites without introducing additional risks in the development and commissioning of the satellite. End-effector and berthing fixtures are being designed in the range of few

Is the Recently Proposed Mars-Sized Perturber at 65–80 AU Ruled Out by the Cassini Ranging Data?

Energy Technology Data Exchange (ETDEWEB)

Iorio, Lorenzo, E-mail: lorenzo.iorio@libero.it [Ministero dell' Istruzione, dell' Università e della Ricerca, Rome (Italy)

2017-10-26

Recently, the existence of a pointlike pertuber PX with 1 m{sub ♂} ≲ m{sub X} ≲ 2.4 m{sub ⊕} (the symbol “♂” denotes Mars) supposedly moving at 65–80 AU along a moderately inclined orbit has been hypothesized in order to explain certain features of the midplane of the Kuiper Belt Objects (KBOs). We preliminarily selected two possible scenarios for such a PX, and numerically simulated its effect on the Earth-Saturn range ρ(t) by varying some of its orbital parameters over a certain time span; then, we compared our results with some existing actual range residuals. By assuming m{sub X} = 1 m{sub ♂} and a circular orbit, such a putative new member of our Solar System would nominally perturb ρ(t) by a few km over Δt = 12 year (2004 − 2016). However, the Cassini spacecraft accurately measured ρ(t) to the level of σ{sub ρ} ≃ 100 m. Nonetheless, such a scenario should not be considered as necessarily ruled out since the Cassini data were reduced so far without explicitly modeling any PX. Indeed, a NASA JPL team recently demonstrated that an extra-signature as large as 4 km affecting the Kronian range would be almost completely absorbed in fitting incomplete dynamical models, i.e., without PX itself, to such simulated data, thus not showing up in the standard post-fit range residuals. Larger anomalous signatures would instead occur for m{sub X} > 1 m{sub ♂}. Their nominal amplitude could be as large as 50 − 150 km for m{sub X} = 2.4 m{sub ⊕}, thus making less plausible their existence.

Is the Recently Proposed Mars-Sized Perturber at 65–80 AU Ruled Out by the Cassini Ranging Data?

Directory of Open Access Journals (Sweden)

Lorenzo Iorio

2017-10-01

Full Text Available Recently, the existence of a pointlike pertuber PX with 1 m♂ ≲ mX ≲ 2.4 m⊕ (the symbol “♂” denotes Mars supposedly moving at 65–80 AU along a moderately inclined orbit has been hypothesized in order to explain certain features of the midplane of the Kuiper Belt Objects (KBOs. We preliminarily selected two possible scenarios for such a PX, and numerically simulated its effect on the Earth-Saturn range ρ(t by varying some of its orbital parameters over a certain time span; then, we compared our results with some existing actual range residuals. By assuming mX = 1 m♂ and a circular orbit, such a putative new member of our Solar System would nominally perturb ρ(t by a few km over Δt = 12 year (2004 − 2016. However, the Cassini spacecraft accurately measured ρ(t to the level of σρ ≃ 100 m. Nonetheless, such a scenario should not be considered as necessarily ruled out since the Cassini data were reduced so far without explicitly modeling any PX. Indeed, a NASA JPL team recently demonstrated that an extra-signature as large as 4 km affecting the Kronian range would be almost completely absorbed in fitting incomplete dynamical models, i.e., without PX itself, to such simulated data, thus not showing up in the standard post-fit range residuals. Larger anomalous signatures would instead occur for mX > 1 m♂. Their nominal amplitude could be as large as 50 − 150 km for mX = 2.4 m⊕, thus making less plausible their existence.

Ongoing Analysis of Jupiter's Equatorial Hotspots and Plumes from Cassini

Science.gov (United States)

Choi, D. S.; Showmwn, A. P.; Vasavada, A. R.; Simon-Miller, A. A.

2012-01-01

We present updated results from our ongoing analysis of Cassini observations of Jupiter's equatorial meteorology. For two months preceding the spacecraft's closest approach of the planet, the ISS instrument onboard Cassini regularly imaged the atmosphere of Jupiter. We created time-lapse movies from this period that show the complex activity and interactions of the equatorial atmosphere. During this period, hot spots exhibited significant variations in size and shape over timescales of days and weeks. Some of these changes appear to be a result of interactions with passing vortex systems in adjacent latitudes. Strong anticyclonic gyres to the southeast of the dark areas converge with flow from the west and appear to circulate into a hot spot at its southwestern corner.

SAMSI: An orbiting spatial interferometer for micro-arc second astronomical observations. [Spacecraft Array for Michelson Spatial Interferometry (SAMSI)

Science.gov (United States)

Stachnik, R. V.; Gezari, D. Y.

1985-01-01

The concept and performance of (SAMSI) Spacecraft Array for Michelson Spatial Interferometry, an orbiting spatial interferometer comprised of three free-flying spacecraft, two collector telescopes and a central mixing station are described. In the one-dimensional interferometry mode orbits exist which provide natural scanning of the baseline. These orbits place extremely small demands on thrusters and fuel consumption. Resolution of 0.00001 arcsecond and magnitude limits of mv = 15 to 20 are achievable in a single orbit. In the imaging mode, SAMSI could synthesize images equivalent to those produced by equal diameter filled apertures in space, making use of the fuel resupply capability of a space station. Simulations indicate that image reconstruction can be performed with milliarcsecond resolution to a visual magnitude 12 in 12 hr of spiral scanning integration time.

High-precision relative position and attitude measurement for on-orbit maintenance of spacecraft

Science.gov (United States)

Zhu, Bing; Chen, Feng; Li, Dongdong; Wang, Ying

2018-02-01

In order to realize long-term on-orbit running of satellites, space stations, etc spacecrafts, in addition to the long life design of devices, The life of the spacecraft can also be extended by the on-orbit servicing and maintenance. Therefore, it is necessary to keep precise and detailed maintenance of key components. In this paper, a high-precision relative position and attitude measurement method used in the maintenance of key components is given. This method mainly considers the design of the passive cooperative marker, light-emitting device and high resolution camera in the presence of spatial stray light and noise. By using a series of algorithms, such as background elimination, feature extraction, position and attitude calculation, and so on, the high precision relative pose parameters as the input to the control system between key operation parts and maintenance equipment are obtained. The simulation results show that the algorithm is accurate and effective, satisfying the requirements of the precision operation technique.

Detecting dust hits at Enceladus, Saturn and beyond using CAPS / ELS data from Cassini

Science.gov (United States)

Vandegriff, J. D.; Stoneberger, P. J.; Jones, G.; Waite, J. H., Jr.

2016-12-01

It has recently been shown (1) that the impact of hypervelocity dust grains on the Cassini spacecraft can be detected by the Cassini Plasma Spectrometer (CAPS) Electron Spectrometer (ELS) instrument. For multiple Enceladus flybys, fine scale features in the lower energy regime of ELS energy spectra can be explained as short-duration, isotropic plasma clouds due to dust impacts. We have developed an algorithm for detecting these hypervelocity dust impacts, and the list of such impacts during Enceladus flybys will be presented. We also present preliminary results obtained when using the algorithm to search for dust impacts in other regions of Saturn's magnetosphere as well as in the solar wind. (1) Jones, Geraint, Hypervelocity dust impact signatures detected by Cassini CAPS-ELS in the Enceladus plume, MOP Meeting, June 1-5, 2015, Atlanta, GA

Cassini RTG's -- Small scale module tests

International Nuclear Information System (INIS)

Kelly, C.E.; Klee, P.M.

1994-01-01

The Cassini spacecraft, scheduled for a 1997 launch to Saturn, will be powered by three GPHS RTGs (General Purpose Heat Source Radioisotope thermoelectric Generators). The RTGs are the same type as those powering the Galileo and Ulysses spacecraft. Three new converters (F-6, F-7, and F-8) are to be built and one converter (F-2) remaining from the GPHS program will be used. F-6 and F-7 are to be fueled and F-8 serves as a spare converter. In addition, the back-up RTG (F-5) from the Ulysses launch, which is still fueled, will serve as the Cassini back-up RTG. The new RTGs will have a lower fuel loading than in the past and will provide a minimum of 276 watts each at B.O.M. (beginning of mission). The mission length is 10.75 years, at which time these RTGs will provide a minimum of 216 watts and a possible extension to 16 years when the power will be 199 watts. This paper discusses tests performed to date to confirm the successful re-establishment of the unicouple production at Martin Marietta. This production line, shut down 10 years ago, has been restarted and over 1,500 unicouples have been produced to date. Confirmation will be primarily obtained by the performance of three small scale converters in comparison with previously tested modules from the Multi Hundred Watt (MHW) (Voyager) and GPHS (Galileo, Ulysses) programs. Test results to date have shown excellent agreement with the data base

A computer graphics system for visualizing spacecraft in orbit

Science.gov (United States)

Eyles, Don E.

1989-01-01

To carry out unanticipated operations with resources already in space is part of the rationale for a permanently manned space station in Earth orbit. The astronauts aboard a space station will require an on-board, spatial display tool to assist the planning and rehearsal of upcoming operations. Such a tool can also help astronauts to monitor and control such operations as they occur, especially in cases where first-hand visibility is not possible. A computer graphics visualization system designed for such an application and currently implemented as part of a ground-based simulation is described. The visualization system presents to the user the spatial information available in the spacecraft's computers by drawing a dynamic picture containing the planet Earth, the Sun, a star field, and up to two spacecraft. The point of view within the picture can be controlled by the user to obtain a number of specific visualization functions. The elements of the display, the methods used to control the display's point of view, and some of the ways in which the system can be used are described.

Semi-Autonomous Telerobotic Manipulation for On-Orbit Spacecraft Servicing and Assembly over Time-Delayed Telemetry

Data.gov (United States)

National Aeronautics and Space Administration — Despite the prevalence of this telerobotic surgery of humans, we still do not have the capability to service (refuel and repair) or assemble spacecraft on-orbit with...

Cassini RTG acceptance test results and RTG performance on Galileo and Ulysses

International Nuclear Information System (INIS)

Kelly, C.E.; Klee, P.M.

1997-01-01

Flight acceptance testing has been completed for the RTGs to be used on the Cassini spacecraft which is scheduled for an October 6, 1997 launch to Saturn. The acceptance test program includes vibration tests, magnetic field measurements, properties (weight and c.g.) and thermal vacuum test. This paper presents The thermal vacuum test results. Three RTGs are to be used, F-2, F-6, and F-7. F-5 is tile back-up RTG, as it was for the Galileo and Ulysses missions launched in 1989 and 1990, respectively. RTG performance measured during the thermal vacuum tests carried out at die Mound Laboratory facility met all specification requirements. Beginning of mission (BOM) and end of mission (EOM) power predictions have been made based on than tests results. BOM power is predicted to be 888 watts compared to the minimum requirement of 826 watts. Degradation models predict the EOM power after 16 years is to be 640 watts compared to a minimum requirement of 596 watts. Results of small scale module tests are also showing. The modules contain couples from the qualification and flight production runs. The tests have exceeded 28,000 hours (3.2 years) and are continuing to provide increased confidence in the predicted long term performance of the Cassini RTGs. All test results indicate that the power requirements of the Cassini spacecraft will be met. BOM and EOM power margins of over five percent are predicted. Power output from telemetry for the two Galileo RTGs are shown from the 1989 launch to the recent Jupiter encounter. Comparisons of predicted, measured and required performance are shown. Telemetry data are also shown for the RTG on the Ulysses spacecraft which completed its planned mission in 1995 and is now in the extended mission

Cassini RTG acceptance test results and RTG performance on Galileo and Ulysses

International Nuclear Information System (INIS)

Kelly, C.E.; Klee, P.M.

1997-01-01

Flight acceptance testing has been completed for the RTGs to be used on the Cassini spacecraft which is scheduled for an October 6, 1997 launch to Saturn. The acceptance test program includes vibration tests, magnetic field measurements, mass properties (weight and c.g.) and thermal vacuum test. This paper presents the thermal vacuum test results. Three RTGs are to be used, F-2, F-6, and F-7. F-5 is the backup RTG, as it was for the Galileo and Ulysses missions launched in 1989 and 1990, respectively. RTG performance measured during the thermal vacuum tests carried out at the Mound Laboratory facility met all specification requirements. Beginning of mission (BOM) and end of mission (EOM) power predictions have been made based on these tests results. BOM power is predicted to be 888 watts compared to the minimum requirement of 826 watts. Degradation models predict the EOM power after 16 years is to be 640 watts compared to a minimum requirement of 596 watts. Results of small scale module tests are also shown. The modules contain couples from the qualification and flight production runs. The tests have exceeded 28,000 hours (3.2 years) and are continuing to provide increased confidence in the predicted long term performance of the Cassini RTGs. All test results indicate that the power requirements of the Cassini spacecraft will be met. BOM and EOM power margins of over 5% are predicted. Power output from telemetry for the two Galileo RTGs are shown from the 1989 launch to the recent Jupiter encounter. Comparisons of predicted, measured and required performance are shown. Telemetry data are also shown for the RTG on the Ulysses spacecraft which completed its planned mission in 1995 and is now in the extended mission

Jupiter's magnetosphere and aurorae observed by the Juno spacecraft during its first polar orbits

DEFF Research Database (Denmark)

Connerney, J. E. P.; Adriani, Alberto; Allegrini, F.

2017-01-01

The Juno spacecraft acquired direct observations of the jovian magnetosphere and auroral emissions from a vantage point above the poles. Juno's capture orbit spanned the jovian magnetosphere from bow shock to the planet, providing magnetic field, charged particle, and wave phenomena context...

Health Physics Innovations Developed During Cassini for Future Space Applications

Science.gov (United States)

Nickell, Rodney E.; Rutherford, Theresa M.; Marmaro, George M.

1999-01-01

The long history of space flight includes missions that used Space Nuclear Auxiliary Power devices, starting with the Transit 4A Spacecraft (1961), continuing through the Apollo, Pioneer, Viking, Voyager, Galileo, Ulysses, Mars Pathfinder, and most recently, Cassini (1997). All Major Radiological Source (MRS) missions were processed at Kennedy Space Center/Cape Canaveral Air Station (KSC/CCAS) Launch Site in full compliance with program and regulatory requirements. The cumulative experience gained supporting these past missions has led to significant innovations which will be useful for benchmarking future MRS mission ground processing. Innovations developed during ground support for the Cassini mission include official declaration of sealed-source classifications, utilization of a mobile analytical laboratory, employment of a computerized dosimetry record management system, and cross-utilization of personnel from related disciplines.

Analyses of the Behavior of Spokes in Saturn's B Ring as Observed in Cassini ISS Images

Science.gov (United States)

Mitchell, Colin; Porco, C.; Dones, L.; Spitale, J.

2008-09-01

We report on analyses of the spokes in Saturn's B ring as observed by the Cassini spacecraft, from the first sighting in September 2005 to the present. Following Porco and Danielson (1982), we calculate as a function of time the spoke activity level, defined as the area-integrated optical depth of the spokes. We convert the spoke I/F into optical depth, using a radiative transfer "doubling code" and assuming that the presence of microscopic particles in the spokes is the only change in the optical properties of the ring region within a spoke. We search for periodicities in the variation of spoke activity and also correlations with magnetic longitude using a magnetic longitude system derived from the emission of the Saturn Kilometric Radiation (SKR), the rotation of which varies slightly from a constant rate (Kurth et al. 2008). Additionally, we track the activity over a period of years in order to characterize the seasonal nature of this phenomenon. We also report on the photometric profiles of spokes during different phases of their evolution. We present an analysis of spoke kinematics, measuring the motion on timescales of tens of minutes of the leading and trailing edges of spokes that appear in multiple consecutive images. Assuming that the small ice particles which comprise the spokes are in circular orbits, the azimuthal motion is a measure of their charge-to-mass ratio. While most spoke edges have exhibited normal Keplerian orbital motion and shear, some spokes were observed during their active phase in which the spoke's optical depth increases and its edges move at different rates, broadening the spoke. We acknowledge the financial support of the Cassini Project.

Solar Array Disturbances to Spacecraft Pointing During the Lunar Reconnaissance Orbiter (LRO) Mission

Science.gov (United States)

Calhoun, Philip

2010-01-01

The Lunar Reconnaissance Orbiter (LRO), the first spacecraft to support NASA s return to the Moon, launched on June 18, 2009 from the Cape Canaveral Air Force Station aboard an Atlas V launch vehicle. It was initially inserted into a direct trans-lunar trajectory to the Moon. After a five day transit to the Moon, LRO was inserted into the Lunar orbit and successfully lowered to a low altitude elliptical polar orbit for spacecraft commissioning. Successful commissioning was completed in October 2009 when LRO was placed in its near circular mission orbit with an approximate altitude of 50km. LRO will spend at least one year orbiting the Moon, collecting lunar environment science and mapping data, utilizing a suite of seven instruments to enable future human exploration. The objective is to provide key science data necessary to facilitate human return to the Moon as well as identification of opportunities for future science missions. LRO's instrument suite will provide the high resolution imaging data with sub-meter accuracy, highly accurate lunar cartographic maps, mineralogy mapping, amongst other science data of interest. LRO employs a 3-axis stabilized attitude control system (ACS) whose primary control mode, the "Observing Mode", provides Lunar nadir, off-nadir, and inertial fine pointing for the science data collection and instrument calibration. This controller combines the capability of fine pointing with on-demand large angle full-sky attitude reorientation. It provides simplicity of spacecraft operation as well as additional flexibility for science data collection. A conventional suite of ACS components is employed in the Observing Mode to meet the pointing and control objectives. Actuation is provided by a set of four reaction wheels developed in-house at NASA Goddard Space Flight Center (GSFC). Attitude feedback is provided by a six state Kalman filter which utilizes two SELEX Galileo Star Trackers for attitude updates, and a single Honeywell Miniature

Seasonal Evolution of the North and South Polar Vortex on Titan From 2004 to 2017 as Seen by Cassini/VIMS

Science.gov (United States)

Le Mouelic, S.; Robidel, R.; Rousseau, B.; Rodriguez, S.; Cornet, T.; Sotin, C.; Barnes, J. W.; Brown, R. H.; Buratti, B. J.; Baines, K. H.; Clark, R. N.; Nicholson, P. D.

2017-12-01

Cassini entered in Saturn's orbit in July 2004. In thirteen years, 127 targeted flybys of Titan have been performed. We focus our study on the analysis of the complete Visual and Infrared Mapping Spectrometer data set, with a particular emphasis on the evolving features on both poles. We have computed individual global maps of the north and south poles for each of the 127 targeted flybys, using VIMS wavelengths sensitive both to clouds and surface features. First evidences for a vast ethane cloud covering the North Pole is seen as soon as the first and second targeted flyby in October 2004 and December 2005 [1]. The first detailed imaging of this north polar feature with VIMS was obtained in December 2006, thanks to a change in inclination of the spacecraft orbit [2]. At this time, the northern lakes and seas of Titan were totally masked to the optical instruments by the haze and clouds, whereas the southern pole was well illuminated and mostly clear of haze and vast clouds. The vast north polar feature progressively vanished around the equinox in 2009 [2,3,4], in agreement with the predictions of Global Circulation Models [5]. It revealed progressively the underlying lakes to the ISS and VIMS instruments, which show up very nicely in VIMS in a series of flybys between T90 and T100. First evidences of an atmospheric vortex growing over the south pole occurred in May 2012 (T82), with a high altitude feature being detected consistently at each flyby up to the last T126 targeted flyby, and also appearing in more distant observations up to the end of the Cassini mission. Cassini has covered almost half a titanian year, corresponding to two seasons. The situation observed at the South Pole in the last images may correspond to what was observed in the north as Cassini just arrived. [1] Griffith et al., Science, 2006. [2] Le Mouélic et al., PSS, 2012. [3] Rodriguez et al., Nature, 2009. [4] Rodriguez et al., Icarus 2011. [4] Hirtzig et al., Icarus, 2013. [5] Rannou et al

Charged Particle In-Situ Measurements in the Inner Saturnian Magnetosphere during the "grand Finale" of Cassini in 2016/2017

Science.gov (United States)

Krupp, N.; Roussos, E.; Mitchell, D. G.; Kollmann, P.; Paranicas, C.; Krimigis, S. M.; Hedman, M. M.; Dougherty, M. K.

2017-12-01

After 13 years in orbit around Saturn Cassini came to an end on 15 September 2017. The last phase of the mission was called the "Grand Finale" and consisted of high latitude orbits crossing the F-Ring 22 times between Nov 2016 and April 2017 followed by the so called proximal orbits passing the ring plane inside the D-ring. The roughly 7-day long F-ring orbits with periapsis at nearly the same local time allowed to study temporal variations of the particle distributions in the inner part of Saturn's magnetosphere while during the proximal orbits Cassini measured for the first time the charged particle environment in-situ inside the D-ring up to 2500 km above the 1-bar cloud level of the planet. In this presentation first results of the Low Energy Magnetospheric Measurement System LEMMS, part of the Magnetosphere Imaging Instrument MIMI during the "Grand Finale" will be summarized in detail, including the discovery of MeV particles close to Saturn, higher intensities of charged particles when Cassini was magnetically connected to the D-Ring, sharp dropouts at the inner edge of the D-ring as well as unexpected features and asymmetries in the particle measurements related to newly discovered ring arcs in the inner magnetosphere.

Identifying Cassini's Magnetospheric Location Using Magnetospheric Imaging Instrument (MIMI) Data and Machine Learning

Science.gov (United States)

Vandegriff, J. D.; Smith, G. L.; Edenbaum, H.; Peachey, J. M.; Mitchell, D. G.

2017-12-01

We analyzed data from Cassini's Magnetospheric Imaging Instrument (MIMI) and Magnetometer (MAG) and attempted to identify the region of Saturn's magnetosphere that Cassini was in at a given time using machine learning. MIMI data are from the Charge-Energy-Mass Spectrometer (CHEMS) instrument and the Low-Energy Magnetospheric Measurement System (LEMMS). We trained on data where the region is known based on a previous analysis of Cassini Plasma Spectrometer (CAPS) plasma data. Three magnetospheric regions are considered: Magnetosphere, Magnetosheath, and Solar Wind. MIMI particle intensities, magnetic field values, and spacecraft position are used as input attributes, and the output is the CAPS-based region, which is available from 2004 to 2012. We then use the trained classifier to identify Cassini's magnetospheric regions for times after 2012, when CAPS data is no longer available. Training accuracy is evaluated by testing the classifier performance on a time range of known regions that the classifier has never seen. Preliminary results indicate a 68% accuracy on such test data. Other techniques are being tested that may increase this performance. We present the data and algorithms used, and will describe the latest results, including the magnetospheric regions post-2012 identified by the algorithm.

Estimating the Backup Reaction Wheel Orientation Using Reaction Wheel Spin Rates Flight Telemetry from a Spacecraft

Science.gov (United States)

Rizvi, Farheen

2013-01-01

A report describes a model that estimates the orientation of the backup reaction wheel using the reaction wheel spin rates telemetry from a spacecraft. Attitude control via the reaction wheel assembly (RWA) onboard a spacecraft uses three reaction wheels (one wheel per axis) and a backup to accommodate any wheel degradation throughout the course of the mission. The spacecraft dynamics prediction depends upon the correct knowledge of the reaction wheel orientations. Thus, it is vital to determine the actual orientation of the reaction wheels such that the correct spacecraft dynamics can be predicted. The conservation of angular momentum is used to estimate the orientation of the backup reaction wheel from the prime and backup reaction wheel spin rates data. The method is applied in estimating the orientation of the backup wheel onboard the Cassini spacecraft. The flight telemetry from the March 2011 prime and backup RWA swap activity on Cassini is used to obtain the best estimate for the backup reaction wheel orientation.

Cassini Scientist for a Day: a tactile experience

Science.gov (United States)

Canas, L.; Altobelli, N.

2012-09-01

In September 2011, the Cassini spacecraft took images of three targets and a challenge was launched to all students: to choose the one target they thought would provide the best science and to write an essay explaining their reasons (more information on the "Cassini Scientist for a Day" essay contest official webpage in: http://saturn.jpl.nasa.gov/education/scientistforaday10thedition/, run by NASA/JPL) The three targets presented were: Hyperion, Rhea and Titan, and Saturn. The idea behind "Cassini Scientist for a Day: a tactile experience" was to transform each of these images into schematic tactile images, highlighting relevant features apprehended through a tactile key, accompanied by a small text in Braille with some additional information. This initial approach would allow reach a broader community of students, more specifically those with visual impairment disabilities. Through proper implementation and careful study cases the adapted images associated with an explanatory key provide more resources in tactile astronomy. As the 2012 edition approaches a new set of targeted objet images will be once again transformed and adapted to visually impaired students and will aim to reach more students into participate in this international competition and to engage them in a quest to expand their knowledge in the amazing Cassini discoveries and the wonders of Saturn and its moons. As the winning essays will be published on the Cassini website and contest winners invited to participate in a dedicated teleconference with Cassini scientists from NASA's Jet Propulsion Laboratory, this initiative presents a great chance to all visually impaired students and teachers to participate in an exciting experience. These initiatives must be complemented with further information to strengthen the learning experience. However they stand as a good starting point to tackle further astronomical concepts in the classroom, especially this field that sometimes lacks the resources. Although

Estimating Torque Imparted on Spacecraft Using Telemetry

Science.gov (United States)

Lee, Allan Y.; Wang, Eric K.; Macala, Glenn A.

2013-01-01

There have been a number of missions with spacecraft flying by planetary moons with atmospheres; there will be future missions with similar flybys. When a spacecraft such as Cassini flies by a moon with an atmosphere, the spacecraft will experience an atmospheric torque. This torque could be used to determine the density of the atmosphere. This is because the relation between the atmospheric torque vector and the atmosphere density could be established analytically using the mass properties of the spacecraft, known drag coefficient of objects in free-molecular flow, and the spacecraft velocity relative to the moon. The density estimated in this way could be used to check results measured by science instruments. Since the proposed methodology could estimate disturbance torque as small as 0.02 N-m, it could also be used to estimate disturbance torque imparted on the spacecraft during high-altitude flybys.

Nuclear electric propulsion /NEP/ spacecraft for the outer planet orbiter mission

International Nuclear Information System (INIS)

Garrison, P.W.; Nock, K.T.

1982-01-01

The design, operating features, and a possible Neptune orbit for the spacecraft powered by the SP-100 nuclear electric propulsion (NEP) system under study by NASA and the DOE are described. The system features a reactor and a payload situated on opposite ends of a 0.5 m diam, 11 m long astromast. Mercury-ion thrusters are located beneath the reactor for side thrusting, and no contamination of the payload or obstruction of the viewing angles for scientific objectives occurs with the system, which would not degrade in performance even under high insolation during near-sun maneuvers. Results of a theoretical study of earth escapes are presented to show that an NEP powered spiral trajectory out of a 700 km Shuttle orbit and using a Triton gravity assist would be superior to departing from a 300 km orbit with a Centaur boost. The mission profile includes a 1249 kg Galileo payload. The SP-100 has a 1.4 MWth reactor with UO2 fuel tiles and weighs 19,904 kg

Contingency Trajectory Design for a Lunar Orbit Insertion Maneuver Failure by the Lunar Atmosphere Dust Environment Explorer (LADEE) Spacecraft

Science.gov (United States)

Genova, Anthony L.; Loucks, Michael; Carrico, John

2014-01-01

The purpose of this extended abstract is to present results from a failed lunar-orbit insertion (LOI) maneuver contingency analysis for the Lunar Atmosphere Dust Environment Explorer (LADEE) mission, managed and operated by NASA Ames Research Center in Moffett Field, CA. The LADEE spacecrafts nominal trajectory implemented multiple sub-lunar phasing orbits centered at Earth before eventually reaching the Moon (Fig. 1) where a critical LOI maneuver was to be performed [1,2,3]. If this LOI was missed, the LADEE spacecraft would be on an Earth-escape trajectory, bound for heliocentric space. Although a partial mission recovery is possible from a heliocentric orbit (to be discussed in the full paper), it was found that an escape-prevention maneuver could be performed several days after a hypothetical LOI-miss, allowing a return to the desired science orbit around the Moon without leaving the Earths sphere-of-influence (SOI).

Statistics of Langmuir wave amplitudes observed inside Saturn's foreshock by the Cassini spacecraft

Czech Academy of Sciences Publication Activity Database

Píša, David; Hospodarsky, G. B.; Kurth, W. S.; Santolík, Ondřej; Souček, Jan; Gurnett, D. A.; Masters, A.; Hill, M. E.

2015-01-01

Roč. 120, č. 4 (2015), s. 2531-2542 ISSN 2169-9380 R&D Projects: GA ČR GAP205/10/2279; GA ČR(CZ) GAP209/12/2394 Institutional support: RVO:68378289 Keywords : Langmuir waves * foreshock * Saturn * Cassini Subject RIV: BL - Plasma and Gas Discharge Physics Impact factor: 3.318, year: 2015 http://onlinelibrary.wiley.com/doi/10.1002/2014JA020560/abstract

Tidal Control of Jet Eruptions Observed by Cassini ISS

Science.gov (United States)

Hurford, T. A.; Helfenstein, P.; Spitale, J. N.

2012-01-01

Observations by Cassini's Imaging Science Subsystem (ISS) of Enceladus' south polar region at high phase angles has revealed jets of material venting into space. Observations by Cassini's Composite Infrared Spectrometer (CIRS) have also shown that the south polar region is anomalously warm with hotspots associated with geological features called the Tiger Stripes. The Tiger Stripes are large rifts near the south pole of Enceladus, which are typically about 130 km in length, 2 km wide, with a trough 500 m deep, and are l1anked on each side by 100m tall ridges. Preliminary triangulation of jets as viewed at different times and with different viewing geometries in Cassini ISS images taken between 2005 and 2007 have constrained the locations of eight major eruptions of material and found all of them associated with the south polar fractures unofficially the 'Tiger Stripes', and found four of them coincident with the hotspots reported in 2006 by CIRS. While published ISS observations of jet activity suggest that individual eruption sites stay active on the timescale of years, any shorter temporal variability (on timescales of an orbital period, or 1.3 Earth days, for example) is more difficult to establish because of the spotty temporal coverage and the difficulty of visually isolating one jet from the forest of many seen in a typical image. Consequently, it is not known whether individual jets are continuously active, randomly active, or if they erupt on a predictable, periodic schedule. One mechanism that may control the timing of eruptions is diurnal tidal stress, which oscillates between compression/tension as well as right and left lateral shear at any given location throughout Enceladus' orbit and may allow the cracks to open and close regularly. We examine the stresses on the Tiger Stripe regions to see how well diurnal tidal stress caused by Enceladus' orbital eccentricity may possibly correlate with and thus control the observed eruptions. We then identify

Fast forward modeling of Titan’s infrared spectra to invert VIMS/CASSINI hyperspectral images

Science.gov (United States)

Rodriguez, S.; Le Mouélic, S.; Rannou, P.; Combe, J.; Le Corre, L.; Griffith, C. A.; Tobie, G.; Barnes, J. W.; Sotin, C.; Brown, R. H.; Baines, K. H.; Buratti, B. J.; Clark, R. N.

2009-12-01

The surface of Titan, the largest icy moon of Saturn, is veiled by a very thick and hazy atmosphere. The Visual and Infrared Mapping Spectrometer onboard the Cassini spacecraft, in orbit around Saturn since July 2004, has been conducting an intensive survey of Titan with the objective of understanding the complex nature and interaction of the atmosphere and surface of this mysterious moon. Retrieving and separating contributions from the surface and the atmosphere in Titan’s infrared spectra requires accurate radiative transfer modeling, which is often very demanding of computer resources. As Cassini has gathered hitherto millions of spectra of Titan and will continue to observe it until at least 2010, we report here on the development of a new rapid, simple and versatile radiative transfer model specially designed to process VIMS datacubes. Currently, our model accounts for gas absorption, haze scattering and surface reflectance and can be implemented in an inversion scheme. First results of forward modeling provide spectral shapes that are consistent with VIMS measurements, as well as surface and aerosol properties in the range of validity for Titan. Further inversion tests will be carried on VIMS hyperspectral images for the estimate of spatial coherence of the results, accuracy of the surface reflectance within the atmospheric windows, and potential needs for improved input data and modeling. This work was partly performed at the Jet Propulsion Laboratory, California Institute of Technology, under contract to the National Aeronautics and Space Administration. Calibrated VIMS data appear courtesy of the VIMS team. We thank the CNES French agency for its financial support.

Long-term orbit prediction for Tiangong-1 spacecraft using the mean atmosphere model

Science.gov (United States)

Tang, Jingshi; Liu, Lin; Cheng, Haowen; Hu, Songjie; Duan, Jianfeng

2015-03-01

China is planning to complete its first space station by 2020. For the long-term management and maintenance, the orbit of the space station needs to be predicted for a long period of time. Since the space station is expected to work in a low-Earth orbit, the error in the a priori atmosphere model contributes significantly to the rapid increase of the predicted orbit error. When the orbit is predicted for 20 days, the error in the a priori atmosphere model, if not properly corrected, could induce a semi-major axis error of up to a few kilometers and an overall position error of several thousand kilometers respectively. In this work, we use a mean atmosphere model averaged from NRLMSISE00. The a priori reference mean density can be corrected during the orbit determination. For the long-term orbit prediction, we use sufficiently long period of observations and obtain a series of the diurnal mean densities. This series contains the recent variation of the atmosphere density and can be analyzed for various periodic components. After being properly fitted, the mean density can be predicted and then applied in the orbit prediction. Here we carry out the test with China's Tiangong-1 spacecraft at the altitude of about 340 km and we show that this method is simple and flexible. The densities predicted with this approach can serve in the long-term orbit prediction. In several 20-day prediction tests, most predicted orbits show semi-major axis errors better than 700 m and overall position errors better than 400 km.

On-orbit assembly of a team of flexible spacecraft using potential field based method

Science.gov (United States)

Chen, Ti; Wen, Hao; Hu, Haiyan; Jin, Dongping

2017-04-01

In this paper, a novel control strategy is developed based on artificial potential field for the on-orbit autonomous assembly of four flexible spacecraft without inter-member collision. Each flexible spacecraft is simplified as a hub-beam model with truncated beam modes in the floating frame of reference and the communication graph among the four spacecraft is assumed to be a ring topology. The four spacecraft are driven to a pre-assembly configuration first and then to the assembly configuration. In order to design the artificial potential field for the first step, each spacecraft is outlined by an ellipse and a virtual leader of circle is introduced. The potential field mainly depends on the attitude error between the flexible spacecraft and its neighbor, the radial Euclidian distance between the ellipse and the circle and the classical Euclidian distance between the centers of the ellipse and the circle. It can be demonstrated that there are no local minima for the potential function and the global minimum is zero. If the function is equal to zero, the solution is not a certain state, but a set. All the states in the set are corresponding to the desired configurations. The Lyapunov analysis guarantees that the four spacecraft asymptotically converge to the target configuration. Moreover, the other potential field is also included to avoid the inter-member collision. In the control design of the second step, only small modification is made for the controller in the first step. Finally, the successful application of the proposed control law to the assembly mission is verified by two case studies.

Cassini-Huygens maneuver automation for navigation

Science.gov (United States)

Goodson, Troy; Attiyah, Amy; Buffington, Brent; Hahn, Yungsun; Pojman, Joan; Stavert, Bob; Strange, Nathan; Stumpf, Paul; Wagner, Sean; Wolff, Peter;

Cassini's Test Methodology for Flight Software Verification and Operations

Science.gov (United States)

Wang, Eric; Brown, Jay

2007-01-01

The Cassini spacecraft was launched on 15 October 1997 on a Titan IV-B launch vehicle. The spacecraft is comprised of various subsystems, including the Attitude and Articulation Control Subsystem (AACS). The AACS Flight Software (FSW) and its development has been an ongoing effort, from the design, development and finally operations. As planned, major modifications to certain FSW functions were designed, tested, verified and uploaded during the cruise phase of the mission. Each flight software upload involved extensive verification testing. A standardized FSW testing methodology was used to verify the integrity of the flight software. This paper summarizes the flight software testing methodology used for verifying FSW from pre-launch through the prime mission, with an emphasis on flight experience testing during the first 2.5 years of the prime mission (July 2004 through January 2007).

Simultaneous spacecraft orbit estimation and control based on GPS measurements via extended Kalman filter

Directory of Open Access Journals (Sweden)

Tamer Mekky Ahmed Habib

2013-06-01

Full Text Available The primary aim of this work is to provide simultaneous spacecraft orbit estimation and control based on the global positioning system (GPS measurements suitable for application to the next coming Egyptian remote sensing satellites. Disturbance resulting from earth’s oblateness till the fourth order (i.e., J4 is considered. In addition, aerodynamic drag and random disturbance effects are taken into consideration.

Jovian atmospheric dynamics: an update after Galileo and Cassini

International Nuclear Information System (INIS)

Vasavada, Ashwin R; Showman, Adam P

2005-01-01

The Galileo and Cassini spacecrafts have greatly enhanced the observational record of Jupiter's tropospheric dynamics, particularly through returning high spatial resolution, multi-spectral and global imaging data with episodic coverage over periods of months to years. These data, along with those from Earth-based telescopes, have revealed the stability of Jupiter's zonal jets, captured the evolution of vortices and equatorial waves, and mapped the distributions of lightning and moist convection. Because no observations of Jupiter's interior exist, a forward modelling approach has been used to relate observations at cloud level to models of shallow or deep jet structure, shallow or deep jet forcing and energy transfer between turbulence, vortices and jets. A range of observed phenomena can be reproduced in shallow models, though the Galileo probe winds and jet stability arguments hint at the presence of deep jets. Many deep models, however, fail to reproduce Jupiter-like non-zonal features (e.g. vortices). Jupiter's dynamics likely include both deep and shallow processes, requiring an integrated approach to future modelling-an important goal for the post-Galileo and Cassini era

Cassini RTG program. Monthly technical progress report, September 29, 1997--October 26, 1997

Energy Technology Data Exchange (ETDEWEB)

NONE

1997-11-24

This report describes work on the contract to provide Radioisotope Thermoelectric Generators (RTG) and Ancillary Activities in support of the Cassini Spacecraft launch. The craft was successfully launched on October 15, 1997. Early telemetry results show excellent performance from the three launched RTG modules. A major share of this report describes safety analyses for contamination radii in the event of launch failures and generator destruction, as well as launch related activities.

Advanced Solar-propelled Cargo Spacecraft for Mars Missions

Science.gov (United States)

Auziasdeturenne, Jacqueline; Beall, Mark; Burianek, Joseph; Cinniger, Anna; Dunmire, Barbrina; Haberman, Eric; Iwamoto, James; Johnson, Stephen; Mccracken, Shawn; Miller, Melanie

1989-01-01

Three concepts for an unmanned, solar powered, cargo spacecraft for Mars support missions were investigated. These spacecraft are designed to carry a 50,000 kg payload from a low Earth orbit to a low Mars orbit. Each design uses a distinctly different propulsion system: A Solar Radiation Absorption (SRA) system, a Solar-Pumped Laser (SPL) system and a solar powered magnetoplasmadynamic (MPD) arc system. The SRA directly converts solar energy to thermal energy in the propellant through a novel process. In the SPL system, a pair of solar-pumped, multi-megawatt, CO2 lasers in sunsynchronous Earth orbit converts solar energy to laser energy. The MPD system used indium phosphide solar cells to convert sunlight to electricity, which powers the propulsion system. Various orbital transfer options are examined for these concepts. In the SRA system, the mother ship transfers the payload into a very high Earth orbit and a small auxiliary propulsion system boosts the payload into a Hohmann transfer to Mars. The SPL spacecraft and the SPL powered spacecraft return to Earth for subsequent missions. The MPD propelled spacecraft, however, remains at Mars as an orbiting space station. A patched conic approximation was used to determine a heliocentric interplanetary transfer orbit for the MPD propelled spacecraft. All three solar-powered spacecraft use an aerobrake procedure to place the payload into a low Mars parking orbit. The payload delivery times range from 160 days to 873 days (2.39 years).

A Brief History of Meteoroid and Orbital Debris Shielding Technology for US Manned Spacecraft

Science.gov (United States)

Bjorkman, Michael D.; Hyde, James L.

2008-01-01

Meteoroid and orbital debris shielding has played an important role from the beginning of manned spaceflight. During the early 60 s, meteoroid protection drove requirements for new meteor and micrometeoroid impact science. Meteoroid protection also stimulated advances in the technology of hypervelocity impact launchers and impact damage assessment methodologies. The first phase of meteoroid shielding assessments closed in the early 70 s with the end of the Apollo program. The second phase of meteoroid protection technology began in the early 80 s when it was determined that there is a manmade Earth orbital debris belt that poses a significant risk to LEO manned spacecraft. The severity of the Earth orbital debris environment has dictated changes in Space Shuttle and ISS operations as well as driven advances in shielding technology and assessment methodologies. A timeline of shielding technology and assessment methodology advances is presented along with a summary of risk assessment results.

Application of Space Environmental Observations to Spacecraft Pre-Launch Engineering and Spacecraft Operations

Science.gov (United States)

Barth, Janet L.; Xapsos, Michael

2008-01-01

This presentation focuses on the effects of the space environment on spacecraft systems and applying this knowledge to spacecraft pre-launch engineering and operations. Particle radiation, neutral gas particles, ultraviolet and x-rays, as well as micrometeoroids and orbital debris in the space environment have various effects on spacecraft systems, including degradation of microelectronic and optical components, physical damage, orbital decay, biasing of instrument readings, and system shutdowns. Space climate and weather must be considered during the mission life cycle (mission concept, mission planning, systems design, and launch and operations) to minimize and manage risk to both the spacecraft and its systems. A space environment model for use in the mission life cycle is presented.

Deployable Brake for Spacecraft

Science.gov (United States)

Rausch, J. R.; Maloney, J. W.

1987-01-01

Aerodynamic shield that could be opened and closed proposed. Report presents concepts for deployable aerodynamic brake. Brake used by spacecraft returning from high orbit to low orbit around Earth. Spacecraft makes grazing passes through atmosphere to slow down by drag of brake. Brake flexible shield made of woven metal or ceramic withstanding high temperatures created by air friction. Stored until needed, then deployed by set of struts.

Long-term orbit prediction for China's Tiangong-1 spacecraft based on mean atmosphere model

Science.gov (United States)

Tang, Jingshi; Liu, Lin; Miao, Manqian

Tiangong-1 is China's test module for future space station. It has gone through three successful rendezvous and dockings with Shenzhou spacecrafts from 2011 to 2013. For the long-term management and maintenance, the orbit sometimes needs to be predicted for a long period of time. As Tiangong-1 works in a low-Earth orbit with an altitude of about 300-400 km, the error in the a priori atmosphere model contributes significantly to the rapid increase of the predicted orbit error. When the orbit is predicted for 10-20 days, the error in the a priori atmosphere model, if not properly corrected, could induce the semi-major axis error and the overall position error up to a few kilometers and several thousand kilometers respectively. In this work, we use a mean atmosphere model averaged from NRLMSIS00. The a priori reference mean density can be corrected during precise orbit determination (POD). For applications in the long-term orbit prediction, the observations are first accumulated. With sufficiently long period of observations, we are able to obtain a series of the diurnal mean densities. This series bears the recent variation of the atmosphere density and can be analyzed for various periods. After being properly fitted, the mean density can be predicted and then applied in the orbit prediction. We show that the densities predicted with this approach can serve to increase the accuracy of the predicted orbit. In several 20-day prediction tests, most predicted orbits show semi-major axis errors better than 700m and overall position errors better than 600km.

NANODUST DETECTION BETWEEN 1 AND 5 AU USING CASSINI WAVE MEASUREMENTS

Energy Technology Data Exchange (ETDEWEB)

Schippers, P.; Vernet, N. Meyer-; Lecacheux, A.; Belheouane, S.; Moncuquet, M. [LESIA—CNRS—Observatoire de Paris, 5 place Jules Janssen, F-92195 Meudon (France); Kurth, W. S. [Department of Physics and Astronomy, University of Iowa, Iowa City, IA (United States); Mann, I. [EISCAT Scientific Association, Kiruna, Sweden and Department of Physics Umeå University (Sweden); Mitchell, D. G. [Applied Physics Laboratory, John Hopkins University, Laurel, MD (United States); André, N. [IRAP, 9 Avenue du Colonel Roche, F-31028 Toulouse (France)

2015-06-10

The solar system contains solids of all sizes, ranging from kilometer-sized bodies to nano-sized particles. Nanograins have been detected in situ in the Earth's atmosphere, near cometary and giant planet environments, and more recently in the solar wind at 1 AU. The latter nanograins are thought to be formed in the inner solar system dust cloud, mainly through the collisional break-up of larger grains, and are then picked up and accelerated by the magnetized solar wind because of their large charge-to-mass ratio. In the present paper, we analyze the low frequency bursty noise identified in the Cassini radio and plasma wave data during the spacecraft cruise phase inside Jupiter's orbit. The magnitude, spectral shape, and waveform of this broadband noise are consistent with the signatures of the nano particles that traveled at solar wind speed and impinged on the spacecraft surface. Nanoparticles were observed whenever the radio instrument was turned on and able to detect them at different heliocentric distances between Earth and Jupiter, suggesting their ubiquitous presence in the heliosphere. We analyzed the radial dependence of the nanodust flux with heliospheric distance and found that it is consistent with the dynamics of nanodust originating from the inner heliosphere and picked up by the solar wind. The contribution of the nanodust produced in the asteroid belt appears to be negligible compared to the trapping region in the inner heliosphere. In contrast, further out, nanodust is mainly produced by the volcanism of active moons such as Io and Enceladus.

Statistical Analysis of Interchange Injection Events from Over a Decade of Cassini Data

Science.gov (United States)

Azari, A.; Jia, X.; Liemohn, M. W.; Sergis, N.; Thomsen, M. F.; Mitchell, D. G.; Rymer, A. M.; Paranicas, C.; Provan, G.; Ye, S.; Cowley, S. W. H.; Hospodarsky, G. B.; Vandegriff, J. D.; Kurth, W. S.

2017-12-01

The Cassini spacecraft has routinely observed interchange injection events with multiple instruments since arriving at Saturn in 2004. Interchange injection events are thought to initiate from a Rayleigh-Taylor like plasma instability sourced from Saturn's rapid rotation (period 10.8 hours) and dense plasma population outgassing primarily from Enceladus, and are the primary source of mass transport in the inner/middle magnetosphere. This dense plasma must be transported outward, and to conserve magnetic flux, inward moving flux tubes of low density, energetic (> keV) plasma from the outer reaches of the Saturnian system also occur. These inward-bound flux tubes are referred to as interchange injections. We will present a statistical evaluation of the occurrence rates of interchange injections at Saturn demonstrating seasonal dependence of interchange over the entirety of the Cassini mission's equatorial orbits between 2005 and 2016. We identify interchange events from CHarge Energy Mass Spectrometer (CHEMS) H+ data using a trained and tested automated algorithm. Our event identification compares well with manual identification and previous surveys of injections by L-shell and local time (Chen and Hill, 2008, Lai et al., 2016, Kennelly et al., 2013). We find that peak rates of interchange events occur between 7 - 9 Saturn radii, in agreement with previous surveys. We also evaluate interchange by preferred local time sector and season, splitting our events into pre-equinox, equinox, and post - equinox time periods. We determine that over all seasons, nightside occurrence dominated as compared to dayside, but the preferred dayside sector shifts from pre-noon during equinox, to post-noon during post-equinox. We will further investigate seasonal dependence by presenting occurrence organized by the phase systems derived based on Saturn kilometric radiation (SKR) and magnetic field perturbations (PPO).

10 Years at Saturn, and More Excitement to Come!

Science.gov (United States)

Edgington, S. G.; Spilker, L. J.; Altobelli, N.

2014-04-01

After 10 years in orbit, the Cassini-Huygens Mission to Saturn, a collaboration of NASA, ESA, and ASI, continues to wow the imagination. Every year Cassini produces answers to questions raised by the Voyager flybys, while at the same time posing new questions that can only be answered with a long duration mission using a flagship-class spacecraft. In this talk, we sample a few of Cassini's discoveries from the past decade and give an overview of what comes next.

Absolute orbit determination using line-of-sight vector measurements between formation flying spacecraft

Science.gov (United States)

Ou, Yangwei; Zhang, Hongbo; Li, Bin

2018-04-01

The purpose of this paper is to show that absolute orbit determination can be achieved based on spacecraft formation. The relative position vectors expressed in the inertial frame are used as measurements. In this scheme, the optical camera is applied to measure the relative line-of-sight (LOS) angles, i.e., the azimuth and elevation. The LIDAR (Light radio Detecting And Ranging) or radar is used to measure the range and we assume that high-accuracy inertial attitude is available. When more deputies are included in the formation, the formation configuration is optimized from the perspective of the Fisher information theory. Considering the limitation on the field of view (FOV) of cameras, the visibility of spacecraft and the installation of cameras are investigated. In simulations, an extended Kalman filter (EKF) is used to estimate the position and velocity. The results show that the navigation accuracy can be enhanced by using more deputies and the installation of cameras significantly affects the navigation performance.

Production of activation products in space-craft components by protons in low earth orbit

International Nuclear Information System (INIS)

Normand, E.; Johnson, M.L.

1986-01-01

A spacecraft orbiting the Earth through trapped radiation belts will be subject to an induced effect as well as to the direct irradiation by the protons and electrons of the trapped belts. This induced effect is activation of the spacecraft materials by the trapped belt protons. This activation will produce many radioisotopes having half-lives ranging from seconds to millions of years, and emitting various types of radiation. Of primary concern are radioisotopes that emit gamma rays and have half-lives of several years or less. Cross-section data sets are currently being compiled for proton-induced activation products by the Los Alamos National Laboratory. Despite uncertainties in cross-section data, it is instructive to illustrate the magnitude of activation levels and the resulting dose rates calculated in an approximate manner. A number of simplifying assumptions are made

Exact Cavity Perturbation Technique to Determine Complex Permittivity of Dielectric Materials

Science.gov (United States)

Starr, Charles H.; Barmatz, Martin B.

2011-01-01

Cassini is an international spacecraft mission facilitated by NASA and ESA which seeks to understand the Saturn planetary system, including rings and moons. Launched in 1997, the Cassini spacecraft contains two major components: the Cassini orbiter that has been orbiting Saturn since October 2004, and the European-built Huygens probe that landed on Titan's surface in December 2004 to study its geology and atmosphere. Titan, Saturn's largest moon and the second largest moon in the solar system, possesses surface and atmospheric features similar to those of Earth, including lakes, seas, and mountains. A physical characterization of these features is critical to understanding the origin and evolution of Titan, whose surface composition reflects its geological history. Because Titan's atmosphere is largely composed of methane, it is believed that surfaces lakes are filled with mixtures of liquid hydrocarbons. The Cassini orbiter's RADAR instrument has been scanning Titan's surface at the atmosphere-penetrating microwave frequency of 13.8 Gigahertz since 2004. However, accurate interpretation of these data is limited by a lack of knowledge regarding dielectric properties of liquid hydrocarbons at cryogenic temperatures. Therefore, it is of specific interest to experimentally determine values for the complex permittivities of various liquid hydrocarbon mixtures at the surface conditions of Titan. In particular, more accurate values for complex permittivity would improve estimates of lake depth and surface composition obtained from the instrument's altimetry and backscatter modes.

Equinoctial Activity Over Titan Dune Fields Revealed by Cassini/vims

Science.gov (United States)

Rodriguez, S.; Le Mouelic, S.; Barnes, J. W.; Hirtzig, M.; Rannou, P.; Sotin, C.; Brown, R. H.; Bow, J.; Vixie, G.; Cornet, T.; Bourgeois, O.; Narteau, C.; Courrech Du Pont, S.; Le Gall, A.; Reffet, E.; Griffith, C. A.; Jaumann, R.; Stephan, K.; Buratti, B. J.; Clark, R. N.; Baines, K. H.; Nicholson, P. D.; Coustenis, A.

2012-12-01

Titan, the largest satellite of Saturn, is the only satellite in the solar system with a dense atmosphere. The close and continuous observations of Titan by the Cassini spacecraft, in orbit around Saturn since July 2004, bring us evidences that Titan troposphere and low stratosphere experience an exotic, but complete meteorological cycle similar to the Earth hydrological cycle, with hydrocarbons evaporation, condensation in clouds, and rainfall. Cassini monitoring campaigns also demonstrate that Titan's cloud coverage and climate vary with latitude. Titan's tropics, with globally weak meteorological activity and widespread dune fields, seem to be slightly more arid than the poles, where extensive and numerous liquid reservoirs and sustained cloud activity have been discovered. Only a few tropo-spheric clouds have been observed at Titan's tropics during the southern summer. As equinox was approaching (in August 2009), they occurred more frequently and appeared to grow in strength and size. We present here the observation of intense brightening at Titan's tropics, very close to the equinox. These detections were conducted with the Visual and Infrared Mapping Spectrometer (VIMS) onboard Cassini. We will discuss the VIMS images of the three individual events detected so far, observed during the Titan's flybys T56 (22 May 2009), T65 (13 January 2010) and T70 (21 June 2010). T56, T65 and T70 observations show an intense and transient brighten-ing of large regions very close to the equator, right over the extensive dune fields of Senkyo, Belet and Shangri-La. They all appear spectrally and morphologically different from all transient surface features or atmospheric phenomena previously reported. Indeed, these events share in particular a strong brightening at wavelengths greater than 2 μm (especially at 5 μm), making them spectrally distinct from the small tropical clouds observed before the equinox and the large storms observed near the equator in September and October

Iapetus: First data from the Cassini Visual Infrared Mapping Spectrometer

Science.gov (United States)

Buratti, B. J.; Cruikshank, D. P.; Clark, R.; Brown, R. H.; Bauer, J. M.; Simonelli, D. P.; Jaumann, R.; Hibbitts, K.; McCord, T. B.; Soderlund, K.; Baines, K. H.; Bellucci, G.; Bibring, J. P.; Capaccioni, F.; Cerroni, P.; Coradini, A.; Drossart, P.; Formisano, V.; Langevin, Y.; Matson, D. L.; Mennella, V.; Nelson, R.; Nicholson, P. D.; Sicardy, B.; Sotin, C.

2004-11-01

Iapetus is perhaps the most enigmatic body in the solar system: One hemisphere is as dark as lampblack, and the other is almost as bright as snow. The models that have been offered to explain this dichotomy range from endogenously placed material (Smith et al., 1982, Science 215, 504), to material exogenously placed from Phoebe (Soter, 1974, IAU Colloq. 28), or other bodies (Owen et al., 2001, Icarus 149, 160; Buratti et al., 2002, Icarus 155, 375; Buratti et al., 2003, B.A.A.S, 915). No mechanism for the darkening process or purported source for the exogenic particles is entirely satisfactory. One key question is whether the process that led to the formation of the low-albedo hemisphere of Iapetus is unique, or whether the satellite has been subjected to a satellite alteration process in a more extreme form. Both Callisto and the outer satellites of Uranus show evidence for exogenic accretion of particles onto their leading sides. A targeted flyby of Iapetus by Cassini, during which the spacecraft will approach the satellite to within 1000 km, is scheduled to occur in September 2007. An untargeted approach of 65,000 km to the satellite will occur on New Year's day 2005, and observations are planned for the period around closest approach. However, a "sneak peak" of the satellite was afforded by Cassini on July 19, 2004, during which the spacecraft approached to less than three million miles (the Voyager closest approach was 909,070 km). The first disk resolved spectra of Iapetus in the 0.4 to 5 micron region were obtained by the Cassini Visual Infrared Mapping Spectrometer (VIMS). We report the tentative identification of carbon dioxide on the low-albedo portion of the surface. A comparison of the spectrum of Iapetus to that obtained by VIMS during its flyby of Phoebe on June 11, 2004 will be made. Mixing models incorporating water ice, minerals, and organics can replicate the spectrum of the dark hemisphere. Work performed at the Jet Propulsion Laboratory

Spacecraft Charging and the Microwave Anisotropy Probe Spacecraft

Science.gov (United States)

Timothy, VanSant J.; Neergaard, Linda F.

1998-01-01

The Microwave Anisotropy Probe (MAP), a MIDEX mission built in partnership between Princeton University and the NASA Goddard Space Flight Center (GSFC), will study the cosmic microwave background. It will be inserted into a highly elliptical earth orbit for several weeks and then use a lunar gravity assist to orbit around the second Lagrangian point (L2), 1.5 million kilometers, anti-sunward from the earth. The charging environment for the phasing loops and at L2 was evaluated. There is a limited set of data for L2; the GEOTAIL spacecraft measured relatively low spacecraft potentials (approx. 50 V maximum) near L2. The main area of concern for charging on the MAP spacecraft is the well-established threat posed by the "geosynchronous region" between 6-10 Re. The launch in the autumn of 2000 will coincide with the falling of the solar maximum, a period when the likelihood of a substorm is higher than usual. The likelihood of a substorm at that time has been roughly estimated to be on the order of 20% for a typical MAP mission profile. Because of the possibility of spacecraft charging, a requirement for conductive spacecraft surfaces was established early in the program. Subsequent NASCAP/GEO analyses for the MAP spacecraft demonstrated that a significant portion of the sunlit surface (solar cell cover glass and sunshade) could have nonconductive surfaces without significantly raising differential charging. The need for conductive materials on surfaces continually in eclipse has also been reinforced by NASCAP analyses.

Spacecraft 3D Augmented Reality Mobile App

Science.gov (United States)

Hussey, Kevin J.; Doronila, Paul R.; Kumanchik, Brian E.; Chan, Evan G.; Ellison, Douglas J.; Boeck, Andrea; Moore, Justin M.

2013-01-01

The Spacecraft 3D application allows users to learn about and interact with iconic NASA missions in a new and immersive way using common mobile devices. Using Augmented Reality (AR) techniques to project 3D renditions of the mission spacecraft into real-world surroundings, users can interact with and learn about Curiosity, GRAIL, Cassini, and Voyager. Additional updates on future missions, animations, and information will be ongoing. Using a printed AR Target and camera on a mobile device, users can get up close with these robotic explorers, see how some move, and learn about these engineering feats, which are used to expand knowledge and understanding about space. The software receives input from the mobile device's camera to recognize the presence of an AR marker in the camera's field of view. It then displays a 3D rendition of the selected spacecraft in the user's physical surroundings, on the mobile device's screen, while it tracks the device's movement in relation to the physical position of the spacecraft's 3D image on the AR marker.

Spacecraft Charging: Hazard Causes, Hazard Effects, Hazard Controls

Science.gov (United States)

Koontz, Steve.

2018-01-01

Spacecraft flight environments are characterized both by a wide range of space plasma conditions and by ionizing radiation (IR), solar ultraviolet and X-rays, magnetic fields, micrometeoroids, orbital debris, and other environmental factors, all of which can affect spacecraft performance. Dr. Steven Koontz's lecture will provide a solid foundation in the basic engineering physics of spacecraft charging and charging effects that can be applied to solving practical spacecraft and spacesuit engineering design, verification, and operations problems, with an emphasis on spacecraft operations in low-Earth orbit, Earth's magnetosphere, and cis-Lunar space.

Cassini Mission Sequence Subsystem (MSS)

Science.gov (United States)

Alland, Robert

2011-01-01

This paper describes my work with the Cassini Mission Sequence Subsystem (MSS) team during the summer of 2011. It gives some background on the motivation for this project and describes the expected benefit to the Cassini program. It then introduces the two tasks that I worked on - an automatic system auditing tool and a series of corrections to the Cassini Sequence Generator (SEQ_GEN) - and the specific objectives these tasks were to accomplish. Next, it details the approach I took to meet these objectives and the results of this approach, followed by a discussion of how the outcome of the project compares with my initial expectations. The paper concludes with a summary of my experience working on this project, lists what the next steps are, and acknowledges the help of my Cassini colleagues.

Analysis of Hot Ions Detected during Equatorial Orbits of the Cassini Spacecraft at Saturn using the Convected Kappa Distribution Function and a Comparison to Voyager and Galileo Measurements at Jupiter

Science.gov (United States)

Kane, M.; Mitchell, D. G.; Carbary, J. F.; Hill, M. E.; Dialynas, K.; Mauk, B.; Krimigis, S. M.

2017-12-01

An extensive analysis of Cassini INCA and CHEMS measurements of 5-149 keV ions acquired during all equatorial orbits has been completed using a 3-D convected kappa distribution model. The computed plasma azimuthal speed, expressed as a fraction of the local corotation speed, decreases sharply with increasing distance from Saturn. The oxygen ion profile follows the hydrogen ion trend. For both species, the polar convection speed is the smallest of the 3 velocity components, and is centered about zero, but the radial speed has a significant radially outward component. Further, the radial component is enhanced in the pre-dawn sector. The hydrogen and oxygen temperatures increase with decreasing distance to Saturn. The calculated pattern of convection is consistent with an empirical model of plasma convection that includes outward radial transport and escape of plasma in a dawnside boundary layer of plasma entrained by the dawn magnetosheath flow. When the model convection pattern is scaled to the sub-solar magnetopause distance and to the sizes of Jupiter and Saturn, the pattern agrees with that derived from analysis of hot ions detected by the LECP detector on Voyager and the EPD instrument on Galileo. This and previous analysis of hot ion distributions has shown that the convected kappa distribution, with isotropy assumed in the plasma rest frame, has well described hot ion observed fluxes within a limited range of ion energies and has produced meaningful and ordered physical plasma parameters including plasma bulk velocity vectors, kappa distribution temperature profiles, and the general magnetospheric convection pattern at Jupiter and Saturn.

The Saturnian satellite Rhea as seen by Cassini VIMS

Science.gov (United States)

Stephan, K.; Jaumann, R.; Wagner, R.; Clark, R.N.; Cruikshank, D.P.; Giese, B.; Hibbitts, C.A.; Roatsch, T.; Matz, K.-D.; Brown, R.H.; Filacchione, G.; Cappacioni, F.; Scholten, F.; Buratti, B.J.; Hansen, G.B.; Nicholson, P.D.; Baines, K.H.; Nelson, R.M.; Matson, D.L.

2012-01-01

Since the arrival of the Cassini spacecraft at Saturn in June 2004, the Visual and Infrared Mapping Spectrometer has obtained new spectral data of the icy satellites of Saturn in the spectral range from 0.35 to 5.2 ??m. Numerous flybys were performed at Saturn's second largest satellite Rhea, providing a nearly complete coverage with pixel-ground resolutions sufficient to analyze variations of spectral properties across Rhea's surface in detail. We present an overview of the VIMS observations obtained so far, as well as the analysis of the spectral properties identified in the VIMS spectra and their variations across its surface compared with spatially highly resolved Cassini ISS images and digital elevation models. Spectral variations measured across Rhea's surface are similar to the variations observed in the VIMS observations of its neighbor Dione, implying similar processes causing or at least inducing their occurrence. Thus, magnetospheric particles and dust impacting onto the trailing hemisphere appear to be responsible for the concentration of dark rocky/organic material and minor amounts of CO 2 in the cratered terrain on the trailing hemisphere. Despite the prominent spectral signatures of Rhea's fresh impact crater Inktomi, radiation effects were identified that also affect the H 2O ice-rich cratered terrain of the leading hemisphere. The concentration of H 2O ice in the vicinity of steep tectonic scarps near 270??W and geologically fresh impact craters implies that Rhea exhibits an icy crust at least in the upper few kilometers. Despite the evidence for past tectonic events, no indications of recent endogenically powered processes could be identified in the Cassini data. ?? 2011 Elsevier Ltd. All rights reserved.

Cassini-VIMS at Jupiter: Solar occultation measurements using Io

Science.gov (United States)

Formisano, V.; D'Aversa, E.; Bellucci, G.; Baines, K.H.; Bibring, J.-P.; Brown, R.H.; Buratti, B.J.; Capaccioni, F.; Cerroni, P.; Clark, R.N.; Coradini, A.; Cruikshank, D.P.; Drossart, P.; Jaumann, R.; Langevin, Y.; Matson, D.L.; McCord, T.B.; Mennella, V.; Nelson, R.M.; Nicholson, P.D.; Sicardy, B.; Sotin, Christophe; Chamberlain, M.C.; Hansen, G.; Hibbits, K.; Showalter, M.; Filacchione, G.

2003-01-01

We report unusual and somewhat unexpected observations of the jovian satellite Io, showing strong methane absorption bands. These observations were made by the Cassini VIMS experiment during the Jupiter flyby of December/January 2000/2001. The explanation is straightforward: Entering or exiting from Jupiter's shadow during an eclipse, Io is illuminated by solar light which has transited the atmosphere of Jupiter. This light, therefore becomes imprinted with the spectral signature of Jupiter's upper atmosphere, which includes strong atmospheric methane absorption bands. Intercepting solar light refracted by the jovian atmosphere, Io essentially becomes a "miffor" for solar occultation events of Jupiter. The thickness of the layer where refracted solar light is observed is so large (more than 3000 km at Io's orbit), that we can foresee a nearly continuous multi-year period of similar events at Saturn, utilizing the large and bright ring system. During Cassini's 4-year nominal mission, this probing tecnique should reveal information of Saturn's atmosphere over a large range of southern latitudes and times. ?? 2003 Elsevier Inc. All rights reserved.

A Survey of Cassini CAPS Ion Observations During Titan Flybys TA-T83

Science.gov (United States)

Woodson, A. K.; Johnson, R. E.; Smith, H. T.; Crary, F. J.

2015-12-01

The Cassini Plasma Spectrometer (CAPS) sampled Titan's plasma environment during each of 83 encounters with the moon between orbit insertion on June 30, 2004 and June 1, 2012. The CAPS Ion Mass Spectrometer (IMS) acquired energy- and mass-per-charge-discriminated time-of-flight (TOF) spectra associated with ionospheric H+, H2+, H3+, CHx+, and C2Hx+ during at least 68 of those encounters. Herein we discuss ion energy distributions extracted from these spectra, each accumulated over an ~4 minute interval along the spacecraft trajectory. This is accomplished by fitting calibration peak models to TOF spectra in order to determine the TOF range associated with each aforementioned ion group, and then summing counts over each TOF range to obtain well-resolved energy peaks for each group. Energy distributions are determined by fitting the logistic power peak function to each of the resulting energy spectra. We then plot the resulting distribution parameters (peak energy, peak amplitude, and peak width or temperature) for each species and each encounter against Titan latitude, longitude, and altitude to generate a map of ion parameters. In addition, the encounters are grouped according to ambient plasma and magnetic field measurements in order to characterize the ion distribution parameters in different regions of Saturn's magnetosphere.

Light weight radioisotope heater unit (LWRHU) production for the Cassini mission

International Nuclear Information System (INIS)

Rinehart, G.H.

1997-01-01

The Light-Weight Radioisotope Heater Unit (LWRHU) is a [sup 238]PuO[sub 2] fueled heat source designed to provide one thermal watt in each of various locations on a spacecraft. The heat sources are required to maintain the temperature of specific components within normal operating ranges. The heat source consists of a hot- pressed [sup 238]PuO[sub 2] fuel pellet, a Pt-3ORh vented capsule, a pyrolytic graphite insulator, and a woven graphite aeroshell assembly. Los Alamos National Laboratory has fabricated 180 heat sources, 157 of which will be used on the Cassini mission

How Spacecraft Fly Spaceflight Without Formulae

CERN Document Server

Swinerd, Graham

2009-01-01

About half a century ago a small satellite, Sputnik 1, was launched. The satellite did very little other than to transmit a radio signal to announce its presence in orbit. However, this humble beginning heralded the dawn of the Space Age. Today literally thousands of robotic spacecraft have been launched, many of which have flown to far-flung regions of the Solar System carrying with them the human spirit of scientific discovery and exploration. Numerous other satellites have been launched in orbit around the Earth providing services that support our technological society on the ground. How Spacecraft Fly: Spaceflight Without Formulae by Graham Swinerd focuses on how these spacecraft work. The book opens with a historical perspective of how we have come to understand our Solar System and the Universe. It then progresses through orbital flight, rocket science, the hostile environment within which spacecraft operate, and how they are designed. The concluding chapters give a glimpse of what the 21st century may ...

GPHS-RTGs in support of the Cassini Mission. Semi-annual technical progress report, April 3, 1995--October 1, 1995

International Nuclear Information System (INIS)

1995-01-01

This document is the April-October 1995 Progress Report on the Cassini RTG Program. Nine tasks are summarized; (1) Spacecraft integration and liason, (2) Engineering support, (3) Safety, (4) Unicouple fabrication, (5) ETG fabrication, assembly, and test, (6) Ground support equipment, (7) RTG shipping and launch support, (8) Design, reviews, and mission applications, and (9) Project management, QA, contract changes, and material acquisitions

OPUS: A Comprehensive Search Tool for Remote Sensing Observations of the Outer Planets. Now with Enhanced Geometric Metadata for Cassini and New Horizons Optical Remote Sensing Instruments.

Science.gov (United States)

Gordon, M. K.; Showalter, M. R.; Ballard, L.; Tiscareno, M.; French, R. S.; Olson, D.

2017-06-01

The PDS RMS Node hosts OPUS - an accurate, comprehensive search tool for spacecraft remote sensing observations. OPUS supports Cassini: CIRS, ISS, UVIS, VIMS; New Horizons: LORRI, MVIC; Galileo SSI; Voyager ISS; and Hubble: ACS, STIS, WFC3, WFPC2.

Cassini/VIMS hyperspectral observations of the HUYGENS landing site on Titan

Science.gov (United States)

Rodriguez, S.; Le, Mouelic S.; Sotin, Christophe; Clenet, H.; Clark, R.N.; Buratti, B.; Brown, R.H.; McCord, T.B.; Nicholson, P.D.; Baines, K.H.

2006-01-01

Titan is one of the primary scientific objectives of the NASA-ESA-ASI Cassini-Huygens mission. Scattering by haze particles in Titan's atmosphere and numerous methane absorptions dramatically veil Titan's surface in the visible range, though it can be studied more easily in some narrow infrared windows. The Visual and Infrared Mapping Spectrometer (VIMS) instrument onboard the Cassini spacecraft successfully imaged its surface in the atmospheric windows, taking hyperspectral images in the range 0.4-5.2 ??m. On 26 October (TA flyby) and 13 December 2004 (TB flyby), the Cassini-Huygens mission flew over Titan at an altitude lower than 1200 km at closest approach. We report here on the analysis of VIMS images of the Huygens landing site acquired at TA and TB, with a spatial resolution ranging from 16 to14.4 km/pixel. The pure atmospheric backscattering component is corrected by using both an empirical method and a first-order theoretical model. Both approaches provide consistent results. After the removal of scattering, ratio images reveal subtle surface heterogeneities. A particularly contrasted structure appears in ratio images involving the 1.59 and 2.03 ??m images north of the Huygens landing site. Although pure water ice cannot be the only component exposed at Titan's surface, this area is consistent with a local enrichment in exposed water ice and seems to be consistent with DISR/Huygens images and spectra interpretations. The images show also a morphological structure that can be interpreted as a 150 km diameter impact crater with a central peak. ?? 2006 Elsevier Ltd. All rights reserved.

A post-Cassini view of Titan's methane-based hydrologic cycle

Science.gov (United States)

Hayes, Alexander G.; Lorenz, Ralph D.; Lunine, Jonathan I.

2018-05-01

The methane-based hydrologic cycle on Saturn's largest moon, Titan, is an extreme analogue to Earth's water cycle. Titan is the only planetary body in the Solar System, other than Earth, that is known to have an active hydrologic cycle. With a surface pressure of 1.5 bar and temperatures of 90 to 95 K, methane and ethane condense out of a nitrogen-based atmosphere and flow as liquids on the moon's surface. Exchange processes between atmospheric, surface and subsurface reservoirs produce methane and ethane cloud systems, as well as erosional and depositional landscapes that have strikingly similar forms to their terrestrial counterparts. Over its 13-year exploration of the Saturn system, the Cassini-Huygens mission revealed that Titan's hydrocarbon-based hydrology is driven by nested methane cycles that operate over a range of timescales, including geologic, orbital (for example, Croll-Milankovitch cycles), seasonal and that of a single convective storm. In this Review Article, we describe the dominant exchange processes that operate over these timescales and present a post-Cassini view of Titan's methane-based hydrologic system.

Spacecraft Charge Monitor

Science.gov (United States)

Goembel, L.

2003-12-01

We are currently developing a flight prototype Spacecraft Charge Monitor (SCM) with support from NASA's Small Business Innovation Research (SBIR) program. The device will use a recently proposed high energy-resolution electron spectroscopic technique to determine spacecraft floating potential. The inspiration for the technique came from data collected by the Atmosphere Explorer (AE) satellites in the 1970s. The data available from the AE satellites indicate that the SCM may be able to determine spacecraft floating potential to within 0.1 V under certain conditions. Such accurate measurement of spacecraft charge could be used to correct biases in space plasma measurements. The device may also be able to measure spacecraft floating potential in the solar wind and in orbit around other planets.

Internet Technology on Spacecraft

Science.gov (United States)

Rash, James; Parise, Ron; Hogie, Keith; Criscuolo, Ed; Langston, Jim; Powers, Edward I. (Technical Monitor)

2000-01-01

The Operating Missions as Nodes on the Internet (OMNI) project has shown that Internet technology works in space missions through a demonstration using the UoSAT-12 spacecraft. An Internet Protocol (IP) stack was installed on the orbiting UoSAT-12 spacecraft and tests were run to demonstrate Internet connectivity and measure performance. This also forms the basis for demonstrating subsequent scenarios. This approach provides capabilities heretofore either too expensive or simply not feasible such as reconfiguration on orbit. The OMNI project recognized the need to reduce the risk perceived by mission managers and did this with a multi-phase strategy. In the initial phase, the concepts were implemented in a prototype system that includes space similar components communicating over the TDRS (space network) and the terrestrial Internet. The demonstration system includes a simulated spacecraft with sample instruments. Over 25 demonstrations have been given to mission and project managers, National Aeronautics and Space Administration (NASA), Department of Defense (DoD), contractor technologists and other decisions makers, This initial phase reached a high point with an OMNI demonstration given from a booth at the Johnson Space Center (JSC) Inspection Day 99 exhibition. The proof to mission managers is provided during this second phase with year 2000 accomplishments: testing the use of Internet technologies onboard an actual spacecraft. This was done with a series of tests performed using the UoSAT-12 spacecraft. This spacecraft was reconfigured on orbit at very low cost. The total period between concept and the first tests was only 6 months! On board software was modified to add an IP stack to support basic IP communications. Also added was support for ping, traceroute and network timing protocol (NTP) tests. These tests show that basic Internet functionality can be used onboard spacecraft. The performance of data was measured to show no degradation from current

Titan from Cassini-Huygens

CERN Document Server

Brown, Robert H; Waite, J. Hunter

2010-01-01

This book reviews our current knowledge of Saturn's largest moon Titan featuring the latest results obtained by the Cassini-Huygens mission. A global author team addresses Titan’s origin and evolution, internal structure, surface geology, the atmosphere and ionosphere as well as magnetospheric interactions. The book closes with an outlook beyond the Cassini-Huygens mission. Colorfully illustrated, this book will serve as a reference to researchers as well as an introduction for students.

Retrieval of RTG'S in earth orbit

International Nuclear Information System (INIS)

Raab, B.; Frieder, M.A.; Skrabek, A.

1982-01-01

Since 1961, some ten Radioisotope Thermoelectric Generators (RTG's) have been placed into a variety of spacecraft which are now in earth orbit. All of these spacecraft are in orbits with lifetimes in excess of 100 years and pose no risk. However, since most of these spacecraft are no longer being actively used, these may be subject to an active removal program to reduce the population of objects in space. Therefore, a study was undertaken to evaluate the feasibility of retrieving or disposing of spacecraft with RTGs on board. Intervention scenarios are developed and an orbital rendezvous vehicle is conceptualized. The costs of RTG retrieval are derived and compared to the costs of RTG disposal, i.e., boost to a higher, multi-millenium-lifetime orbit, and are found to be not significantly different

AN ATLAS OF BRIGHT STAR SPECTRA IN THE NEAR-INFRARED FROM CASSINI-VIMS

International Nuclear Information System (INIS)

Stewart, Paul N.; Tuthill, Peter G.; Nicholson, Philip D.; Sloan, G. C.; Hedman, Matthew M.

2015-01-01

We present the Cassini Atlas Of Stellar Spectra (CAOSS), comprised of near-infrared, low-resolution spectra of bright stars recovered from space-based observations by the Cassini spacecraft. The 65 stellar targets in the atlas are predominately M, K, and S giants. However, it also contains spectra of other bright nearby stars including carbon stars and main-sequence stars from A to F. The spectra presented are free of all spectral contamination caused by the Earth's atmosphere, including the detrimental telluric molecular bands which put parts of the near-infrared spectrum out of reach of terrestrial observations. With a single instrument, a spectro-photometric data set is recovered that spans the near-infrared from 0.8 to 5.1 μm with spectral resolution ranging from R = 53.5 to R = 325. Spectra have been calibrated into absolute flux units after careful characterization of the instrumental spectral efficiency. Spectral energy distributions for most stars match closely with literature values. All final data products have been made available online

AN ATLAS OF BRIGHT STAR SPECTRA IN THE NEAR-INFRARED FROM CASSINI-VIMS

Energy Technology Data Exchange (ETDEWEB)

Stewart, Paul N.; Tuthill, Peter G. [Sydney Institute for Astronomy, School of Physics, The University of Sydney, NSW 2006 (Australia); Nicholson, Philip D. [Department of Astronomy, Cornell University, Ithaca, NY 14853 (United States); Sloan, G. C. [Cornell Center for Astrophyics and Planetary Science, Cornell University, Ithaca, NY 14853 (United States); Hedman, Matthew M., E-mail: p.stewart@physics.usyd.edu.au [Department of Physics, University of Idaho, Moscow, ID 83844 (United States)

2015-12-15

We present the Cassini Atlas Of Stellar Spectra (CAOSS), comprised of near-infrared, low-resolution spectra of bright stars recovered from space-based observations by the Cassini spacecraft. The 65 stellar targets in the atlas are predominately M, K, and S giants. However, it also contains spectra of other bright nearby stars including carbon stars and main-sequence stars from A to F. The spectra presented are free of all spectral contamination caused by the Earth's atmosphere, including the detrimental telluric molecular bands which put parts of the near-infrared spectrum out of reach of terrestrial observations. With a single instrument, a spectro-photometric data set is recovered that spans the near-infrared from 0.8 to 5.1 μm with spectral resolution ranging from R = 53.5 to R = 325. Spectra have been calibrated into absolute flux units after careful characterization of the instrumental spectral efficiency. Spectral energy distributions for most stars match closely with literature values. All final data products have been made available online.

Design feasibility via ascent optimality for next-generation spacecraft

Science.gov (United States)

Miele, A.; Mancuso, S.

This paper deals with the optimization of the ascent trajectories for single-stage-sub-orbit (SSSO), single-stage-to-orbit (SSTO), and two-stage-to-orbit (TSTO) rocket-powered spacecraft. The maximum payload weight problem is studied for different values of the engine specific impulse and spacecraft structural factor. The main conclusions are that: feasibility of SSSO spacecraft is guaranteed for all the parameter combinations considered; feasibility of SSTO spacecraft depends strongly on the parameter combination chosen; not only feasibility of TSTO spacecraft is guaranteed for all the parameter combinations considered, but the TSTO payload is several times the SSTO payload. Improvements in engine specific impulse and spacecraft structural factor are desirable and crucial for SSTO feasibility; indeed, aerodynamic improvements do not yield significant improvements in payload. For SSSO, SSTO, and TSTO spacecraft, simple engineering approximations are developed connecting the maximum payload weight to the engine specific impulse and spacecraft structural factor. With reference to the specific impulse/structural factor domain, these engineering approximations lead to the construction of zero-payload lines separating the feasibility region (positive payload) from the unfeasibility region (negative payload).

Vibration and Acoustic Testing for Mars Micromission Spacecraft

Science.gov (United States)

Kern, Dennis L.; Scharton, Terry D.

1999-01-01

The objective of the Mars Micromission program being managed by the Jet Propulsion Laboratory (JPL) for NASA is to develop a common spacecraft that can carry telecommunications equipment and a variety of science payloads for exploration of Mars. The spacecraft will be capable of carrying robot landers and rovers, cameras, probes, balloons, gliders or aircraft, and telecommunications equipment to Mars at much lower cost than recent NASA Mars missions. The lightweight spacecraft (about 220 Kg mass) will be launched in a cooperative venture with CNES as a TWIN auxiliary payload on the Ariane 5 launch vehicle. Two or more Mars Micromission launches are planned for each Mars launch opportunity, which occur every 26 months. The Mars launch window for the first mission is November 1, 2002 through April 2003, which is planned to be a Mars airplane technology demonstration mission to coincide with the 100 year anniversary of the Kittyhawk flight. Several subsequent launches will create a telecommunications network orbiting Mars, which will provide for continuous communication with lenders and rovers on the Martian surface. Dedicated science payload flights to Mars are slated to start in 2005. This new cheaper and faster approach to Mars exploration calls for innovative approaches to the qualification of the Mars Micromission spacecraft for the Ariane 5 launch vibration and acoustic environments. JPL has in recent years implemented new approaches to spacecraft testing that may be effectively applied to the Mars Micromission. These include 1) force limited vibration testing, 2) combined loads, vibration and modal testing, and 3) direct acoustic testing. JPL has performed nearly 200 force limited vibration tests in the past 9 years; several of the tests were on spacecraft and large instruments, including the Cassini and Deep Space One spacecraft. Force limiting, which measures and limits the spacecraft base reaction force using triaxial force gages sandwiched between the

GPHS-RTGs in support of the Cassini RTG Program. Final technical report, January 11, 1991--April 30, 1998

Energy Technology Data Exchange (ETDEWEB)

NONE

1998-08-01

As noted in the historical summary, this program encountered a number of changes in direction, schedule, and scope over the period 11 January 1991 to 31 December 1998. The report provides a comprehensive summary of all the varied aspects of the program over its seven and a quarter years, and highlights those aspects that provide information beneficial to future radioisotope programs. In addition to summarizing the scope of the Cassini GPHS-RTG Program provided as background, the introduction includes a discussion of the scope of the final report and offers reference sources for information on those topics not covered. Much of the design heritage of the GPHS-RTG comes from the Multi-Hundred Watt (MHW) RTGs used on the Lincoln Experimental Satellites (LES) 8/9 and Voyager spacecraft. The design utilized for the Cassini program was developed, in large part, under the GPHS-RTG program which produced the Galileo and Ulysses RTGs. Reports from those programs included detailed documentation of the design, development, and testing of converter components and full converters that were identical to, or similar to, components used in the Cassini program. Where such information is available in previous reports, it is not repeated here.

GPHS-RTGs in support of the Cassini RTG Program. Final technical report, January 11, 1991 - April 30, 1998

International Nuclear Information System (INIS)

1998-08-01

As noted in the historical summary, this program encountered a number of changes in direction, schedule, and scope over the period 11 January 1991 to 31 December 1998. The report provides a comprehensive summary of all the varied aspects of the program over its seven and a quarter years, and highlights those aspects that provide information beneficial to future radioisotope programs. In addition to summarizing the scope of the Cassini GPHS-RTG Program provided as background, the introduction includes a discussion of the scope of the final report and offers reference sources for information on those topics not covered. Much of the design heritage of the GPHS-RTG comes from the Multi-Hundred Watt (MHW) RTGs used on the Lincoln Experimental Satellites (LES) 8/9 and Voyager spacecraft. The design utilized for the Cassini program was developed, in large part, under the GPHS-RTG program which produced the Galileo and Ulysses RTGs. Reports from those programs included detailed documentation of the design, development, and testing of converter components and full converters that were identical to, or similar to, components used in the Cassini program. Where such information is available in previous reports, it is not repeated here

Analytical thermal model validation for Cassini radioisotope thermoelectric generator

International Nuclear Information System (INIS)

Lin, E.I.

1997-01-01

The Saturn-bound Cassini spacecraft is designed to rely, without precedent, on the waste heat from its three radioisotope thermoelectric generators (RTGs) to warm the propulsion module subsystem, and the RTG end dome temperature is a key determining factor of the amount of waste heat delivered. A previously validated SINDA thermal model of the RTG was the sole guide to understanding its complex thermal behavior, but displayed large discrepancies against some initial thermal development test data. A careful revalidation effort led to significant modifications and adjustments of the model, which result in a doubling of the radiative heat transfer from the heat source support assemblies to the end domes and bring up the end dome and flange temperature predictions to within 2 C of the pertinent test data. The increased inboard end dome temperature has a considerable impact on thermal control of the spacecraft central body. The validation process offers an example of physically-driven analytical model calibration with test data from not only an electrical simulator but also a nuclear-fueled flight unit, and has established the end dome temperatures of a flight RTG where no in-flight or ground-test data existed before

JSC Orbital Debris Website Description

Science.gov (United States)

Johnson, Nicholas L.

2006-01-01

Purpose: The website provides information about the NASA Orbital Debris Program Office at JSC, which is the lead NASA center for orbital debris research. It is recognized world-wide for its leadership in addressing orbital debris issues. The NASA Orbital Debris Program Office has taken the international lead in conducting measurements of the environment and in developing the technical consensus for adopting mitigation measures to protect users of the orbital environment. Work at the center continues with developing an improved understanding of the orbital debris environment and measures that can be taken to control its growth. Major Contents: Orbital Debris research is divided into the following five broad efforts. Each area of research contains specific information as follows: 1) Modeling - NASA scientists continue to develop and upgrade orbital debris models to describe and characterize the current and future debris environment. Evolutionary and engineering models are described in detail. Downloadable items include a document in PDF format and executable software. 2) Measurements - Measurements of near-Earth orbital debris are accomplished by conducting ground-based and space-based observations of the orbital debris environment. The data from these sources provide validation of the environment models and identify the presence of new sources. Radar, optical and surface examinations are described. External links to related topics are provided. 3) Protection - Orbital debris protection involves conducting hypervelocity impact measurements to assess the risk presented by orbital debris to operating spacecraft and developing new materials and new designs to provide better protection from the environment with less weight penalty. The data from this work provides the link between the environment defined by the models and the risk presented by that environment to operating spacecraft and provides recommendations on design and operations procedures to reduce the risk as

Artist concept of Galileo spacecraft

Science.gov (United States)

1988-01-01

Galileo spacecraft is illustrated in artist concept. Gallileo, named for the Italian astronomer, physicist and mathematician who is credited with construction of the first complete, practical telescope in 1620, will make detailed studies of Jupiter. A cooperative program with the Federal Republic of Germany the Galileo mission will amplify information acquired by two Voyager spacecraft in their brief flybys. Galileo is a two-element system that includes a Jupiter-orbiting observatory and an entry probe. Jet Propulsion Laboratory (JPL) is Galileo project manager and builder of the main spacecraft. Ames Research Center (ARC) has responsibility for the entry probe, which was built by Hughes Aircraft Company and General Electric. Galileo will be deployed from the payload bay (PLB) of Atlantis, Orbiter Vehicle (OV) 104, during mission STS-34.

A user's guide to the Flexible Spacecraft Dynamics and Control Program

Science.gov (United States)

Fedor, J. V.

1984-01-01

A guide to the use of the Flexible Spacecraft Dynamics Program (FSD) is presented covering input requirements, control words, orbit generation, spacecraft description and simulation options, and output definition. The program can be used in dynamics and control analysis as well as in orbit support of deployment and control of spacecraft. The program is applicable to inertially oriented spinning, Earth oriented or gravity gradient stabilized spacecraft. Internal and external environmental effects can be simulated.

Guidance and control of swarms of spacecraft

Science.gov (United States)

Morgan, Daniel James

There has been considerable interest in formation flying spacecraft due to their potential to perform certain tasks at a cheaper cost than monolithic spacecraft. Formation flying enables the use of smaller, cheaper spacecraft that distribute the risk of the mission. Recently, the ideas of formation flying have been extended to spacecraft swarms made up of hundreds to thousands of 100-gram-class spacecraft known as femtosatellites. The large number of spacecraft and limited capabilities of each individual spacecraft present a significant challenge in guidance, navigation, and control. This dissertation deals with the guidance and control algorithms required to enable the flight of spacecraft swarms. The algorithms developed in this dissertation are focused on achieving two main goals: swarm keeping and swarm reconfiguration. The objectives of swarm keeping are to maintain bounded relative distances between spacecraft, prevent collisions between spacecraft, and minimize the propellant used by each spacecraft. Swarm reconfiguration requires the transfer of the swarm to a specific shape. Like with swarm keeping, minimizing the propellant used and preventing collisions are the main objectives. Additionally, the algorithms required for swarm keeping and swarm reconfiguration should be decentralized with respect to communication and computation so that they can be implemented on femtosats, which have limited hardware capabilities. The algorithms developed in this dissertation are concerned with swarms located in low Earth orbit. In these orbits, Earth oblateness and atmospheric drag have a significant effect on the relative motion of the swarm. The complicated dynamic environment of low Earth orbits further complicates the swarm-keeping and swarm-reconfiguration problems. To better develop and test these algorithms, a nonlinear, relative dynamic model with J2 and drag perturbations is developed. This model is used throughout this dissertation to validate the algorithms

Jovian decametric radiation seen from Juno, Cassini, STEREO A, WIND, and Earth-based radio observatories

Science.gov (United States)

Imai, M.; Kurth, W. S.; Hospodarsky, G. B.; Bolton, S. J.; Connerney, J. E. P.; Levin, S. M.; Lecacheux, A.; Lamy, L.; Zarka, P.; Clarke, T. E.; Higgins, C. A.

2017-09-01

Jupiter's decametric (DAM) radiation is generated very close to the local gyrofrequency by the electron cyclotron maser instability (CMI). The first two-point common detections of Jovian DAM radiation were made using the Voyager spacecraft and ground-based radio observatories in early 1979, but, due to geometrical constraints and limited flyby duration, a full understanding of the latitudinal beaming of Jovian DAM radiation remains elusive. The stereoscopic DAM radiation viewed from Juno, Cassini, STEREO A, WIND, and Earth-based radio observatories provides a unique opportunity to analyze the CMI emission mechanism and beaming properties.

Navigating the MESSENGER Spacecraft through End of Mission

Science.gov (United States)

Bryan, C. G.; Williams, B. G.; Williams, K. E.; Taylor, A. H.; Carranza, E.; Page, B. R.; Stanbridge, D. R.; Mazarico, E.; Neumann, G. A.; O'Shaughnessy, D. J.; McAdams, J. V.; Calloway, A. B.

2015-12-01

The MErcury Surface, Space ENvironment, GEochemistry, and Ranging (MESSENGER) spacecraft orbited the planet Mercury from March 2011 until the end of April 2015, when it impacted the planetary surface after propellant reserves used to maintain the orbit were depleted. This highly successful mission was led by the principal investigator, Sean C. Solomon, of Columbia University. The Johns Hopkins University Applied Physics Laboratory (JHU/APL) designed and assembled the spacecraft and served as the home for spacecraft operations. Spacecraft navigation for the entirety of the mission was provided by the Space Navigation and Flight Dynamics Practice (SNAFD) of KinetX Aerospace. Orbit determination (OD) solutions were generated through processing of radiometric tracking data provided by NASA's Deep Space Network (DSN) using the MIRAGE suite of orbital analysis tools. The MESSENGER orbit was highly eccentric, with periapsis at a high northern latitude and periapsis altitude in the range 200-500 km for most of the orbital mission phase. In a low-altitude "hover campaign" during the final two months of the mission, periapsis altitudes were maintained within a narrow range between about 35 km and 5 km. Navigating a spacecraft so near a planetary surface presented special challenges. Tasks required to meet those challenges included the modeling and estimation of Mercury's gravity field and of solar and planetary radiation pressure, and the design of frequent orbit-correction maneuvers. Superior solar conjunction also presented observational modeling issues. One key to the overall success of the low-altitude hover campaign was a strategy to utilize data from an onboard laser altimeter as a cross-check on the navigation team's reconstructed and predicted estimates of periapsis altitude. Data obtained from the Mercury Laser Altimeter (MLA) on a daily basis provided near-real-time feedback that proved invaluable in evaluating alternative orbit estimation strategies, and

Multiple spacecraft Michelson stellar interferometer

Science.gov (United States)

Stachnik, R. V.; Arnold, D.; Melroy, P.; Mccormack, E. F.; Gezari, D. Y.

1984-01-01

Results of an orbital analysis and performance assessment of SAMSI (Spacecraft Array for Michelson Spatial Interferometry) are presented. The device considered includes two one-meter telescopes in orbits which are identical except for slightly different inclinations; the telescopes achieve separations as large as 10 km and relay starlight to a central station which has a one-meter optical delay line in one interferometer arm. It is shown that a 1000-km altitude, zero mean inclination orbit affords natural scanning of the 10-km baseline with departures from optical pathlength equality which are well within the corrective capacity of the optical delay line. Electric propulsion is completely adequate to provide the required spacecraft motions, principally those needed for repointing. Resolution of 0.00001 arcsec and magnitude limits of 15 to 20 are achievable.

An Artificial Gravity Spacecraft Approach which Minimizes Mass, Fuel and Orbital Assembly Reg

Science.gov (United States)

Bell, L.

2002-01-01

The Sasakawa International Center for Space Architecture (SICSA) is undertaking a multi-year research and design study that is exploring near and long-term commercial space development opportunities. Space tourism in low-Earth orbit (LEO), and possibly beyond LEO, comprises one business element of this plan. Supported by a financial gift from the owner of a national U.S. hotel chain, SICSA has examined opportunities, requirements and facility concepts to accommodate up to 100 private citizens and crewmembers in LEO, as well as on lunar/planetary rendezvous voyages. SICSA's artificial gravity Science Excursion Vehicle ("AGSEV") design which is featured in this presentation was conceived as an option for consideration to enable round-trip travel to Moon and Mars orbits and back from LEO. During the course of its development, the AGSEV would also serve other important purposes. An early assembly stage would provide an orbital science and technology testbed for artificial gravity demonstration experiments. An ultimate mature stage application would carry crews of up to 12 people on Mars rendezvous missions, consuming approximately the same propellant mass required for lunar excursions. Since artificial gravity spacecraft that rotate to create centripetal accelerations must have long spin radii to limit adverse effects of Coriolis forces upon inhabitants, SICSA's AGSEV design embodies a unique tethered body concept which is highly efficient in terms of structural mass and on-orbit assembly requirements. The design also incorporates "inflatable" as well as "hard" habitat modules to optimize internal volume/mass relationships. Other important considerations and features include: maximizing safety through element and system redundancy; means to avoid destabilizing mass imbalances throughout all construction and operational stages; optimizing ease of on-orbit servicing between missions; and maximizing comfort and performance through careful attention to human needs. A

Man with a Mission: Jean-Dominique Cassini

Science.gov (United States)

Belkora, Leila

2004-03-01

Jean-Dominique Cassini, for whom the Cassini mission to Saturn is named, is best known for his early understanding of that planet's rings. This article is an overview of his influential career in astronomy and other scientific fields.= Born in Italy in1625 and formally educated at an early age, he was a professor of astronomy at the University of Bologna, a leading center of learning in Europe of the time. He was an early observer of Jupiter, Mars, and Venus. He is best known for constructing a giant pinhole camera in a cathedral that he used with a meridian line on the floor to track the Sun's image through the year, thus providing the Catholic Church with a reliable calendar. Cassini also used the pinhole camera observations to calculate the variation in the distance between the Sun and Earth, thus lending support to the Copernican (Sun-centered) view of the solar system. Cassini moved to Paris at the request of King Louis XIV, originally to oversee the surveying needed for a new map system of France, but ultimately he took over as the director of the Paris Observatory. Cassini's descendants ran the observatory there for the following century.

Science Planning and Orbit Classification for Solar Probe Plus

Science.gov (United States)

Kusterer, M. B.; Fox, N. J.; Rodgers, D. J.; Turner, F. S.

2016-12-01

There are a number of challenges for the Science Planning Team (SPT) of the Solar Probe Plus (SPP) Mission. Since SPP is using a decoupled payload operations approach, tight coordination between the mission operations and payload teams will be required. The payload teams must manage the volume of data that they write to the spacecraft solid-state recorders (SSR) for their individual instruments for downlink to the ground. Making this process more difficult, the geometry of the celestial bodies and the spacecraft during some of the SPP mission orbits cause limited uplink and downlink opportunities. The payload teams will also be required to coordinate power on opportunities, command uplink opportunities, and data transfers from instrument memory to the spacecraft SSR with the operation team. The SPT also intend to coordinate observations with other spacecraft and ground based systems. To solve these challenges, detailed orbit activity planning is required in advance for each orbit. An orbit planning process is being created to facilitate the coordination of spacecraft and payload activities for each orbit. An interactive Science Planning Tool is being designed to integrate the payload data volume and priority allocations, spacecraft ephemeris, attitude, downlink and uplink schedules, spacecraft and payload activities, and other spacecraft ephemeris. It will be used during science planning to select the instrument data priorities and data volumes that satisfy the orbit data volume constraints and power on, command uplink and data transfer time periods. To aid in the initial stages of science planning we have created an orbit classification scheme based on downlink availability and significant science events. Different types of challenges arise in the management of science data driven by orbital geometry and operational constraints, and this scheme attempts to identify the patterns that emerge.

A Sampling Based Approach to Spacecraft Autonomous Maneuvering with Safety Specifications

Science.gov (United States)

Starek, Joseph A.; Barbee, Brent W.; Pavone, Marco

2015-01-01

This paper presents a methods for safe spacecraft autonomous maneuvering that leverages robotic motion-planning techniques to spacecraft control. Specifically the scenario we consider is an in-plan rendezvous of a chaser spacecraft in proximity to a target spacecraft at the origin of the Clohessy Wiltshire Hill frame. The trajectory for the chaser spacecraft is generated in a receding horizon fashion by executing a sampling based robotic motion planning algorithm name Fast Marching Trees (FMT) which efficiently grows a tree of trajectories over a set of probabillistically drawn samples in the state space. To enforce safety the tree is only grown over actively safe samples for which there exists a one-burn collision avoidance maneuver that circularizes the spacecraft orbit along a collision-free coasting arc and that can be executed under potential thrusters failures. The overall approach establishes a provably correct framework for the systematic encoding of safety specifications into the spacecraft trajectory generations process and appears amenable to real time implementation on orbit. Simulation results are presented for a two-fault tolerant spacecraft during autonomous approach to a single client in Low Earth Orbit.

Thermal elastic shock and its effect on TOPEX spacecraft attitude control

Science.gov (United States)

Zimbelman, Darrell F.

1991-01-01

Thermal elastic shock (TES) is a twice per orbit impulsive disturbance torque experienced by low-Earth orbiting spacecraft. The fundamental equations used to model the TES disturbance torque for typical spacecraft appendages (e.g., solar arrays and antenna booms) are derived in detail. In particular, the attitude-pointing performance of the TOPEX spacecraft, when subjected to the TES disturbance, is analyzed using a three-axis nonlinear time-domain simulation. Results indicate that the TOPEX spacecraft could exceed its roll-axis attitude-control requirement during penumbral transitions, and remain in violation for approximately 150 sec each orbit until the umbra collapses. A localized active-control system is proposed as a solution to minimize and/or eliminate the degrading effects of the TES disturbance.

Attitude control for on-orbit servicing spacecraft using hybrid actuator

Science.gov (United States)

Wu, Yunhua; Han, Feng; Zheng, Mohong; He, Mengjie; Chen, Zhiming; Hua, Bing; Wang, Feng

2018-03-01

On-orbit servicing is one of the research hotspots of space missions. A small satellite equipped with multiple robotic manipulators is expected to carry out device replacement task for target large spacecraft. Attitude hyperstable control of a small satellite platform under rotations of the manipulators is a challenging problem. A hybrid momentum exchanging actuator consists of Control Moment Gyro (CMG) and Reaction Wheel (RW) is proposed to tackle the above issue, due to its huge amount of momentum storage capacity of the CMG and high control accuracy of the RW, in which the CMG produces large command torque while the RW offers additional control degrees. The constructed dynamic model of the servicing satellite advises that it's feasible for attitude hyperstable control of the platform with arbitrary manipulators through compensating the disturbance generated by rapid rotation of the manipulators. Then, null motion between the CMG and RW is exploited to drive the system to the expected target with favorable performance, and to overcome the CMG inherent geometric singularity and RW saturation. Simulations with different initial situations, including CMG hyperbolic and elliptic singularities and RW saturation, are executed. Compared to the scenarios where the CMG or RW fails stabilizing the platform, large control torque, precise control effect and escape of singularity are guaranteed by the introduced hybrid actuator, CMGRW (CMGRW refers to the hybrid momentum exchanging devices in this paper, consisting of 4 CMGs in classical pyramid cluster and 3 RWs in an orthogonal group (specific description can been found in Section 4)). The feasible performance of the satellite, CMG and RW under large disturbance demonstrates that the control architecture proposed is capable of attitude control for on-orbit servicing satellite with multiple robotic manipulators.

Summary of the Orbiter mechanical systems

Science.gov (United States)

Kiker, J.; Hinson, K.

1979-01-01

Major mechanical systems of the Orbiter space vehicle are summarized with respect to general design details, manner of operation, expected performance, and, where applicable, unique features. A synopsis of data obtained during the five atmospheric flight tests of spacecraft OV-101 and status of the systems for the first orbital spacecraft STS-1 are presented.

Investigating fundamental physics and space environment with a dedicated Earth-orbiting spacecraft

Science.gov (United States)

Peron, Roberto

-year requirement and thus they need specific arrangements for deorbiting at the end of life or they can simply rely on mother nature for reentry. The goal of this proposed approach is to utilize existing technology developed for acceleration measurement in space and state-of-the-art satellite tracking to precisely determine the orbit of a satellite with well-defined geometrical and mass characteristics (i.e., (A/m) ratio), at the same time accurately measuring over a long period of time the drag deceleration (as well as others non-gravitational effects) acting on the satellite. This will result in a virtually drag-free object that can be exploited to: 1. perform fundamental physics tests by verifying the equation of motion of a test mass in the general relativistic context and placing limits to alternative theories of gravitation; 2. improve the knowledge of selected tidal terms; 3. map, through acceleration measurements, the atmospheric density in the orbital region of interest. In its preliminary incarnation, the satellite would be cylindrical in shape and spinning about its cylinder axis that would be also orthogonal to the orbital plane. The satellite should be placed on a dawn-dusk, sun-synchronous, elliptical orbit spanning the orbital altitudes of interest (e.g., between 500 and 1200 km of altitude). The satellite should be equipped with a 3-axis accelerometer package with an acceleration resolution better than (10^{-11} g) (with (g) the acceleration at the Earth's surface). The expected measurement range is (10^{-8} - 10^{-11} g) considering estimates of drag forces at minimum and maximum solar activity conditions in the altitude range of interest and a preliminary estimate of the satellite (A/m) ratio. The overall concept of the mission will be discussed, concentrating on the fundamental aspects and main scientific return. The main instrumentation to be hosted on-board the spacecraft will be then reviewed, with a focus on current and projected capabilities.

Definition of the topological structure of the automatic control system of spacecrafts

International Nuclear Information System (INIS)

KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" data-affiliation=" (Siberian State Aerospace University named after Academician M.F.Reshetnev 31 KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" >Zelenkov, P V; KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" data-affiliation=" (Siberian State Aerospace University named after Academician M.F.Reshetnev 31 KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" >Karaseva, M V; KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" data-affiliation=" (Siberian State Aerospace University named after Academician M.F.Reshetnev 31 KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" >Tsareva, E A; Tsarev, R Y

2015-01-01

The paper considers the problem of selection the topological structure of the automated control system of spacecrafts. The integer linear model of mathematical programming designed to define the optimal topological structure for spacecraft control is proposed. To solve the determination problem of topological structure of the control system of spacecrafts developed the procedure of the directed search of some structure variants according to the scheme 'Branch and bound'. The example of the automated control system of spacecraft development included the combination of ground control stations, managing the spacecraft of three classes with a geosynchronous orbit with constant orbital periods is presented

Orbit-attitude coupled motion around small bodies: Sun-synchronous orbits with Sun-tracking attitude motion

Science.gov (United States)

Kikuchi, Shota; Howell, Kathleen C.; Tsuda, Yuichi; Kawaguchi, Jun'ichiro

2017-11-01

The motion of a spacecraft in proximity to a small body is significantly perturbed due to its irregular gravity field and solar radiation pressure. In such a strongly perturbed environment, the coupling effect of the orbital and attitude motions exerts a large influence that cannot be neglected. However, natural orbit-attitude coupled dynamics around small bodies that are stationary in both orbital and attitude motions have yet to be observed. The present study therefore investigates natural coupled motion that involves both a Sun-synchronous orbit and Sun-tracking attitude motion. This orbit-attitude coupled motion enables a spacecraft to maintain its orbital geometry and attitude state with respect to the Sun without requiring active control. Therefore, the proposed method can reduce the use of an orbit and attitude control system. This paper first presents analytical conditions to achieve Sun-synchronous orbits and Sun-tracking attitude motion. These analytical solutions are then numerically propagated based on non-linear coupled orbit-attitude equations of motion. Consequently, the possibility of implementing Sun-synchronous orbits with Sun-tracking attitude motion is demonstrated.

Effects of Space Weather on Geosynchronous Electromagnetic Spacecraft Perturbations Using Statistical Fluxes

Science.gov (United States)

Hughes, J.; Schaub, H.

2017-12-01

Spacecraft can charge to very negative voltages at GEO due to interactions with the space plasma. This can cause arcing which can damage spacecraft electronics or solar panels. Recently, it has been suggested that spacecraft charging may lead to orbital perturbations which change the orbits of lightweight uncontrolled debris orbits significantly. The motions of High Area to Mass Ratio objects are not well explained with just perturbations from Solar Radiation Pressure (SRP) and earth, moon, and sun gravity. A charged spacecraft will experience a Lorentz force as the spacecraft moves relative to Earth's magnetic field, as well as a Lorentz torque and eddy current torques if the object is rotating. Prior work assuming a constant "worst case" voltage has shown that Lorentz and eddy torques can cause quite large orbital changes by rotating the object to experience more or less SRP. For some objects, including or neglecting these electromagnetic torques can lead to differences of thousands of kilometers after only two orbits. This paper will further investigate the effects of electromagnetic perturbations by using a charging model that uses measured flux distributions to better simulate natural charging. This differs from prior work which used a constant voltage or Maxwellian distributions. This is done to a calm space weather case of Kp = 2 and a stormy case where Kp = 8. Preliminary analysis suggests that electrostatics will still cause large orbital changes even with the more realistic charging model.

Spacecraft Environmental Interactions Technology, 1983

Science.gov (United States)

1985-01-01

State of the art of environment interactions dealing with low-Earth-orbit plasmas; high-voltage systems; spacecraft charging; materials effects; and direction of future programs are contained in over 50 papers.

Testing programs for the Multimission Modular Spacecraft

Science.gov (United States)

Greenwell, T. J.

1978-01-01

The Multimission Modular Spacecraft (MMS) provides a standard spacecraft bus to a user for a variety of space missions ranging from near-earth to synchronous orbits. The present paper describes the philosophy behind the MMS module test program and discusses the implementation of the test program. It is concluded that the MMS module test program provides an effective and comprehensive customer buy-off at the subsystem contractor's plant, is an optimum approach for checkout of the subsystems prior to use for on-orbit servicing in the Shuttle Cargo Bay, and is a cost-effective technique for environmental testing.

The Coupled Orbit-Attitude Dynamics and Control of Electric Sail in Displaced Solar Orbits

Directory of Open Access Journals (Sweden)

Mingying Huo

2017-01-01

Full Text Available Displaced solar orbits for spacecraft propelled by electric sails are investigated. Since the propulsive thrust is induced by the sail attitude, the orbital and attitude dynamics of electric-sail-based spacecraft are coupled and required to be investigated together. However, the coupled dynamics and control of electric sails have not been discussed in most published literatures. In this paper, the equilibrium point of the coupled dynamical system in displaced orbit is obtained, and its stability is analyzed through a linearization. The results of stability analysis show that only some of the orbits are marginally stable. For unstable displaced orbits, linear quadratic regulator is employed to control the coupled attitude-orbit system. Numerical simulations show that the proposed strategy can control the coupled system and a small torque can stabilize both the attitude and orbit. In order to generate the control force and torque, the voltage distribution problem is studied in an optimal framework. The numerical results show that the control force and torque of electric sail can be realized by adjusting the voltage distribution of charged tethers.

Viking orbiter and its Mariner inheritance

Science.gov (United States)

1984-01-01

Improvements to the design of the Mariner spacecraft resulted in the Viking spacecraft. The Viking spacecraft would consist of two major systems - an orbiter and a lander, while the lander would provide the means for safely delivering the scientific instruments to the surface, house, and provide the necessary power source and communication links for those experiments, the orbiter would transport the lander to Mars, rovide a platform for the Viking imaging system so that proposed landing sites could be surveyed and certified, relay lander science information back to Earth, and conduct scientific observations in its own right.

Dynamics and control of Lorentz-augmented spacecraft relative motion

CERN Document Server

Yan, Ye; Yang, Yueneng

2017-01-01

This book develops a dynamical model of the orbital motion of Lorentz spacecraft in both unperturbed and J2-perturbed environments. It explicitly discusses three kinds of typical space missions involving relative orbital control: spacecraft hovering, rendezvous, and formation flying. Subsequently, it puts forward designs for both open-loop and closed-loop control schemes propelled or augmented by the geomagnetic Lorentz force. These control schemes are entirely novel and represent a significantly departure from previous approaches.

A dusty road connecting Saturn and its rings - preliminary results from Cassini Cosmic Dust Analyser during the Grand Finale Mission

Science.gov (United States)

Hsu, S.; Burton, M. E.; Horanyi, M.; Kempf, S.; Khawaja, N.; Moragas-Klostermeyer, G.; Postberg, F.; Schirdenwahn, D.; Seiss, M.; Schmidt, J.; Spahn, F.; Srama, R.

2017-12-01

The Cosmic Dust Analyzer observations during the Cassini Grand Finale Orbits were designed for the in situ characterization of Saturn's ring composition and to study their interaction with the host planet. It is found that the gap between the inner most D ring and Saturn is almost free of larger, micron-sized dust grains but rich in nanodust particles (radius smaller than 100 nm) that only become detectable by CDA because of the high spacecraft speed of 30 km/s through this region. Regarding the grain composition, while the majority of CDA mass spectra recorded during this phase are too faint to be individually calibrated, two types of mass spectra have been identified - water ice and silicates. These two types of grains were detected at different locations with respect to the ring plane, indicating that there are compositional differences across the rings. As for the dynamics, the observations confirm the transport of charged nanodust from the main rings along magnetic field lines to the planet, as previously proposed. The agreement between the simulated density profile and the observation strongly suggests nanodust as a pathway of ring-planet interaction associated with both exogenous (e.g., impactor ejecta) and endogenous (ionospheric plasma charging) processes. CDA measurements do not indicate significant temporal variation during the the Grand Finale orbits. The measured flux corresponds to a mass transport of < 0.1 kg/sec from the main rings to Saturn in the form of nanodust, with most of the deposition occurring near the equator.

Cassini Spacecraft Uncertainty Analysis Data and Methodology Review and Update/Volume 1: Updated Parameter Uncertainty Models for the Consequence Analysis

Energy Technology Data Exchange (ETDEWEB)

WHEELER, TIMOTHY A.; WYSS, GREGORY D.; HARPER, FREDERICK T.

2000-11-01

Uncertainty distributions for specific parameters of the Cassini General Purpose Heat Source Radioisotope Thermoelectric Generator (GPHS-RTG) Final Safety Analysis Report consequence risk analysis were revised and updated. The revisions and updates were done for all consequence parameters for which relevant information exists from the joint project on Probabilistic Accident Consequence Uncertainty Analysis by the United States Nuclear Regulatory Commission and the Commission of European Communities.

Spacecraft electrical power subsystem: Failure behavior, reliability, and multi-state failure analyses

International Nuclear Information System (INIS)

Kim, So Young; Castet, Jean-Francois; Saleh, Joseph H.

2012-01-01

This article investigates the degradation and failure behavior of spacecraft electrical power subsystem (EPS) on orbit. First, this work provides updated statistical reliability and multi-state failure analyses of spacecraft EPS and its different constituents, namely the batteries, the power distribution, and the solar arrays. The EPS is shown to suffer from infant mortality and to be a major driver of spacecraft unreliability. Over 25% of all spacecraft failures are the result of EPS failures. As a result, satellite manufacturers may wish to pursue targeted improvement to this subsystem, either through better testing or burn-in procedures, better design or parts selection, or additional redundancy. Second, this work investigates potential differences in the EPS degradation and failure behavior for spacecraft in low earth orbits (LEO) and geosynchronous orbits (GEO). This analysis was motivated by the recognition that the power/load cycles and the space environment are significantly different in LEO and GEO, and as such, they may result in different failure behavior for the EPS in these two types of orbits. The results indicate, and quantify the extent to which, the EPS fails differently in LEO and GEO, both in terms of frequency and severity of failure events. A casual summary of the findings can be stated as follows: the EPS fails less frequently but harder (with fatal consequences to the spacecraft) in LEO than in GEO.

Pursuit/evasion in orbit

Science.gov (United States)

Kelley, H. J.; Cliff, E. M.; Lutze, F. H.

1981-01-01

Maneuvers available to a spacecraft having sufficient propellant to escape an antisatellite satellite (ASAT) attack are examined. The ASAT and the evading spacecraft are regarded as being in circular orbits, and equations of motion are developed for the ASAT to commence a two-impulse maneuver sequence. The ASAT employs thrust impulses which yield a minimum-time-to-rendezvous, considering available fuel. Optimal evasion is shown to involve only in-plane maneuvers, and begins as soon as the ASAT launch information is gathered and thrust activation can be initiated. A closest approach, along with a maximum evasion by the target spacecraft, is calculated to be 14,400 ft. Further research to account for ASATs in parking orbit and for generalization of a continuous control-modeled differential game is indicated.

Propulsion Trade Studies for Spacecraft Swarm Mission Design

Science.gov (United States)

Dono, Andres; Plice, Laura; Mueting, Joel; Conn, Tracie; Ho, Michael

2018-01-01

Spacecraft swarms constitute a challenge from an orbital mechanics standpoint. Traditional mission design involves the application of methodical processes where predefined maneuvers for an individual spacecraft are planned in advance. This approach does not scale to spacecraft swarms consisting of many satellites orbiting in close proximity; non-deterministic maneuvers cannot be preplanned due to the large number of units and the uncertainties associated with their differential deployment and orbital motion. For autonomous small sat swarms in LEO, we investigate two approaches for controlling the relative motion of a swarm. The first method involves modified miniature phasing maneuvers, where maneuvers are prescribed that cancel the differential delta V of each CubeSat's deployment vector. The second method relies on artificial potential functions (APFs) to contain the spacecraft within a volumetric boundary and avoid collisions. Performance results and required delta V budgets are summarized, indicating that each method has advantages and drawbacks for particular applications. The mini phasing maneuvers are more predictable and sustainable. The APF approach provides a more responsive and distributed performance, but at considerable propellant cost. After considering current state of the art CubeSat propulsion systems, we conclude that the first approach is feasible, but the modified APF method of requires too much control authority to be enabled by current propulsion systems.

A geometric model of a V-slit Sun sensor correcting for spacecraft wobble

Science.gov (United States)

Mcmartin, W. P.; Gambhir, S. S.

1994-01-01

A V-Slit sun sensor is body-mounted on a spin-stabilized spacecraft. During injection from a parking or transfer orbit to some final orbit, the spacecraft may not be dynamically balanced. This may result in wobble about the spacecraft spin axis as the spin axis may not be aligned with the spacecraft's axis of symmetry. While the widely used models in Spacecraft Attitude Determination and Control, edited by Wertz, correct for separation, elevation, and azimuthal mounting biases, spacecraft wobble is not taken into consideration. A geometric approach is used to develop a method for measurement of the sun angle which corrects for the magnitude and phase of spacecraft wobble. The algorithm was implemented using a set of standard mathematical routines for spherical geometry on a unit sphere.

Cassini MIMI Close-Up of Saturn Energetic Particles: Low Altitude Trapped Radiation, Auroral Ion Acceleration, and Interchange Flow Channels

Science.gov (United States)

Mitchell, D. G.; Krimigis, S. M.; Krupp, N.; Paranicas, C.; Roussos, E.; Kollmann, P.

2017-12-01

We present observations from the final orbits of the Cassini Mission at Saturn by the Magnetospheric Imaging Instrument (MIMI). Crossing inside the D-Ring at the equator and just above Saturn's atmosphere, these orbits covered regions never visited previously in the mission. Highlights include the confirmation of an inner radiation belt analogous to the inner radiation belt at Earth by the Low Energy Magnetospheric Measurement System (LEMMS), with surprising twists—Saturn's D-ring material appears to be a source for these particles. Details will be presented in another session. The Grand Finale orbits also afforded a close-up view of the auroral ion acceleration regions by the Ion and Neutral Camera (INCA). Ionospheric ions at the base of auroral field lines are accelerated perpendicular to the magnetic field to 10's and 100's of keV, and charge exchange with exospheric neutrals to be emitted as energetic neutral atoms and images by INCA. We show that this acceleration region lies at about 0.1 Rs. Another feature seen previously in the mission but imaged with greater resolution is a flow channel associated with interchange motion in the middle magnetosphere. We show this feature to extend over several Saturn radii in the radial direction, and over about 2 Saturn radii azimuthally. Additional data have been received since the writing of this abstract and before Cassini's plunge into the atmosphere on September 15, so additional features may be presented.

Charging in the environment of large spacecraft

International Nuclear Information System (INIS)

Lai, S.T.

1993-01-01

This paper discusses some potential problems of spacecraft charging as a result of interactions between a large spacecraft, such as the Space Station, and its environment. Induced electric field, due to VXB effect, may be important for large spacecraft at low earth orbits. Differential charging, due to different properties of surface materials, may be significant when the spacecraft is partly in sunshine and partly in shadow. Triple-root potential jump condition may occur because of differential charging. Sudden onset of severe differential charging may occur when an electron or ion beam is emitted from the spacecraft. The beam may partially return to the ''hot spots'' on the spacecraft. Wake effects, due to blocking of ambient ion trajectories, may result in an undesirable negative potential region in the vicinity of a large spacecraft. Outgassing and exhaust may form a significant spacecraft induced environment; ionization may occur. Spacecraft charging and discharging may affect the electronic components on board

The tides of Titan.

Science.gov (United States)

Iess, Luciano; Jacobson, Robert A; Ducci, Marco; Stevenson, David J; Lunine, Jonathan I; Armstrong, John W; Asmar, Sami W; Racioppa, Paolo; Rappaport, Nicole J; Tortora, Paolo

2012-07-27

We have detected in Cassini spacecraft data the signature of the periodic tidal stresses within Titan, driven by the eccentricity (e = 0.028) of its 16-day orbit around Saturn. Precise measurements of the acceleration of Cassini during six close flybys between 2006 and 2011 have revealed that Titan responds to the variable tidal field exerted by Saturn with periodic changes of its quadrupole gravity, at about 4% of the static value. Two independent determinations of the corresponding degree-2 Love number yield k(2) = 0.589 ± 0.150 and k(2) = 0.637 ± 0.224 (2σ). Such a large response to the tidal field requires that Titan's interior be deformable over time scales of the orbital period, in a way that is consistent with a global ocean at depth.

LDEF materials results for spacecraft applications: Executive summary

Science.gov (United States)

Whitaker, A. F.; Dooling, D.

1995-03-01

To address the challenges of space environmental effects, NASA designed the Long Duration Exposure Facility (LDEF) for an 18-month mission to expose thousands of samples of candidate materials that might be used on a space station or other orbital spacecraft. LDEF was launched in April 1984 and was to have been returned to Earth in 1985. Changes in mission schedules postponed retrieval until January 1990, after 69 months in orbit. Analyses of the samples recovered from LDEF have provided spacecraft designers and managers with the most extensive data base on space materials phenomena. Many LDEF samples were greatly changed by extended space exposure. Among even the most radially altered samples, NASA and its science teams are finding a wealth of surprising conclusions and tantalizing clues about the effects of space on materials. Many were discussed at the first two LDEF results conferences and subsequent professional papers. The LDEF Materials Results for Spacecraft Applications Conference was convened in Huntsville to discuss implications for spacecraft design. Already, paint and thermal blanket selections for space station and other spacecraft have been affected by LDEF data. This volume synopsizes those results.

Orbital debris: a technical assessment

National Research Council Canada - National Science Library

Committee on Space Debris, National Research Council

..., and other debris created as a byproduct of space operations. Orbital Debris examines the methods we can use to characterize orbital debris, estimates the magnitude of the debris population, and assesses the hazard that this population poses to spacecraft...

Spacecraft Thermal Management

Science.gov (United States)

Hurlbert, Kathryn Miller

2009-01-01

In the 21st century, the National Aeronautics and Space Administration (NASA), the Russian Federal Space Agency, the National Space Agency of Ukraine, the China National Space Administration, and many other organizations representing spacefaring nations shall continue or newly implement robust space programs. Additionally, business corporations are pursuing commercialization of space for enabling space tourism and capital business ventures. Future space missions are likely to include orbiting satellites, orbiting platforms, space stations, interplanetary vehicles, planetary surface missions, and planetary research probes. Many of these missions will include humans to conduct research for scientific and terrestrial benefits and for space tourism, and this century will therefore establish a permanent human presence beyond Earth s confines. Other missions will not include humans, but will be autonomous (e.g., satellites, robotic exploration), and will also serve to support the goals of exploring space and providing benefits to Earth s populace. This section focuses on thermal management systems for human space exploration, although the guiding principles can be applied to unmanned space vehicles as well. All spacecraft require a thermal management system to maintain a tolerable thermal environment for the spacecraft crew and/or equipment. The requirements for human rating and the specified controlled temperature range (approximately 275 K - 310 K) for crewed spacecraft are unique, and key design criteria stem from overall vehicle and operational/programatic considerations. These criteria include high reliability, low mass, minimal power requirements, low development and operational costs, and high confidence for mission success and safety. This section describes the four major subsystems for crewed spacecraft thermal management systems, and design considerations for each. Additionally, some examples of specialized or advanced thermal system technologies are presented

Special Semaphore Scheme for UHF Spacecraft Communications

Science.gov (United States)

Butman, Stanley; Satorius, Edgar; Ilott, Peter

2006-01-01

A semaphore scheme has been devised to satisfy a requirement to enable ultrahigh- frequency (UHF) radio communication between a spacecraft descending from orbit to a landing on Mars and a spacecraft, in orbit about Mars, that relays communications between Earth and the lander spacecraft. There are also two subsidiary requirements: (1) to use UHF transceivers, built and qualified for operation aboard the spacecraft that operate with residual-carrier binary phase-shift-keying (BPSK) modulation at a selectable data rate of 8, 32, 128, or 256 kb/s; and (2) to enable low-rate signaling even when received signals become so weak as to prevent communication at the minimum BPSK rate of 8 kHz. The scheme involves exploitation of Manchester encoding, which is used in conjunction with residual-carrier modulation to aid the carrier-tracking loop. By choosing various sequences of 1s, 0s, or 1s alternating with 0s to be fed to the residual-carrier modulator, one would cause the modulator to generate sidebands at a fundamental frequency of 4 or 8 kHz and harmonics thereof. These sidebands would constitute the desired semaphores. In reception, the semaphores would be detected by a software demodulator.

Lifetime of a spacecraft around a synchronous system of asteroids using a dipole model

Science.gov (United States)

dos Santos, Leonardo Barbosa Torres; de Almeida Prado, Antonio F. Bertachini; Sanchez, Diogo Merguizo

2017-11-01

Space missions allow us to expand our knowledge about the origin of the solar system. It is believed that asteroids and comets preserve the physical characteristics from the time that the solar system was created. For this reason, there was an increase of missions to asteroids in the past few years. To send spacecraft to asteroids or comets is challenging, since these objects have their own characteristics in several aspects, such as size, shape, physical properties, etc., which are often only discovered after the approach and even after the landing of the spacecraft. These missions must be developed with sufficient flexibility to adjust to these parameters, which are better determined only when the spacecraft reaches the system. Therefore, conducting a dynamic investigation of a spacecraft around a multiple asteroid system offers an extremely rich environment. Extracting accurate information through analytical approaches is quite challenging and requires a significant number of restrictive assumptions. For this reason, a numerical approach to the dynamics of a spacecraft in the vicinity of a binary asteroid system is offered in this paper. In the present work, the equations of the Restricted Synchronous Four-Body Problem (RSFBP) are used to model a binary asteroid system. The main objective of this work is to construct grids of initial conditions, which relates semi-major axis and eccentricity, in order to quantify the lifetime of a spacecraft when released close to the less massive body of the binary system (modeled as a rotating mass dipole). We performed an analysis of the lifetime of the spacecraft considering several mass ratios of a binary system of asteroids and investigating the behavior of a spacecraft in the vicinity of this system. We analyze direct and retrograde orbits. This study investigated orbits that survive for at least 500 orbital periods of the system (which is approximately one year), then not colliding or escaping from the system during this

Formation flying for electric sails in displaced orbits. Part I: Geometrical analysis

Science.gov (United States)

Wang, Wei; Mengali, Giovanni; Quarta, Alessandro A.; Yuan, Jianping

2017-09-01

We present a geometrical methodology for analyzing the formation flying of electric solar wind sail based spacecraft that operate in heliocentric, elliptic, displaced orbits. The spacecraft orbit is maintained by adjusting its propulsive acceleration modulus, whose value is estimated using a thrust model that takes into account a variation of the propulsive performance with the sail attitude. The properties of the relative motion of the spacecraft are studied in detail and a geometrical solution is obtained in terms of relative displaced orbital elements, assumed to be small quantities. In particular, for the small eccentricity case (i.e. for a near-circular displaced orbit), the bounds characterized by the extreme values of relative distances are analytically calculated, thus providing an useful mathematical tool for preliminary design of the spacecraft formation structure.

Mars Molniya Orbit Atmospheric Resource Mining

Science.gov (United States)

Mueller, Robert P.; Braun, Robert D.; Sibille, Laurent; Sforzo, Brandon; Gonyea, Keir; Ali, Hisham

2016-01-01

This NIAC (NASA Advanced Innovative Concepts) work will focus on Mars and will build on previous efforts at analyzing atmospheric mining at Earth and the outer solar system. Spacecraft systems concepts will be evaluated and traded, to assess feasibility. However the study will primarily examine the architecture and associated missions to explore the closure, constraints and critical parameters through sensitivity studies. The Mars atmosphere consists of 95.5 percent CO2 gas which can be converted to methane fuel (CH4) and Oxidizer (O2) for chemical rocket propulsion, if hydrogen is transported from electrolyzed water on the Mars surface or from Earth. By using a highly elliptical Mars Molniya style orbit, the CO2 atmosphere can be scooped, ram-compressed and stored while the spacecraft dips into the Mars atmosphere at periapsis. Successive orbits result in additional scooping of CO2 gas, which also serves to aerobrake the spacecraft, resulting in a decaying Molniya orbit.

Gravity Probe B spacecraft description

International Nuclear Information System (INIS)

Bennett, Norman R; Burns, Kevin; Katz, Russell; Kirschenbaum, Jon; Mason, Gary; Shehata, Shawky

2015-01-01

The Gravity Probe B spacecraft, developed, integrated, and tested by Lockheed Missiles and Space Company and later Lockheed Martin Corporation, consisted of structures, mechanisms, command and data handling, attitude and translation control, electrical power, thermal control, flight software, and communications. When integrated with the payload elements, the integrated system became the space vehicle. Key requirements shaping the design of the spacecraft were: (1) the tight mission timeline (17 months, 9 days of on-orbit operation), (2) precise attitude and translational control, (3) thermal protection of science hardware, (4) minimizing aerodynamic, magnetic, and eddy current effects, and (5) the need to provide a robust, low risk spacecraft. The spacecraft met all mission requirements, as demonstrated by dewar lifetime meeting specification, positive power and thermal margins, precision attitude control and drag-free performance, reliable communications, and the collection of more than 97% of the available science data. (paper)

Project Overview of the Naval Postgraduate School Spacecraft Architecture and Technology Demonstration Experiment

National Research Council Canada - National Science Library

Reuer, Charles

2001-01-01

The Naval Postgraduate School's current attempt at getting another spacecraft into orbit is focusing on Naval Postgraduate School Spacecraft Architecture and Technology Demonstration Experiment (NPSAT1...

A Shaftless Magnetically Levitated Multifunctional Spacecraft Flywheel Storage System

Science.gov (United States)

Stevens, Ken; Thornton, Richard; Clark, Tracy; Beaman, Bob G.; Dennehy, Neil; Day, John H. (Technical Monitor)

2002-01-01

Presently many types of spacecraft use a Spacecraft Attitude Control System (ACS) with momentum wheels for steering and electrochemical batteries to provide electrical power for the eclipse period of the spacecraft orbit. Future spacecraft will use Flywheels for combined use in ACS and Energy Storage. This can be done by using multiple wheels and varying the differential speed for ACS and varying the average speed for energy storage and recovery. Technology in these areas has improved since the 1990s so it is now feasible for flywheel systems to emerge from the laboratory for spacecraft use. This paper describes a new flywheel system that can be used for both ACS and energy storage. Some of the possible advantages of a flywheel system are: lower total mass and volume, higher efficiency, less thermal impact, improved satellite integration schedule and complexity, simplified satellite orbital operations, longer life with lower risk, less pointing jitter, and greater capability for high-rate slews. In short, they have the potential to enable new types of missions and provide lower cost. Two basic types of flywheel configurations are the Flywheel Energy Storage System (FESS) and the Integrated Power and Attitude Control System (IPACS).

An algorithm for enhanced formation flying of satellites in low earth orbit

Science.gov (United States)

Folta, David C.; Quinn, David A.

1998-01-01

With scientific objectives for Earth observation programs becoming more ambitious and spacecraft becoming more autonomous, the need for innovative technical approaches on the feasibility of achieving and maintaining formations of spacecraft has come to the forefront. The trend to develop small low-cost spacecraft has led many scientists to recognize the advantage of flying several spacecraft in formation to achieve the correlated instrument measurements formerly possible only by flying many instruments on a single large platform. Yet, formation flying imposes additional complications on orbit maintenance, especially when each spacecraft has its own orbit requirements. However, advances in automation and technology proposed by the Goddard Space Flight Center (GSFC) allow more of the burden in maneuver planning and execution to be placed onboard the spacecraft, mitigating some of the associated operational concerns. The purpose of this paper is to present GSFC's Guidance, Navigation, and Control Center's (GNCC) algorithm for Formation Flying of the low earth orbiting spacecraft that is part of the New Millennium Program (NMP). This system will be implemented as a close-loop flight code onboard the NMP Earth Orbiter-1 (EO-1) spacecraft. Results of this development can be used to determine the appropriateness of formation flying for a particular case as well as operational impacts. Simulation results using this algorithm integrated in an autonomous `fuzzy logic' control system called AutoCon™ are presented.

Spacecraft attitude determination using the earth's magnetic field

Science.gov (United States)

Simpson, David G.

1989-01-01

A method is presented by which the attitude of a low-Earth orbiting spacecraft may be determined using a vector magnetometer, a digital Sun sensor, and a mathematical model of the Earth's magnetic field. The method is currently being implemented for the Solar Maximum Mission spacecraft (as a backup for the failing star trackers) as a way to determine roll gyro drift.

The Effect of Air Drag in Optimal Power-Limited Rendezvous Between Coplanar Low-Earth Orbits

Directory of Open Access Journals (Sweden)

Gil-Young Maeng

1998-06-01

Full Text Available The effect of air drag was researched when a low-earth orbit spacecraft using power-limited thruster rendezvoused another low-earth orbit spacecraft. The air density was assumed to decrease exponentially. The radius of parking orbit was 6655.935 km and that of target orbit was 7321.529 km. From the trajectories of active vehicles, the fuel consumption and the magnitude of thrust acceleration, we could conclude that the effect of air drag had to be considered in fuel optimal rendezvous problem between low-earth orbit spacecrafts. In multiple-revolution rendezvous case, the air drag was more effective.

Orbital Express fluid transfer demonstration system

Science.gov (United States)

Rotenberger, Scott; SooHoo, David; Abraham, Gabriel

2008-04-01

Propellant resupply of orbiting spacecraft is no longer in the realm of high risk development. The recently concluded Orbital Express (OE) mission included a fluid transfer demonstration that operated the hardware and control logic in space, bringing the Technology Readiness Level to a solid TRL 7 (demonstration of a system prototype in an operational environment). Orbital Express (funded by the Defense Advanced Research Projects Agency, DARPA) was launched aboard an Atlas-V rocket on March 9th, 2007. The mission had the objective of demonstrating technologies needed for routine servicing of spacecraft, namely autonomous rendezvous and docking, propellant resupply, and orbital replacement unit transfer. The demonstration system used two spacecraft. A servicing vehicle (ASTRO) performed multiple dockings with the client (NextSat) spacecraft, and performed a variety of propellant transfers in addition to exchanges of a battery and computer. The fluid transfer and propulsion system onboard ASTRO, in addition to providing the six degree-of-freedom (6 DOF) thruster system for rendezvous and docking, demonstrated autonomous transfer of monopropellant hydrazine to or from the NextSat spacecraft 15 times while on orbit. The fluid transfer system aboard the NextSat vehicle was designed to simulate a variety of client systems, including both blowdown pressurization and pressure regulated propulsion systems. The fluid transfer demonstrations started with a low level of autonomy, where ground controllers were allowed to review the status of the demonstration at numerous points before authorizing the next steps to be performed. The final transfers were performed at a full autonomy level where the ground authorized the start of a transfer sequence and then monitored data as the transfer proceeded. The major steps of a fluid transfer included the following: mate of the coupling, leak check of the coupling, venting of the coupling, priming of the coupling, fluid transfer, gauging

Displaced Electric Sail Orbits Design and Transition Trajectory Optimization

Directory of Open Access Journals (Sweden)

Naiming Qi

2014-01-01

Full Text Available Displaced orbits for spacecraft propelled by electric sails are investigated as an alternative to the use of solar sails. The orbital dynamics of electric sails based spacecraft are studied within a spherical coordinate system, which permits finding the solutions of displaced electric sail orbits and optimize transfer trajectory. Transfer trajectories from Earth's orbit to displaced orbit are also studied in an optimal framework, by using genetic algorithm and Gauss pseudospectral method. The initial guesses for the state and control histories used in the Gauss pseudospectral method are interpolated from the best solution of a genetic algorithm. Numerical simulations show that the electric sail is able to perform the transfer from Earth’s orbit to displaced orbit in acceptable time, and the hybrid optimization method has the capability to search the feasible and optimal solution without any initial value guess.

Spacecraft Charging Modeling -- Nascap-2k 2014 Annual Report

Science.gov (United States)

2014-09-19

appears to work similarly in Internet Explorer, FireFox , and Opera, but fails in Safari and Chrome. Note that the SEE Spacecraft Charging Handbook is... Characteristics of Spacecraft Charging in Low Earth Orbit, J Geophys Res. 11 7, doi: 10.1029/20 11JA016875, 2012. 2 M. Cho, K. Saito, T. Hamanaga, Data

Mercury Orbiter: Report of the Science Working Team

Science.gov (United States)

Belcher, John W.; Slavin, James A.; Armstrong, Thomas P.; Farquhar, Robert W.; Akasofu, Syun I.; Baker, Daniel N.; Cattell, Cynthia A.; Cheng, Andrew F.; Chupp, Edward L.; Clark, Pamela E.

1991-01-01

The results are presented of the Mercury Orbiter Science Working Team which held three workshops in 1988 to 1989 under the auspices of the Space Physics and Planetary Exploration Divisions of NASA Headquarters. Spacecraft engineering and mission design studies at the Jet Propulsion Lab were conducted in parallel with this effort and are detailed elsewhere. The findings of the engineering study, summarized herein, indicate that spin stabilized spacecraft carrying comprehensive particles and fields experiments and key planetology instruments in high elliptical orbits can survive and function in Mercury orbit without costly sun shields and active cooling systems.

Evaluation of spacecraft technology programs (effects on communication satellite business ventures), volume 1

Science.gov (United States)

Greenburg, J. S.; Gaelick, C.; Kaplan, M.; Fishman, J.; Hopkins, C.

1985-01-01

Commercial organizations as well as government agencies invest in spacecraft (S/C) technology programs that are aimed at increasing the performance of communications satellites. The value of these programs must be measured in terms of their impacts on the financial performane of the business ventures that may ultimately utilize the communications satellites. An economic evaluation and planning capability was developed and used to assess the impact of NASA on-orbit propulsion and space power programs on typical fixed satellite service (FSS) and direct broadcast service (DBS) communications satellite business ventures. Typical FSS and DBS spin and three-axis stabilized spacecraft were configured in the absence of NASA technology programs. These spacecraft were reconfigured taking into account the anticipated results of NASA specified on-orbit propulsion and space power programs. In general, the NASA technology programs resulted in spacecraft with increased capability. The developed methodology for assessing the value of spacecraft technology programs in terms of their impact on the financial performance of communication satellite business ventures is described. Results of the assessment of NASA specified on-orbit propulsion and space power technology programs are presented for typical FSS and DBS business ventures.

On the atmospheric drag in orbit determination for low Earth orbit

Science.gov (United States)

Tang, Jingshi; Liu, Lin; Miao, Manqian

2012-07-01

The atmosphere model is always a major limitation for low Earth orbit (LEO) in orbit prediction and determination. The accelerometer can work around the non-gravitational perturbations in orbit determination, but it helps little to improve the atmosphere model or to predict the orbit. For certain satellites, there may be some specific software to handle the orbit problem. This solution can improve the orbit accuracy for both prediction and determination, yet it always contains empirical terms and is exclusive for certain satellites. This report introduces a simple way to handle the atmosphere drag for LEO, which does not depend on instantaneous atmosphere conditions and improves accuracy of predicted orbit. This approach, which is based on mean atmospheric density, is supported by two reasons. One is that although instantaneous atmospheric density is very complicated with time and height, the major pattern is determined by the exponential variation caused by hydrostatic equilibrium and periodic variation caused by solar radiation. The mean density can include the major variations while neglect other minor details. The other reason is that the predicted orbit is mathematically the result from integral and the really determinant factor is the mean density instead of instantaneous density for every time and spot. Using the mean atmospheric density, which is mainly determined by F10.7 solar flux and geomagnetic index, can be combined into an overall parameter B^{*} = C_{D}(S/m)ρ_{p_{0}}. The combined parameter contains several less accurate parameters and can be corrected during orbit determination. This approach has been confirmed in various LEO computations and an example is given below using Tiangong-1 spacecraft. Precise orbit determination (POD) is done using one-day GPS positioning data without any accurate a-priori knowledge on spacecraft or atmosphere conditions. Using the corrected initial state vector of the spacecraft and the parameter B^* from POD, the

REQUIREMENTS FOR IMAGE QUALITY OF EMERGENCY SPACECRAFTS

Directory of Open Access Journals (Sweden)

A. I. Altukhov

2015-05-01

Full Text Available The paper deals with the method for formation of quality requirements to the images of emergency spacecrafts. The images are obtained by means of remote sensing of near-earth space orbital deployment in the visible range. of electromagnetic radiation. The method is based on a joint taking into account conditions of space survey, characteristics of surveillance equipment, main design features of the observed spacecrafts and orbital inspection tasks. Method. Quality score is the predicted linear resolution image that gives the possibility to create a complete view of pictorial properties of the space image obtained by electro-optical system from the observing satellite. Formulation of requirements to the numerical value of this indicator is proposed to perform based on the properties of remote sensing system, forming images in the conditions of outer space, and the properties of the observed emergency spacecraft: dimensions, platform construction of the satellite, on-board equipment placement. For method implementation the authors have developed a predictive model of requirements to a linear resolution for images of emergency spacecrafts, making it possible to select the intervals of space shooting and get the satellite images required for quality interpretation. Main results. To verify the proposed model functionality we have carried out calculations of the numerical values for the linear resolution of the image, ensuring the successful task of determining the gross structural damage of the spacecrafts and identifying changes in their spatial orientation. As input data were used with dimensions and geometric primitives corresponding to the shape of deemed inspected spacecrafts: Resurs-P", "Canopus-B", "Electro-L". Numerical values of the linear resolution images have been obtained, ensuring the successful task solution for determining the gross structural damage of spacecrafts.

Earth to Orbit Beamed Energy Experiment

Science.gov (United States)

Johnson, Les; Montgomery, Edward E.

2017-01-01

As a means of primary propulsion, beamed energy propulsion offers the benefit of offloading much of the propulsion system mass from the vehicle, increasing its potential performance and freeing it from the constraints of the rocket equation. For interstellar missions, beamed energy propulsion is arguably the most viable in the near- to mid-term. A near-term demonstration showing the feasibility of beamed energy propulsion is necessary and, fortunately, feasible using existing technologies. Key enabling technologies are large area, low mass spacecraft and efficient and safe high power laser systems capable of long distance propagation. NASA is currently developing the spacecraft technology through the Near Earth Asteroid Scout solar sail mission and has signed agreements with the Planetary Society to study the feasibility of precursor laser propulsion experiments using their LightSail-2 solar sail spacecraft. The capabilities of Space Situational Awareness assets and the advanced analytical tools available for fine resolution orbit determination now make it possible to investigate the practicalities of an Earth-to-orbit Beamed Energy eXperiment (EBEX) - a demonstration at delivered power levels that only illuminate a spacecraft without causing damage to it. The degree to which this can be expected to produce a measurable change in the orbit of a low ballistic coefficient spacecraft is investigated. Key system characteristics and estimated performance are derived for a near term mission opportunity involving the LightSail-2 spacecraft and laser power levels modest in comparison to those proposed previously. While the technology demonstrated by such an experiment is not sufficient to enable an interstellar precursor mission, if approved, then it would be the next step toward that goal.

ExoMars Trace Gas Orbiter provides atmospheric data during Aerobraking into its final orbit

Science.gov (United States)

Svedhem, Hakan; Vago, Jorge L.; Bruinsma, Sean; Müller-Wodarg, Ingo; ExoMars 2016 Team

2017-10-01

After the arrival of the Trace Gas Orbiter (TGO) at Mars on 19 October 2016 a number of initial orbit change manoeuvres were executed and the spacecraft was put in an orbit with a 24 hour period and 74 degrees inclination. The spacecraft and its four instruments were thoroughly checked out after arrival and a few measurements and images were taken in November 2016 and in Feb-March 2017. The solar occultation observations have however not yet been possible due to lack of the proper geometry.On 15 March a long period of aerobraking to reach the final 400km semi-circular frozen orbit (370x430km, with a fixed pericentre latitude). This orbit is optimised for the payload observations and for the communication relay with the ExoMars Rover, due to arrive in 2021.The aerobraking is proceeding well and the final orbit is expected to be reached in April 2018. A large data set is being acquired for the upper atmosphere of Mars, from the limit of the sensitivity of the accelerometer, down to lowest altitude of the aerobraking at about 105km. Initial analysis has shown a highly variable atmosphere with a slightly lower density then predicted by existing models. Until the time of the abstract writing no dust storms have been observed.The ExoMars programme is a joint activity by the European Space Agency(ESA) and ROSCOSMOS, Russia. ESA is providing the TGO spacecraft and Schiaparelli (EDM) and two of the TGO instruments and ROSCOSMOS is providing the Proton launcher and the other two TGO instruments. After the arrival of the ExoMars 2020 mission, consisting of a Rover and a Surface platform also launched by a Proton rocket, the TGO will handle the communication between the Earth and the Rover and Surface Platform through its (NASA provided) UHF communication system.

System concepts and design examples for optical communication with planetary spacecraft

Science.gov (United States)

Lesh, James R.

Systems concepts for optical communication with future deep-space (planetary) spacecraft are described. These include not only the optical transceiver package aboard the distant spacecraft, but the earth-vicinity optical-communications receiving station as well. Both ground-based, and earth-orbiting receivers are considered. Design examples for a number of proposed or potential deep-space missions are then presented. These include an orbital mission to Saturn, a Lander and Rover mission to Mars, and an astronomical mission to a distance of 1000 astronomical units.

Tidally modulated eruptions on Enceladus: Cassini ISS observations and models

International Nuclear Information System (INIS)

Nimmo, Francis; Porco, Carolyn; Mitchell, Colin

2014-01-01

We use images acquired by the Cassini Imaging Science Subsystem (ISS) to investigate the temporal variation of the brightness and height of the south polar plume of Enceladus. The plume's brightness peaks around the moon's apoapse, but with no systematic variation in scale height with either plume brightness or Enceladus' orbital position. We compare our results, both alone and supplemented with Cassini near-infrared observations, with predictions obtained from models in which tidal stresses are the principal control of the eruptive behavior. There are three main ways of explaining the observations: (1) the activity is controlled by right-lateral strike slip motion; (2) the activity is driven by eccentricity tides with an apparent time delay of about 5 hr; (3) the activity is driven by eccentricity tides plus a 1:1 physical libration with an amplitude of about 0.°8 (3.5 km). The second hypothesis might imply either a delayed eruptive response, or a dissipative, viscoelastic interior. The third hypothesis requires a libration amplitude an order of magnitude larger than predicted for a solid Enceladus. While we cannot currently exclude any of these hypotheses, the third, which is plausible for an Enceladus with a subsurface ocean, is testable by using repeat imaging of the moon's surface. A dissipative interior suggests that a regional background heat source should be detectable. The lack of a systematic variation in plume scale height, despite the large variations in plume brightness, is plausibly the result of supersonic flow; the details of the eruption process are yet to be understood.

Dione and Rhea seasonal exospheres revealed by Cassini CAPS and INMS

Science.gov (United States)

Teolis, B. D.; Waite, J. H.

2016-07-01

A Dione O2 and CO2 exosphere of similar composition and density to Rhea's is confirmed by Cassini spacecraft Ion Neutral Mass Spectrometer (INMS) flyby data. INMS results from three Dione and two Rhea flybys show exospheric spatial and temporal variability indicative of seasonal exospheres, modulated by winter polar gas adsorption and desorption at the equinoxes. Cassini Plasma Spectrometer (CAPS) pickup ion fluxes also show exospheric structure and evolution at Rhea consistent with INMS, after taking into consideration the anticipated charge exchange, electron impact, and photo-ionization rates. Data-model comparisons show the exospheric evolution to be consistent with polar frost diffusion into the surface regolith, which limits surface exposure and loss of the winter frost cap by sputtering. Implied O2 source rates of ∼45(7) × 1021 s-1 at Dione(Rhea) are ∼50(300) times less than expected from known O2 radiolysis yields from ion-irradiated pure water ice measured in the laboratory, ruling out secondary sputtering as a major exospheric contributor, and implying a nanometer scale surface refractory lag layer consisting of concentrated carbonaceous impurities. We estimate ∼30:1(2:1) relative O2:CO2 source rates at Dione(Rhea), consistent with a stoichiometric bulk composition below the lag layer of 0.01(0.13) C atoms per H2O molecule, deriving from endogenic constituents, implanted micrometeoritic organics, and (in particular at Dione) exogenous H2O delivery by E-ring grains. Impact deposition, gardening and vaporization may thereby control the global O2 source rates by fresh H2O ice exposure to surface radiolysis and trapped oxidant ejection.

Short rendezvous missions for advanced Russian human spacecraft

Science.gov (United States)

Murtazin, Rafail F.; Budylov, Sergey G.

2010-10-01

The two-day stay of crew in a limited inhabited volume of the Soyuz-TMA spacecraft till docking to ISS is one of the most stressful parts of space flight. In this paper a number of possible ways to reduce the duration of the free flight phase are considered. The duration is defined by phasing strategy that is necessary for reduction of the phase angle between the chaser and target spacecraft. Some short phasing strategies could be developed. The use of such strategies creates more comfortable flight conditions for crew thanks to short duration and additionally it allows saving spacecraft's life support resources. The transition from the methods of direct spacecraft rendezvous using one orbit phasing (first flights of " Vostok" and " Soyuz" vehicles) to the currently used methods of two-day rendezvous mission can be observed in the history of Soviet manned space program. For an advanced Russian human rated spacecraft the short phasing strategy is recommended, which can be considered as a combination between the direct and two-day rendezvous missions. The following state of the art technologies are assumed available: onboard accurate navigation; onboard computations of phasing maneuvers; launch vehicle with high accuracy injection orbit, etc. Some operational requirements and constraints for the strategies are briefly discussed. In order to provide acceptable phase angles for possible launch dates the experience of the ISS altitude profile control can be used. As examples of the short phasing strategies, the following rendezvous missions are considered: direct ascent, short mission with the phasing during 3-7 orbits depending on the launch date (nominal or backup). For each option statistical modeling of the rendezvous mission is fulfilled, as well as an admissible phase angle range, accuracy of target state vector and addition fuel consumption coming out of emergency is defined. In this paper an estimation of pros and cons of all options is conducted.

Global Precipitation Measurement (GPM) Spacecraft Lithium Ion Battery Micro-Cycling Investigation

Science.gov (United States)

Dakermanji, George; Lee, Leonine; Spitzer, Thomas

2016-01-01

The Global Precipitation Measurement (GPM) spacecraft was jointly developed by NASA and JAXA. It is a Low Earth Orbit (LEO) spacecraft launched on February 27, 2014. The power system is a Direct Energy Transfer (DET) system designed to support 1950 watts orbit average power. The batteries use SONY 18650HC cells and consist of three 8s by 84p batteries operated in parallel as a single battery. During instrument integration with the spacecraft, large current transients were observed in the battery. Investigation into the matter traced the cause to the Dual-Frequency Precipitation Radar (DPR) phased array radar which generates cyclical high rate current transients on the spacecraft power bus. The power system electronics interaction with these transients resulted in the current transients in the battery. An accelerated test program was developed to bound the effect, and to assess the impact to the mission.

Secure communications with low-orbit spacecraft using quantum cryptography

Science.gov (United States)

Hughes, Richard J.; Buttler, William T.; Kwiat, Paul G.; Luther, Gabriel G.; Morgan, George L; Nordholt, Jane E.; Peterson, Charles G.; Simmons, Charles M.

1999-01-01

Apparatus and method for secure communication between an earth station and spacecraft. A laser outputs single pulses that are split into preceding bright pulses and delayed attenuated pulses, and polarized. A Pockels cell changes the polarization of the polarized delayed attenuated pulses according to a string of random numbers, a first polarization representing a "1," and a second polarization representing a "0." At the receiving station, a beamsplitter randomly directs the preceding bright pulses and the polarized delayed attenuated pulses onto longer and shorter paths, both terminating in a beamsplitter which directs the preceding bright pulses and a first portion of the polarized delayed attenuated pulses to a first detector, and a second portion of the polarized delayed attenuated pulses to a second detector to generate a key for secure communication between the earth station and the spacecraft.

Return to Venus of AKATSUKI, the Japanese Venus Orbiter

Science.gov (United States)

Nakamura, M.; Iwagami, N.; Satoh, T.; Taguchi, M.; Watanabe, S.; Takahashi, Y.; Imamura, T.; Suzuki, M.; Ueno, M.; Yamazaki, A.; Fukuhara, T.; Yamada, M.; Ishii, N.; Ogohara, K.

2011-12-01

Japanese Venus Climate Orbiter 'AKATSUKI' (PLANET-C) was proposed in 2001 with strong support by international Venus science community and approved as an ISAS mission soon after the proposal. AKATSUKI and ESA's Venus Express complement each other in Venus climate study. Various coordinated observations using the two spacecraft have been planned. Also participating scientists from US have been selected. Its science target is to understand the climate of Venus. The mission life we expected was more than 2 Earth years in Venus orbit. AKATSUKI was successfully launched at 06:58:22JST on May 21, by H-IIA F17. After the separation from H-IIA, the telemetry from AKATSUKI was normally detected by DSN Goldstone station (10:00JST) and the solar cell paddles' expansion was confirmed. AKATSUKI was put into the 3-axis stabilized mode in the initial operation from Uchinoura station and the critical operation was finished at 20:00JST on the same day. The malfunction, which happened during the Venus Orbit Insertion (VOI) on7 Dec, 2010 is as follows. We set all commands on Dec. 5. Attitude control for Venus orbit insertion (VOI) was automatically done on Dec. 6. Orbital maneuver engine (OME) was fired 08:49 JST on Dec. 7. 1min. after firing the spacecraft went into the occultation region and we had no telemetry, but we expected to continuous firing for 12min. Recording on the spacecraft told us later that, unfortunately the firing continued just 152sec. and stopped. The reason of the malfunction of the OME was the blocking of check valve of the gas pressure line to push the fuel to the engine. We failed to make the spacecraft the Venus orbiter, and it is rotating the sun with the orbital period of 203 days. As the Venus orbit the sun with the period of 225 days, AKATSUKI has a chance to meet Venus again in 5 or 6 years depending on the orbit correction plan. Let us summarize the present situation of AKATSUKI. Most of the fuel still remains. But the condition of the propulsion

Detection of gravitational radiation by the Doppler tracking of spacecraft

International Nuclear Information System (INIS)

Mashhoon, B.

1979-01-01

It has been suggested that the residual Doppler shift in the precision electromagnetic tracking of spacecraft be used to search for gravitational radiation that may be incident on the Earth-spacecraft system. The influence of a gravitational wave on the Doppler shift is calculated, and it is found that the residual shift is dominated by two terms: one is due to the passage of electromagnetic waves through the gravitational radiation field, and the other depends on the change in the relative velocity of the Earth and the spacecraft caused by the external wave. A detailed analysis is given of the influence of gravitational radiation on a binary system with an orbital size small compared to the wavelength of the incident radiation. It is shown that, as a consequence of the interaction with the external wave, the system makes a transition from one Keplerian orbit into another which, in general, has a different energy and angular momentum. It is therefore proposed to search for such effects in the solar system. Observations of the orbit of an artificial Earth satellite, the lunar orbit, and especially the planetary orbits offer exciting possibilities for the detection of gravitational waves of various wavelengths. From the results of the lunar laser ranging experiment and the range measurement to Mars, certain interesting limits may be established on the frequency of incidence of gravitational waves of a given flux on the Earth-Moon and the Earth-Mars systems. This is followed by a brief and preliminary analysis of the possibility of detecting gravitational radiation by measuring a residual secular Doppler shift in the satellite-to-satellite Doppler tracking of two counterorbiting drag-free spacecraft around the Earth as in the Van Patten-Everitt experiment

The Plasma Proton Environment within Saturn's inner magnetosphere as Observed by the Cassini Plasma Spectrometer (CAPS) during Saturn Orbit Insertion

Science.gov (United States)

Sittler, E. C., Jr.; Elrod, M. K.; Johnson, R. E.; Cooper, J. F.; Tseng, W. L.; Smith, H. T.; Chornay, D. J.; Shappirio, M.; Simpson, D. G.

2017-12-01

In analyzing the Cassini data between Saturn's A-ring outer edge and Mimas' L shell numerous inconsistencies have been reported in estimates of total ionic charge and electron density. The primary focus of our work is to understand these inconsistencies. We present our recent discovery of plasma protons during Saturn Orbit Insertion (SOI) outbound pass of the magnetospheric region between the F and G rings. We also searched for H2+ ions but no such events were found. The discovery of protons was made possible by a recent analysis of the CAPS Ion Mass Spectrometer's (IMS's) time-of-flight (TOF) composition data in a mode of reduced post-acceleration voltage at 6 kV instead of the usual 14.6 kV. All previous work for this region had not considered the TOF data. The new proton analysis was enabled by minimum scattering of 6 kV protons in the instrument's ultrathin carbon foils (CF), in comparison to larger scattering for the heavier ions such as for O+ and O2+. We use a SIMION model of the CAPS IMS including the effects of energy straggling and scattering by the instrument's CFs in an attempt to understand the TOF composition data for the heavier ions. This analysis within the uncertainties of the instrument allows us to estimate the relative abundances of the heavier ions and thus run our 2D velocity ion moments code to get ion densities, temperatures and velocities during the SOI outbound pass through the F-ring and G-ring gap. Comparisons with other data sets will be made.

Effects of DeOrbitSail as applied to Lifetime predictions of Low Earth Orbit Satellites

Science.gov (United States)

Afful, Andoh; Opperman, Ben; Steyn, Herman

2016-07-01

Orbit lifetime prediction is an important component of satellite mission design and post-launch space operations. Throughout its lifetime in space, a spacecraft is exposed to risk of collision with orbital debris or operational satellites. This risk is especially high within the Low Earth Orbit (LEO) region where the highest density of space debris is accumulated. This paper investigates orbital decay of some LEO micro-satellites and accelerating orbit decay by using a deorbitsail. The Semi-Analytical Liu Theory (SALT) and the Satellite Toolkit was employed to determine the mean elements and expressions for the time rates of change. Test cases of observed decayed satellites (Iridium-85 and Starshine-1) are used to evaluate the predicted theory. Results for the test cases indicated that the theory fitted observational data well within acceptable limits. Orbit decay progress of the SUNSAT micro-satellite was analysed using relevant orbital parameters derived from historic Two Line Element (TLE) sets and comparing with decay and lifetime prediction models. This paper also explored the deorbit date and time for a 1U CubeSat (ZACUBE-01). The use of solar sails as devices to speed up the deorbiting of LEO satellites is considered. In a drag sail mode, the deorbitsail technique significantly increases the effective cross-sectional area of a satellite, subsequently increasing atmospheric drag and accelerating orbit decay. The concept proposed in this study introduced a very useful technique of orbit decay as well as deorbiting of spacecraft.

Present status of the Japanese Venus climate orbiter

Science.gov (United States)

Nakamura, M.; Imamura, T.; Abe, T.; Ishii, N.

The code name of 24th science spacecraft of ISAS/JAXA is Planet-C. It is the first Venus Climate Orbiter (VCO) of Japan. The ministry of finance of Japan finally agreed to start phase B study of VCO from this April, 2004. We plan 1-2 years phase B study followed by 2 years of flight model integration. The spacecraft will be launched between 2009 and 2010. After arriving Venus, 2 years of operation is expected. VCO will complemet the ESA's Venus Express mission which have several spectrometers and will reveal the composition of the Venusian atmosphere. On the other hand, VCO is designed to reveal the details of the atmospheric motion on Venus and approach the dynamics of the Venusian climate. Cooperation between Japanese VCO and ESA's Venus Express, in the colaboration framework of U.S., Europian, and Japanese scienctist is very important. To elucidate the driving mechanism of the 4-days super-rotation is one of our main targets. We have 4 cameras to take snap shots of the planets in different wave lengths. They are the IR1 camera (1 micron-meter), the IR2 camera (2.4 micron-meter), the LIR camera (10-12 micron-meter), and the UVI camera (340nm). They are attached to the side panel of the 3-axis stabilized spacecraft, and are directed to Venus with the spacecraft's attitude control. Snap shots are expected to be taken every 2 hours. The spacecraft has an orbit of 300km x 13Rv (Venusian radii) with 172 degrees inclination. Orbital period is 30 hours. The angular position of the spacecraft on this orbit is synchronized for 20 hours at its apoapsis with the global atmospheric circulation at the altitude of 50km, thus the snap shots of every 2 hours will be the images of the same side of the atmosphere. In addition to these 4 cameras, we have a Lightning and Airglow camera (LAC) in visible range. This will be operated when the orbiter is close to the planet.

GOES-R active vibration damping controller design, implementation, and on-orbit performance

Science.gov (United States)

Clapp, Brian R.; Weigl, Harald J.; Goodzeit, Neil E.; Carter, Delano R.; Rood, Timothy J.

2018-01-01

GOES-R series spacecraft feature a number of flexible appendages with modal frequencies below 3.0 Hz which, if excited by spacecraft disturbances, can be sources of undesirable jitter perturbing spacecraft pointing. To meet GOES-R pointing stability requirements, the spacecraft flight software implements an Active Vibration Damping (AVD) rate control law which acts in parallel with the nadir point attitude control law. The AVD controller commands spacecraft reaction wheel actuators based upon Inertial Measurement Unit (IMU) inputs to provide additional damping for spacecraft structural modes below 3.0 Hz which vary with solar wing angle. A GOES-R spacecraft dynamics and attitude control system identified model is constructed from pseudo-random reaction wheel torque commands and IMU angular rate response measurements occurring over a single orbit during spacecraft post-deployment activities. The identified Fourier model is computed on the ground, uplinked to the spacecraft flight computer, and the AVD controller filter coefficients are periodically computed on-board from the Fourier model. Consequently, the AVD controller formulation is based not upon pre-launch simulation model estimates but upon on-orbit nadir point attitude control and time-varying spacecraft dynamics. GOES-R high-fidelity time domain simulation results herein demonstrate the accuracy of the AVD identified Fourier model relative to the pre-launch spacecraft dynamics and control truth model. The AVD controller on-board the GOES-16 spacecraft achieves more than a ten-fold increase in structural mode damping for the fundamental solar wing mode while maintaining controller stability margins and ensuring that the nadir point attitude control bandwidth does not fall below 0.02 Hz. On-orbit GOES-16 spacecraft appendage modal frequencies and damping ratios are quantified based upon the AVD system identification, and the increase in modal damping provided by the AVD controller for each structural mode is

Cassini data assessment report

International Nuclear Information System (INIS)

1998-08-01

On October 15, 1997, the Cassini spacecraft was launched from Cape Canaveral Air Station (CCAS) and is now on its way to the planet Saturn. The functional support provided to NASA by DOE included the Advance Launch Support Group (ALSG). If there had been a launch anomaly, the ALSG would have provided a level of radiological emergency response support adequate to transition into a Federal Radiological Monitoring and Assessment Center (FRMAC). Additional functional radiological emergency response support, as part of the ALSG, included the: (1) Aerial Measurement System (AMS); (2) Atmospheric Release Advisory Capability (ARAC); (3) Geographic Information System (GIS); (4) Emergency Response Data System (ERDS); (5) Radiation Emergency Assistance Center and Training Site (REAC/TS); (6) Field monitoring and sampling; (7) Radioanalysis via RASCAL; (8) Source recovery; and (9) Neutron dosimetry and communications support. This functional support provided the capability to rapidly measure and assess radiological impacts from a launch anomaly. The Radiological Control Officer (RCO) on KSC established a Radiological Control Center (RADCC) as the focal point for all on-site and off-site radiological data and information flow. Scientists and radiological response personnel located at the RADCC managed the field monitoring team on the KSC/CCAS federal properties. Off-site radiological emergency response activities for all public lands surrounding the KSC/CCAS complex were coordinated through the Off-site ALSG located at the National Guard Armory in Cocoa, Florida. All of the in situ measurement data of good quality gathered during the dry run, the first launch attempt and the launch day are listed in this document. The RASCAL analysis results of the air filters and impactor planchets are listed

Spacecraft Software Maintenance: An Effective Approach to Reducing Costs and Increasing Science Return

Science.gov (United States)

Shell, Elaine M.; Lue, Yvonne; Chu, Martha I.

1999-01-01

Flight software is a mission critical element of spacecraft functionality and performance. When ground operations personnel interface to a spacecraft, they are typically dealing almost entirely with the capabilities of onboard software. This software, even more than critical ground/flight communications systems, is expected to perform perfectly during all phases of spacecraft life. Due to the fact that it can be reprogrammed on-orbit to accommodate degradations or failures in flight hardware, new insights into spacecraft characteristics, new control options which permit enhanced science options, etc., the on- orbit flight software maintenance team is usually significantly responsible for the long term success of a science mission. Failure of flight software to perform as needed can result in very expensive operations work-around costs and lost science opportunities. There are three basic approaches to maintaining spacecraft software--namely using the original developers, using the mission operations personnel, or assembling a center of excellence for multi-spacecraft software maintenance. Not planning properly for flight software maintenance can lead to unnecessarily high on-orbit costs and/or unacceptably long delays, or errors, in patch installations. A common approach for flight software maintenance is to access the original development staff. The argument for utilizing the development staff is that the people who developed the software will be the best people to modify the software on-orbit. However, it can quickly becomes a challenge to obtain the services of these key people. They may no longer be available to the organization. They may have a more urgent job to perform, quite likely on another project under different project management. If they havn't worked on the software for a long time, they may need precious time for refamiliarization to the software, testbeds and tools. Further, a lack of insight into issues related to flight software in its on-orbit

Overview of SDCM - The Spacecraft Design and Cost Model

Science.gov (United States)

Ferebee, Melvin J.; Farmer, Jeffery T.; Andersen, Gregory C.; Flamm, Jeffery D.; Badi, Deborah M.

1988-01-01

The Spacecraft Design and Cost Model (SDCM) is a computer-aided design and analysis tool for synthesizing spacecraft configurations, integrating their subsystems, and generating information concerning on-orbit servicing and costs. SDCM uses a bottom-up method in which the cost and performance parameters for subsystem components are first calculated; the model then sums the contributions from individual components in order to obtain an estimate of sizes and costs for each candidate configuration within a selected spacecraft system. An optimum spacraft configuration can then be selected.

The Mars Reconnaissance Orbiter Mission: 10 Years of Exploration from Mars Orbit

Science.gov (United States)

Johnston, M. Daniel; Zurek, Richard W.

2016-01-01

The Mars Reconnaissance Orbiter ( MRO ) entered Mars orbit on March 10, 2006. After five months of aerobraking, a series of propulsive maneuvers were used to establish the desired low -altitude science orbit. The spacecraft has been on station in its 255 x 320 k m, sun -synchronous (approximately 3 am -pm ), primary science orbit since September 2006 performing both scientific and Mars programmatic support functions. This paper will provide a summary of the major achievements of the mission to date and the major flight activities planned for the remainder of its third Extended Mission (EM3). Some of the major flight challenges the flight team has faced are also discussed.

Fundamentals of the orbit and response for TianQin

Science.gov (United States)

Hu, Xin-Chun; Li, Xiao-Hong; Wang, Yan; Feng, Wen-Fan; Zhou, Ming-Yue; Hu, Yi-Ming; Hu, Shou-Cun; Mei, Jian-Wei; Shao, Cheng-Gang

2018-05-01

TianQin is a space-based laser interferometric gravitational wave detector aimed at detecting gravitational waves at low frequencies (0.1 mHz–1 Hz). It is formed by three identical drag-free spacecrafts in an equilateral triangular constellation orbiting around the Earth. The distance between each pair of spacecrafts is approximately 1.7 × 105 ~km . The spacecrafts are interconnected by infrared laser beams forming up to three Michelson-type interferometers. The detailed mission design and the study of science objectives for the TianQin project depend crucially on the orbit and the response of the detector. In this paper, we provide the analytic expressions for the coordinates of the orbit for each spacecraft in the heliocentric-ecliptic coordinate system to the leading orders. This enables a sufficiently accurate study of science objectives and data analysis, and serves as a first step to further orbit design and optimization. We calculate the response of a single Michelson detector to plane gravitational waves in arbitrary waveform which is valid in the full range of the sensitive frequencies. It is then used to generate the more realistic sensitivity curve of TianQin. We apply this model on a reference white-dwarf binary as a proof of principle.

Magnus Effect on a Spinning Satellite in Low Earth Orbit

Science.gov (United States)

Ramjatan, Sahadeo; Fitz-Coy, Norman; Yew, Alvin Garwai

2016-01-01

A spinning body in a flow field generates an aerodynamic lift or Magnus effect that displaces the body in a direction normal to the freestream flow. Earth orbiting satellites with substantial body rotation in appreciable atmospheric densities may generate a Magnus force to perturb orbital dynamics. We investigate the feasibility of using this effect for spacecraft at a perigee of 80km using the Systems Tool Kit (STK). Results show that for a satellite of reasonable properties, the Magnus effect doubles the amount of time in orbit. Orbital decay was greatly mitigated for satellites spinning at 10000 and 15000RPM. This study demonstrates that the Magnus effect has the potential to sustain a spacecraft's orbit at a low perigee altitude and could also serve as an orbital maneuver capability.

Deep-Space Ka-Band Flight Experience

Science.gov (United States)

Morabito, D. D.

2017-11-01

Lower frequency bands have become more congested in allocated bandwidth as there is increased competition between flight projects and other entities. Going to higher frequency bands offers significantly more bandwidth, allowing for the use of much higher data rates. However, Ka-band is more susceptible to weather effects than lower frequency bands currently used for most standard downlink telemetry operations. Future or prospective flight projects considering deep-space Ka-band (32-GHz) telemetry data links have expressed an interest in understanding past flight experience with received Ka-band downlink performance. Especially important to these flight projects is gaining a better understanding of weather effects from the experience of current or past missions that operated Ka-band radio systems. We will discuss the historical flight experience of several Ka-band missions starting from Mars Observer in 1993 up to present-day deep-space missions such as Kepler. The study of historical Ka-band flight experience allows one to recommend margin policy for future missions. Of particular interest, we will review previously reported-on flight experience with the Cassini spacecraft Ka-band radio system that has been used for radio science investigations as well as engineering studies from 2004 to 2015, when Cassini was in orbit around the planet Saturn. In this article, we will focus primarily on the Kepler spacecraft Ka-band link, which has been used for operational telemetry downlink from an Earth trailing orbit where the spacecraft resides. We analyzed the received Ka-band signal level data in order to characterize link performance over a wide range of weather conditions and as a function of elevation angle. Based on this analysis of Kepler and Cassini flight data, we found that a 4-dB margin with respect to adverse conditions ensures that we achieve at least a 95 percent data return.

Cassini/CIRS Observations of Water Vapor in Titan's Stratosphere

Science.gov (United States)

Bjoraker, Gordon L.; Achterberg, R. K.; Anderson, C. M.; Samuelson, R. E.; Carlson, R. C.; Jennings, D. E.

2008-01-01

The Composite Infrared Spectrometer (CIRS) on the Cassini spacecraft has obtained spectra of Titan during most of the 44 flybys of the Cassini prime mission. Water vapor on Titan was first detected using whole-disk observations from the Infrared Space Observatory (Coustenis et al 1998, Astron. Astrophys. 336, L85-L89). CIRS data permlt the retrieval of the latitudinal variation of water on Titan and some limited information on its vertical profile. Emission lines of H2O on Titan are very weak in the CIRS data. Thus, large spectral averages as well as improvements in calibration are necessary to detect water vapor. Water abundances were retrieved in nadir spectra at 55 South, the Equator, and at 19 North. Limb spectra of the Equator were also modeled to constrain the vertical distribution of water. Stratospheric temperatures in the 0.5 - 4.0 mbar range were obtained by inverting spectra of CH4 in the v4 band centered at 1304/cm. The temperature in the lower stratosphere (4 - 20 mbar) was derived from fitting pure rotation lines of CH4 between 80 and 160/cm. The origin of H2O and CO2 is believed to be from the ablation of micrometeorites containing water ice, followed by photochemistry. This external source of water originates either within the Saturn system or from the interplanetary medium. Recently, Horst et al (J. Geophys. Res. 2008, in press) developed a photochemical model of Titan in which there are two external sources of oxygen. Oxygen ions (probably from Enceladus) precipitate into Titan's atmosphere to form CO at very high altitudes (1100 km). Water ice ablation at lower altitudes (700 km) forms H2O and subsequent chemistry produces CO2. CIRS measurements of CO, CO2, and now of H2O will provide valuable constraints to these photochemical models and - improve our understanding of oxygen chemistry on Titan.

Link Analysis of High Throughput Spacecraft Communication Systems for Future Science Missions

Science.gov (United States)

Simons, Rainee N.

2015-01-01

NASA's plan to launch several spacecrafts into low Earth Orbit (LEO) to support science missions in the next ten years and beyond requires down link throughput on the order of several terabits per day. The ability to handle such a large volume of data far exceeds the capabilities of current systems. This paper proposes two solutions, first, a high data rate link between the LEO spacecraft and ground via relay satellites in geostationary orbit (GEO). Second, a high data rate direct to ground link from LEO. Next, the paper presents results from computer simulations carried out for both types of links taking into consideration spacecraft transmitter frequency, EIRP, and waveform; elevation angle dependent path loss through Earths atmosphere, and ground station receiver GT.

Mission operations for unmanned nuclear electric propulsion outer planet exploration with a thermionic reactor spacecraft.

Science.gov (United States)

Spera, R. J.; Prickett, W. Z.; Garate, J. A.; Firth, W. L.

1971-01-01

Mission operations are presented for comet rendezvous and outer planet exploration NEP spacecraft employing in-core thermionic reactors for electric power generation. The selected reference missions are the Comet Halley rendezvous and a Jupiter orbiter at 5.9 planet radii, the orbit of the moon Io. The characteristics of the baseline multi-mission NEP spacecraft are presented and its performance in other outer planet missions, such as Saturn and Uranus orbiters and a Neptune flyby, are discussed. Candidate mission operations are defined from spacecraft assembly to mission completion. Pre-launch operations are identified. Shuttle launch and subsequent injection to earth escape by the Centaur D-1T are discussed, as well as power plant startup and the heliocentric mission phases. The sequence and type of operations are basically identical for all missions investigated.

High Gain Antenna Calibration on Three Spacecraft

Science.gov (United States)

Hashmall, Joseph A.

2011-01-01

This paper describes the alignment calibration of spacecraft High Gain Antennas (HGAs) for three missions. For two of the missions (the Lunar Reconnaissance Orbiter and the Solar Dynamics Observatory) the calibration was performed on orbit. For the third mission (the Global Precipitation Measurement core satellite) ground simulation of the calibration was performed in a calibration feasibility study. These three satellites provide a range of calibration situations-Lunar orbit transmitting to a ground antenna for LRO, geosynchronous orbit transmitting to a ground antenna fer SDO, and low Earth orbit transmitting to TDRS satellites for GPM The calibration results depend strongly on the quality and quantity of calibration data. With insufficient data the calibration Junction may give erroneous solutions. Manual intervention in the calibration allowed reliable parameters to be generated for all three missions.

A Technology Program that Rescues Spacecraft

Science.gov (United States)

Deutsch, Leslie J.; Lesh, J. R.

2004-03-01

There has never been a long-duration deep space mission that did not have unexpected problems during operations. JPL's Interplanetary Network Directorate (IND) Technology Program was created to develop new and improved methods of communication, navigation, and operations. A side benefit of the program is that it maintains a cadre of human talent and experimental systems that can be brought to bear on unexpected problems that may occur during mission operations. Solutions fall into four categories: applying new technology during operations to enhance science performance, developing new operational strategies, providing domain experts to help find solutions, and providing special facilities to trouble-shoot problems. These are illustrated here using five specific examples of spacecraft anomalies that have been solved using, at least in part, expertise or facilities from the IND Technology Program: Mariner 10, Voyager, Galileo, SOHO, and Cassini/Huygens. In this era of careful cost management, and emphasis on returns-on-investment, it is important to recognize this crucial additional benefit from such technology program investments.

Determining Spacecraft Reaction Wheel Friction Parameters

Science.gov (United States)

Sarani, Siamak

2009-01-01

Software was developed to characterize the drag in each of the Cassini spacecraft's Reaction Wheel Assemblies (RWAs) to determine the RWA friction parameters. This tool measures the drag torque of RWAs for not only the high spin rates (greater than 250 RPM), but also the low spin rates (less than 250 RPM) where there is a lack of an elastohydrodynamic boundary layer in the bearings. RWA rate and drag torque profiles as functions of time are collected via telemetry once every 4 seconds and once every 8 seconds, respectively. Intermediate processing steps single-out the coast-down regions. A nonlinear model for the drag torque as a function of RWA spin rate is incorporated in order to characterize the low spin rate regime. The tool then uses a nonlinear parameter optimization algorithm based on the Nelder-Mead simplex method to determine the viscous coefficient, the Dahl friction, and the two parameters that account for the low spin-rate behavior.

Drag De-Orbit Device: A New Standard Re-Entry Actuator for CubeSats

Science.gov (United States)

Guglielmo, David; Omar, Sanny R.; Bevilacqua, Riccardo

2017-01-01

With the advent of CubeSats, research in Low Earth Orbit (LEO) becomes possible for universities and small research groups. Only a handful of launch sites can be used, due to geographical and political restrictions. As a result, common orbits in LEO are becoming crowded due to the additional launches made possible by low-cost access to space. CubeSat design principles require a maximum of a 25-year orbital lifetime in an effort to reduce the total number of spacecraft in orbit at any time. Additionally, since debris may survive re-entry, it is ideal to de-orbit spacecraft over unpopulated areas to prevent casualties. The Drag Deorbit Device (D3) is a self-contained targeted re-entry subsystem intended for CubeSats. By varying the cross-wind area, the atmospheric drag can be varied in such a way as to produce desired maneuvers. The D3 is intended to be used to remove spacecraft from orbit to reach a desired target interface point. Additionally, attitude stabilization is performed by the D3 prior to deployment and can replace a traditional ADACS on many missions.This paper presents the hardware used in the D3 and operation details. Four stepper-driven, repeatedly retractable booms are used to modify the cross-wind area of the D3 and attached spacecraft. Five magnetorquers (solenoids) over three axes are used to damp rotational velocity. This system is expected to be used to improve mission flexibility and allow additional launches by reducing the orbital lifetime of spacecraft.The D3 can be used to effect a re-entry to any target interface point, with the orbital inclination limiting the maximum latitude. In the chance that the main spacecraft fails, a timer will automatically deploy the booms fully, ensuring the spacecraft will at the minimum reenter the atmosphere in the minimum possible time, although not necessarily at the desired target interface point. Although this does not reduce the risk of casualties, the 25-year lifetime limit is still respected, allowing

Cassini Radar EQM Model: Instrument Description and Performance Status

Science.gov (United States)

Borgarelli, L.; Faustini, E. Zampolini; Im, E.; Johnson, W. T. K.

1996-01-01

The spaeccraft of the Cassini Mission is planned to be launched towards Saturn in October 1997. The mission is designed to study the physical structure and chemical composition of Titan. The results of the tests performed on the Cassini radar engineering qualification model (EQM) are summarized. The approach followed in the verification and evaluation of the performance of the radio frequency subsystem EQM is presented. The results show that the instrument satisfies the relevant mission requirements.

Indirect evidences for a gas/dust torus along the Phobos orbit

International Nuclear Information System (INIS)

Dubinin, E.M.; Lundin, R.; Pissarenko, N.F.; Barabash, S.V.; Zakharov, A.V.; Koskinen, H.; Schwingenshuh, K.; Yeroshenko, Ye.G.

1990-01-01

Observations from the PHOBOS-2 space-craft of plasma and magnetic field effects in the solar wind near Mars suggest that a neutral gas (dust?)torus/ring resides along the orbit of the Martian satellite Phobos. Magnetic cavities (strong decreases of the magnetic field strength) coincident with strong plasma density increases (up to a factor of ten) are observed during the first elliptic transition orbits when the spacecraft approached the Phobos orbits. The characteristic transverse dimension of the structures along the spacecraft orbit is in the range 100-1,000 km. Torus effects also have characteristics similar to the formation of a bow shock with increases of plasma density and ion temperature, and a characteristic deflection of the ion flow. This suggests a rather strong interaction between the solar wind plasma and plasma near Phobos orbit. The interaction appears quite similar to that of the solar wind with a comet. The outgassing of matter from Phobos (and Deimos) is also suggested by plasma observations in the wake/tail of the Martian satellites. Altogether, the authors observations imply that a neutral gas cloud - possibly also associated with a faint dust ring - exists along the Phobos orbit

Streamlined Modeling for Characterizing Spacecraft Anomalous Behavior

Science.gov (United States)

Klem, B.; Swann, D.

2011-09-01

Anomalous behavior of on-orbit spacecraft can often be detected using passive, remote sensors which measure electro-optical signatures that vary in time and spectral content. Analysts responsible for assessing spacecraft operational status and detecting detrimental anomalies using non-resolved imaging sensors are often presented with various sensing and identification issues. Modeling and measuring spacecraft self emission and reflected radiant intensity when the radiation patterns exhibit a time varying reflective glint superimposed on an underlying diffuse signal contribute to assessment of spacecraft behavior in two ways: (1) providing information on body component orientation and attitude; and, (2) detecting changes in surface material properties due to the space environment. Simple convex and cube-shaped spacecraft, designed to operate without protruding solar panel appendages, may require an enhanced level of preflight characterization to support interpretation of the various physical effects observed during on-orbit monitoring. This paper describes selected portions of the signature database generated using streamlined signature modeling and simulations of basic geometry shapes apparent to non-imaging sensors. With this database, summarization of key observable features for such shapes as spheres, cylinders, flat plates, cones, and cubes in specific spectral bands that include the visible, mid wave, and long wave infrared provide the analyst with input to the decision process algorithms contained in the overall sensing and identification architectures. The models typically utilize baseline materials such as Kapton, paints, aluminum surface end plates, and radiators, along with solar cell representations covering the cylindrical and side portions of the spacecraft. Multiple space and ground-based sensors are assumed to be located at key locations to describe the comprehensive multi-viewing aspect scenarios that can result in significant specular reflection

Preliminary thermal design of the COLD-SAT spacecraft

Science.gov (United States)

Arif, Hugh

1991-01-01

The COLD-SAT free-flying spacecraft was to perform experiments with LH2 in the cryogenic fluid management technologies of storage, supply and transfer in reduced gravity. The Phase A preliminary design of the Thermal Control Subsystem (TCS) for the spacecraft exterior and interior surfaces and components of the bus subsystems is described. The TCS was composed of passive elements which were augmented with heaters. Trade studies to minimize the parasitic heat leakage into the cryogen storage tanks are described. Selection procedure for the thermally optimum on-orbit spacecraft attitude was defined. TRASYS-2 and SINDA'85 verification analysis was performed on the design and the results are presented.

A Preliminary Formation Flying Orbit Dynamics Analysis for Leonardo-BRDF

Science.gov (United States)

Hughes, Steven P.; Mailhe, Laurie M.

2001-01-01

Leonardo-BRDF is a NASA mission concept proposed to allow the investigation of radiative transfer and its effect on the Earth's climate and atmospheric phenomenon. Enabled by the recent developments in small-satellite and formation flying technology, the mission is envisioned to be composed of an array of spacecraft in carefully designed orbits. The different perspectives provided by a distributed array of spacecraft offer a unique advantage to study the Earth's albedo. This paper presents the orbit dynamics analysis performed in the context of the Leonardo-BRDF science requirements. First, the albedo integral is investigated and the effect of viewing geometry on science return is studied. The method used in this paper, based on Gauss quadrature, provides the optimal formation geometry to ensure that the value of the integral is accurately approximated. An orbit design approach is presented to achieve specific relative orbit geometries while simultaneously satisfying orbit dynamics constraints to reduce formation-keeping fuel expenditure. The relative geometry afforded by the design is discussed in terms of mission requirements. An optimal two-burn initialization scheme is presented with the required delta-V to distribute all spacecraft from a common parking orbit into their appropriate orbits in the formation. Finally, formation-keeping strategies are developed and the associated delta-V's are calculated to maintain the formation in the presence of perturbations.

Estimating spacecraft attitude based on in-orbit sensor measurements

DEFF Research Database (Denmark)

Jakobsen, Britt; Lyn-Knudsen, Kevin; Mølgaard, Mathias

2014-01-01

of 2014/15. To better evaluate the performance of the payload, it is desirable to couple the payload data with the satellite's orientation. With AAUSAT3 already in orbit it is possible to collect data directly from space in order to evaluate the performance of the attitude estimation. An extended kalman...... filter (EKF) is used for quaternion-based attitude estimation. A Simulink simulation environment developed for AAUSAT3, containing a "truth model" of the satellite and the orbit environment, is used to test the performance The performance is tested using different sensor noise parameters obtained both...... from a controlled environment on Earth as well as in-orbit. By using sensor noise parameters obtained on Earth as the expected parameters in the attitude estimation, and simulating the environment using the sensor noise parameters from space, it is possible to assess whether the EKF can be designed...

Envisioning a 21st Century, National, Spacecraft Servicing and Protection Infrastructure and Demand Potential: A Logical Development of the Earth Orbit Economy

Science.gov (United States)

Horsham, Gary A.

2003-01-01

The modern world is extremely dependent on thin strings of several hundred civil, military, and commercial spacecraft/satellites currently stationed in space. They provide a steady stream of commerce, defense, and knowledge data. This dependency will in all likelihood increase significantly during this century. A major disruption of any kind in these essential systems and networks could be socially, economically, and politically catastrophic, on a global scale. The development of a space-based, robotic services economy could be useful in mitigating this growing risk, from an efficiency and security standpoint. This paper attempts to suggest what makes sense to invest in next for the logical, economic development of Earth orbit i.e., after ISS completion. It expands on the results of an advanced market research and analysis study that sampled the opinions of several satellite industry executives and presents these results within a broad policy context. The concept of a spacecraft carrier that serves as the nucleus of a national, space-based or on-orbit, robotic services infrastructure is introduced as the next logical step for United States leadership in space. This is viewed as a reasonable and appropriate followon to the development of ELVs and satellites in the 1950s and 1960s, the Space Shuttle/PRLV in the 1970s and 1980s, and the International Space Station (ISS) in the 1980s, 1990s and 2000s. Large-scale experience in LEO-to-GEO spacecraft/satellite servicing and protection by robotic means is assumed to be an indispensable prerequisite or stepping-stone toward the development and preservation of the large scientific exploration facilities that are envisioned by NASA for operation beyond GEO. A balanced, return on national investment (RONI) strategy for space, focused on the provision of enhanced national/homeland security for increased protection, national economic/industrial expansion for increased revenue, and national scientific exploration for increased

Spacecraft navigation at Mars using earth-based and in situ radio tracking techniques

Science.gov (United States)

Thurman, S. W.; Edwards, C. D.; Kahn, R. D.; Vijayaraghavan, A.; Hastrup, R. C.; Cesarone, R. J.

1992-08-01

A survey of earth-based and in situ radiometric data types and results from a number of studies investigating potential radio navigation performance for spacecraft approaching/orbiting Mars and for landed spacecraft and rovers on the surface of Mars are presented. The performance of Doppler, ranging and interferometry earth-based data types involving single or multiple spacecraft is addressed. This evaluation is conducted with that of in situ data types, such as Doppler and ranging measurements between two spacecraft near Mars, or between a spacecraft and one or more surface radio beacons.

Orbital Maneuvers for Spacecrafts Travelling to/from the Lagrangian Points

Science.gov (United States)

Bertachini, A.

The well-known Lagrangian points that appear in the planar restricted three-body problem (Szebehely, 1967) are very important for astronautical applications. They are five points of equilibrium in the equations of motion, what means that a particle located at one of those points with zero velocity will remain there indefinitely. The collinear points (L1, L2 and L3) are always unstable and the triangular points (L4 and L5) are stable in the present case studied (Sun-Earth system). They are all very good points to locate a space-station, since they require a small amount of V (and fuel), the control to be used for station-keeping. The triangular points are specially good for this purpose, since they are stable equilibrium points. In this paper, the planar restricted three-body problem is regularized (using Lemaître regularization) and combined with numerical integration and gradient methods to solve the two point boundary value problem (the Lambert's three-body problem). This combination is applied to the search of families of transfer orbits between the Lagrangian points and the Earth, in the Sun-Earth system, with the minimum possible cost of the control used. So, the final goal of this paper is to find the magnitude and direction of the two impulses to be applied in the spacecraft to complete the transfer: the first one when leaving/arriving at the Lagrangian point and the second one when arriving/living at the Earth. This paper is a continuation of two previous papers that studied transfers in the Earth-Moon system: Broucke (1979), that studied transfer orbits between the Lagrangian points and the Moon and Prado (1996), that studied transfer orbits between the Lagrangian points and the Earth. So, the equations of motion are: whereis the pseudo-potential given by: To solve the TPBVP in the regularized variables the following steps are used: i) Guess a initial velocity Vi, so together with the initial prescribed position ri the complete initial state is known; ii

Autonomous orbit determination and its error analysis for deep space using X-ray pulsar

International Nuclear Information System (INIS)

Feng, Dongzhu; Yuan, Xiaoguang; Guo, Hehe; Wang, Xin

2014-01-01

Autonomous orbit determination (OD) is a complex process using filtering method to integrate observation and orbit dynamic model effectively and estimate the position and velocity of a spacecraft. As a novel technology for autonomous interplanetary OD, X-ray pulsar holds great promise for deep space exploration. The position and velocity of spacecraft should be estimated accurately during the OD process. However, under the same condition, the accuracy of OD can be greatly reduced by the error of the initial orbit value and the orbit mutation. To resolve this problem, we propose a novel OD method, which is based on the X-ray pulsar measurement and Adaptive Unscented Kalman Filter (AUKF). The accuracy of OD can be improved obviously because the AUKF estimates the orbit of spacecraft using measurement residual. During the simulation, the orbit of Phoenix Mars Lander, Deep Impact Probe, and Voyager 1 are selected. Compared with Unscented Kalman Filter (UKF) and Extended Kalman Filter (EKF), the simulation results demonstrate that the proposed OD method based on AUKF can accurately determinate the velocity and position and effectively decrease the orbit estimated errors which is caused by the orbit mutation and orbit initial errors. (authors)

Analysis on coverage ability of BeiDou navigation satellite system for manned spacecraft

Science.gov (United States)

Zhao, Sihao; Yao, Zheng; Zhuang, Xuebin; Lu, Mingquan

2014-12-01

To investigate the service ability of the BeiDou Navigation Satellite System (BDS) for manned spacecraft, both the current regional and the future-planned global constellations of BDS are introduced and simulated. The orbital parameters of the International Space Station and China's Tiangong-1 spacelab are used to create the simulation scenario and evaluate the performance of the BDS constellations. The number of visible satellites and the position dilution (PDOP) of precision at the spacecraft-based receiver are evaluated. Simulation and analysis show quantitative results on the coverage ability and time percentages of both the current BDS regional and future global constellations for manned-space orbits which can be a guideline to the applications and mission design of BDS receivers on manned spacecraft.

Lifetime predictions for the Solar Maximum Mission (SMM) and San Marco spacecraft

Science.gov (United States)

Smith, E. A.; Ward, D. T.; Schmitt, M. W.; Phenneger, M. C.; Vaughn, F. J.; Lupisella, M. L.

1989-01-01

Lifetime prediction techniques developed by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) are described. These techniques were developed to predict the Solar Maximum Mission (SMM) spacecraft orbit, which is decaying due to atmospheric drag, with reentry predicted to occur before the end of 1989. Lifetime predictions were also performed for the Long Duration Exposure Facility (LDEF), which was deployed on the 1984 SMM repair mission and is scheduled for retrieval on another Space Transportation System (STS) mission later this year. Concepts used in the lifetime predictions were tested on the San Marco spacecraft, which reentered the Earth's atmosphere on December 6, 1988. Ephemerides predicting the orbit evolution of the San Marco spacecraft until reentry were generated over the final 90 days of the mission when the altitude was less than 380 kilometers. The errors in the predicted ephemerides are due to errors in the prediction of atmospheric density variations over the lifetime of the satellite. To model the time dependence of the atmospheric densities, predictions of the solar flux at the 10.7-centimeter wavelength were used in conjunction with Harris-Priester (HP) atmospheric density tables. Orbital state vectors, together with the spacecraft mass and area, are used as input to the Goddard Trajectory Determination System (GTDS). Propagations proceed in monthly segments, with the nominal atmospheric drag model scaled for each month according to the predicted monthly average value of F10.7. Calibration propagations are performed over a period of known orbital decay to obtain the effective ballistic coefficient. Progagations using plus or minus 2 sigma solar flux predictions are also generated to estimate the despersion in expected reentry dates. Definitive orbits are compared with these predictions as time expases. As updated vectors are received, these are also propagated to reentryto continually update the lifetime predictions.

Automated Spacecraft Conjunction Assessment at Mars and the Moon

Science.gov (United States)

Berry, David; Guinn, Joseph; Tarzi, Zahi; Demcak, Stuart

2012-01-01

Conjunction assessment and collision avoidance are areas of current high interest in space operations. Most current conjunction assessment activity focuses on the Earth orbital environment. Several of the world's space agencies have satellites in orbit at Mars and the Moon, and avoiding collisions there is important too. Smaller number of assets than Earth, and smaller number of organizations involved, but consequences similar to Earth scenarios.This presentation will examine conjunction assessment processes implemented at JPL for spacecraft in orbit at Mars and the Moon.

Spacecraft formation control using analytical finite-duration approaches

Science.gov (United States)

Ben Larbi, Mohamed Khalil; Stoll, Enrico

2018-03-01

This paper derives a control concept for formation flight (FF) applications assuming circular reference orbits. The paper focuses on a general impulsive control concept for FF which is then extended to the more realistic case of non-impulsive thrust maneuvers. The control concept uses a description of the FF in relative orbital elements (ROE) instead of the classical Cartesian description since the ROE provide a direct insight into key aspects of the relative motion and are particularly suitable for relative orbit control purposes and collision avoidance analysis. Although Gauss' variational equations have been first derived to offer a mathematical tool for processing orbit perturbations, they are suitable for several different applications. If the perturbation acceleration is due to a control thrust, Gauss' variational equations show the effect of such a control thrust on the Keplerian orbital elements. Integrating the Gauss' variational equations offers a direct relation between velocity increments in the local vertical local horizontal frame and the subsequent change of Keplerian orbital elements. For proximity operations, these equations can be generalized from describing the motion of single spacecraft to the description of the relative motion of two spacecraft. This will be shown for impulsive and finite-duration maneuvers. Based on that, an analytical tool to estimate the error induced through impulsive maneuver planning is presented. The resulting control schemes are simple and effective and thus also suitable for on-board implementation. Simulations show that the proposed concept improves the timing of the thrust maneuver executions and thus reduces the residual error of the formation control.

Distributed Operations for the Cassini/Huygens Mission

Science.gov (United States)

Lock, P.; Sarrel, M.

1998-01-01

The cassini project employs a concept known as distributed operations which allows independent instrument operations from diverse locations, provides full empowerment of all participants and maximizes use of limited resources.

Scientific and synergistic lessons learned from the Cassini-Huygens mission

Science.gov (United States)

Coustenis, Athena

The Cassini-Huygens mission to the Saturnian system has been an extraordinary success for the planetary community since the Saturn-Orbit-Insertion (SOI) in July 2004 and again the very successful probe descent and landing of Huygens on January 14, 2005. One of its main targets was Titan. Titan, Saturn's largest satellite, is the only other object in our Solar system to possess an extensive nitrogen atmosphere, host to an active organic chemistry, based on the interaction of N2 with methane (CH4). Titan was revealed to be a complex world more like the Earth than any other: it has a dense mostly nitrogen atmosphere and active climate and meteorological cycles where the working fluid, methane, behaves under Titan conditions the way that water does on Earth. Its geology, from lakes and seas to broad river valleys and mountains, while carved in ice is, in its balance of processes, again most like Earth. Beneath this panoply of Earth-like processes an ice crust floats atop what appears to be a liquid water ocean. Titan is also rich in organic molecules—more so in its surface and atmosphere than anyplace in the solar system, including Earth [2-4]. These molecules were formed in the atmosphere, deposited on the surface and, in coming into contact with liquid water may undergo an aqueous chemistry that could replicate aspects of life's origins. I will discuss our current understanding of Titan's complex environment in view of the Cassini-Huygens mission results [1-8], which demonstrated the power of synergistic remote and in situ exploration. I will focus on the atmospheric structure (temperature and composition), and the surface nature. I will show how these and other elements can give us clues as to the origin and evolution of the satellite, and how they connect to the observations of the other satellites, Enceladus in particular. Future space missions to Titan can help us understand the kronian and also our Solar System as a whole. Finally, I will describe the future

Deadly Sunflower Orbits

Science.gov (United States)

Hamilton, Douglas P.

2018-04-01

Solar radiation pressure is usually very effective at removing hazardous millimeter-sized debris from distant orbits around asteroidsand other small solar system bodies (Hamilton and Burns 1992). Theprimary loss mechanism, driven by the azimuthal component of radiationpressure, is eccentricity growth followed by a forced collision withthe central body. One large class of orbits, however, neatly sidestepsthis fate. Orbits oriented nearly perpendicular to the solar directioncan maintain their face-on geometry, oscillating slowly around a stableequilibrium orbit. These orbits, designated sunflower orbits, arerelated to terminator orbits studied by spacecraft mission designers(Broschart etal. 2014).Destabilization of sunflower orbits occurs only for particles smallenough that radiation pressure is some tens of percent the strength ofthe central body's direct gravity. This greatly enhanced stability,which follows from the inability of radiation incident normal to theorbit to efficiently drive eccentricities, presents a threat tospacecraft missions, as numerous dangerous projectiles are potentiallyretained in orbit. We have investigated sunflower orbits insupport of the New Horizons, Aida, and Lucy missions and find thatthese orbits are stable for hazardous particle sizes at asteroids,comets, and Kuiper belt objects of differing dimensions. Weinvestigate the sources and sinks for debris that might populate suchorbits, estimate timescales and equilibrium populations, and willreport on our findings.

Discovery Of B Ring Propellers In Cassini UVIS, And ISS

Science.gov (United States)

Sremcevic, Miodrag; Stewart, G. R.; Albers, N.; Esposito, L. W.

2012-10-01

We present evidence for the existence of propellers in Saturn's B ring by combining data from Cassini Ultraviolet Imaging Spectrograph (UVIS) and Imaging Science Subsystem (ISS) experiments. We identify two propeller populations: (1) tens of degrees wide propellers in the dense B ring core, and (2) smaller, more A ring like, propellers populating the inner B ring. The prototype of the first population is an object observed at 18 different epochs between 2005 and 2010. The ubiquitous propeller "S" shape is seen both in UVIS occultations as an optical depth depletion and in ISS as a 40 degrees wide bright stripe in unlit geometries and dark in lit geometries. Combining the available Cassini data we infer that the object is a partial gap embedded in the high optical depth region of the B ring. The gap moves at orbital speed consistent with its radial location. From the radial separation of the propeller wings we estimate that the embedded body, which causes the propeller structure, is about 1.5km in size located at a=112,921km. The UVIS occultations indicate an asymmetric propeller "S" shape. Since the object is located at an edge between high and relatively low optical depth, this asymmetry is most likely a consequence of the strong surface mass density gradient. We estimate that there are possibly dozen up to 100 other propeller objects in Saturn's B ring. The location of the discovered body, at an edge of a dense ringlet within the B ring, suggests a novel mechanism for the up to now illusive B ring irregular large-scale structure of alternating high and low optical depth ringlets. We propose that this B ring irregular structure may have its cause in the presence of many embedded bodies that shepherd the individual B ring ringlets.

Differential Evolution Optimization for Targeting Spacecraft Maneuver Plans

Science.gov (United States)

Mattern, Daniel

2016-01-01

Previous analysis identified specific orbital parameters as being safer for conjunction avoidance for the TDRS fleet. With TDRS-9 being considered an at-risk spacecraft, a potential conjunction concern was raised should TDRS-9 fail while at a longitude of 12W. This document summarizes the analysis performed to identify if these specific orbital parameters could be targeted using the remaining drift-termination maneuvers for the relocation of TDRS-9 from 41W longitude to 12W longitude.

Titan Orbiter with Aerorover Mission (TOAM)

Science.gov (United States)

Sittler, Edward C.; Cooper, J. F.; Mahaffey, P.; Esper, J.; Fairbrother, D.; Farley, R.; Pitman, J.; Kojiro, D. R.; TOAM Team

2006-12-01

We propose to develop a new mission to Titan called Titan Orbiter with Aerorover Mission (TOAM). This mission is motivated by the recent discoveries of Titan, its atmosphere and its surface by the Huygens Probe, and a combination of in situ, remote sensing and radar mapping measurements of Titan by the Cassini orbiter. Titan is a body for which Astrobiology (i.e., prebiotic chemistry) will be the primary science goal of any future missions to it. TOAM is planned to use an orbiter and balloon technology (i.e., aerorover). Aerobraking will be used to put payload into orbit around Titan. The Aerorover will probably use a hot air balloon concept using the waste heat from the MMRTG 500 watts. Orbiter support for the Aerorover is unique to our approach for Titan. Our strategy to use an orbiter is contrary to some studies using just a single probe with balloon. Autonomous operation and navigation of the Aerorover around Titan will be required, which will include descent near to the surface to collect surface samples for analysis (i.e., touch and go technique). The orbiter can provide both relay station and GPS roles for the Aerorover. The Aerorover will have all the instruments needed to sample Titan’s atmosphere, surface, possible methane lakes-rivers, use multi-spectral imagers for surface reconnaissance; to take close up surface images; take core samples and deploy seismometers during landing phase. Both active and passive broadband remote sensing techniques will be used for surface topography, winds and composition measurements.

Dynamics and control of robotic spacecrafts for the transportation of flexible elements

International Nuclear Information System (INIS)

Wen, Hao; Chen, Ti; Yu, Bensong; Jin, Dongping

2016-01-01

The technology of robotic spacecrafts has been identified as one of the most appealing solutions to the on-orbit construction of large space structures in future space missions. As a prerequisite of a successful on-orbit construction, it is needed to use small autonomous spacecrafts for the transportation of flexible elements. To this end, the paper presents an energy-based scheme to control a couple of robotic spacecrafts carrying a flexible slender structure to its desired position. The flexible structure is modelled as a long beam based on the formulation of absolute nodal coordinates to account for the geometrical nonlinearity due to large displacement. Meanwhile, the robotic spacecrafts are actuated on their rigid-body degrees of freedom and modelled as two rigid bodies attached to the flexible beam. The energy-based controller is designed using the technique of energy shaping and damping injection such that translational and rotational maneuvers can be achieved with the suppression of the flexible vibrations of the beam. Finally, numerical case studies are performed to demonstrate the proposed schemes. (paper)

Multi-spacecraft observations of ICMEs propagating beyond Earth orbit during MSL/RAD flight and surface phases

Science.gov (United States)

von Forstner, J.; Guo, J.; Wimmer-Schweingruber, R. F.; Hassler, D.; Temmer, M.; Vrsnak, B.; Čalogović, J.; Dumbovic, M.; Lohf, H.; Appel, J. K.; Heber, B.; Steigies, C. T.; Zeitlin, C.; Ehresmann, B.; Jian, L. K.; Boehm, E.; Boettcher, S. I.; Burmeister, S.; Martin-Garcia, C.; Brinza, D. E.; Posner, A.; Reitz, G.; Matthiae, D.; Rafkin, S. C.; weigle, G., II; Cucinotta, F.

2017-12-01

The propagation of interplanetary coronal mass ejections (ICMEs) between Earth's orbit (1 AU) and Mars ( 1.5 AU) has been studied with their propagation speed estimated from both measurements and simulations. The enhancement of the magnetic fields related to ICMEs and their shock fronts cause so-called Forbush decreases, which can be detected as a reduction of galactic cosmic rays measured on-ground or on a spacecraft. We have used galactic cosmic ray (GCR) data from in-situ measurements at Earth, from both STEREO A and B as well as the GCR measurement by the Radiation Assessment Detector (RAD) instrument onboard Mars Science Laboratory (MSL) on the surface of Mars as well as during its flight to Mars in 2011-2012. A set of ICME events has been selected during the periods when Earth (or STEREO A or B) and MSL locations were nearly aligned on the same side of the Sun in the ecliptic plane (so-called opposition phase). Such lineups allow us to estimate the ICMEs' transit times between 1 AU and the MSL location by estimating the delay time of the corresponding Forbush decreases measured at each location. We investigate the evolution of their propagation speeds after passing Earth's orbit and find that the deceleration of ICMEs due to their interaction with the ambient solar wind continues beyond 1 AU. The results are compared to simulation data obtained from two CME propagation models, namely the Drag-Based Model (DBM) and the WSA-ENLIL plus cone model.

CASSINI V/E/J/S/SS RPWS EDITED WAVEFORM FULL RES V1.0

Data.gov (United States)

National Aeronautics and Space Administration — The Cassini Radio and Plasma Wave Science (RPWS) edited full resolution data set includes all waveform data for the entire Cassini mission. This data set includes...

SOLAR OCCULTATION BY TITAN MEASURED BY CASSINI/UVIS

Energy Technology Data Exchange (ETDEWEB)

Capalbo, Fernando J.; Benilan, Yves [Laboratoire Interuniversitaire des Systemes Atmospheriques (LISA), UMR 7583 du CNRS, Universites Paris Est Creteil (UPEC) and Paris Diderot - UPD, 61 avenue du General de Gaulle, 94010 Creteil Cedex (France); Yelle, Roger V.; Koskinen, Tommi T.; Sandel, Bill R. [Lunar and Planetary Laboratory, University of Arizona, 1629 E. University Blvd., Tucson, AZ 85721 (United States); Holsclaw, Gregory M.; McClintock, William E., E-mail: fernando.capalbo@lisa.u-pec.fr [Laboratory for Atmospheric and Space Physics, University of Colorado, 3665 Discovery Drive, Boulder, CO 80303 (United States)

2013-04-01

We present the first published analysis of a solar occultation by Titan's atmosphere measured by the Ultraviolet Imaging Spectrograph on board Cassini. The data were measured during flyby T53 in 2009 April and correspond to latitudes between 21 Degree-Sign and 28 Degree-Sign south. The analysis utilizes the absorption of two solar emission lines (584 A and 630 A) in the ionization continuum of the N{sub 2} absorption cross section and solar emission lines around 1085 A where absorption is due to CH{sub 4}. The measured transmission at these wavelengths provides a direct estimate of the N{sub 2} and CH{sub 4} column densities along the line of sight from the spacecraft to the Sun, which we inverted to obtain the number densities. The high signal-to-noise ratio of the data allowed us to retrieve density profiles in the altitude range 1120-1400 km for nitrogen and 850-1300 km for methane. We find an N{sub 2} scale height of {approx}76 km and a temperature of {approx}153 K. Our results are in general agreement with those from previous work, although there are some differences. Particularly, our profiles agree, considering uncertainties, with the density profiles derived from the Voyager 1 Ultraviolet Spectrograph data, and with in situ measurements by the Ion Neutral Mass Spectrometer with revised calibration.

CASSINI MAGNETOMETER RAW DATA V1.0

Data.gov (United States)

National Aeronautics and Space Administration — This data set contains magnetic-field data acquired during the cruise and tour phases of the Cassini mission to Saturn. The data set begins with data collected on 16...

Comparison of the Cloud Morphology Spatial Structure Between Jupiter and Saturn Using JunoCam and Cassini ISS

Science.gov (United States)

Garland, Justin; Sayanagi, Kunio M.; Blalock, John J.; Gunnarson, Jacob; McCabe, Ryan M.; Gallego, Angelina; Hansen, Candice; Orton, Glenn S.

2017-10-01

We present an analysis of the spatial-scales contained in the cloud morphology of Jupiter’s southern high latitudes using images captured by JunoCam in 2016 and 2017, and compare them to those on Saturn using images captured using the Imaging Science Subsystem (ISS) on board the Cassini orbiter. For Jupiter, the characteristic spatial scale of cloud morphology as a function of latitude is calculated from images taken in three visual (600-800, 500-600, 420-520 nm) bands and a near-infrared (880- 900 nm) band. In particular, we analyze the transition from the banded structure characteristic of Jupiter’s mid-latitudes to the chaotic structure of the polar region. We apply similar analysis to Saturn using images captured using Cassini ISS. In contrast to Jupiter, Saturn maintains its zonally organized cloud morphology from low latitudes up to the poles, culminating in the cyclonic polar vortices centered at each of the poles. By quantifying the differences in the spatial scales contained in the cloud morphology, our analysis will shed light on the processes that control the banded structures on Jupiter and Saturn. Our work has been supported by the following grants: NASA PATM NNX14AK07G, NASA MUREP NNX15AQ03A, and NSF AAG 1212216.

ATS-6 engineering performance report. Volume 2: Orbit and attitude controls

Science.gov (United States)

Wales, R. O. (Editor)

1981-01-01

Attitude control is reviewed, encompassing the attitude control subsystem, spacecraft attitude precision pointing and slewing adaptive control experiment, and RF interferometer experiment. The spacecraft propulsion system (SPS) is discussed, including subsystem, SPS design description and validation, orbital operations and performance, in-orbit anomalies and contingency operations, and the cesium bombardment ion engine experiment. Thruster failure due to plugging of the propellant feed passages, a major cause for mission termination, are considered among the critical generic failures on the satellite.

A Survey of LIDAR Technology and Its Use in Spacecraft Relative Navigation

Science.gov (United States)

Christian, John A.; Cryan, Scott P.

2013-01-01

This paper provides a survey of modern LIght Detection And Ranging (LIDAR) sensors from a perspective of how they can be used for spacecraft relative navigation. In addition to LIDAR technology commonly used in space applications today (e.g. scanning, flash), this paper reviews emerging LIDAR technologies gaining traction in other non-aerospace fields. The discussion will include an overview of sensor operating principles and specific pros/cons for each type of LIDAR. This paper provides a comprehensive review of LIDAR technology as applied specifically to spacecraft relative navigation. HE problem of orbital rendezvous and docking has been a consistent challenge for complex space missions since before the Gemini 8 spacecraft performed the first successful on-orbit docking of two spacecraft in 1966. Over the years, a great deal of effort has been devoted to advancing technology associated with all aspects of the rendezvous, proximity operations, and docking (RPOD) flight phase. After years of perfecting the art of crewed rendezvous with the Gemini, Apollo, and Space Shuttle programs, NASA began investigating the problem of autonomous rendezvous and docking (AR&D) to support a host of different mission applications. Some of these applications include autonomous resupply of the International Space Station (ISS), robotic servicing/refueling of existing orbital assets, and on-orbit assembly.1 The push towards a robust AR&D capability has led to an intensified interest in a number of different sensors capable of providing insight into the relative state of two spacecraft. The present work focuses on exploring the state-of-the-art in one of these sensors - LIght Detection And Ranging (LIDAR) sensors. It should be noted that the military community frequently uses the acronym LADAR (LAser Detection And Ranging) to refer to what this paper calls LIDARs. A LIDAR is an active remote sensing device that is typically used in space applications to obtain the range to one or more

NASA's New Orbital Debris Engineering Model, ORDEM2010

Science.gov (United States)

Krisko, Paula H.

2010-01-01

This paper describes the functionality and use of ORDEM2010, which replaces ORDEM2000, as the NASA Orbital Debris Program Office (ODPO) debris engineering model. Like its predecessor, ORDEM2010 serves the ODPO mission of providing spacecraft designers/operators and debris observers with a publicly available model to calculate orbital debris flux by current-state-of-knowledge methods. The key advance in ORDEM2010 is the input file structure of the yearly debris populations from 1995-2035 of sizes 10 micron - 1 m. These files include debris from low-Earth orbits (LEO) through geosynchronous orbits (GEO). Stable orbital elements (i.e., those that do not randomize on a sub-year timescale) are included in the files as are debris size, debris number, material density, random error and population error. Material density is implemented from ground-test data into the NASA breakup model and assigned to debris fragments accordingly. The random and population errors are due to machine error and uncertainties in debris sizes. These high-fidelity population files call for a much higher-level model analysis than what was possible with the populations of ORDEM2000. Population analysis in the ORDEM2010 model consists of mapping matrices that convert the debris population elements to debris fluxes. One output mode results in a spacecraft encompassing 3-D igloo of debris flux, compartmentalized by debris size, velocity, pitch, and yaw with respect to spacecraft ram direction. The second output mode provides debris flux through an Earth-based telescope/radar beam from LEO through GEO. This paper compares the new ORDEM2010 with ORDEM2000 in terms of processes and results with examples of specific orbits.

Laser Mode Behavior of the Cassini CIRS Fourier Transform Spectrometer at Saturn

Science.gov (United States)

Brasunas, John C.

2012-01-01

The CIRS Fourier transform spectrometer aboard the NASA/ESA/ASI Cassini orbiter has been acquiring spectra of the Saturnian system since 2004. The CIRS reference interferometer employs a laser diode to trigger the interferogram sampling. Although the control of laser diode drive current and operating temperature are stringent enough to restrict laser wavelength variation to a small fraction of CIRS finest resolution element, the CIRS instrument does need to be restarted every year or two, at which time it may start in a new laser mode. By monitoring the Mylar absorption features in uncalibrated spectra due to the beam splitter Mylar substrate, it can be shown that these jumps are to adjacent modes and that most of the eight-year operation so far is restricted to three adjacent modes. For a given mode, the wavelength stability appears consistent with the stability of the laser diode drive curren.t and operating temperature.

Comprehensive evaluation of attitude and orbit estimation using real earth magnetic field data

Science.gov (United States)

Deutschmann, Julie; Bar-Itzhack, Itzhack

1997-01-01

A single, augmented extended Kalman filter (EKF) which simultaneously and autonomously estimates spacecraft attitude and orbit was developed and tested with simulated and real magnetometer and rate data. Since the earth's magnetic field is a function of time and position, and since time is accurately known, the differences between the computed and measured magnetic field components, as measured by the magnetometers throughout the entire spacecraft's orbit, are a function of orbit and attitude errors. These differences can be used to estimate the orbit and attitude. The test results of the EKF with magnetometer and gyro data from three NASA satellites are presented and evaluated.

STS-49 Endeavour, Orbiter Vehicle (OV) 105, Orbit Team O1 in MCC Bldg 30 FCR

Science.gov (United States)

1992-01-01

STS-49 Endeavour, Orbiter Vehicle (OV) 105, Orbit Team 1 (O1) poses in front of large display screens in JSC's Mission Control Center (MCC) Bldg 30 Flight Control Room (FCR) for group portrait. Lead Flight Director (FD) Granvil A. Pennington stands next to a model of the James Cook's ship, the Endeavour (left). Astronaut and Spacecraft Communicator (CAPCOM) John H. Casper stands at the right of the model.

Robust Spacecraft Component Detection in Point Clouds

Directory of Open Access Journals (Sweden)

Quanmao Wei

2018-03-01

Full Text Available Automatic component detection of spacecraft can assist in on-orbit operation and space situational awareness. Spacecraft are generally composed of solar panels and cuboidal or cylindrical modules. These components can be simply represented by geometric primitives like plane, cuboid and cylinder. Based on this prior, we propose a robust automatic detection scheme to automatically detect such basic components of spacecraft in three-dimensional (3D point clouds. In the proposed scheme, cylinders are first detected in the iteration of the energy-based geometric model fitting and cylinder parameter estimation. Then, planes are detected by Hough transform and further described as bounded patches with their minimum bounding rectangles. Finally, the cuboids are detected with pair-wise geometry relations from the detected patches. After successive detection of cylinders, planar patches and cuboids, a mid-level geometry representation of the spacecraft can be delivered. We tested the proposed component detection scheme on spacecraft 3D point clouds synthesized by computer-aided design (CAD models and those recovered by image-based reconstruction, respectively. Experimental results illustrate that the proposed scheme can detect the basic geometric components effectively and has fine robustness against noise and point distribution density.

Robust Spacecraft Component Detection in Point Clouds.

Science.gov (United States)

Wei, Quanmao; Jiang, Zhiguo; Zhang, Haopeng

2018-03-21

Automatic component detection of spacecraft can assist in on-orbit operation and space situational awareness. Spacecraft are generally composed of solar panels and cuboidal or cylindrical modules. These components can be simply represented by geometric primitives like plane, cuboid and cylinder. Based on this prior, we propose a robust automatic detection scheme to automatically detect such basic components of spacecraft in three-dimensional (3D) point clouds. In the proposed scheme, cylinders are first detected in the iteration of the energy-based geometric model fitting and cylinder parameter estimation. Then, planes are detected by Hough transform and further described as bounded patches with their minimum bounding rectangles. Finally, the cuboids are detected with pair-wise geometry relations from the detected patches. After successive detection of cylinders, planar patches and cuboids, a mid-level geometry representation of the spacecraft can be delivered. We tested the proposed component detection scheme on spacecraft 3D point clouds synthesized by computer-aided design (CAD) models and those recovered by image-based reconstruction, respectively. Experimental results illustrate that the proposed scheme can detect the basic geometric components effectively and has fine robustness against noise and point distribution density.

Reconciling Electrical Properties of Titan's Surface Derived from Cassini RADAR Scatterometer and Radiometer Measurements

Science.gov (United States)

Zebker, H. A.; Wye, L. C.; Janssen, M.; Paganelli, F.; Cassini RADAR Team

2006-12-01

We observe Titan, Saturn's largest moon, using active and passive microwave instruments carried on board the Cassini spacecraft. The 2.2-cm wavelength penetrates the thick atmosphere and provides surface measurements at resolutions from 10-200 km over much of the satellite's surface. The emissivity and reflectivity of surface features are generally anticorrelated, and both values are fairly high. Inversion of either set of data alone yields dielectric constants ranging from 1.5 to 3 or 4, consistent with an icy hydrocarbon or water ice composition. However, the dielectric constants retrieved from radiometric data alone are usually less than those inferred from backscatter measurements, a discrepancy consistent with similar analyses dating back to lunar observations in the 1960's. Here we seek to reconcile Titan's reflectivity and emissivity observations using a single physical model of the surface. Our approach is to calculate the energy scattered by Titan's surface and near subsurface, with the remainder absorbed. In equilibrium the absorption equals the emission, so that both the reflectivity and emissivity are described by the model. We use a form of the Kirchhoff model for modeling surface scatter, and a model based on weak localization of light for the volume scatter. With this model we present dielectric constant and surface roughness parameters that match both sets of Cassini RADAR observations over limited regions on Titan's surface, helping to constrain the composition and roughness of the surface. Most regions display electrical properties consistent with solid surfaces, however some of the darker "lake-like" features at higher latitudes can be modeled as either solid or liquid materials. The ambiguity arises from the limited set of observational angles available.

Huygens begins its final journey into the unknown

Science.gov (United States)

2004-12-01

“Today’s release is another successful milestone in the Cassini/Huygens odyssey”, said Dr David Southwood, ESA’s director of science programmes. “This was an amicable separation after seven years of living together. Our thanks to our partners at NASA for the lift. Each spacecraft will now continue on its own but we expect they’ll keep in touch to complete this amazing mission. Now all our hopes and expectations are focused on getting the first in-situ data from a new world we’ve been dreaming of exploring for decades”. Final stage of a 7-year odyssey The Cassini/Huygens mission, jointly developed by NASA, ESA and the Italian space agency (ASI), began on 15 October 1997, when the composite spacecraft were launched from Cape Canaveral, Florida, atop a Titan 4B/Centaur vehicle. Together, the two probes weighed 5548 kg at launch and became the largest space mission ever sent to the outer planets. To gain sufficient velocity to reach Saturn, they had to conduct four gravity-assist manoeuvres by flying twice by Venus, once by the Earth and once by Jupiter. On 1 July Cassini/Huygens eventually became the first spacecraft to enter an orbit around Saturn. On 17 December, while on its third orbit around the ringed planet, the Cassini orbiter performed a manoeuvre to enter a controlled collision trajectory towards Titan. As planned, a fine tuning of the trajectory took place on 22 December to place Huygens on its nominal entry trajectrory. While Huygens will remain on this trajectory till it plunges into Titan’s atmosphere on 14 January, the orbiter will perform a deflection manoeuvre on 28 December to avoid crashing onto the moon. Today’s separation was achieved by the firing of pyrotechnic devices. Under the action of push-off springs, ramps and rollers, the probe was released at a relative velocity of about 0.3 m/s with a spin rate of 7 rpm. Telemetry data confirming the separation were collected by NASA’s Deep Space Network stations in Madrid, Spain

Investigation of electrodynamic stabilization and control of long orbiting tethers

Science.gov (United States)

Colombo, G.; Arnold, D.

1984-01-01

The state-of-the-art in tether modelling among participants in the Tethered Satellite System (TSS) Program, the slack tether and its behavior, and certain advanced applications of the tether to problems in orbital mechanics are identified. The features and applications of the TSS software set are reviewed. Modelling the slack tether analytically with as many as 50 mass points and the application of this new model to a study of the behavior of a broken tether near the Shuttle are described. A reel control algorithm developed by SAO and examples of its use are described, including an example which also demonstrates the use of the tether in transferring a heavy payload from a low-orbiting Shuttle to a high circular orbit. Capture of a low-orbiting payload by a Space Station in high circular orbit is described. Energy transfer within a dumbbell-type spacecraft by cyclical reeling operations or gravitational effects on the natural elasticity of the connecting tether, it is shown, can circularize the orbit of the spacecraft.

Spacecraft Tests of General Relativity

Science.gov (United States)

Anderson, John D.

1997-01-01

Current spacecraft tests of general relativity depend on coherent radio tracking referred to atomic frequency standards at the ground stations. This paper addresses the possibility of improved tests using essentially the current system, but with the added possibility of a space-borne atomic clock. Outside of the obvious measurement of the gravitational frequency shift of the spacecraft clock, a successor to the suborbital flight of a Scout D rocket in 1976 (GP-A Project), other metric tests would benefit most directly by a possible improved sensitivity for the reduced coherent data. For purposes of illustration, two possible missions are discussed. The first is a highly eccentric Earth orbiter, and the second a solar-conjunction experiment to measure the Shapiro time delay using coherent Doppler data instead of the conventional ranging modulation.

Stability Analysis of Spacecraft Motion in the Vicinity of Asteroids

Data.gov (United States)

National Aeronautics and Space Administration — The objective of my proposal is to determine the stability of a spacecraft when in the vicinity of an asteroid. Orbiting an asteroid is a difficult task. The unique...

ROGER a potential orbital space debris removal system

Science.gov (United States)

Starke, Juergen; Bischof, Bernd; Foth, W.-O.; -J., J.; Günther

The previous activities in the field of On Orbit Servicing studied in the 1990's included in partic-ular the capability of vehicles in GEO to capture and support satellites (mainly communication satellites) to enable repair and continuation of operations, and finally the controlled transfer the target into a permanent graveyard orbit. The specific capture tools for these applications were mostly based on robotic systems to capture and fix the target under specific dynamic constraints (e.g. slowly tumbling target) without damage, and to allow the stabilization, re-orientation and potential repair of the target and subsequent release or transport to the final disposal orbit. Due to the drastically increasing number of debris particularly in the Low Earth Orbits (SSO) the active debris removal is now necessary to counteract to the predicted debris production cascade (Kessler Syndrome), which means the pollution of the total sphere in low earth orbit and not only the SSO area. In most of the debris congresses it was recommended to start removal with the still integrated systems as soon as possible. In the case of large debris objects, the soft capture system can be replaced by a simpler and robust system able to operate from a safe distance to the target and flexible enough to capture and hold different types of targets such as deactivated and/or defective satellites, upper stages and big fragments. These nominally non -cooperative targets might be partially destroyed by the capture process, but the production of additional debris shall be avoided. A major argument for the commercial applications is a multi-target mission potential, which is possible at GEO because the transfer propellant requirement to the disposal orbit and the return to the orbit of the next potential target is relative low (orbits with similar inclination and altitude). The proposed ROGER system is designed as a spacecraft with rendezvous capabilities including inspection in the vicinity of the

Determination of Eros Physical Parameters for Near Earth Asteroid Rendezvous Orbit Phase Navigation

Science.gov (United States)

Miller, J. K.; Antreasian, P. J.; Georgini, J.; Owen, W. M.; Williams, B. G.; Yeomans, D. K.

1995-01-01

Navigation of the orbit phase of the Near Earth steroid Rendezvous (NEAR) mission will re,quire determination of certain physical parameters describing the size, shape, gravity field, attitude and inertial properties of Eros. Prior to launch, little was known about Eros except for its orbit which could be determined with high precision from ground based telescope observations. Radar bounce and light curve data provided a rough estimate of Eros shape and a fairly good estimate of the pole, prime meridian and spin rate. However, the determination of the NEAR spacecraft orbit requires a high precision model of Eros's physical parameters and the ground based data provides only marginal a priori information. Eros is the principal source of perturbations of the spacecraft's trajectory and the principal source of data for determining the orbit. The initial orbit determination strategy is therefore concerned with developing a precise model of Eros. The original plan for Eros orbital operations was to execute a series of rendezvous burns beginning on December 20,1998 and insert into a close Eros orbit in January 1999. As a result of an unplanned termination of the rendezvous burn on December 20, 1998, the NEAR spacecraft continued on its high velocity approach trajectory and passed within 3900 km of Eros on December 23, 1998. The planned rendezvous burn was delayed until January 3, 1999 which resulted in the spacecraft being placed on a trajectory that slowly returns to Eros with a subsequent delay of close Eros orbital operations until February 2001. The flyby of Eros provided a brief glimpse and allowed for a crude estimate of the pole, prime meridian and mass of Eros. More importantly for navigation, orbit determination software was executed in the landmark tracking mode to determine the spacecraft orbit and a preliminary shape and landmark data base has been obtained. The flyby also provided an opportunity to test orbit determination operational procedures that will be

Method and associated apparatus for capturing, servicing, and de-orbiting earth satellites using robotics

Science.gov (United States)

Cepollina, Frank J. (Inventor); Burns, Richard D. (Inventor); Holz, Jill M. (Inventor); Corbo, James E. (Inventor); Jedhrich, Nicholas M. (Inventor)

2009-01-01

This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot.In one embodiment, a vehicle used for carrying out the method of this invention comprises an ejection module, which includes the robot, and a de-orbit module. Once servicing is completed by the robot, the ejection module separates from the de-orbit module, leaving the de-orbit module attached to the satellite for de-orbiting the same at a future time. Upon separation, the ejection module can either de-orbit itself or rendezvous with another satellite for servicing. The ability to de-orbit a spacecraft further allows the opportunity to direct the landing of the spent satellite in a safe location away from population centers, such as the ocean.

Operational factors affecting microgravity levels in orbit

Science.gov (United States)

Olsen, R. E.; Mockovciak, J., Jr.

1980-01-01

Microgravity levels desired for proposed materials processing payloads are fundamental considerations in the design of future space platforms. Disturbance sources, such as aerodynamic drag, attitude control torques, crew motion and orbital dynamics, influence the microgravity levels attainable in orbit. The nature of these effects are assessed relative to platform design parameters such as orbital altitude and configuration geometry, and examples are presented for a representative spacecraft configuration. The possible applications of control techniques to provide extremely low acceleration levels are also discussed.

On the Mitigation of Solar Index Variability for High Precision Orbit Determination in Low Earth Orbit

Science.gov (United States)

2016-09-16

spacecraft state, or solve for an orbit using a Kalman Filter -Smoother (KFS) or Weighted Least Squares Orbit Determination (WLS-OD) process. Early...1 Researchers at the NRL developed the NRLMSISE-00 model in 2002 to better calculate at- mospheric temperature and density profiles for a number of...spectrometer and incoherent scatter data MSIS, 1. N2 density and temperature ,” Journal of Geophysical Research, Vol. 82, No. 16, 1977, pp. 2139–2147

Revisit the modeling of the Saturnian ring atmosphere and ionosphere from the "Cassini Grand Finale" results

Science.gov (United States)

Tseng, W. L.; Johnson, R. E.; Tucker, O. J.; Perry, M. E.; Ip, W. H.

2017-12-01

During the Cassini Grand Finale mission, this spacecraft, for the first time, has done the in-situ measurements of Saturn's upper atmosphere and its rings and provides critical information for understanding the coupling dynamics between the main rings and the Saturnian system. The ring atmosphere is the source of neutrals (i.e., O2, H2, H; Tseng et al., 2010; 2013a), which is primarily generated by photolytic decomposition of water ice (Johnson et al., 2006), and plasma (i.e., O2+ and H2+; Tseng et al., 2011) in the Saturnian magnetosphere. In addition, the main rings have strong interaction with Saturn's atmosphere and ionosphere (i.e., a source of oxygen into Saturn's upper atmosphere and/or the "ring rain" in O'Donoghue et al., 2013). Furthermore, the near-ring plasma environment is complicated by the neutrals from both the seasonally dependent ring atmosphere and Enceladus torus (Tseng et al., 2013b), and, possibly, from small grains from the main and tenuous F and G rings (Johnson et al.2017). The data now coming from Cassini Grand Finale mission already shed light on the dominant physics and chemistry in this region of Saturn's magnetosphere, for example, the presence of carbonaceous material from meteorite impacts in the main rings and each gas species have similar distribution in the ring atmosphere. We will revisit the details in our ring atmosphere/ionosphere model to study, such as the source mechanism for the organic material and the neutral-grain-plasma interaction processes.

Tidal Control of Jet Eruptions on Enceladus as Observed by Cassini ISS between 2005 and 2007

Science.gov (United States)

Hurford, T. A.; Helfenstein, P.; Spitale, J. N.

2012-01-01

Observations of Enceladus have revealed active jets of material erupting from cracks on its south polar surface. It has previously been proposed that diurnal tidal stress, driven by Enceladus' orbital eccentricity, may actively produce surface movement along these cracks daily and thus may regulate when eruptions occur. Our analysis of the stress on jet source regions identified in Cassini ISS images reveals tidal stress as a plausible controlling mechanism of jet activity. However, the evidence available in the published and preliminary observations of jet activity between 2005 and 2007 may not be able to solidify the link between tidal stress and eruptions from fissures. Ongoing, far more comprehensive analyses based on recent, much higher resolution jetting observations have the potential to prove otherwise.

Multiple spacecraft configuration designs for coordinated flight missions

Science.gov (United States)

Fumenti, Federico; Theil, Stephan

2018-06-01

Coordinated flight allows the replacement of a single monolithic spacecraft with multiple smaller ones, based on the principle of distributed systems. According to the mission objectives and to ensure a safe relative motion, constraints on the relative distances need to be satisfied. Initially, differential perturbations are limited by proper orbit design. Then, the induced differential drifts can be properly handled through corrective maneuvers. In this work, several designs are surveyed, defining the initial configuration of a group of spacecraft while counteracting the differential perturbations. For each of the investigated designs, focus is placed upon the number of deployable spacecraft and on the possibility to ensure safe relative motion through station keeping of the initial configuration, with particular attention to the required Δ V budget and the constraints violations.

Interior properties of the inner Saturnian moons from space astrometry data

Science.gov (United States)

Lainey, Valery; Noyelles, Benoît; Cooper, Nick; Murray, Carl; Park, Ryan; Rambaux, Nicolas

2018-04-01

During thirteen years in orbit around Saturn before its final plunge, the Cassini spacecraft provided more than ten thousand astrometric measurements. Such large amounts of accurate data enable the search for extremely faint signals in the orbital motion of the moons. Among those, the detection of the dynamical feedback of the rotation of the inner moons of Saturn on their respective orbits becomes possible. Using all the currently available astrometric data associated with Atlas, Prometheus, Pandora, Janus and Epimetheus, we provide a detailed analysis of the ISS data, with special emphasis on their statistical behavior and source of biases. Then, we try quantifying the physical librations of Prometheus, Pandora, Epimetheus and Janus from the monitoring of their orbits. Last, we show how introducing measurements directly derived from imaging can provide tighter constraints on these quantities.

Stability and evolution of orbits around the binary asteroid 175706 (1996 FG3): Implications for the MarcoPolo-R mission

Science.gov (United States)

Hussmann, Hauke; Oberst, Jürgen; Wickhusen, Kai; Shi, Xian; Damme, Friedrich; Lüdicke, Fabian; Lupovka, Valery; Bauer, Sven

2012-09-01

In support of the MarcoPolo-R mission, we have carried out numerical simulations of spacecraft trajectories about the binary asteroid 175706 (1996 FG3) under the influence of solar radiation pressure. We study the effects of (1) the asteroid's mass, shape, and rotational parameters, (2) the secondary's mass, shape, and orbit parameters, (3) the spacecraft's mass, surface area, and reflectivity, and (4) the time of arrival, and therefore the relative position to the sun and planets. We have considered distance regimes between 5 and 20 km, the typical range for a detailed characterization of the asteroids - primary and secondary - with imaging systems, spectrometers and by laser altimetry. With solar radiation pressure and gravity forces of the small asteroid competing, orbits are found to be unstable, in general. However, limited orbital stability can be found in the so-called Self-Stabilized Terminator Orbits (SSTO), where initial orbits are circular, orbital planes are oriented approximately perpendicular to the solar radiation pressure, and where the orbital plane of the spacecraft is shifted slightly (between 0.2 and 1 km) from the asteroid in the direction away from the sun. Under the effect of radiation pressure, the vector perpendicular to the orbit plane is observed to follow the sun direction. Shape and rotation parameters of the asteroid as well as gravitational perturbations by the secondary (not to mention sun and planets) were found not to affect the results. Such stable orbits may be suited for long radio tracking runs, which will allow for studying the gravity field. As the effect of the solar radiation pressure depends on the spacecraft mass, shape, and albedo, good knowledge of the spacecraft model and persistent monitoring of the spacecraft orientation are required.

Effort to recover SOHO spacecraft continue as investigation board focuses on most likely causes

Science.gov (United States)

1998-07-01

in its orbit around the Sun, the orientation of the panels with respect to the Sun should gradually change. The orbit of the spacecraft and the seasonal change in the spacecraft-Sun alignment should result in the increased solar illumination of the spacecraft solar arrays over the next few months. The engineers predict that in late September 1998, illumination of the solar arrays and, consequently, power supplied to the spacecraft, should approach a maximum. The probability of successfully establishing contact reaches a maximum at this point. After this time, illumination of the solar arrays gradually diminishes as the spacecraft-Sun alignment continues to change. In an attempt to recover SOHO as soon as possible, the Flight Operations Team is uplinking commands to the spacecraft via NASA's Deep Space Network, managed by NASA's Jet Propulsion Laboratory, Pasadena, CA, approximately 12 hours per day with no success to date. A recovery plan is under development by ESA and NASA to provide for orderly restart of the spacecraft and to mitigate risks involved. The recovery of the Olympus spacecraft by ESA in 1991 under similar conditions leads to optimism that the SOHO spacecraft may be recoverable once contact is re-established. In May 1991, ESA's Olympus telecommunications satellite experienced a similar major anomaly which resulted in the loss of attitude, leading to intermittent power availability. As a consequence, there was inadequate communication, and the batteries and fuel froze. From analysis of the data available prior to the loss, there was confidence that the power situation would improve over the coming months. A recovery plan was prepared, supported by laboratory tests, to assess the characteristics of thawing batteries and propellants. Telecommand access of Olympus was regained four weeks later, and batteries and propellant tanks were thawed out progressively over the next four weeks. The attitude was then fully recovered and the payload switched back on

Noncircular features in Saturn's rings IV: Absolute radius scale and Saturn's pole direction

Science.gov (United States)

French, Richard G.; McGhee-French, Colleen A.; Lonergan, Katherine; Sepersky, Talia; Jacobson, Robert A.; Nicholson, Philip D.; Hedman, Mathew M.; Marouf, Essam A.; Colwell, Joshua E.

2017-07-01

We present a comprehensive solution for the geometry of Saturn's ring system, based on orbital fits to an extensive set of occultation observations of 122 individual ring edges and gaps. We begin with a restricted set of very high quality Cassini VIMS, UVIS, and RSS measurements for quasi-circular features in the C and B rings and the Cassini Division, and then successively add suitably weighted additional Cassini and historical occultation measurements (from Voyager, HST and the widely-observed 28 Sgr occultation of 3 Jul 1989) for additional non-circular features, to derive an absolute radius scale applicable across the entire classical ring system. As part of our adopted solution, we determine first-order corrections to the spacecraft trajectories used to determine the geometry of individual occultation chords. We adopt a simple linear model for Saturn's precession, and our favored solution yields a precession rate on the sky n^˙P = 0.207 ± 0 .006‧‧yr-1 , equivalent to an angular rate of polar motion ΩP = 0.451 ± 0 .014‧‧yr-1 . The 3% formal uncertainty in the fitted precession rate is approaching the point where it can provide a useful constraint on models of Saturn's interior, although realistic errors are likely to be larger, given the linear approximation of the precession model and possible unmodeled systematic errors in the spacecraft ephemerides. Our results are largely consistent with independent estimates of the precession rate based on historical RPX times (Nicholson et al., 1999 AAS/Division for Planetary Sciences Meeting Abstracts #31 31, 44.01) and from theoretical expectations that account for Titan's 700-yr precession period (Vienne and Duriez 1992, Astronomy and Astrophysics 257, 331-352). The fitted precession rate based on Cassini data only is somewhat lower, which may be an indication of unmodeled shorter term contributions to Saturn's polar motion from other satellites, or perhaps the result of inconsistencies in the assumed

Manifold dynamics in the Earth-Moon system via isomorphic mapping with application to spacecraft end-of-life strategies

Science.gov (United States)

Pontani, Mauro; Giancotti, Marco; Teofilatto, Paolo

2014-12-01

Recently, manifold dynamics has assumed an increasing relevance for analysis and design of low-energy missions, both in the Earth-Moon system and in alternative multibody environments. With regard to lunar missions, exterior and interior transfers, based on the transit through the regions where the collinear libration points L1 and L2 are located, have been studied for a long time and some space missions have already taken advantage of the results of these studies. This paper is focused on the definition and use of a special isomorphic mapping for low-energy mission analysis. A convenient set of cylindrical coordinates is employed to describe the spacecraft dynamics (i.e. position and velocity), in the context of the circular restricted three-body problem, used to model the spacecraft motion in the Earth-Moon system. This isomorphic mapping of trajectories allows the identification and intuitive representation of periodic orbits and of the related invariant manifolds, which correspond to tubes that emanate from the curve associated with the periodic orbit. Heteroclinic connections, i.e. the trajectories that belong to both the stable and the unstable manifolds of two distinct periodic orbits, can be easily detected by means of this representation. This paper illustrates the use of isomorphic mapping for finding (a) periodic orbits, (b) heteroclinic connections between trajectories emanating from two Lyapunov orbits, the first at L1, and the second at L2, and (c) heteroclinic connections between trajectories emanating from the Lyapunov orbit at L1 and from a particular unstable lunar orbit. Heteroclinic trajectories are asymptotic trajectories that travels at zero-propellant cost. In practical situations, a modest delta-v budget is required to perform transfers along the manifolds. This circumstance implies the possibility of performing complex missions, by combining different types of trajectory arcs belonging to the manifolds. This work studies also the possible

Link calibrations for the TDRSS orbiting VLBI experiment

International Nuclear Information System (INIS)

Edwards, C.D.

1987-01-01

The first successful interferometric observations of extragalactic radio sources using an orbiting antenna as one of the observing stations were achieved in July and August 1986 using the TDRSS. The technical obstacles to maintaining phase coherence between the orbiting antenna and the ground stations are reviewed, with an emphasis on the effects of spacecraft motion. An analysis of the interferometric delay and phase reveals the signature of errors in the spacecraft ephemeris. Various calibration schemes are discussed, including the use of a ground beacon at White Sands to calibrate the communications link between White Sands and the TDRSE satellite. Using all available calibrations, coherence of 84 percent was achieved for 700-sec integrations

Earth Orbiting Support Systems for commercial low Earth orbit data relay: Assessing architectures through tradespace exploration

Science.gov (United States)

Palermo, Gianluca; Golkar, Alessandro; Gaudenzi, Paolo

2015-06-01

As small satellites and Sun Synchronous Earth Observation systems are assuming an increased role in nowadays space activities, including commercial investments, it is of interest to assess how infrastructures could be developed to support the development of such systems and other spacecraft that could benefit from having a data relay service in Low Earth Orbit (LEO), as opposed to traditional Geostationary relays. This paper presents a tradespace exploration study of the architecture of such LEO commercial satellite data relay systems, here defined as Earth Orbiting Support Systems (EOSS). The paper proposes a methodology to formulate architectural decisions for EOSS constellations, and enumerate the corresponding tradespace of feasible architectures. Evaluation metrics are proposed to measure benefits and costs of architectures; lastly, a multicriteria Pareto criterion is used to downselect optimal architectures for subsequent analysis. The methodology is applied to two case studies for a set of 30 and 100 customer-spacecraft respectively, representing potential markets for LEO services in Exploration, Earth Observation, Science, and CubeSats. Pareto analysis shows how increased performance of the constellation is always achieved by an increased node size, as measured by the gain of the communications antenna mounted on EOSS spacecraft. On the other hand, nonlinear trends in optimal orbital altitude, number of satellites per plane, and number of orbital planes, are found in both cases. An upward trend in individual node memory capacity is found, although never exceeding 256 Gbits of onboard memory for both cases that have been considered, assuming the availability of a polar ground station for EOSS data downlink. System architects can use the proposed methodology to identify optimal EOSS constellations for a given service pricing strategy and customer target, thus identifying alternatives for selection by decision makers.

In-Flight Operation of the Dawn Ion Propulsion System Through Survey Science Orbit at Ceres

Science.gov (United States)

Garner, Charles E.; Rayman, Marc D.

2015-01-01

The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt objects, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H- 9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide a total delta V of 11 km/s for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer between Ceres science orbits. Full-power thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional delta V of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. From July 2011 through September 2012 the IPS was used to transfer to all the different science orbits at Vesta and to escape from Vesta orbit. Cruise for a rendezvous with Ceres began in September 2012 and concluded with the start of the approach to Ceres phase on December 26, 2015, leading to orbit capture on March 6, 2015. Deterministic thrusting continued during approach to place the spacecraft in its first science orbit, called RC3, which was achieved on April 23, 2015. Following science operations at RC3 ion thrusting was resumed for twenty-five days leading to arrival to the next science orbit, called survey orbit, on June 3, 2015. The IPS will be used for all subsequent orbit transfers and trajectory correction maneuvers until completion of the primary mission in approximately June 2016. To date the IPS has been operated for over 46,774 hours, consumed approximately 393 kg of xenon, and provided

Collision risk investigation for an operational spacecraft caused by space debris

Science.gov (United States)

Zhang, Binbin; Wang, Zhaokui; Zhang, Yulin

2017-04-01

The collision probability between an operational spacecraft and a population of space debris is investigated. By dividing the 3-dimensional operational space of the spacecraft into several space volume cells (SVC) and proposing a boundary selection method to calculate the collision probability in each SVC, the distribution of the collision risk, as functions of the time, the orbital height, the declination, the impact elevation, the collision velocity, etc., can be obtained. Thus, the collision risk could be carefully evaluated over a time span for the general orbital configurations of the spacecraft and the space debris. As an application, the collision risk for the Tiangong-2 space laboratory caused by the cataloged space debris is discussed and evaluated. Results show that most of the collision threat comes from the front left and front right in Tiangong-2's local, quasi-horizontal plane. And the collision probability will also accumulate when Tiangong-2 moves to the largest declinations (about {±} 42°). As a result, the manned space activities should be avoided at those declinations.

Modeling Attitude Dynamics in Simulink: A Study of the Rotational and Translational Motion of a Spacecraft Given Torques and Impulses Generated by RMS Hand Controllers

Science.gov (United States)

Mauldin, Rebecca H.

2010-01-01

In order to study and control the attitude of a spacecraft, it is necessary to understand the natural motion of a body in orbit. Assuming a spacecraft to be a rigid body, dynamics describes the complete motion of the vehicle by the translational and rotational motion of the body. The Simulink Attitude Analysis Model applies the equations of rigid body motion to the study of a spacecraft?s attitude in orbit. Using a TCP/IP connection, Matlab reads the values of the Remote Manipulator System (RMS) hand controllers and passes them to Simulink as specified torque and impulse profiles. Simulink then uses the governing kinematic and dynamic equations of a rigid body in low earth orbit (LE0) to plot the attitude response of a spacecraft for five seconds given known applied torques and impulses, and constant principal moments of inertia.

The History of Orbiter Corrosion Control (1981 - 2011)

Science.gov (United States)

Russell, Richard W.

2014-01-01

After 135 missions and 30 years the Orbiter fleet was retired in 2011. Working with Orbiter project management and a world class engineering team the CCRB was successful in providing successful sustaining engineering support for approximately 20 years. Lessons learned from the Orbiter program have aided NASA and contractor engineers in the design and manufacture of new spacecraft so that exploration of space can continue. The Orbiters are proudly being displayed for all the public to see in New York City, Washington D.C., Los Angeles, and at the Kennedy Space Center in Florida.

Investigation of nickel hydrogen battery technology for the RADARSAT spacecraft

Science.gov (United States)

Mccoy, D. A.; Lackner, J. L.

1986-01-01

The low Earth orbit (LEO) operations of the RADARSAT spacecraft require high performance batteries to provide energy to the payload and platform during eclipse period. Nickel Hydrogen cells are currently competing with the more traditional Nickel Cadmium cells for high performance spacecraft applications at geostationary Earth orbit (GEO) and Leo. Nickel Hydrogen cells appear better suited for high power applications where high currents and high Depths of Discharge are required. Although a number of GEO missions have flown with Nickel Hydrogen batteries, it is not readily apparent that the LEO version of the Nickel Hydrogen cell is able to withstand the extended cycle lifetime (5 years) of the RADARSAT mission. The problems associated with Nickel Hydrogen cells are discussed in the contex of RADARSAT mission and a test program designed to characterize cell performance is presented.

Spacecraft computer technology at Southwest Research Institute

Science.gov (United States)

Shirley, D. J.

1993-01-01

Southwest Research Institute (SwRI) has developed and delivered spacecraft computers for a number of different near-Earth-orbit spacecraft including shuttle experiments and SDIO free-flyer experiments. We describe the evolution of the basic SwRI spacecraft computer design from those weighing in at 20 to 25 lb and using 20 to 30 W to newer models weighing less than 5 lb and using only about 5 W, yet delivering twice the processing throughput. Because of their reduced size, weight, and power, these newer designs are especially applicable to planetary instrument requirements. The basis of our design evolution has been the availability of more powerful processor chip sets and the development of higher density packaging technology, coupled with more aggressive design strategies in incorporating high-density FPGA technology and use of high-density memory chips. In addition to reductions in size, weight, and power, the newer designs also address the necessity of survival in the harsh radiation environment of space. Spurred by participation in such programs as MSTI, LACE, RME, Delta 181, Delta Star, and RADARSAT, our designs have evolved in response to program demands to be small, low-powered units, radiation tolerant enough to be suitable for both Earth-orbit microsats and for planetary instruments. Present designs already include MIL-STD-1750 and Multi-Chip Module (MCM) technology with near-term plans to include RISC processors and higher-density MCM's. Long term plans include development of whole-core processors on one or two MCM's.

CASSINI SCALAR MAGNETOMETER CALIB DATA V1.0

Data.gov (United States)

National Aeronautics and Space Administration — This data set contains magnetic-field data acquired during the cruise and tour phases of the Cassini mission to Saturn. Data collection began on 16 August (day 228),...

Ascent performance feasibility for next-generation spacecraft

Science.gov (United States)

Mancuso, Salvatore Massimo

This thesis deals with the optimization of the ascent trajectories for single-stage suborbital (SSSO), single-stage-to-orbit (SSTO), and two-stage-to-orbit (TSTO) rocket-powered spacecraft. The maximum payload weight problem has been solved using the sequential gradient-restoration algorithm. For the TSTO case, some modifications to the original version of the algorithm have been necessary in order to deal with discontinuities due to staging and the fact that the functional being minimized depends on interface conditions. The optimization problem is studied for different values of the initial thrust-to-weight ratio in the range 1.3 to 1.6, engine specific impulse in the range 400 to 500 sec, and spacecraft structural factor in the range 0.08 to 0.12. For the TSTO configuration, two subproblems are studied: uniform structural factor between stages and nonuniform structural factor between stages. Due to the regular behavior of the results obtained, engineering approximations have been developed which connect the maximum payload weight to the engine specific impulse and spacecraft structural factor; in turn, this leads to useful design considerations. Also, performance sensitivity to the scale of the aerodynamic drag is studied, and it is shown that its effect on payload weight is relatively small, even for drag changes approaching ± 50%. The main conclusions are that: the design of a SSSO configuration appears to be feasible; the design of a SSTO configuration might be comfortably feasible, marginally feasible, or unfeasible, depending on the parameter values assumed; the design of a TSTO configuration is not only feasible, but its payload appears to be considerably larger than that of a SSTO configuration. Improvements in engine specific impulse and spacecraft structural factor are desirable and crucial for SSTO feasibility; indeed, it appears that aerodynamic improvements do not yield significant improvements in payload weight.

Multi-spacecraft observations of solar hard X-ray bursts

International Nuclear Information System (INIS)

Kane, S.R.

1981-01-01

The role of multi-spacecraft observations in solar flare research is examined from the point of view of solar hard X-ray bursts and their implications with respect to models of the impulsive phase. Multi-spacecraft measurements provide a stereoscopic view of the flare region, and hence represent the only direct method of measuring directivity of X-rays. In absence of hard X-ray imaging instruments with high spatial and temporal resolution, multi-spacecraft measurements provide the only means of determining the radial (vertical) structure of the hard X-ray source. This potential of the multi-spacecraft observations is illustrated with an analysis of the presently available observations of solar hard X-ray bursts made simultaneously by two or more of the following spacecraft: International Sun Earth Explorer-3 (ISEE-3), Pioneer Venus Orbiter (PVO), Helios-B and High Energy Astrophysical Observatory-A (HEAO-A). In particular, some conclusions have been drawn about the spatial structure and directivity of 50-100 keV X-rays from impulsive flares. Desirable features of future multi-spacecraft missions are briefly discussed followed by a short description of the hard X-ray experiment on the International Solar Polar Mission which has been planned specifically for multi-spacecraft observations of the Sun. (orig.)

CONSTRAINING SATURN'S CORE PROPERTIES BY A MEASUREMENT OF ITS MOMENT OF INERTIA-IMPLICATIONS TO THE CASSINI SOLSTICE MISSION

International Nuclear Information System (INIS)

Helled, R.

2011-01-01

Knowledge of Saturn's axial moment of inertia can provide valuable information on its internal structure. We suggest that Saturn's angular momentum may be determined by the Solstice Mission (Cassini XXM) by measuring Saturn's pole precession rate and the Lense-Thirring acceleration on the spacecraft, and therefore put constraints on Saturn's moment of inertia. It is shown that Saturn's moment of inertia can change up to ∼2% due to different core properties. However, a determination of Saturn's rotation rate is required to constrain its axial moment of inertia. A change of about seven minutes in rotation period leads to a similar uncertainty in the moment of inertia value as different core properties (mass, radius). A determination of Saturn's angular momentum and rotation period by the Solstice Mission could reveal important information on Saturn's internal structure, in particular, its core properties.

Risk of spacecraft on-orbit obsolescence: Novel framework, stochastic modeling, and implications

Science.gov (United States)

Dubos, Gregory F.; Saleh, Joseph H.

2010-07-01

The Government Accountability Office (GAO) has repeatedly noted the difficulties encountered by the Department of Defense (DOD) in keeping its acquisition of space systems on schedule and within budget. Among the recommendations provided by GAO, a minimum Technology Readiness Level (TRL) for technologies to be included in the development of a space system is advised. The DOD considers this recommendation impractical arguing that if space systems were designed with only mature technologies (high TRL), they would likely become obsolete on-orbit fairly quickly. The risk of on-orbit obsolescence is a key argument in the DOD's position for dipping into low technology maturity for space acquisition programs, but this policy unfortunately often results in the cost growth and schedule slippage criticized by the GAO. The concept of risk of on-orbit obsolescence has remained qualitative to date. In this paper, we formulate a theory of risk of on-orbit obsolescence by building on the traditional notion of obsolescence and adapting it to the specificities of space systems. We develop a stochastic model for quantifying and analyzing the risk of on-orbit obsolescence, and we assess, in its light, the appropriateness of DOD's rationale for maintaining low TRL technologies in its acquisition of space assets as a strategy for mitigating on-orbit obsolescence. Our model and results contribute one step towards the resolution of the conceptual stalemate on this matter between the DOD and the GAO, and we hope will inspire academics to further investigate the risk of on-orbit obsolescence.

Ultra High Resolution Imaging of Enceladus Tiger Stripe Thermal Emission with Cassini CIRS

Science.gov (United States)

Spencer, John R.; Gorius, Nicolas; Howett, Carly; Verbiscer, Anne J.; Cassini CIRS Team

2017-10-01

In October 2015, Cassini flew within 48 km of Enceladus’ south pole. The spacecraft attitude was fixed during the flyby, but the roll angle of the spacecraft was chosen so that the remote sensing instrument fields of view passed over Damascus, Baghdad, and Cairo Sulci. The Composite Infrared Spectrometer (CIRS) instrument obtained a single interferometer scan during the flyby, using a special mode, enabled by a flight software update, which bypassed numerical filters to improve the fidelity of the interferograms. This generated a total of 11 interferograms, at 5 contiguous spatial locations for each of the 7 - 9 micron (FP4) and 9 - 17 micron (FP3) focal planes, and a single larger field of view for the 17 - 500 micron focal plane (FP1). Strong spikes were seen in the interferograms when crossing each of the sulci, due to the rapid passage of warm material through the field of view. For FP3 and FP4, the temporal variations of the signals from the 5 contiguous detectors can be used to generated 5-pixel-wide images of the thermal emission, which show excellent agreement between the two focal planes. FP3 and FP4 spatial resolution, limited along track by the 5 msec time sampling of the interferogram, and across track by the CIRS field of view, is a remarkable 40 x 40 meters. At this resolution, the tiger stripe thermal emission shows a large amount of structure, including both continuous emission along the fractures, discrete hot spots less than 100 meters across, and extended emission with complex structure.

Solar Power System Options for the Radiation and Technology Demonstration Spacecraft

Science.gov (United States)

Kerslake, Thomas W.; Haraburda, Francis M.; Riehl, John P.

2000-01-01

The Radiation and Technology Demonstration (RTD) Mission has the primary objective of demonstrating high-power (10 kilowatts) electric thruster technologies in Earth orbit. This paper discusses the conceptual design of the RTD spacecraft photovoltaic (PV) power system and mission performance analyses. These power system studies assessed multiple options for PV arrays, battery technologies and bus voltage levels. To quantify performance attributes of these power system options, a dedicated Fortran code was developed to predict power system performance and estimate system mass. The low-thrust mission trajectory was analyzed and important Earth orbital environments were modeled. Baseline power system design options are recommended on the basis of performance, mass and risk/complexity. Important findings from parametric studies are discussed and the resulting impacts to the spacecraft design and cost.

Dynamic and reduced-dynamic precise orbit determination of satellites in low earth orbits

International Nuclear Information System (INIS)

Swatschina, P.

2009-01-01

The precise positioning of satellites in Low Earth Orbits (LEO) has become a key technology for advanced space missions. Dedicated satellite missions, such as CHAMP, GRACE and GOCE, that aim to map the Earths gravity field and its variation over time with unprecedented accuracy, initiated the demand for highly precise orbit solutions of LEO satellites. Furthermore, a wide range of additional science opportunities opens up with the capability to generate accurate LEO orbits. For all considered satellite missions, the primary measurement system for navigation is a spaceborne GPS receiver. The goal of this thesis is to establish and implement methods for Precise Orbit Determination (POD) of LEO satellites using GPS. Striving for highest precision using yet efficient orbit generation strategies, the attained orbit solutions are aimed to be competitive with the most advanced solutions of other institutions. Dynamic and reduced-dynamic orbit models provide the basic concepts of this work. These orbit models are subsequently adjusted to the highly accurate GPS measurements. The GPS measurements are introduced at the zero difference level in the ionosphere free linear combination. Appropriate procedures for GPS data screening and editing are established to detect erroneous data and to employ measurements of good quality only. For the dynamic orbit model a sophisticated force model, especially designed for LEO satellites, has been developed. In order to overcome the limitations that are induced by the deficiencies of the purely dynamical model, two different types of empirical parameters are introduced into the force model. These reduced-dynamic orbit models allow for the generation of much longer orbital arcs while preserving the spacecraft dynamics to the most possible extent. The two methods for reduced-dynamic orbit modeling are instantaneous velocity changes (pulses) or piecewise constant accelerations. For both techniques highly efficient modeling algorithms are

Leo Spacecraft Charging Design Guidelines: A Proposed NASA Standard

Science.gov (United States)

Hillard, G. B.; Ferguson, D. C.

2004-01-01

Over the past decade, Low Earth Orbiting (LEO) spacecraft have gradually required ever-increasing power levels. As a rule, this has been accomplished through the use of high voltage systems. Recent failures and anomalies on such spacecraft have been traced to various design practices and materials choices related to the high voltage solar arrays. NASA Glenn has studied these anomalies including plasma chamber testing on arrays similar to those that experienced difficulties on orbit. Many others in the community have been involved in a comprehensive effort to understand the problems and to develop practices to avoid them. The NASA Space Environments and Effects program, recognizing the timeliness of this effort, commissioned and funded a design guidelines document intended to capture the current state of understanding. This document, which was completed in the spring of 2003, has been submitted as a proposed NASA standard. We present here an overview of this document and discuss the effort to develop it as a NASA standard.

Stochastic orbital migration of small bodies in Saturn's rings

Science.gov (United States)

Rein, H.; Papaloizou, J. C. B.

2010-12-01

Many small moonlets that create propeller structures have been found in Saturn's rings by the Cassini spacecraft. We study the dynamical evolution of such 20-50 m sized bodies, which are embedded in Saturn's rings. We estimate the importance of various interaction processes with the ring particles on the moonlet's eccentricity and semi-major axis analytically. For low ring surface densities, the main effects on the evolution of the eccentricity and the semi-major axis are found to be caused by collisions and the gravitational interaction with particles in the vicinity of the moonlet. For high surface densities, the gravitational interaction with self-gravity wakes becomes important. We also perform realistic three-dimensional, collisional N-body simulations with up to a quarter of a million particles. A new set of pseudo shear periodic boundary conditions is used, which reduces the computational costs by an order of magnitude compared to previous studies. Our analytic estimates are confirmed to within a factor of two. On short timescales the evolution is always dominated by stochastic effects caused by collisions and gravitational interaction with self-gravitating ring particles. These result in a random walk of the moonlet's semi-major axis. The eccentricity of the moonlet quickly reaches an equilibrium value owing to collisional damping. The average change in semi-major axis of the moonlet after 100 orbital periods is 10-100m. This translates to an offset in the azimuthal direction of several hundred kilometres. We expect that such a shift is easily observable. Two movies are only available in electronic form at http://www.aanda.org

Dynamics of space particles and spacecrafts passing by the atmosphere of the Earth.

Science.gov (United States)

Gomes, Vivian Martins; Prado, Antonio Fernando Bertachini de Almeida; Golebiewska, Justyna

2013-01-01

The present research studies the motion of a particle or a spacecraft that comes from an orbit around the Sun, which can be elliptic or hyperbolic, and that makes a passage close enough to the Earth such that it crosses its atmosphere. The idea is to measure the Sun-particle two-body energy before and after this passage in order to verify its variation as a function of the periapsis distance, angle of approach, and velocity at the periapsis of the particle. The full system is formed by the Sun, the Earth, and the particle or the spacecraft. The Sun and the Earth are in circular orbits around their center of mass and the motion is planar for all the bodies involved. The equations of motion consider the restricted circular planar three-body problem with the addition of the atmospheric drag. The initial conditions of the particle or spacecraft (position and velocity) are given at the periapsis of its trajectory around the Earth.

An Analysis of Cassini Observations Regarding the Structure of Jupiter's Equatorial Atmosphere

Science.gov (United States)

Choi, David S.; Simon-Miller, Amy A.

2012-01-01

A variety of intriguing atmospheric phenomena reside on both sides of Jupiter's equator. 5-micron bright hot spots and opaque plumes prominently exhibit dynamic behavior to the north, whereas compact, dark chevron-shaped features and isolated anticyclonic disturbances periodically occupy the southern equatorial latitudes. All of these phenomena are associated with the vertical and meridional perturbations of Rossby waves disturbing the mean atmospheric state. As previous observational analysis and numerical simulations have investigated the dynamics of the region, an examination of the atmosphere's vertical structure though radiative transfer analysis is necessary for improved understanding of this unique environment. Here we present preliminary analysis of a multispectral Cassini imaging data set acquired during the spacecraft's flyby of Jupiter in 2000. We evaluated multiple methane and continuum spectral channels at available viewing angles to improve constraints on the vertical structure of the haze and cloud layers comprising these interesting features. Our preliminary results indicate distinct differences in the structure for both hemispheres. Upper troposphere hazes and cloud layers are prevalent in the northern equatorial latitudes, but are not present in corresponding southern latitudes. Continued analysis will further constrain the precise structure present in these phenomena and the differences between them.

The Changing Surface of Saturn's Titan: Cassini Observations Suggest Active Cryovolcanism

Science.gov (United States)

Nelson, R. M.

2008-12-01

R. M. Nelson(1), L. Kamp(1), R. M. C. Lopes(1), D. L. Matson(1), S. D. Wall(1), R. L. Kirk(2), K. L Mitchell(1), G. Mitri(1), B. W. Hapke(3), M. D. Boryta(4), F. E. Leader(1) , W. D. Smythe(1), K. H. Baines(1), R. Jauman(5), C. Sotin(1), R. N. Clark(6), D. P. Cruikshank(7) , P. Drossart(9), B. J. Buratti(1) , J.Lunine(8), M. Combes(9), G. Bellucci(10), J.-P. Bibring(11), F. Capaccioni(10), P. Cerroni(10), A. Coradini(10), V. Formisano(10), G Filacchione(10), R. Y. Langevin(11), T. B. McCord(12), V. Mennella(13), P. D. Nicholson(14) , B. Sicardy(8) 1-JPL, 4800 Oak Grove Drive, Pasadena CA 91109, 2-USGS, Flagstaff, 3-U Pittsburgh, 4-Mt. Sac Col, 5- DLR, Berlin, 6-USGS Denver, 7-NASA AMES, 8-U Paris-Meudon, 9-Obs de Paris, 10-ISFI-CNR Rome, 11-U Paris -Sud. Orsay, 12-Bear Flt Cntr Winthrop WA, 13-Obs Capodimonte Naples, 14-Cornell U. Several Instruments on the Cassini Saturn Orbiter have been observing the surface of Saturn's moon Titan since mid 2004. The Visual and Infrared Mapping Spectrometer (VIMS) reports that regions near 26oS, 78oW (region 1) and 7oS, 138oW (region 2) exhibit photometric changes consistent with on-going surface activity. These regions are photometrically variable with time(1). Cassini Synthetic Aperture Rader (SAR) has investigated these regions and reports that both of these regions exhibit morphologies consistent with cryovolcanism (2). VIMS observed region 1 eight times and reported that on two occasions the region brightened two-fold and then decreased again on timescales of several weeks. Region 2 was observed on four occasions (Tb-Dec13/2004 ,T8-Oct27/2005, T10-Jan15/2006, T12-Mar18/2006) and exhibited a pronounced change in I/F betweenT8 and T10. Our photometric analysis finds that both regions do not exhibit photometric properties consistent with atmospheric phenomena such as tropospheric clouds. These changes must be at or very near the surface. Radar images of these regions reveal morphology that is consistent with cryovolcanoes. We

SpaceX's Dragon America's next generation spacecraft

CERN Document Server

Seedhouse, Erik

2016-01-01

This book describes Dragon V2, a futuristic vehicle that not only provides a means for NASA to transport its astronauts to the orbiting outpost but also advances SpaceX’s core objective of reusability. A direct descendant of Dragon, Dragon V2 can be retrieved, refurbished and re-launched. It is a spacecraft with the potential to completely revolutionize the economics of an industry where equipment costing hundreds of millions of dollars is routinely discarded after a single use. It was presented by SpaceX CEO Elon Musk in May 2014 as the spaceship that will carry NASA astronauts to the International Space Station as soon as 2016 SpaceX’s Dragon – America’s Next Generation Spacecraft describes the extraordinary feats of engineering and human achievement that have placed this revolutionary spacecraft at the forefront of the launch industry and positioned it as the precursor for ultimately transporting humans to Mars. It describes the design and development of Dragon, provides mission highlights of the f...

Probing interferometric parallax with interplanetary spacecraft

Science.gov (United States)

Rodeghiero, G.; Gini, F.; Marchili, N.; Jain, P.; Ralston, J. P.; Dallacasa, D.; Naletto, G.; Possenti, A.; Barbieri, C.; Franceschini, A.; Zampieri, L.

2017-07-01

We describe an experimental scenario for testing a novel method to measure distance and proper motion of astronomical sources. The method is based on multi-epoch observations of amplitude or intensity correlations between separate receiving systems. This technique is called Interferometric Parallax, and efficiently exploits phase information that has traditionally been overlooked. The test case we discuss combines amplitude correlations of signals from deep space interplanetary spacecraft with those from distant galactic and extragalactic radio sources with the goal of estimating the interplanetary spacecraft distance. Interferometric parallax relies on the detection of wavefront curvature effects in signals collected by pairs of separate receiving systems. The method shows promising potentialities over current techniques when the target is unresolved from the background reference sources. Developments in this field might lead to the construction of an independent, geometrical cosmic distance ladder using a dedicated project and future generation instruments. We present a conceptual overview supported by numerical estimates of its performances applied to a spacecraft orbiting the Solar System. Simulations support the feasibility of measurements with a simple and time-saving observational scheme using current facilities.

SEQ-POINTER: Next generation, planetary spacecraft remote sensing science observation design tool

Science.gov (United States)

Boyer, Jeffrey S.

1994-11-01

Since Mariner, NASA-JPL planetary missions have been supported by ground software to plan and design remote sensing science observations. The software used by the science and sequence designers to plan and design observations has evolved with mission and technological advances. The original program, PEGASIS (Mariners 4, 6, and 7), was re-engineered as POGASIS (Mariner 9, Viking, and Mariner 10), and again later as POINTER (Voyager and Galileo). Each of these programs were developed under technological, political, and fiscal constraints which limited their adaptability to other missions and spacecraft designs. Implementation of a multi-mission tool, SEQ POINTER, under the auspices of the JPL Multimission Operations Systems Office (MOSO) is in progress. This version has been designed to address the limitations experienced on previous versions as they were being adapted to a new mission and spacecraft. The tool has been modularly designed with subroutine interface structures to support interchangeable celestial body and spacecraft definition models. The computational and graphics modules have also been designed to interface with data collected from previous spacecraft, or on-going observations, which describe the surface of each target body. These enhancements make SEQ POINTER a candidate for low-cost mission usage, when a remote sensing science observation design capability is required. The current and planned capabilities of the tool will be discussed. The presentation will also include a 5-10 minute video presentation demonstrating the capabilities of a proto-Cassini Project version that was adapted to test the tool. The work described in this abstract was performed by the Jet Propulsion Laboratory, California Institute of Technology, under contract to the National Aeronautics and Space Administration.

International Symposium on Spacecraft Ground Control and Flight Dynamics, SCD1, Sao Jose dos Campos, Brazil, Feb. 7-11, 1994

Science.gov (United States)

Rozenfeld, Pawel; Kuga, Helio Koiti; Orlando, Valcir

An international symposium on spacecraft flight dynamics and ground control systems produced 85 papers in the areas of attitude determination and control, orbit control, satellite constellation strategies, stationkeeping, spacecraft maneuvering, orbit determination, astrodynamics, ground command and control systems, and mission operations. Several papers included discussions on the application of artificial intelligence, neural networks, expert systems, and ion propulsion. For individual titles, see A95-89098 through A95-89182.

Weight estimates and packaging techniques for the microwave radiometer spacecraft. [shuttle compatible design

Science.gov (United States)

Jensen, J. K.; Wright, R. L.

1981-01-01

Estimates of total spacecraft weight and packaging options were made for three conceptual designs of a microwave radiometer spacecraft. Erectable structures were found to be slightly lighter than deployable structures but could be packaged in one-tenth the volume. The tension rim concept, an unconventional design approach, was found to be the lightest and transportable to orbit in the least number of shuttle flights.

Titan Orbiter Aerorover Mission

Science.gov (United States)

Sittler Jr., E. C.; Acuna, M.; Burchell, M. J.; Coates, A.; Farrell, W.; Flasar, M.; Goldstein, B. E.; Gorevan, S.; Hartle, R. E.; Johnson, W. T. K.

2001-01-01

We propose a combined Titan orbiter and Titan Aerorover mission with an emphasis on both in situ and remote sensing measurements of Titan's surface, atmosphere, ionosphere, and magnetospheric interaction. The biological aspect of the Titan environment will be emphasized by the mission (i.e., search for organic materials which may include simple organics to 'amono' analogues of amino acids and possibly more complex, lightening detection and infrared, ultraviolet, and charged particle interactions with Titan's surface and atmosphere). An international mission is assumed to control costs. NASA will provide the orbiter, launch vehicle, DSN coverage and operations, while international partners will provide the Aerorover and up to 30% of the cost for the scientific instruments through collaborative efforts. To further reduce costs we propose a single PI for orbiter science instruments and a single PI for Aerorover science instruments. This approach will provide single command/data and power interface between spacecraft and orbiter instruments that will have redundant central DPU and power converter for their instruments. A similar approach could be used for the Aerorover. The mission profile will be constructed to minimize conflicts between Aerorover science, orbiter radar science, orbiter radio science, orbiter imaging science, and orbiter fields and particles (FP) science. Additional information is contained in the original extended abstract.

Development of Large-Scale Spacecraft Fire Safety Experiments

DEFF Research Database (Denmark)

Ruff, Gary A.; Urban, David L.; Fernandez-Pello, A. Carlos

2013-01-01

exploration missions outside of low-earth orbit and accordingly, more complex in terms of operations, logistics, and safety. This will increase the challenge of ensuring a fire-safe environment for the crew throughout the mission. Based on our fundamental uncertainty of the behavior of fires in low...... of the spacecraft fire safety risk. The activity of this project is supported by an international topical team of fire experts from other space agencies who conduct research that is integrated into the overall experiment design. The large-scale space flight experiment will be conducted in an Orbital Sciences...

Gravity and Macro-Model Tuning for the Geosat Follow-on Spacecraft

Science.gov (United States)

Lemoine, Frank G.; Rowlands, David D.; Marr, Gregory C.; Zelensky, Nikita P.; Luthcke, Scott B.; Cox, Christopher M.

1999-01-01

The US Navy's GEOSAT Follow-On (GFO) spacecraft was launched on February 10, 1998 and the primary objective of the mission was to map the oceans using a radar altimeter. The spacecraft tracking complement consisted of GPS receivers, a laser retroreflector and Doppler beacons. Since the GPS receivers have not yet returned reliable data, the only means of providing high-quality precise orbits has been though satellite laser ranging (SLR). The spacecraft has been tracked by the international satellite laser ranging network since April 22, 1998, and an average of 7.4 passes per day have been obtained from US and participating foreign stations. Since the predicted radial orbit error due to the gravity field is two to three cm, the largest contributor to the high SLR residuals (7-10 cm RMS for five day arcs) is the mismodelling of the non-conservative forces, not withstanding the development of a three-dimensional eight-panel model and an analytical attitude model for the GFO spacecraft. The SLR residuals show a clear correlation with beta-prime (solar elevation) angle, peaking in mid-August 1998 when the beta-prime angle reached -80 to -90 degrees. In this paper we discuss the tuning of the non-conservative force model, for GFO and report the subsequent addition of the GFO tracking data to the Earth gravity model solutions.

Application of the NASCAP Spacecraft Simulation Tool to Investigate Electrodynamic Tether Current Collection in LEO

Science.gov (United States)

Adams, Mitzi; HabashKrause, Linda

2012-01-01

Recent interest in using electrodynamic tethers (EDTs) for orbital maneuvering in Low Earth Orbit (LEO) has prompted the development of the Marshall ElectroDynamic Tether Orbit Propagator (MEDTOP) model. The model is comprised of several modules which address various aspects of EDT propulsion, including calculation of state vectors using a standard orbit propagator (e.g., J2), an atmospheric drag model, realistic ionospheric and magnetic field models, space weather effects, and tether librations. The natural electromotive force (EMF) attained during a radially-aligned conductive tether results in electrons flowing down the tether and accumulating on the lower-altitude spacecraft. The energy that drives this EMF is sourced from the orbital energy of the system; thus, EDTs are often proposed as de-orbiting systems. However, when the current is reversed using satellite charged particle sources, then propulsion is possible. One of the most difficult challenges of the modeling effort is to ascertain the equivalent circuit between the spacecraft and the ionospheric plasma. The present study investigates the use of the NASA Charging Analyzer Program (NASCAP) to calculate currents to and from the tethered satellites and the ionospheric plasma. NASCAP is a sophisticated set of computational tools to model the surface charging of three-dimensional (3D) spacecraft surfaces in a time-varying space environment. The model's surface is tessellated into a collection of facets, and NASCAP calculates currents and potentials for each one. Additionally, NASCAP provides for the construction of one or more nested grids to calculate space potential and time-varying electric fields. This provides for the capability to track individual particles orbits, to model charged particle wakes, and to incorporate external charged particle sources. With this study, we have developed a model of calculating currents incident onto an electrodynamic tethered satellite system, and first results are shown

Future spacecraft propulsion systems. Enabling technologies for space exploration. 2. ed.

Energy Technology Data Exchange (ETDEWEB)

Czysz, Paul A. [St. Louis Univ., MO (United States). Oliver L. Parks Endowed Chair in Aerospace Engineering; Bruno, Claudio [Univ. degli Studi di Roma (Italy). Dipt. di Meccanica e Aeronautica

2009-07-01

In this second edition of Future Spacecraft Propulsion Systems, the authors demonstrate the need to break free from the old established concepts of expendable rockets, using chemical propulsion, and to develop new breeds of launch vehicle capable of both launching payloads into orbit at a dramatically reduced cost and for sustained operations in low-Earth orbit. The next steps to establishing a permanent 'presence' in the Solar System beyond Earth are the commercialisation of sustained operations on the Moon and the development of advanced nuclear or high-energy space propulsion systems for Solar System exploration out to the boundary of interstellar space. In the future, high-energy particle research facilities may one day yield a very high-energy propulsion system that will take us to the nearby stars, or even beyond. Space is not quiet: it is a continuous series of nuclear explosions that provide the material for new star systems to form and provide the challenge to explore. This book provides an assessment of the industrial capability required to construct and operate the necessary spacecraft. Time and distance communication and control limitations impose robotic constraints. Space environments restrict human sustained presence and put high demands on electronic, control and materials systems. This comprehensive and authoritative book puts spacecraft propulsion systems in perspective, from earth orbit launchers to astronomical/space exploration vehicles. It includes new material on fusion propulsion, new figures and updates and expands the information given in the first edition. (orig.)

Comprehensive Evaluation of Attitude and Orbit Estimation Using Actual Earth Magnetic Field Data

Science.gov (United States)

Deutschmann, Julie K.; Bar-Itzhack, Itzhack Y.

2000-01-01

A single, augmented Extended Kalman Filter (EKF), which simultaneously and autonomously estimates spacecraft attitude and orbit has been developed and successfully tested with real magnetometer and gyro data only. Because the earth magnetic field is a function of time and position, and because time is known quite precisely, the differences between the computed and measured magnetic field components, as measured by the magnetometers throughout the entire spacecraft orbit, are a function of both orbit and attitude errors. Thus, conceivably these differences could be used to estimate both orbit and attitude; an observability study validated this assumption. The results of testing the EKF with actual magnetometer and gyro data, from four satellites supported by the NASA Goddard Space Flight Center (GSFC) Guidance, Navigation, and Control Center, are presented and evaluated. They confirm the assumption that a single EKF can estimate both attitude and orbit when using gyros and magnetometers only.

Modes of uncontrolled rotational motion of the Progress M-29M spacecraft

Science.gov (United States)

Belyaev, M. Yu.; Matveeva, T. V.; Monakhov, M. I.; Rulev, D. N.; Sazonov, V. V.

2018-01-01

We have reconstructed the uncontrolled rotational motion of the Progress M-29M transport cargo spacecraft in the single-axis solar orientation mode (the so-called sunward spin) and in the mode of the gravitational orientation of a rotating satellite. The modes were implemented on April 3-7, 2016 as a part of preparation for experiments with the DAKON convection sensor onboard the Progress spacecraft. The reconstruction was performed by integral statistical techniques using the measurements of the spacecraft's angular velocity and electric current from its solar arrays. The measurement data obtained in a certain time interval have been jointly processed using the least-squares method by integrating the equations of the spacecraft's motion relative to the center of mass. As a result of processing, the initial conditions of motion and parameters of the mathematical model have been estimated. The motion in the sunward spin mode is the rotation of the spacecraft with an angular velocity of 2.2 deg/s about the normal to the plane of solar arrays; the normal is oriented toward the Sun or forms a small angle with this direction. The duration of the mode is several orbit passes. The reconstruction has been performed over time intervals of up to 1 h. As a result, the actual rotational motion of the spacecraft relative to the Earth-Sun direction was obtained. In the gravitational orientation mode, the spacecraft was rotated about its longitudinal axis with an angular velocity of 0.1-0.2 deg/s; the longitudinal axis executed small oscillated relative to the local vertical. The reconstruction of motion relative to the orbital coordinate system was performed in time intervals of up to 7 h using only the angularvelocity measurements. The measurements of the electric current from solar arrays were used for verification.

Fuzzy based attitude controller for flexible spacecraft with on/off thrusters

Science.gov (United States)

Knapp, Roger Glenn

1993-05-01

A fuzzy-based attitude controller is designed for attitude control of a generic spacecraft with on/off thrusters. The controller is comprised of packages of rules dedicated to addressing different objectives (e.g., disturbance rejection, low fuel consumption, avoiding the excitation of flexible appendages, etc.). These rule packages can be inserted or removed depending on the requirements of the particular spacecraft and are parameterized based on vehicle parameters such as inertia or operational parameters such as the maneuvering rate. Individual rule packages can be 'weighted' relative to each other to emphasize the importance of one objective relative to another. Finally, the fuzzy controller and rule packages are demonstrated using the high-fidelity Space Shuttle Interactive On-Orbit Simulator (IOS) while performing typical on-orbit operations and are subsequently compared with the existing shuttle flight control system performance.

Digital image transformation and rectification of spacecraft and radar images

Science.gov (United States)

Wu, S. S. C.

1985-01-01

The application of digital processing techniques to spacecraft television pictures and radar images is discussed. The use of digital rectification to produce contour maps from spacecraft pictures is described; images with azimuth and elevation angles are converted into point-perspective frame pictures. The digital correction of the slant angle of radar images to ground scale is examined. The development of orthophoto and stereoscopic shaded relief maps from digital terrain and digital image data is analyzed. Digital image transformations and rectifications are utilized on Viking Orbiter and Lander pictures of Mars.

Spatial analysis of galactic cosmic ray particles in low earth orbit/near equator orbit using SPENVIS

International Nuclear Information System (INIS)

Suparta, W; Zulkeple, S K

2014-01-01

The space environment has grown intensively harmful to spacecraft and astronauts. Galactic cosmic rays (GCRs) are one of the radiation sources that composed of high energetic particles originated from space and capable of damaging electronic systems through single event upset (SEU) process. In this paper, we analyzed GCR fluxes at different altitudes by using Space Environment Information System (SPENVIS) software and the results are compared to determine their intensities with respect to distance in the Earth's orbit. The altitudes are set at low earth orbit (400 km and 685 km), medium earth orbit (19,100 km and 20,200 km) and high earth orbit (35,793 km and 1,000,000 km). Then, within Low Earth Orbit (LEO) near the equator (NEqO), we used altitude of 685 km to compare GCRs with the intensities of solar particles and trapped particles in the radiation belt to determine the significance of GCRs in the orbit itself.

In-Flight spacecraft magnetic field monitoring using scalar/vector gradiometry

DEFF Research Database (Denmark)

Primdahl, Fritz; Risbo, Torben; Merayo, José M.G.

2006-01-01

Earth magnetic field mapping from planetary orbiting satellites requires a spacecraft magnetic field environment control program combined with the deployment of the magnetic sensors on a boom in order to reduce the measurement error caused by the local spacecraft field. Magnetic mapping missions...... (Magsat, Oersted, CHAMP, SAC-C MMP and the planned ESA Swarm project) carry a vector magnetometer and an absolute scalar magnetometer for in-flight calibration of the vector magnetometer scale values and for monitoring of the inter-axes angles and offsets over time intervals from months to years...... sensors onboard the Oersted satellite. For Oersted, a large difference between the pre-flight determined spacecraft magnetic field and the in-flight estimate exists causing some concern about the general applicability of the dual sensors technique....

The problem of space nuclear power sources collisions with artificial space objects in near-earth orbits

International Nuclear Information System (INIS)

Gafarov, A.A.

1993-01-01

Practically all space objects with onboard nuclear power sources stay in earth satellite orbits with an orbital lifetime long enough to reduce their radioactivity to levels presenting no danger for the Earth population. One of the reasons for orbit lifetime reduction can be collisions with other space objects in near-earth orbits. The possible consequence of collisions can be partial, or even complete, destruction of the spacecraft with an onboard nuclear power source; as well as delivery of additional impulse both to the spacecraft and its fragments. It is shown that collisions in orbit do not cause increase of radiation hazard for the Earth population if there is aerodynamic breakup of nuclear power sources into fragments of safe sizes during atmospheric reentry

Plasma Interactions with Spacecraft. Volume 2, NASCAP-2K Scientific Documentation for Version 4.1

Science.gov (United States)

2011-04-15

surface is taken as the equipotential surface at  = ±ln2. This choice is made because the attracted species is absorbed by the sheath, so we have only...spacecraft-generated plasma environments on spacecraft systems. This document describes the physics and numeric models used in the surface charging...2 2.1 Surface Charging from Orbit Limited Currents

Wuhan University Deep-space Orbit Determination and Gravity Recovery System(WUDOGS and Its Application Analysis

Directory of Open Access Journals (Sweden)

YE Mao

2017-03-01

Full Text Available WUDOGS(Wuhan University deep-space orbit determination and gravity recovery system is a software system designed for deep spacecraft precise orbit determination and planetary gravity recovery, developed independently at Wuhan University. WUDOGS now has the function for Lunar and Mars spacecraft precision orbit determination. Its design pattern and main function are briefly introduced. The cross verification test(CVT between WUDOGS and state of the art planetary precise orbit determination software GEODYN-Ⅱ are elaborated. The results show that:①for orbit propagation, with all the same forces and other configuration, the predicted orbit difference in R,T,N directions are less than 0.3 mm for one month arc, 5×10-3 mm for 2 days arc, compared with GEODYN-Ⅱ;②the difference RMS of computed values of observables for two-way range and two-way range rate is at levels of 0.06 mm and 0.002 mm/s respectively;③for Chinese Chang'E-1 POD, the reconstructed orbit difference between WUDOGS and GEODYN-Ⅱ is at 2 cm level, for ESA MEX POD, the reconstructed orbit difference between WUDOGS and ESA is at 25 m level. Current developing situation of WUDOGS and comparison with international research level show that WUDOGS has a good application prospect, which will be important for meeting the demand of Chinese future planetary exploration and the development of deep space spacecraft POD software.

Characterization of dust aggregates in the vicinity of the Rosetta spacecraft

Science.gov (United States)

Güttler, C.; Hasselmann, P. H.; Li, Y.; Fulle, M.; Tubiana, C.; Kovacs, G.; Agarwal, J.; Sierks, H.; Fornasier, S.; Hofmann, M.; Gutiérrez Marqués, P.; Ott, T.; Drolshagen, E.; Bertini, I.; Osiris Team

2017-09-01

In a Rosetta/OSIRIS imaging activity in June 2015, we have observed the dynamic motion of particles close to the spacecraft. Due to the focal setting of the OSIRIS Wide Angle Camera (WAC), these particles were blurred, which can be used to measure their distances to the spacecraft. We detected 108 dust aggregates over a 130 minutes long sequence, and find that their sizes are around a millimetre and their distances cluster between 2 m and 40 m from the spacecraft. Their number densities are about a factor 10 higher than expected for the overall coma and highly fluctuating. Their velocities are small compared to the spacecraft orbital motion and directed away from the spacecraft, towards the comet. From this we conclude that they have interacted with the spacecraft and assess three possible scenarios. We prefer a scenario where centimeter-sized aggregates collide with the spacecraft and we would observe the fragments. Ablation of a dust layer on the spacecraft's z panel when rotated towards the sun is a reasonable alternative. We could also measure an acceleration for a subset of 18 aggregates, which is directed away from the sun and can be explain by a rocket effect, which requires a minimum ice fraction in the order of 0.1%

Spacecraft TT&C and information transmission theory and technologies

CERN Document Server

Liu, Jiaxing

2015-01-01

Spacecraft TT&C and Information Transmission Theory and Technologies introduces the basic theory of spacecraft TT&C (telemetry, track and command) and information transmission. Combining TT&C and information transmission, the book presents several technologies for continuous wave radar including measurements for range, range rate and angle, analog and digital information transmissions, telecommand, telemetry, remote sensing and spread spectrum TT&C. For special problems occurred in the channels for TT&C and information transmission, the book represents radio propagation features and its impact on orbit measurement accuracy, and the effects caused by rain attenuation, atmospheric attenuation and multi-path effect, and polarization composition technology. This book can benefit researchers and engineers in the field of spacecraft TT&C and communication systems. Liu Jiaxing is a professor at The 10th Institute of China Electronics Technology Group Corporation.

Lay out, test verification and in orbit performance of HELIOS a temperature control system

Science.gov (United States)

Brungs, W.

1975-01-01

HELIOS temperature control system is described. The main design features and the impact of interactions between experiment, spacecraft system, and temperature control system requirements on the design are discussed. The major limitations of the thermal design regarding a closer sun approach are given and related to test experience and performance data obtained in orbit. Finally the validity of the test results achieved with prototype and flight spacecraft is evaluated by comparison between test data, orbit temperature predictions and flight data.

Modeling the angular motion dynamics of spacecraft with a magnetic attitude control system based on experimental studies and dynamic similarity

Science.gov (United States)

Kulkov, V. M.; Medvedskii, A. L.; Terentyev, V. V.; Firsyuk, S. O.; Shemyakov, A. O.

2017-12-01

The problem of spacecraft attitude control using electromagnetic systems interacting with the Earth's magnetic field is considered. A set of dimensionless parameters has been formed to investigate the spacecraft orientation regimes based on dynamically similar models. The results of experimental studies of small spacecraft with a magnetic attitude control system can be extrapolated to the in-orbit spacecraft motion control regimes by using the methods of the dimensional and similarity theory.

Probing Small Lakes on Titan Using the Cassini RADAR Altimeter

Science.gov (United States)

Mastrogiuseppe, M.; Poggiali, V.; Hayes, A.; Lunine, J. I.; Seu, R.; Lorenz, R. D.; Mitri, G.; Mitchell, K. L.; Janssen, M. A.; Casarano, D.; Notarnicola, C.; Le Gall, A. A.

2017-12-01

The T126 Cassini's final flyby of Titan has offered a unique opportunity to observe an area in the Northern Polar terrain, where several small - medium size (10 - 50 km) hydrocarbon lakes are present and have been previously imaged by Cassini. The successful observation allowed the radar to operate at the closest approach over several small lakes, using its altimetry mode for the investigation of depth and liquid composition. Herein we present the result of a dedicate processing previously applied to altimetric data acquired over Ligeia Mare where the radar revealed the bathymetry and composition of the sea [1,2]. We show that, the optimal geometry condition met during the T126 fly-by allowed the radar to probe Titan's lakes revealing that such small liquid bodies can exceed one-hundred meters of depth. [1] M. Mastrogiuseppe et al. (2014, Mar.). The bathymetry of a Titan Sea. Geophysical Research Letters. [Online]. 41 (5), pp. 1432-1437. Available: http://dx.doi.org/10.1002/2013GL058618 [2] M.Mastrogiuseppe et al. (2016, Oct). Radar Sounding Using the Cassini Altimeter: Waveform Modeling and Monte Carlo Approach for Data Inversion of Observations of Titan's Seas, IEEE Transactions On Geoscience And Remote Sensing, Vol. 54, No. 10, doi: 10.1109/TGRS.2016.2563426.

Passive Plasma Contact Mechanisms for Small-Scale Spacecraft

Science.gov (United States)

McTernan, Jesse K.

Small-scale spacecraft represent a paradigm shift in how entities such as academia, industry, engineering firms, and the scientific community operate in space. However, although the paradigm shift produces unique opportunities to build satellites in unique ways for novel missions, there are also significant challenges that must be addressed. This research addresses two of the challenges associated with small-scale spacecraft: 1) the miniaturization of spacecraft and associated instrumentation and 2) the need to transport charge across the spacecraft-environment boundary. As spacecraft decrease in size, constraints on the size, weight, and power of on-board instrumentation increase--potentially limiting the instrument's functionality or ability to integrate with the spacecraft. These constraints drive research into mechanisms or techniques that use little or no power and efficiently utilize existing resources. One limited resource on small-scale spacecraft is outer surface area, which is often covered with solar panels to meet tight power budgets. This same surface area could also be needed for passive neutralization of spacecraft charging. This research explores the use of a transparent, conductive layer on the solar cell coverglass that is electrically connected to spacecraft ground potential. This dual-purpose material facilitates the use of outer surfaces for both energy harvesting of solar photons as well as passive ion collection. Mission capabilities such as in-situ plasma measurements that were previously infeasible on small-scale platforms become feasible with the use of indium tin oxide-coated solar panel coverglass. We developed test facilities that simulate the space environment in low Earth orbit to test the dual-purpose material and the various application of this approach. Particularly, this research is in support of two upcoming missions: OSIRIS-3U, by Penn State's Student Space Programs Lab, and MiTEE, by the University of Michigan. The purpose of

Orbital motions as gradiometers for post-Newtonian tidal effects

Directory of Open Access Journals (Sweden)

Lorenzo eIorio

2014-08-01

Full Text Available The direct long-term changes occurring in the orbital dynamics of a local gravitationally bound binary system S due to the post-Newtonian tidal acceleration caused by an external massive source are investigated. A class of systems made of a test particle m rapidly orbiting with orbital frequency nb an astronomical body of mass M which, in turn, slowly revolves around a distantobject of mass M with orbital frequency nb'<< □ nb is considered. The characteristic frequenciesof the non-Keplerian orbital variations of m and of M itself are assumed to be negligible withrespect to both nb and nb'. General expressions for the resulting Newtonian and post-Newtoniantidal orbital shifts of m are obtained. The future missions BepiColombo and JUICE to Mercuryand Ganymede, respectively, are considered in view of a possible detection. The largest effects,of the order of □ 0:1 □□ 0:5 milliarcseconds per year (mas yr□□1, occur for the Ganymede orbiterof the JUICE mission. Although future improvements in spacecraft tracking and orbit determina14tion might, perhaps, reach the required sensitivity, the systematic bias represented by the otherknown orbital perturbations of both Newtonian and post-Newtonian origin would be overwhel16ming. The realization of a dedicated artificial mini-planetary system to be carried onboard andEarth-orbiting spacecraft is considered as well. Post-Newtonian tidal precessions as large as1 □□ 102 mas yr□□1 could be obtained, but the quite larger Newtonian tidal effects would be amajor source of systematic bias because of the present-day percent uncertainty in the product of the Earth’s mass times the Newtonian gravitational parameter.

Standardization and Economics of Nuclear Spacecraft, Final Report, Phase I, Sense Study

Energy Technology Data Exchange (ETDEWEB)

1973-03-01

Feasibility and cost benefits of nuclear-powered standardized spacecraft are investigated. The study indicates that two shuttle-launched nuclear-powered spacecraft should be able to serve the majority of unmanned NASA missions anticipated for the 1980's. The standard spacecraft include structure, thermal control, power, attitude control, some propulsion capability and tracking, telemetry, and command subsystems. One spacecraft design, powered by the radioisotope thermoelectric generator, can serve missions requiring up to 450 watts. The other spacecraft design, powered by similar nuclear heat sources in a Brayton-cycle generator, can serve missions requiring up to 21000 watts. Design concepts and trade-offs are discussed. The conceptual designs selected are presented and successfully tested against a variety of missions. The thermal design is such that both spacecraft are capable of operating in any earth orbit and any orientation without modification. Three-axis stabilization is included. Several spacecraft can be stacked in the shuttle payload compartment for multi-mission launches. A reactor-powered thermoelectric generator system, operating at an electric power level of 5000 watts, is briefly studied for applicability to two test missions of divers requirements. A cost analysis indicates that use of the two standardized spacecraft offers sizable savings in comparison with specially designed solar-powered spacecraft. There is a duplicate copy.

Searching for Orbits with Minimum Fuel Consumption for Station-Keeping Maneuvers: An Application to Lunisolar Perturbations

Directory of Open Access Journals (Sweden)

Antonio Fernando Bertachini de Almeida Prado

2013-01-01

Full Text Available The present paper has the goal of developing a new criterion to search for orbits that minimize the fuel consumption for station-keeping maneuvers. This approach is based on the integral over the time of the perturbing forces. This integral measures the total variation of velocity caused by the perturbations in the spacecraft, which corresponds to the equivalent variation of velocity that an engine should deliver to the spacecraft to compensate the perturbations and to keep its orbit Keplerian all the time. This integral is a characteristic of the orbit and the set of perturbations considered and does not depend on the type of engine used. In this sense, this integral can be seen as a criterion to select the orbit of the spacecraft. When this value becomes larger, more consumption of fuel is required for the station keeping, and, in this sense, less interesting is the orbit. This concept can be applied to any perturbation. In the present research, as an example, the perturbation caused by a third body is considered. Then, numerical simulations considering the effects of the Sun and the Moon in a satellite around the Earth are shown to exemplify the method.

Pulling it all together: the self-consistent distribution of neutral tori in Saturn's Magnetosphere based on all Cassini observations

Science.gov (United States)

Smith, H. T.; Richardson, J. D.

2017-12-01

Saturn's magnetosphere is unique in that the plumes from the small icy moon, Enceladus, serve at the primary source for heavy particles in Saturn's magnetosphere. The resulting co-orbiting neutral particles interact with ions, electrons, photons and other neutral particles to generate separate H2O, OH and O tori. Characterization of these toroidal distributions is essential for understanding Saturn magnetospheric sources, composition and dynamics. Unfortunately, limited direct observations of these features are available so modeling is required. A significant modeling challenge involves ensuring that either the plasma and neutral particle populations are not simply input conditions but can provide feedback to each population (i.e. are self-consistent). Jurac and Richardson (2005) executed such a self-consistent model however this research was performed prior to the return of Cassini data. In a similar fashion, we have coupled a 3-D neutral particle model (Smith et al. 2004, 2005, 2006, 2007, 2009, 2010) with a plasma transport model (Richardson 1998; Richardson & Jurac 2004) to develop a self-consistent model which is constrained by all available Cassini observations and current findings on Saturn's magnetosphere and the Enceladus plume source resulting in much more accurate neutral particle distributions. We present a new self-consistent model of the distribution of the Enceladus-generated neutral tori that is validated by all available observations. We also discuss the implications for source rate and variability.

A technician works on the Mars Climate Orbiter in SAEF-2

Science.gov (United States)

1998-01-01

In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), a technician works on the Mars Climate Orbiter which is scheduled to launch on Dec. 10, 1998, aboard a Boeing Delta II rocket. The Mars Climate Orbiter is heading for Mars where it will primarily support its companion Mars Polar Lander spacecraft, planned for launch on Jan. 3, 1999. After that, the Mars Climate Orbiter's instruments will monitor the Martian atmosphere and image the planet's surface on a daily basis for one Martian year (two Earth years). It will observe the appearance and movement of atmospheric dust and water vapor, as well as characterize seasonal changes on the surface. The detailed images of the surface features will provide important clues to the planet's early climate history and give scientists more information about possible liquid water reserves beneath the surface.

Stable low-altitude orbits around Ganymede considering a disturbing body in a circular orbit

Science.gov (United States)

Cardoso dos Santos, J.; Carvalho, J. P. S.; Vilhena de Moraes, R.

2014-10-01

Some missions are being planned to visit Ganymede like the Europa Jupiter System Mission that is a cooperation between NASA and ESA to insert the spacecraft JGO (Jupiter Ganymede Orbiter) into Ganymedes orbit. This comprehension of the dynamics of these orbits around this planetary satellite is essential for the success of this type of mission. Thus, this work aims to perform a search for low-altitude orbits around Ganymede. An emphasis is given in polar orbits and it can be useful in the planning of space missions to be conducted around, with respect to the stability of orbits of artificial satellites. The study considers orbits of artificial satellites around Ganymede under the influence of the third-body (Jupiter's gravitational attraction) and the polygenic perturbations like those due to non-uniform distribution of mass (J_2 and J_3) of the main body. A simplified dynamic model for these perturbations is used. The Lagrange planetary equations are used to describe the orbital motion of the artificial satellite. The equations of motion are developed in closed form to avoid expansions in eccentricity and inclination. The results show the argument of pericenter circulating. However, low-altitude (100 and 150 km) polar orbits are stable. Another orbital elements behaved variating with small amplitudes. Thus, such orbits are convenient to be applied to future space missions to Ganymede. Acknowledgments: FAPESP (processes n° 2011/05671-5, 2012/12539-9 and 2012/21023-6).

Large-Scale Spacecraft Fire Safety Tests

Science.gov (United States)

Urban, David; Ruff, Gary A.; Ferkul, Paul V.; Olson, Sandra; Fernandez-Pello, A. Carlos; T'ien, James S.; Torero, Jose L.; Cowlard, Adam J.; Rouvreau, Sebastien; Minster, Olivier;

Optimal Weighting of Multi-Spacecraft Data to Estimate Gradients of Physical Fields

Science.gov (United States)

Chanteur, G. M.; Le Contel, O.; Sahraoui, F.; Retino, A.; Mirioni, L.

2016-12-01

Multi-spacecraft missions like the ESA mission CLUSTER and the NASA mission MMS are essential to improve our understanding of physical processes in space plasmas. Several methods were designed in the 90's during the preparation phase of the CLUSTER mission to estimate gradients of physical fields from simultaneous multi-points measurements [1, 2]. Both CLUSTER and MMS involve four spacecraft with identical full scientific payloads including various sensors of electromagnetic fields and different type of particle detectors. In the standard methods described in [1, 2], which are presently in use, data from the four spacecraft have identical weights and the estimated gradients are most reliable when the tetrahedron formed by the four spacecraft is regular. There are three types of errors affecting the estimated gradients (see chapter 14 in [1]) : i) truncature errors are due to local non-linearity of spatial variations, ii) physical errors are due to instruments, and iii) geometrical errors are due to uncertainties on the positions of the spacecraft. An assessment of truncature errors for a given observation requires a theoretical model of the measured field. Instrumental errors can easily be taken into account for a given geometry of the cluster but are usually less than the geometrical errors which diverge quite fast when the tetrahedron flattens, a circumstance occurring twice per orbit of the cluster. Hence reliable gradients can be estimated only on part of the orbit. Reciprocal vectors of the tetrahedron were presented in chapter 4 of [1], they have the advantage over other methods to treat the four spacecraft symmetrically and to allow a theoretical analysis of the errors (see chapters 4 of [1] and 4 of [2]). We will present Generalized Reciprocal Vectors for weighted data and an optimization procedure to improve the reliability of the estimated gradients when the tetrahedron is not regular. A brief example using CLUSTER or MMS data will be given. This approach

Orbital motions as gradiometers for post-Newtonian tidal effects

Energy Technology Data Exchange (ETDEWEB)

Iorio, Lorenzo, E-mail: lorenzo.iorio@libero.it [Ministero dell' Istruzione, dell' Università e della Ricerca, Istruzione, Bari (Italy)

2014-08-14

The direct long-term changes occurring in the orbital dynamics of a local gravitationally bound binary system S due to the post-Newtonian tidal acceleration caused by an external massive source are investigated. A class of systems made of a test particle m rapidly orbiting with orbital frequency n{sub b} an astronomical body of mass M which, in turn, slowly revolves around a distant object of mass M′ with orbital frequency n{sub b}′ « n{sub b} is considered. The characteristic frequencies of the non-Keplerian orbital variations of m and of M itself are assumed to be negligible with respect to both n{sub b} and n{sub b}′. General expressions for the resulting Newtonian and post-Newtonian tidal orbital shifts of m are obtained. The future missions BepiColombo and JUICE to Mercury and Ganymede, respectively, are considered in view of a possible detection. The largest effects, of the order of ≈ 0.1-0.5 milliarcseconds per year (mas yr{sup −1}), occur for the Ganymede orbiter of the JUICE mission. Although future improvements in spacecraft tracking and orbit determination might, perhaps, reach the required sensitivity, the systematic bias represented by the other known orbital perturbations of both Newtonian and post-Newtonian origin would be overwhelming. The realization of a dedicated artificial mini-planetary system to be carried onboard and Earth-orbiting spacecraft is considered as well. Post-Newtonian tidal precessions as large as ≈ 1−10{sup 2} mas yr{sup −1} could be obtained, but the quite larger Newtonian tidal effects would be a major source of systematic bias because of the present-day percent uncertainty in the product of the Earth's mass times the Newtonian gravitational parameter.

Trajectories for spacecraft encounters with Comet Honda-Mrkos-Pajdusakova in 1996

Science.gov (United States)

Dunham, David W.; Jen, Shao-Chiang; Farquhar, Robert W.

1989-01-01

Early in 1996, the relatively bright short-period Comet Honda-Mrkos-Pajdusakova (HMP) will pass only 0.17 astronomical unit from the earth, providing both an unusually favorable apparition for ground-based observers and an opportunity for a spacecraft to reach Comet HMP on relatively low-energy trajectories. The Japanense Institute of Space and Astronautical Sciences Sakigake spacecraft is expected to fly by Comet HMP on February 3, 1996, after utilizing four earth swingbys to modify its orbit. If the camera on the ESA Giotto spacecraft is inoperable, Giotto may also be sent to Comet HMP. In addition, 1-year earth-return trajectories to Comet HMP are described, along with some that can be extended to encounter Comet Giacobini-Zinner in 1998.

The Viking Orbiter and its Mariner inheritance

Science.gov (United States)

Wolfe, A. E.; Norris, H. W.

1975-01-01

The orbiter system of the Viking spacecraft performs the functions of transporting the lander into orbit around Mars, surveying the proposed landing sites, relaying lander data to earth, and conducting independent scientific observations of Mars. The orbiter system is a semiautomatic, solar-powered, triaxially stabilized platform capable of making trajectory corrections and communicating with earth on S-band. Its instruments for visual imaging, detecting water vapor, and thermal mapping are mounted on a separate two-degree-of-freedom scan platform. Radio science is conducted at three frequencies, using the main S-band system, a separate X-band derived from the S-band, and the UHF one-way link with the lander.

Spacecraft Trajectory Estimation Using a Sampled-Data Extended Kalman Filter with Range-Only Measurements

National Research Council Canada - National Science Library

Erwin, R. S; Bernstein, Dennis S

2005-01-01

.... In this paper we use a sampled-data extended Kalman Filter to estimate the trajectory or a target satellite when only range measurements are available from a constellation or orbiting spacecraft...

Uncertainty Requirement Analysis for the Orbit, Attitude, and Burn Performance of the 1st Lunar Orbit Insertion Maneuver

Directory of Open Access Journals (Sweden)

Young-Joo Song

2016-12-01

Full Text Available In this study, the uncertainty requirements for orbit, attitude, and burn performance were estimated and analyzed for the execution of the 1st lunar orbit insertion (LOI maneuver of the Korea Pathfinder Lunar Orbiter (KPLO mission. During the early design phase of the system, associate analysis is an essential design factor as the 1st LOI maneuver is the largest burn that utilizes the onboard propulsion system; the success of the lunar capture is directly affected by the performance achieved. For the analysis, the spacecraft is assumed to have already approached the periselene with a hyperbolic arrival trajectory around the moon. In addition, diverse arrival conditions and mission constraints were considered, such as varying periselene approach velocity, altitude, and orbital period of the capture orbit after execution of the 1st LOI maneuver. The current analysis assumed an impulsive LOI maneuver, and two-body equations of motion were adapted to simplify the problem for a preliminary analysis. Monte Carlo simulations were performed for the statistical analysis to analyze diverse uncertainties that might arise at the moment when the maneuver is executed. As a result, three major requirements were analyzed and estimated for the early design phase. First, the minimum requirements were estimated for the burn performance to be captured around the moon. Second, the requirements for orbit, attitude, and maneuver burn performances were simultaneously estimated and analyzed to maintain the 1st elliptical orbit achieved around the moon within the specified orbital period. Finally, the dispersion requirements on the B-plane aiming at target points to meet the target insertion goal were analyzed and can be utilized as reference target guidelines for a mid-course correction (MCC maneuver during the transfer. More detailed system requirements for the KPLO mission, particularly for the spacecraft bus itself and for the flight dynamics subsystem at the ground

FORTE spacecraft vibration mitigation. Final report

International Nuclear Information System (INIS)

Maly, J.R.

1996-02-01

This report documents work that was performed by CSA Engineering, Inc., for Los Alamos National Laboratory (LANL), to reduce vibrations of the FORTE spacecraft by retrofitting damped structural components into the spacecraft structure. The technical objective of the work was reduction of response at the location of payload components when the structure is subjected to the dynamic loading associated with launch and proto-qualification testing. FORTE is a small satellite that will be placed in orbit in 1996. The structure weighs approximately 425 lb, and is roughly 80 inches high and 40 inches in diameter. It was developed and built by LANL in conjunction with Sandia National Laboratories Albuquerque for the United States Department of Energy. The FORTE primary structure was fabricated primarily with graphite epoxy, using aluminum honeycomb core material for equipment decks and solar panel substrates. Equipment decks were bonded and bolted through aluminum mounting blocks to adjoining structure

Spacecraft Conceptual Design for the 8-Meter Advanced Technology Large Aperture Space Telescope (ATLAST)

Science.gov (United States)

Hopkins, Randall C.; Capizzo, Peter; Fincher, Sharon; Hornsby, Linda S.; Jones, David

2010-01-01

The Advanced Concepts Office at Marshall Space Flight Center completed a brief spacecraft design study for the 8-meter monolithic Advanced Technology Large Aperture Space Telescope (ATLAST-8m). This spacecraft concept provides all power, communication, telemetry, avionics, guidance and control, and thermal control for the observatory, and inserts the observatory into a halo orbit about the second Sun-Earth Lagrange point. The multidisciplinary design team created a simple spacecraft design that enables component and science instrument servicing, employs articulating solar panels for help with momentum management, and provides precise pointing control while at the same time fast slewing for the observatory.

CASSINI RSS RAW DATA SET - SROC20 V1.0

Data.gov (United States)

National Aeronautics and Space Administration — The Cassini Radio Science Saturn Ring and Atmospheric Occultation experiments (SROC20) Raw Data Archive is a time-ordered collection of radio science raw data...

Autonomous Spacecraft Communication Interface for Load Planning

Science.gov (United States)

Dever, Timothy P.; May, Ryan D.; Morris, Paul H.

2014-01-01

Ground-based controllers can remain in continuous communication with spacecraft in low Earth orbit (LEO) with near-instantaneous communication speeds. This permits near real-time control of all of the core spacecraft systems by ground personnel. However, as NASA missions move beyond LEO, light-time communication delay issues, such as time lag and low bandwidth, will prohibit this type of operation. As missions become more distant, autonomous control of manned spacecraft will be required. The focus of this paper is the power subsystem. For present missions, controllers on the ground develop a complete schedule of power usage for all spacecraft components. This paper presents work currently underway at NASA to develop an architecture for an autonomous spacecraft, and focuses on the development of communication between the Mission Manager and the Autonomous Power Controller. These two systems must work together in order to plan future load use and respond to unanticipated plan deviations. Using a nominal spacecraft architecture and prototype versions of these two key components, a number of simulations are run under a variety of operational conditions, enabling development of content and format of the messages necessary to achieve the desired goals. The goals include negotiation of a load schedule that meets the global requirements (contained in the Mission Manager) and local power system requirements (contained in the Autonomous Power Controller), and communication of off-plan disturbances that arise while executing a negotiated plan. The message content is developed in two steps: first, a set of rapid-prototyping "paper" simulations are preformed; then the resultant optimized messages are codified for computer communication for use in automated testing.

Advanced Exoplanet Star Tracker for Orbit Self Determination, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — This proposal puts forth an innovative star tracker hardware sensor that allows for autonomous calculation of a spacecraft's orbit by employing Doppler Spectroscopy...

CASSINI S MIMI CHEMS SENSOR CALIBRATED DATA V1.0

Data.gov (United States)

National Aeronautics and Space Administration — The Cassini Magnetospheric Imaging Instrument(MIMI) Charge Energy Mass Spectrometer (CHEMS) contains a deflection system and an overall field of view of 159 x 4 deg....

Optimal Rendezvous and Docking Simulator for Elliptical Orbits, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — It is proposed to develop and implement a simulation of spacecraft rendezvous and docking guidance, navigation, and control in elliptical orbit. The foundation of...

Maintaining Aura's Orbit Requirements While Performing Orbit Maintenance Maneuvers Containing an Orbit Normal Delta-V Component

Science.gov (United States)

Johnson, Megan R.; Petersen, Jeremy D.

2014-01-01

The Earth Observing System (EOS) Afternoon Constellation consists of five member missions (GCOM-W1, Aqua, CALIPSO, CloudSat, and Aura), each of which maintain a frozen, sun-synchronous orbit with a 16-day repeating ground track that follows the Worldwide Reference System-2 (WRS-2). Under nominal science operations for Aura, the propulsion system is oriented such that the resultant thrust vector is aligned 13.493 degrees away from the velocity vector along the yaw axis. When performing orbit maintenance maneuvers, the spacecraft performs a yaw slew to align the thrust vector in the appropriate direction. A new Drag Make Up (DMU) maneuver operations scheme has been implemented for Aura alleviating the need for the 13.493 degree yaw slew. The focus of this investigation is to assess the impact that no-slew DMU maneuver operations will have on Aura's Mean Local Time (MLT) which drives the required along track separation between Aura and the constellation members, as well as Aura's frozen orbit properties, eccentricity and argument of perigee. Seven maneuver strategies were analyzed to determine the best operational approach. A mirror pole strategy, with maneuvers alternating at the North and South poles, was implemented operationally to minimize impact to the MLT. Additional analysis determined that the mirror pole strategy could be further modified to include frozen orbit maneuvers and thus maintain both MLT and the frozen orbit properties under noslew operations.

On-orbit evaluation of the control system/structural mode interactions on OSO-8

Science.gov (United States)

Slafer, L. I.

1980-01-01

The Orbiting Solar Observatory-8 experienced severe structural mode/control loop interaction problems during the spacecraft development. Extensive analytical studies, using the hybrid coordinate modeling approach, and comprehensive ground testing were carried out in order to achieve the system's precision pointing performance requirements. A recent series of flight tests were conducted with the spacecraft in which a wide bandwidth, high resolution telemetry system was utilized to evaluate the on-orbit flexible dynamics characteristics of the vehicle along with the control system performance. This paper describes the results of these tests, reviewing the basic design problem, analytical approach taken, ground test philosophy, and on-orbit testing. Data from the tests was used to determine the primary mode frequency, damping, and servo coupling dynamics for the on-orbit condition. Additionally, the test results have verified analytically predicted differences between the on-orbit and ground test environments. The test results have led to a validation of both the analytical modeling and servo design techniques used during the development of the control system, and also verified the approach taken to vehicle and servo ground testing.

Potentiality of an orbiting interferometer for space-time experiments

International Nuclear Information System (INIS)

Grassi Strini, A.M.; Strini, G.; Tagliaferri, G.

1979-01-01

It is suggested that by putting a Michelson interferometer aboard a spacecraft orbiting around the earth, very substantial progress could be made in space-time experiments. It is estimated that in measurements of e.g. some anisotropy of the light velocity, a spacecraft-borne interferometer of quite small size (0.1 m arm-length) would reach a sensitivity greater by a factor of approximately 10 8 than the best achievements to date of ground-based devices. (author)

Accelerated simulation of near-Earth-orbit polymer degradation

Science.gov (United States)

Laue, Eric

1992-01-01

There is a need to simulate the near-Earth-orbit environmental conditions, and it is useful to be able to monitor the changes in physical properties of spacecraft materials. Two different methods for simulating the vacuum-ultraviolet (VUV) and soft X-ray near-Earth-orbit flux are presented. Also, methods for monitoring the changes in optical ultraviolet transmission and mass loss are presented. The results of exposures to VUV photons and charged particles on these materials are discussed.

Cassini’s Discoveries at Saturn and the Proposed Cassini Solstice Mission

Science.gov (United States)

Pappalardo, R. T.; Spilker, L. J.; Mitchell, R. T.; Cuzzi, J.; Gombosi, T. I.; Ingersoll, A. P.; Lunine, J. I.

2009-12-01

Understanding of the Saturn system has been greatly enhanced by the Cassini-Huygens mission. Fundamental discoveries have altered our views of Saturn, Titan and the other icy satellites, the rings, and magnetosphere of the system. Key discoveries include: water-rich plumes emanating from the south pole of Enceladus; hints of possible activity on Dione and of rings around Rhea; a methane hydrological cycle on Titan complete with fluvial erosion, lakes, and seas of liquid methane and ethane; non-axisymmetric ring microstructure in all moderate optical depth rings; south polar vortices on Saturn; and a unique magnetosphere that shares characteristics with both Earth’s and Jupiter’s magnetospheres. These new discoveries are directly relevant to current Solar System science goals including: planet and satellite formation processes, formation of gas giants, the nature of organic material, the history of volatiles, habitable zones and processes for life, processes that shape planetary bodies, and evolution of exoplanets. The proposed 7-year Cassini Solstice Mission would address new questions that have arisen during the Cassini Prime and Equinox Missions, and would observe seasonal and temporal change in the Saturn system to prepare for future missions to Saturn, Titan, and Enceladus. The proposed Cassini Solstice Mission would provide new science in three ways. First, it would observe seasonally and temporally dependent processes on Saturn, Titan and other icy satellites, and within the rings and magnetosphere, in a hitherto unobserved seasonal phase from equinox to solstice. Second, it would address new questions that have arisen during the mission thus far, providing qualitatively new measurements (e.g. of Enceladus and Titan) which could not be accommodated in the earlier mission phases. Tthird, it would conduct a close-in mission phase at Saturn that would provide unique science including comparison to the Juno observations at Jupiter.

Results of Large-Scale Spacecraft Flammability Tests

Science.gov (United States)

Ferkul, Paul; Olson, Sandra; Urban, David L.; Ruff, Gary A.; Easton, John; T'ien, James S.; Liao, Ta-Ting T.; Fernandez-Pello, A. Carlos; Torero, Jose L.; Eigenbrand, Christian;

Global-scale Observations of the Limb and Disk (GOLD) Mission: Science from Geostationary Orbit on-board a Commercial Communications Satellite

Science.gov (United States)

Eastes, R.; Deaver, T.; Krywonos, A.; Lankton, M. R.; McClintock, W. E.; Pang, R.

2011-12-01

Geostationary orbits are ideal for many science investigations of the Earth system on global scales. These orbits allow continuous observations of the same geographic region, enabling spatial and temporal changes to be distinguished and eliminating the ambiguity inherent to observations from low Earth orbit (LEO). Just as observations from geostationary orbit have revolutionized our understanding of changes in the troposphere, they will dramatically improve our understanding of the space environment at higher altitudes. However, geostationary orbits are infrequently used for science missions because of high costs. Geostationary satellites are large, typically weighing tons. Consequently, devoting an entire satellite to a science mission requires a large financial commitment, both for the spacecraft itself and for sufficient science instrumentation to justify a dedicated spacecraft. Furthermore, the small number of geostationary satellites produced for scientific missions increases the costs of each satellite. For these reasons, it is attractive to consider flying scientific instruments on satellites operated by commercial companies, some of whom have fleets of ~40 satellites. However, scientists' lack of understanding of the capabilities of commercial spacecraft as well as commercial companies' concerns about risks to their primary mission have impeded the cooperation necessary for the shared use of a spacecraft. Working with a commercial partner, the GOLD mission has successfully overcome these issues. Our experience indicates that there are numerous benefits to flying on commercial communications satellites (e.g., it is possible to downlink large amounts of data) and the costs are low if the experimental requirements adequately match the capabilities and available resources of the host spacecraft. Consequently, affordable access to geostationary orbit aboard a communications satellite now appears possible for science payloads.

Comparison of ERBS orbit determination accuracy using batch least-squares and sequential methods

Science.gov (United States)

Oza, D. H.; Jones, T. L.; Fabien, S. M.; Mistretta, G. D.; Hart, R. C.; Doll, C. E.

1991-10-01

The Flight Dynamics Div. (FDD) at NASA-Goddard commissioned a study to develop the Real Time Orbit Determination/Enhanced (RTOD/E) system as a prototype system for sequential orbit determination of spacecraft on a DOS based personal computer (PC). An overview is presented of RTOD/E capabilities and the results are presented of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft obtained using RTOS/E on a PC with the accuracy of an established batch least squares system, the Goddard Trajectory Determination System (GTDS), operating on a mainframe computer. RTOD/E was used to perform sequential orbit determination for the Earth Radiation Budget Satellite (ERBS), and the Goddard Trajectory Determination System (GTDS) was used to perform the batch least squares orbit determination. The estimated ERBS ephemerides were obtained for the Aug. 16 to 22, 1989, timeframe, during which intensive TDRSS tracking data for ERBS were available. Independent assessments were made to examine the consistencies of results obtained by the batch and sequential methods. Comparisons were made between the forward filtered RTOD/E orbit solutions and definitive GTDS orbit solutions for ERBS; the solution differences were less than 40 meters after the filter had reached steady state.

Comparison of ERBS orbit determination accuracy using batch least-squares and sequential methods

Science.gov (United States)

Oza, D. H.; Jones, T. L.; Fabien, S. M.; Mistretta, G. D.; Hart, R. C.; Doll, C. E.

1991-01-01

The Flight Dynamics Div. (FDD) at NASA-Goddard commissioned a study to develop the Real Time Orbit Determination/Enhanced (RTOD/E) system as a prototype system for sequential orbit determination of spacecraft on a DOS based personal computer (PC). An overview is presented of RTOD/E capabilities and the results are presented of a study to compare the orbit determination accuracy for a Tracking and Data Relay Satellite System (TDRSS) user spacecraft obtained using RTOS/E on a PC with the accuracy of an established batch least squares system, the Goddard Trajectory Determination System (GTDS), operating on a mainframe computer. RTOD/E was used to perform sequential orbit determination for the Earth Radiation Budget Satellite (ERBS), and the Goddard Trajectory Determination System (GTDS) was used to perform the batch least squares orbit determination. The estimated ERBS ephemerides were obtained for the Aug. 16 to 22, 1989, timeframe, during which intensive TDRSS tracking data for ERBS were available. Independent assessments were made to examine the consistencies of results obtained by the batch and sequential methods. Comparisons were made between the forward filtered RTOD/E orbit solutions and definitive GTDS orbit solutions for ERBS; the solution differences were less than 40 meters after the filter had reached steady state.

Increased electric sail thrust through removal of trapped shielding electrons by orbit chaotisation due to spacecraft body

Directory of Open Access Journals (Sweden)

P. Janhunen

2009-08-01

Full Text Available An electric solar wind sail is a recently introduced propellantless space propulsion method whose technical development has also started. The electric sail consists of a set of long, thin, centrifugally stretched and conducting tethers which are charged positively and kept in a high positive potential of order 20 kV by an onboard electron gun. The positively charged tethers deflect solar wind protons, thus tapping momentum from the solar wind stream and producing thrust. The amount of obtained propulsive thrust depends on how many electrons are trapped by the potential structures of the tethers, because the trapped electrons tend to shield the charged tether and reduce its effect on the solar wind. Here we present physical arguments and test particle calculations indicating that in a realistic three-dimensional electric sail spacecraft there exist a natural mechanism which tends to remove the trapped electrons by chaotising their orbits and causing them to eventually collide with the conducting tethers. We present calculations which indicate that if these mechanisms were able to remove trapped electrons nearly completely, the electric sail performance could be about five times higher than previously estimated, about 500 nN/m, corresponding to 1 N thrust for a baseline construction with 2000 km total tether length.

Video-Game-Like Engine for Depicting Spacecraft Trajectories

Science.gov (United States)

Upchurch, Paul R.

2009-01-01

GoView is a video-game-like software engine, written in the C and C++ computing languages, that enables real-time, three-dimensional (3D)-appearing visual representation of spacecraft and trajectories (1) from any perspective; (2) at any spatial scale from spacecraft to Solar-system dimensions; (3) in user-selectable time scales; (4) in the past, present, and/or future; (5) with varying speeds; and (6) forward or backward in time. GoView constructs an interactive 3D world by use of spacecraft-mission data from pre-existing engineering software tools. GoView can also be used to produce distributable application programs for depicting NASA orbital missions on personal computers running the Windows XP, Mac OsX, and Linux operating systems. GoView enables seamless rendering of Cartesian coordinate spaces with programmable graphics hardware, whereas prior programs for depicting spacecraft trajectories variously require non-Cartesian coordinates and/or are not compatible with programmable hardware. GoView incorporates an algorithm for nonlinear interpolation between arbitrary reference frames, whereas the prior programs are restricted to special classes of inertial and non-inertial reference frames. Finally, whereas the prior programs present complex user interfaces requiring hours of training, the GoView interface provides guidance, enabling use without any training.

Towards the integration of orbital space use in Life Cycle Impact Assessment.

Science.gov (United States)

Maury, Thibaut; Loubet, Philippe; Ouziel, Jonathan; Saint-Amand, Maud; Dariol, Ludovic; Sonnemann, Guido

2017-10-01

A rising sustainability concern is occurring in the space sector: 29,000 human-made objects, larger than 10cm are orbiting the Earth but only 6% are operational spacecrafts. Today, space debris is today a significant and constant danger to all space missions. Consequently, it becomes compelled to design new space missions considering End-of-Life requirements in order to ensure the sustainable use of space orbits. Furthermore, Life Cycle Assessment (LCA) has been identified by the European Space Agency as an adequate tool to measure the environmental impact of spacecraft missions. Hence, our challenge is to integrate orbital space use into Life Cycle Impact Assessment (LCIA) to broaden the scope of LCA for space systems. The generation of debris in the near-Earth's orbital regions leads to a decrease in volume availability. The Area-of-Protection (AoP) 'resources' seems to be the most relevant reflection of this depletion. To address orbital space use in a comprehensive way, we propose a first attempt at establishing an impact pathway linking outer space use to resources. This framework will be the basis for defining new indicator(s) related to orbital space use. Copyright © 2017 Elsevier B.V. All rights reserved.

Detection of the lunar body tide by the Lunar Orbiter Laser Altimeter.

Science.gov (United States)

Mazarico, Erwan; Barker, Michael K; Neumann, Gregory A; Zuber, Maria T; Smith, David E

2014-04-16

The Lunar Orbiter Laser Altimeter instrument onboard the Lunar Reconnaissance Orbiter spacecraft collected more than 5 billion measurements in the nominal 50 km orbit over ∼10,000 orbits. The data precision, geodetic accuracy, and spatial distribution enable two-dimensional crossovers to be used to infer relative radial position corrections between tracks to better than ∼1 m. We use nearly 500,000 altimetric crossovers to separate remaining high-frequency spacecraft trajectory errors from the periodic radial surface tidal deformation. The unusual sampling of the lunar body tide from polar lunar orbit limits the size of the typical differential signal expected at ground track intersections to ∼10 cm. Nevertheless, we reliably detect the topographic tidal signal and estimate the associated Love number h 2 to be 0.0371 ± 0.0033, which is consistent with but lower than recent results from lunar laser ranging. Altimetric data are used to create radial constraints on the tidal deformationThe body tide amplitude is estimated from the crossover dataThe estimated Love number is consistent with previous estimates but more precise.

MIDN: A spacecraft Micro-dosimeter mission

International Nuclear Information System (INIS)

Pisacane, V. L.; Ziegler, J. F.; Nelson, M. E.; Caylor, M.; Flake, D.; Heyen, L.; Youngborg, E.; Rosenfeld, A. B.; Cucinotta, F.; Zaider, M.; Dicello, J. F.

2006-01-01

MIDN (Micro-dosimetry instrument) is a payload on the MidSTAR-I spacecraft (Midshipman Space Technology Applications Research) under development at the United States Naval Academy. MIDN is a solid-state system being designed and constructed to measure Micro-dosimetric spectra to determine radiation quality factors for space environments. Radiation is a critical threat to the health of astronauts and to the success of missions in low-Earth orbit and space exploration. The system will consist of three separate sensors, one external to the spacecraft, one internal and one embedded in polyethylene. Design goals are mass <3 kg and power <2 W. The MidSTAR-I mission in 2006 will provide an opportunity to evaluate a preliminary version of this system. Its low power and mass makes it useful for the International Space Station and manned and unmanned interplanetary missions as a real-time system to assess and alert astronauts to enhanced radiation environments. (authors)

Initial On-Orbit Engineering Results from the O/OREOS Nanosatellite

OpenAIRE

Kitts, Christopher; Rasay, Mike; Bica, Laura; Mas, Ignacio; Neumann, Michael; Young, Anthony; Minelli, Giovanni; Ricco, Antonio; Stackpole, Eric; Agasid, Elwood; Beasley, Christopher; Friedericks, Charlie; Squires, David; Ehrenfreund, Pascale; Nicholson, Wayne

2011-01-01

The Organism/Organics Exposure to Orbital Stresses (O/OREOS) nanosatellite mission successfully launched on November 19, 2010 from Kodiak, AK aboard a Minotaur IV launch vehicle. The principal goals for this 5.5 kg spacecraft include conducting astrobiologically-relevant experiments in two separate payloads within the 3U cubesat form factor and demonstrating in-situ measurement technology in a small satellite. Developed by the Small Spacecraft Payloads and Technology Team at NASA Ames Researc...

The Gas-Surface Interaction of a Human-Occupied Spacecraft with a Near-Earth Object

Science.gov (United States)

Farrell, W. M.; Hurley, D. M.; Poston, M. J.; Zimmerman, M. I.; Orlando, T. M.; Hibbitts, C. A.; Killen, R. M.

2016-01-01

NASA's asteroid redirect mission (ARM) will feature an encounter of the human-occupied Orion spacecraft with a portion of a near- Earth asteroid (NEA) previously placed in orbit about the Moon by a capture spacecraft. Applying a shuttle analog, we suggest that the Orion spacecraft should have a dominant local water exosphere, and that molecules from this exosphere can adsorb onto the NEA. The amount of adsorbed water is a function of the defect content of the NEA surface, with retention of shuttle-like water levels on the asteroid at 10(exp 15) H2O's/m2 for space weathered regolith at T approximately 300 K.

Steady state obliquity of a rigid body in the spin-orbit resonant problem: application to Mercury

Science.gov (United States)

Lhotka, Christoph

2017-12-01

We investigate the stable Cassini state 1 in the p : q spin-orbit resonant problem. Our study includes the effect of the gravitational potential up to degree and order 4 and p : q spin-orbit resonances with p,q≤ 8 and p≥ q. We derive new formulae that link the gravitational field coefficients with its secular orbital elements and its rotational parameters. The formulae can be used to predict the orientation of the spin axis and necessary angular momentum at exact resonance. We also develop a simple pendulum model to approximate the dynamics close to resonance and make use of it to predict the libration periods and widths of the oscillatory regime of motions in phase space. Our analytical results are based on averaging theory that we also confirm by means of numerical simulations of the exact dynamical equations. Our results are applied to a possible rotational history of Mercury.

Trajectories of inner and outer heliospheric spacecraft: Predicted through 1999

Science.gov (United States)

Parthasarathy, R.; King, Joseph H.

1991-01-01

Information is presented in tabular and graphical form on the trajectories of the international fleet of spacecraft that will be probing the far reaches of the heliosphere during the 1990s. In particular, the following spacecraft are addressed: Pioneer 10 and 11, Pioneer Venus Orbiter (PVO), Voyager 1 and 2, Galileo, Ulysses, Suisei, Sakigake, Giotto, International Cometary Explorer (ICE), and Interplanetary Monitoring Platform 8 (IMP 8). Yearly resolution listing of position information in inertial space are given for Pioneer and Voyager spacecraft from the times of their launches in the 1970s. One series of plots shows the radial distances, latitudes, and longitudes of the Pioneers and Voyagers. The solar ecliptic inertial coordinate system is used. In this system, the Z axis is normal to the ecliptic plane and the X axis is towards the first point of Aries (from Sun to Earth on the vernal equinox).

Omni-directional Particle Detector (ODPD) on Tiangong-2 Spacecraft

Science.gov (United States)

Guohong, S.; Zhang, S.; Yang, X.; Wang, C.

2017-12-01

Tiangong-2 spacecraft is the second space laboratory independently developed by china after Tiangong-1, which was launched on 15 September 2016. It is also the first real space laboratory in china, which will be used to further validate the space rendezvous and docking technology and to carry out a series of space tests. The spacecraft's orbit is 350km height and 42° inclination. The omni-directional particle detector (ODPD) on Tiangong-2 spacecraft is a new instrument developed by China. Its goal is the anisotropy and energy spectra of space particles on manned space flight orbit. The ODPD measures the energy spectra and pitch angle distributions of high energy electrons and protons. It consists of one electron spectrum telescope, one proton spectrum telescope and sixteen directional flux telescopes. The ODPD is designed to measure the protons spectrum from 2.5MeV to 150MeV, electrons spectrum from 0.2MeV to 1.5MeV, the flux of electrons energy >200keV and protons energy>1.5MeV on 2∏ space, also the ODPD has a small sensor to measure the LET spectrum from 1-100MeV/cm2sr. The primary advantage can give the particle's pitch angle distributions at any time because of the sixteen flux telescopes arrange form 0 to 180 degree. This is the first paper dealing with ODPD data, so we firstly spend some time describing the instrument, its theory of operation and its calibration. Then we give the preliminary detecting results.

A Three-Dimensional View of Titan's Surface Features from Cassini RADAR Stereogrammetry

Science.gov (United States)

Kirk, R. L.; Howington-Kraus, E.; Redding, B. L.; Becker, T. L.; Lee, E. M.; Stiles, B. W.; Hensley, S.; Hayes, A.; Lopes, R. M.; Lorenz, R. D.; Mitchell, K. L.; Radebaugh, J.; Paganelli, F.; Soderblom, L. A.; Stofan, E. R.; Wood, C. A.; Wall, S. D.; Cassini RADAR Team

2008-12-01

As of the end of its four-year Prime Mission, Cassini has obtained 300-1500 m resolution synthetic aperture radar images of the surface of Titan during 19 flybys. The elongated image swaths overlap extensively, and ~2% of the surface has now been imaged two or more times. The majority of image pairs have different viewing directions, and thus contain stereo parallax that encodes information about Titan's surface relief over distances of ~1 km and greater. As we have previously reported, the first step toward extracting quantitative topographic information was the development of rigorous "sensor models" that allowed the stereo systems previously used at the USGS and JPL to map Venus with Magellan images to be used for Titan mapping. The second major step toward extensive topomapping of Titan has been the reprocessing of the RADAR images based on an improved model of the satellite's rotation. Whereas the original images (except for a few pairs obtained at similar orbital phase, some of which we have mapped previously) were offset by as much as 30 km, the new versions align much better. The remaining misalignments, typically carbono)logic" cycle of precipitation, evaporation, and surface and subsurface fluid flow?

METER-SIZED MOONLET POPULATION IN SATURN'S C RING AND CASSINI DIVISION

International Nuclear Information System (INIS)

Baillié, Kévin; Colwell, Joshua E.; Esposito, Larry W.; Lewis, Mark C.

2013-01-01

Stellar occultations observed by the Cassini Ultraviolet Imaging Spectrograph reveal the presence of transparent holes a few meters to a few tens of meters in radial extent in otherwise optically thick regions of the C ring and the Cassini Division. We attribute the holes to gravitational disturbances generated by a population of ∼10 m boulders in the rings that is intermediate in size between the background ring particle size distribution and the previously observed ∼100 m propeller moonlets in the A ring. The size distribution of these boulders is described by a shallower power-law than the one that describes the ring particle size distribution. The number and size distribution of these boulders could be explained by limited accretion processes deep within Saturn's Roche zone.

Review of current activities to model and measure the orbital debris environment in low-earth orbit

Science.gov (United States)

Reynolds, R. C.

A very active orbital debris program is currently being pursued at the NASA/Johnson Space Center (JSC), with projects designed to better define the current environment, to project future environments, to model the processes contributing to or constraining the growth of debris in the environment, and to gather supporting data needed to improve the understanding of the orbital debris problem and the hazard it presents to spacecraft. This paper is a review of the activity being conducted at JSC, by NASA, Lockheed Engineering and Sciences Company, and other support contractors, and presents a review of current activity, results of current research, and a discussion of directions for future development.

A Study of Single- and Double-Averaged Second-Order Models to Evaluate Third-Body Perturbation Considering Elliptic Orbits for the Perturbing Body

Directory of Open Access Journals (Sweden)

R. C. Domingos

2013-01-01

Full Text Available The equations for the variations of the Keplerian elements of the orbit of a spacecraft perturbed by a third body are developed using a single average over the motion of the spacecraft, considering an elliptic orbit for the disturbing body. A comparison is made between this approach and the more used double averaged technique, as well as with the full elliptic restricted three-body problem. The disturbing function is expanded in Legendre polynomials up to the second order in both cases. The equations of motion are obtained from the planetary equations, and several numerical simulations are made to show the evolution of the orbit of the spacecraft. Some characteristics known from the circular perturbing body are studied: circular, elliptic equatorial, and frozen orbits. Different initial eccentricities for the perturbed body are considered, since the effect of this variable is one of the goals of the present study. The results show the impact of this parameter as well as the differences between both models compared to the full elliptic restricted three-body problem. Regions below, near, and above the critical angle of the third-body perturbation are considered, as well as different altitudes for the orbit of the spacecraft.

Design of relative trajectories for in orbit proximity operations

Science.gov (United States)

Opromolla, Roberto; Fasano, Giancarmine; Rufino, Giancarlo; Grassi, Michele

2018-04-01

This paper presents an innovative approach to design relative trajectories suitable for close-proximity operations in orbit, by assigning high-level constraints regarding their stability, shape and orientation. Specifically, this work is relevant to space mission scenarios, e.g. formation flying, on-orbit servicing, and active debris removal, which involve either the presence of two spacecraft carrying out coordinated maneuvers, or a servicing/recovery spacecraft (chaser) performing monitoring, rendezvous and docking with respect to another space object (target). In the above-mentioned scenarios, an important aspect is the capability of reducing collision risks and of providing robust and accurate relative navigation solutions. To this aim, the proposed approach exploits a relative motion model relevant to two-satellite formations, and developed in mean orbit parameters, which takes the perturbation effect due to secular Earth oblateness, as well as the motion of the target along a small-eccentricity orbit, into account. This model is used to design trajectories which ensure safe relative motion, to minimize collision risks and relax control requirements, providing at the same time favorable conditions, in terms of target-chaser relative observation geometry for pose determination and relative navigation with passive or active electro-optical sensors on board the chaser. Specifically, three design strategies are proposed in the context of a space target monitoring scenario, considering as design cases both operational spacecraft and debris, characterized by highly variable shape, size and absolute rotational dynamics. The effectiveness of the proposed design approach in providing favorable observation conditions for target-chaser relative pose estimation is demonstrated within a simulation environment which reproduces the designed target-chaser relative trajectory, the operation of an active LIDAR installed on board the chaser, and pose estimation algorithms.

SEL2 servicing: increased science return via on-orbit propellant replenishment

Science.gov (United States)

Reed, Benjamin B.; DeWeese, Keith; Kienlen, Michael; Aranyos, Thomas; Pellegrino, Joseph; Bacon, Charles; Qureshi, Atif

2016-07-01

Spacecraft designers are driving observatories to the distant Sun-Earth Lagrange Point 2 (SEL2) to meet ever-increasing science requirements. The mass fraction dedicated to propellant for these observatories to reach and operate at SEL2 will be allocated with the upmost care, as it comes at the expense of optics and instrument masses. As such, these observatories could benefit from on-orbit refueling, allowing greater dry-to-wet mass ratio at launch and/or longer mission life. NASA is developing technologies, capabilities and integrated mission designs for multiple servicing applications in low Earth orbit (LEO), geosynchronous Earth orbit (GEO) and cisluner locations. Restore-L, a mission officially in formulation, will launch a free-flying robotic servicer to refuel a government-owned satellite in LEO by mid 2020. This paper will detail the results of a point design mission study to extend Restore-L servicing technologies from LEO to SEL2. This SEL2 mission would launch an autonomous, robotic servicer spacecraft equipped to extend the life of two space assets through refueling. Two space platforms were chosen to 1) drive the requirements for achieving SEL2 orbit and rendezvous with a spacecraft, and 2) to drive the requirements to translate within SEL2 to conduct a follow-on servicing mission. Two fuels, xenon and hydrazine, were selected to assess a multiple delivery system. This paper will address key mission drivers, such as servicer autonomy (necessitated due to communications latency at L2). Also discussed will be the value of adding cooperative servicing elements to the client observatories to reduce mission risk.

Interactive Spacecraft Trajectory Design Strategies Featuring Poincare Map Topology

Science.gov (United States)

Schlei, Wayne R.

Space exploration efforts are shifting towards inexpensive and more agile vehicles. Versatility regarding spacecraft trajectories refers to the agility to correct deviations from an intended path or even the ability to adapt the future path to a new destination--all with limited spaceflight resources (i.e., small DeltaV budgets). Trajectory design methods for such nimble vehicles incorporate equally versatile procedures that allow for rapid and interactive decision making while attempting to reduce Delta V budgets, leading to a versatile trajectory design platform. A versatile design paradigm requires the exploitation of Poincare map topology , or the interconnected web of dynamical structures, existing within the chaotic dynamics of multi-body gravitational models to outline low-Delta V transfer options residing nearby to a current path. This investigation details an autonomous procedure to extract the periodic orbits (topology nodes) and correlated asymptotic flow structures (or the invariant manifolds representing topology links). The autonomous process summarized in this investigation (termed PMATE) overcomes discontinuities on the Poincare section that arise in the applied multi-body model (the planar circular restricted three-body problem) and detects a wide variety of novel periodic orbits. New interactive capabilities deliver a visual analytics foundation for versatile spaceflight design, especially for initial guess generation and manipulation. Such interactive strategies include the selection of states and arcs from Poincare section visualizations and the capabilities to draw and drag trajectories to remove dependency on initial state input. Furthermore, immersive selection is expanded to cull invariant manifold structures, yielding low-DeltaV or even DeltaV-free transfers between periodic orbits. The application of interactive design strategies featuring a dense extraction of Poincare map topology is demonstrated for agile spaceflight with a simple

Protection of surface assets on Mars from wind blown jettisoned spacecraft components

Science.gov (United States)

Paton, Mark

2017-07-01

Jettisoned Entry, Descent and Landing System (EDLS) hardware from landing spacecraft have been observed by orbiting spacecraft, strewn over the Martian surface. Future Mars missions that land spacecraft close to prelanded assets will have to use a landing architecture that somehow minimises the possibility of impacts from these jettisoned EDLS components. Computer modelling is used here to investigate the influence of wind speed and direction on the distribution of EDLS components on the surface. Typical wind speeds encountered in the Martian Planetary Boundary Layer (PBL) were found to be of sufficient strength to blow items having a low ballistic coefficient, i.e. Hypersonic Inflatable Aerodynamic Decelerators (HIADs) or parachutes, onto prelanded assets even when the lander itself touches down several kilometres away. Employing meteorological measurements and careful characterisation of the Martian PBL, e.g. appropriate wind speed probability density functions, may then benefit future spacecraft landings, increase safety and possibly help reduce the delta v budget for Mars landers that rely on aerodynamic decelerators.

The O/OREOS Mission - Astrobiology in Low Earth Orbit. [Astrobiology in Low Earth Orbit

Science.gov (United States)

Ehrenfreund, P.; Ricco, A. J.; Squires, D.; Kitts, C.; Agasid, E.; Bramall, N.; Bryson, K.; Chittenden, J.; Conley, C.; Cook, A.;

Space-to-Space Power Beaming Enabling High Performance Rapid Geocentric Orbit Transfer

Science.gov (United States)

Dankanich, John W.; Vassallo, Corinne; Tadge, Megan

2015-01-01

The use of electric propulsion is more prevalent than ever, with industry pursuing all electric orbit transfers. Electric propulsion provides high mass utilization through efficient propellant transfer. However, the transfer times become detrimental as the delta V transitions from near-impulsive to low-thrust. Increasing power and therefore thrust has diminishing returns as the increasing mass of the power system limits the potential acceleration of the spacecraft. By using space-to-space power beaming, the power system can be decoupled from the spacecraft and allow significantly higher spacecraft alpha (W/kg) and therefore enable significantly higher accelerations while maintaining high performance. This project assesses the efficacy of space-to-space power beaming to enable rapid orbit transfer while maintaining high mass utilization. Concept assessment requires integrated techniques for low-thrust orbit transfer steering laws, efficient large-scale rectenna systems, and satellite constellation configuration optimization. This project includes the development of an integrated tool with implementation of IPOPT, Q-Law, and power-beaming models. The results highlight the viability of the concept, limits and paths to infusion, and comparison to state-of-the-art capabilities. The results indicate the viability of power beaming for what may be the only approach for achieving the desired transit times with high specific impulse.

Satellite orbits in Levi-Civita space

Science.gov (United States)

Humi, Mayer

2018-03-01

In this paper we consider satellite orbits in central force field with quadratic drag using two formalisms. The first using polar coordinates in which the satellite angular momentum plays a dominant role. The second is in Levi-Civita coordinates in which the energy plays a central role. We then merge these two formalisms by introducing polar coordinates in Levi-Civita space and derive a new equation for satellite orbits which unifies these two paradigms. In this equation energy and angular momentum appear on equal footing and thus characterize the orbit by its two invariants. Using this formalism we show that equatorial orbits around oblate spheroids can be expressed analytically in terms of Elliptic functions. In the second part of the paper we derive in Levi-Civita coordinates a linearized equation for the relative motion of two spacecrafts whose trajectories are in the same plane. We carry out also a numerical verification of these equations.

Gravity and Nonconservative Force Model Tuning for the GEOSAT Follow-On Spacecraft

Science.gov (United States)

Lemoine, Frank G.; Zelensky, Nikita P.; Rowlands, David D.; Luthcke, Scott B.; Chinn, Douglas S.; Marr, Gregory C.; Smith, David E. (Technical Monitor)

2000-01-01

The US Navy's GEOSAT Follow-On spacecraft was launched on February 10, 1998 and the primary objective of the mission was to map the oceans using a radar altimeter. Three radar altimeter calibration campaigns have been conducted in 1999 and 2000. The spacecraft is tracked by satellite laser ranging (SLR) and Doppler beacons and a limited amount of data have been obtained from the Global Positioning Receiver (GPS) on board the satellite. Even with EGM96, the predicted radial orbit error due to gravity field mismodelling (to 70x70) remains high at 2.61 cm (compared to 0.88 cm for TOPEX). We report on the preliminary gravity model tuning for GFO using SLR, and altimeter crossover data. Preliminary solutions using SLR and GFO/GFO crossover data from CalVal campaigns I and II in June-August 1999, and January-February 2000 have reduced the predicted radial orbit error to 1.9 cm and further reduction will be possible when additional data are added to the solutions. The gravity model tuning has improved principally the low order m-daily terms and has reduced significantly the geographically correlated error present in this satellite orbit. In addition to gravity field mismodelling, the largest contributor to the orbit error is the non-conservative force mismodelling. We report on further nonconservative force model tuning results using available data from over one cycle in beta prime.

High-resolution CASSINI-VIMS mosaics of Titan and the icy Saturnian satellites

Science.gov (United States)

Jaumann, R.; Stephan, K.; Brown, R.H.; Buratti, B.J.; Clark, R.N.; McCord, T.B.; Coradini, A.; Capaccioni, F.; Filacchione, G.; Cerroni, P.; Baines, K.H.; Bellucci, G.; Bibring, J.-P.; Combes, M.; Cruikshank, D.P.; Drossart, P.; Formisano, V.; Langevin, Y.; Matson, D.L.; Nelson, R.M.; Nicholson, P.D.; Sicardy, B.; Sotin, Christophe; Soderbloom, L.A.; Griffith, C.; Matz, K.-D.; Roatsch, Th.; Scholten, F.; Porco, C.C.

2006-01-01

The Visual Infrared Mapping Spectrometer (VIMS) onboard the CASSINI spacecraft obtained new spectral data of the icy satellites of Saturn after its arrival at Saturn in June 2004. VIMS operates in a spectral range from 0.35 to 5.2 ??m, generating image cubes in which each pixel represents a spectrum consisting of 352 contiguous wavebands. As an imaging spectrometer VIMS combines the characteristics of both a spectrometer and an imaging instrument. This makes it possible to analyze the spectrum of each pixel separately and to map the spectral characteristics spatially, which is important to study the relationships between spectral information and geological and geomorphologic surface features. The spatial analysis of the spectral data requires the determination of the exact geographic position of each pixel on the specific surface and that all 352 spectral elements of each pixel show the same region of the target. We developed a method to reproject each pixel geometrically and to convert the spectral data into map projected image cubes. This method can also be applied to mosaic different VIMS observations. Based on these mosaics, maps of the spectral properties for each Saturnian satellite can be derived and attributed to geographic positions as well as to geological and geomorphologic surface features. These map-projected mosaics are the basis for all further investigations. ?? 2006 Elsevier Ltd. All rights reserved.

Early Assessment of VIIRS On-Orbit Calibration and Support Activities

Science.gov (United States)

Xiong, Xiaoxiong; Chiang, Kwofu; McIntire, Jeffrey; Oudrari, Hassan; Wu, Aisheng; Schwaller, Mathew; Butler, James

2012-01-01

The Suomi National Polar-orbiting Partnership (S-NPP) satellite, formally the National Polar-orbiting Operational Environmental Satellite System (NPOESS) Preparatory Project (NPP), provides a bridge between current and future low-Earth orbiting weather and environmental observation satellite systems. The NASA s NPP VIIRS Characterization Support Team (VCST) is designed to assess the long term geometric and radiometric performance of the Visible Infrared Imaging Radiometer Suite (VIIRS) instrument onboard the S-NPP spacecraft and to support NPP Science Team Principal Investigators (PI) for their independent evaluation of VIIRS Environmental Data Records (EDRs). This paper provides an overview of Suomi NPP VIIRS on-orbit calibration activities and examples of sensor initial on-orbit performance. It focuses on the radiometric calibration support activities and capabilities provided by the NASA VCST.

From Data to Equations: Inferring the Laws governing Saturn's Ring Temperature

Science.gov (United States)

Altobelli, N.; Lopez-Paz, D.; Spilker, L.; Pilorz, S.

2011-10-01

Six years after Saturn Orbit Insertion (SOI), the Composite Infrared Spectrometer (CIRS) on-board the Cassini Spacecraft has been performing a thermal mapping of Saturn's main rings, by measuring the thermal radiance in the far-infrared ( [10-600] cm-1 ) for different viewing geometries. So far, more than 2.5 millions individual spectra have been recorded, from Saturn's northern winter solstice till Saturn's northern spring. We present a first attempt of treating the data set globally by applying numerical data mining techniques inherited from the field of artificial intelligence, such as neural networks and genetic programing.

Last Looks at the Eye of Saturn by Cassini/VIMS During the Grand Finale

Science.gov (United States)

Momary, Thomas W.; Baines, Kevin H.; Badman, Sarah; Brown, Robert H.; Buratti, Bonnie J.; Clark, Roger Nelson; Nicholson, Philip D.; Sotin, Christophe

2017-10-01

A lasting remnant of the Great Storm that erupted on Saturn in late 2010 has been a massive lone anticyclone persisting to the present time in a NH3-dry 5-µm-bright “desert” zone that spans the entire Saturnian globe at 34o N. We have been observing this oval storm with Cassini/VIMS since 2011 and, in 2017, as Cassini performs its Grand Finale orbits close to the planet, have captured it at our highest resolution since January 2012 at 260 km/pixel - enough to resolve spiral structure inside the oval at 5 µm. The spot drifts latitudinally in Saturn’s zonal currents: it was at 35.9o planetocentric latitude in May 2011, wandered northward to 37.8o in 2012, hovered near 37o through 2013, meandered as far south as 36.5o in 2014, drifted northward to 37o in 2015, and then returned back to about 36.3o in 2016, where it remains presently. It has also periodically bumped up against the dark band above it, spinning off material in 2013, 2015, and 2017. We measured a prograde zonal drift speed of 22 m/s in 2012, increasing as much as 60% through 2013, then relaxing to a more moderate 15 m/s in 2014 and 2015. It slowed considerably in 2016 to 4.7 m/s and is currently drifting slightly faster at 8.5 m/s. The spot has varied in size over time as it spins, spanning 4.9o x 3.2o in 2011, elongating to 7.3o x 2.9o by 2013, contracting to 5.5o x 2.9o in 2014, enlarging again to 9o x 4o in 2015, and contracting currently to 7.0o x 3.2o (6100 x 3200 km) in 2017, symmetrically oval in shape. It has varied in terms of cloudiness, being 90% 5-µm dark (obscured) in 2011, whereas by 2013 it was mostly bright (clear) with a thin dark edge. It was 90% dark in 2015, and in 2017 is about 65% obscured, with a bright central eye. Utilizing night observations to isolate thermal flux, we have found that the mean 5-µm flux coming from the anticyclone has diminished steadily by about 75% since 2013. The entire storm latitude of ~34o N itself has remained persistently 5-µm bright since 2011

Orbital Normalization of MESSENGER Gamma-Ray Spectrometer Data

Science.gov (United States)

Rhodes, E. A.; Peplowski, P. N.; Evans, L. G.; Hamara, D. K.; Boynton, W. V.; Solomon, S. C.

2011-12-01

The MESSENGER Gamma-Ray Spectrometer (GRS) measures energy spectra of gamma rays emanating from the surface of Mercury. Analysis of these spectra provides elemental abundances of surface material. The MESSENGER mission necessarily provides some data normalization challenges for GRS analysis. So as to keep the spacecraft cool while orbiting the dayside of the planet, the orbits are highly eccentric, with altitudes varying from 200-500 km to ~ 15,000 km. A small fraction of time is spent at the low altitudes where gamma-ray signals are largest, requiring a large number of orbits to yield sufficient counting statistics for elemental analysis. Also, the sunshade must always shield the spacecraft from the Sun, which causes the orientation of the GRS often to be far from nadir-pointing, so the detector efficiency and attenuation of gamma rays from the planet must be known for a wide range of off-nadir orientations. An efficiency/attenuation map for the expected ranges of orientations and energies was constructed in a ground calibration experiment for a limited range of orientations using a nuclear reactor and radioisotope sources, and those results were extended to other orientations by radiation transport computations using as input a computer-aided design model of the spacecraft and its composition. This normalization has allowed abundance determinations of elements K, Th, and U from radioisotopes of these elements in the Mercury regolith during the first quarter of the year-long mission. These results provide constraints on models of Mercury's chemical and thermal evolution. The normalization of gamma-ray spectra for surface elements not having radioisotopes is considerably more complex; these gamma rays come from neutron inelastic-scatter and capture reactions in the regolith, where the neutrons are generated by cosmic ray impact onto the planet. A radiation transport computation was performed to generate the expected count rates in the neutron-generated gamma

Using Solar Radiation Pressure to Control L2 Orbits

Science.gov (United States)

Tene, Noam; Richon, Karen; Folta, David

1998-01-01

The main perturbations at the Sun-Earth Lagrange points L1 and L2 are from solar radiation pressure (SRP), the Moon and the planets. Traditional approaches to trajectory design for Lagrange-point orbits use maneuvers every few months to correct for these perturbations. The gravitational effects of the Moon and the planets are small and periodic. However, they cannot be neglected because small perturbations in the direction of the unstable eigenvector are enough to cause exponential growth within a few months. The main effect of a constant SRP is to shift the center of the orbit by a small distance. For spacecraft with large sun-shields like the Microwave Anisotropy Probe (MAP) and the Next Generation Space Telescope (NGST), the SRP effect is larger than all other perturbations and depends mostly on spacecraft attitude. Small variations in the spacecraft attitude are large enough to excite or control the exponential eigenvector. A closed-loop linear controller based on the SRP variations would eliminate one of the largest errors to the orbit and provide a continuous acceleration for use in controlling other disturbances. It is possible to design reference trajectories that account for the periodic lunar and planetary perturbations and still satisfy mission requirements. When such trajectories are used the acceleration required to control the unstable eigenvector is well within the capabilities of a continuous linear controller. Initial estimates show that by using attitude control it should be possible to minimize and even eliminate thruster maneuvers for station keeping.

Feasibility Study for a Near Term Demonstration of Laser-Sail Propulsion from the Ground to Low Earth Orbit

Science.gov (United States)

Montgomery, Edward E., IV; Johnson, Les; Thomas, Herbert D.

2016-01-01

This paper adds to the body of research related to the concept of propellant-less in-space propulsion utilizing an external high energy laser (HEL) to provide momentum to an ultra-lightweight (gossamer) spacecraft. It has been suggested that the capabilities of Space Situational Awareness assets and the advanced analytical tools available for fine resolution orbit determination make it possible to investigate the practicalities of a ground to Low Earth Orbit (LEO) demonstration at delivered power levels that only illuminate a spacecraft without causing damage to it. The degree to which this can be expected to produce a measurable change in the orbit of a low ballistic coefficient spacecraft is investigated. Key system characteristics and estimated performance are derived for a near term mission opportunity involving the LightSail 2 spacecraft and laser power levels modest in comparison to those proposed previously by Forward, Landis, or Marx. [1,2,3] A more detailed investigation of accessing LightSail 2 from Santa Rosa Island on Eglin Air Force Base on the United States coast of the Gulf of Mexico is provided to show expected results in a specific case.

Cassini ISS Observations of Jupiter: An Exoplanet Perspective

Science.gov (United States)

West, Robert A.; Knowles, Benjamin

2017-10-01

Understanding the optical and physical properties of planets in our solar system can guide our approach to the interpretation of observations of exoplanets. Although some work has already been done along these lines, there remain low-hanging fruit. During the Cassini Jupiter encounter, the Imaging Science Subsystem (ISS) obtained an extensive set of images over a large range of phase angles (near-zero to 140 degrees) and in filters from near-UV to near-IR, including three methane bands and nearby continuum. The ISS also obtained images using polarizers. Much later in the mission we also obtained distant images while in orbit around Saturn. Some of these data have already been studied to reveal phase behavior (Dyudina et al., Astrophys. J.822, DOI: 10.3847/0004-637X/822/2/76; Mayorga et al., 2016, Astron. J. 152, DOI: 10.3847/0004-6256/152/6/209). Here we examine rotational modulation to determine wavelength and phase angle dependence, and how these may depend on cloud and haze vertical structure and optical properties. The existence of an optically thin forward-scattering and longitudinally-homogeneous haze overlying photometrically-variable cloud fields tends to suppress rotational modulation as phase angle increases, although in the strong 890-nm methane band cloud vertical structure is important. Cloud particles (non-spherical ammonia ice, mostly) have very small polarization signatures at intermediate phase angles and rotational modulation is not apparent above the noise level of our instrument. Part of this work was performed by the Jet Propulsion Lab, Cal. Inst. Of Technology.

Cancellation of differential accelerations for the LISA spacecraft

International Nuclear Information System (INIS)

Bender, Peter L

2006-01-01

The three spacecraft of the Laser Interferometer Space Antenna will form a nearly equilateral triangle with nominal side lengths of 5 million km. However, the arm lengths and the corner angles will vary by very roughly 1% over 5-10 years. Part of this variation is due to the nature of Kepler orbits around the Sun. But Sweetser (2006 Astrodynamics 2005, Advances in the Astronautical Sciences vol 123 (San Diego, CA: Univelt Inc.) pp 693-712) has shown recently that differential secular accelerations due to the Earth for the three spacecraft prevent the minimal variations from being preserved for more than 2 or 3 years. Based on Sweetser's results, it appears possible to cancel out the differential secular acceleration due to the Earth by applying forces to the two proof masses in each spacecraft. The applied acceleration is at most 2.1 x 10 -9 m s -2 . However, the directions of the required accelerations would have substantial components along the sides of the triangle, and thus the amplitudes of the applied forces would have to have very low noise, even at frequencies down to below 0.1 mHz

MESSENGER at Mercury: Early Orbital Operations

Science.gov (United States)

McNutt, Ralph L., Jr; Solomon, Sean C.; Bedini, Peter D.; Anderson, Brian J.; Blewett, David T.; Evans, Larry G.; Gold, Robert E.; Krimigis, Stamatios M.; Murchie, Scott L.; Nittler, Larry R.;

Energy deposition and ion production from thermal oxygen ion precipitation during Cassini's T57 flyby

Science.gov (United States)

Snowden, Darci; Smith, Michael; Jimson, Theodore; Higgins, Alex

2018-05-01

Cassini's Radio Science Investigation (RSS) and Langmuir Probe observed abnormally high electron densities in Titan's ionosphere during Cassini's T57 flyby. We have developed a three-dimensional model to investigate how the precipitation of thermal magnetospheric O+ may have contributed to enhanced ion production in Titan's ionosphere. The three-dimensional model builds on previous work because it calculates both the flux of oxygen through Titan's exobase and the energy deposition and ion production rates in Titan's atmosphere. We find that energy deposition rates and ion production rates due to thermal O+ precipitation have a similar magnitude to the rates from magnetospheric electron precipitation and that the simulated ionization rates are sufficient to explain the abnormally high electron densities observed by RSS and Cassini's Langmuir Probe. Globally, thermal O+ deposits less energy in Titan's atmosphere than solar EUV, suggesting it has a smaller impact on the thermal structure of Titan's neutral atmosphere. However, our results indicate that thermal O+ precipitation can have a significant impact on Titan's ionosphere.

An Integrated Vision-Based System for Spacecraft Attitude and Topology Determination for Formation Flight Missions

Science.gov (United States)

Rogers, Aaron; Anderson, Kalle; Mracek, Anna; Zenick, Ray

2004-01-01

With the space industry's increasing focus upon multi-spacecraft formation flight missions, the ability to precisely determine system topology and the orientation of member spacecraft relative to both inertial space and each other is becoming a critical design requirement. Topology determination in satellite systems has traditionally made use of GPS or ground uplink position data for low Earth orbits, or, alternatively, inter-satellite ranging between all formation pairs. While these techniques work, they are not ideal for extension to interplanetary missions or to large fleets of decentralized, mixed-function spacecraft. The Vision-Based Attitude and Formation Determination System (VBAFDS) represents a novel solution to both the navigation and topology determination problems with an integrated approach that combines a miniature star tracker with a suite of robust processing algorithms. By combining a single range measurement with vision data to resolve complete system topology, the VBAFDS design represents a simple, resource-efficient solution that is not constrained to certain Earth orbits or formation geometries. In this paper, analysis and design of the VBAFDS integrated guidance, navigation and control (GN&C) technology will be discussed, including hardware requirements, algorithm development, and simulation results in the context of potential mission applications.

Resonant Orbital Dynamics in LEO Region: Space Debris in Focus

Directory of Open Access Journals (Sweden)

J. C. Sampaio

2014-01-01

Full Text Available The increasing number of objects orbiting the earth justifies the great attention and interest in the observation, spacecraft protection, and collision avoidance. These studies involve different disturbances and resonances in the orbital motions of these objects distributed by the distinct altitudes. In this work, objects in resonant orbital motions are studied in low earth orbits. Using the two-line elements (TLE of the NORAD, resonant angles and resonant periods associated with real motions are described, providing more accurate information to develop an analytical model that describes a certain resonance. The time behaviors of the semimajor axis, eccentricity, and inclination of some space debris are studied. Possible irregular motions are observed by the frequency analysis and by the presence of different resonant angles describing the orbital dynamics of these objects.

Cassini Scientist for a Day: an international contest in Greece

Science.gov (United States)

Solomonidou, Anezina; Moussas, Xenophon; Xystouris, Georgios; Coustenis, Athena; Lebreton, Jean-Pierre; Katsavrias, Christos; Bampasidis, Georgios; Kyriakopoulos, Konstantinos; Kouloumvakos, Athanasios; Patsou, Ioanna

2013-04-01

The Cassini Outreach Team of NASA's Jet Propulsion Laboratory is being organizing a brilliant school contest in Astronomy focusing in the Saturnian system. This essay contest provides school students all around the worlds with the opportunity to get involved in astronomy and astrophysics and planetary sciences in particular. From 2010 the 'Cassini Scientist for a Day' contest has being one of the most successful as well as important outreach activities of ESA and NASA in Greece with hundreds of participants all over Greece. The number of participants is growing rapidly every year. This type of school competition in Greece is particularly important since Astronomy and Astrophysics and Space Sciences, although very popular, are not included in the school curricula and thus students rarely have the opportunity to experience and participate actively in these subjects. For the years 2010 and 2011, the Space Physics Group of the Astronomy, Astrophysics and Mechanics section of the University of Athens in association with external colleagues has been selected as the co-ordinator of NASA for the competition in Greece. Under the guidance of Cassini Outreach team, the members of the Space Physics Group have informed, explained and spread the rules of the competition at primary, secondary and high schools all over Greece. In general, the students have the option to choose Cassini monitoring between three targets of the Saturnian system, which the participants show that will bring the best scientific result. Their arguments should be summarized in an essay of 500 words more or less. They also have the option to do team work through groups of maximum three students. The participation in the contest for 2010 was unexpectedly high and thoroughly satisfied. The winners awarded through a ceremony which was held in the largest amphitheater at the central building of the University of Athens, that was fully packed. The following year 2011 the participation increased up to 300% while

Project ECHO: Electronic Communications from Halo Orbit

Science.gov (United States)

Borrelli, Jason; Cooley, Bryan; Debole, Marcy; Hrivnak, Lance; Nielsen, Kenneth; Sangmeister, Gary; Wolfe, Matthew

1994-01-01

The design of a communications relay to provide constant access between the Earth and the far side of the Moon is presented. Placement of the relay in a halo orbit about the L2 Earth-Moon Lagrange point allows the satellite to maintain constant simultaneous communication between Earth and scientific payloads on the far side of the Moon. The requirements of NASA's Discovery-class missions adopted and modified for this design are: total project cost should not exceed $150 million excluding launch costs, launch must be provided by Delta-class vehicle, and the satellite should maintain an operational lifetime of 10 to 15 years. The spacecraft will follow a transfer trajectory to the L2 point, after launch by a Delta II 7925 vehicle in 1999. Low-level thrust is used for injection into a stationkeeping-free halo orbit once the spacecraft reaches the L2 point. The shape of this halo orbit is highly elliptical with the maximum excursion from the L2 point being 35000 km. A spun section and despun section connected through a bearing and power transfer assembly (BAPTA) compose the structure of the spacecraft. Communications equipment is placed on the despun section to provide for a stationary dual parabolic offset-feed array antenna system. The dual system is necessary to provide communications coverage during portions of maximum excursion on the halo orbit. Transmissions to the NASA Deep Space Network 34 m antenna include six channels (color video, two voice, scientific data from lunar payloads, satellite housekeeping and telemetry and uplinked commands) using the S- and X-bands. Four radioisotope thermoelectric generators (RTG's) provide a total of 1360 W to power onboard systems and any two of the four Hughes 13 cm ion thrusters at once. Output of the ion thrusters is approximately 17.8 mN each with xenon as the propellant. Presence of torques generated by solar pressure on the antenna dish require the addition of a 'skirt' extending from the spun section of the satellite

Automating Trend Analysis for Spacecraft Constellations

Science.gov (United States)

Davis, George; Cooter, Miranda; Updike, Clark; Carey, Everett; Mackey, Jennifer; Rykowski, Timothy; Powers, Edward I. (Technical Monitor)

2001-01-01

Spacecraft trend analysis is a vital mission operations function performed by satellite controllers and engineers, who perform detailed analyses of engineering telemetry data to diagnose subsystem faults and to detect trends that may potentially lead to degraded subsystem performance or failure in the future. It is this latter function that is of greatest importance, for careful trending can often predict or detect events that may lead to a spacecraft's entry into safe-hold. Early prediction and detection of such events could result in the avoidance of, or rapid return to service from, spacecraft safing, which not only results in reduced recovery costs but also in a higher overall level of service for the satellite system. Contemporary spacecraft trending activities are manually intensive and are primarily performed diagnostically after a fault occurs, rather than proactively to predict its occurrence. They also tend to rely on information systems and software that are oudated when compared to current technologies. When coupled with the fact that flight operations teams often have limited resources, proactive trending opportunities are limited, and detailed trend analysis is often reserved for critical responses to safe holds or other on-orbit events such as maneuvers. While the contemporary trend analysis approach has sufficed for current single-spacecraft operations, it will be unfeasible for NASA's planned and proposed space science constellations. Missions such as the Dynamics, Reconnection and Configuration Observatory (DRACO), for example, are planning to launch as many as 100 'nanospacecraft' to form a homogenous constellation. A simple extrapolation of resources and manpower based on single-spacecraft operations suggests that trending for such a large spacecraft fleet will be unmanageable, unwieldy, and cost-prohibitive. It is therefore imperative that an approach to automating the spacecraft trend analysis function be studied, developed, and applied to

Large-Scale Spacecraft Fire Safety Experiments in ISS Resupply Vehicles

Science.gov (United States)

Ruff, Gary A.; Urban, David

2013-01-01

Our understanding of the fire safety risk in manned spacecraft has been limited by the small scale of the testing we have been able to conduct in low-gravity. Fire growth and spread cannot be expected to scale linearly with sample size so we cannot make accurate predictions of the behavior of realistic scale fires in spacecraft based on the limited low-g testing to date. As a result, spacecraft fire safety protocols are necessarily very conservative and costly. Future crewed missions are expected to be longer in duration than previous exploration missions outside of low-earth orbit and accordingly, more complex in terms of operations, logistics, and safety. This will increase the challenge of ensuring a fire-safe environment for the crew throughout the mission. Based on our fundamental uncertainty of the behavior of fires in low-gravity, the need for realistic scale testing at reduced gravity has been demonstrated. To address this concern, a spacecraft fire safety research project is underway to reduce the uncertainty and risk in the design of spacecraft fire safety systems by testing at nearly full scale in low-gravity. This project is supported by the NASA Advanced Exploration Systems Program Office in the Human Exploration and Operations Mission Directorate. The activity of this project is supported by an international topical team of fire experts from other space agencies to maximize the utility of the data and to ensure the widest possible scrutiny of the concept. The large-scale space flight experiment will be conducted on three missions; each in an Orbital Sciences Corporation Cygnus vehicle after it has deberthed from the ISS. Although the experiment will need to meet rigorous safety requirements to ensure the carrier vehicle does not sustain damage, the absence of a crew allows the fire products to be released into the cabin. The tests will be fully automated with the data downlinked at the conclusion of the test before the Cygnus vehicle reenters the

Failure analysis of satellite subsystems to define suitable de-orbit devices

Science.gov (United States)

Palla, Chiara; Peroni, Moreno; Kingston, Jennifer

2016-11-01

Space missions in Low Earth Orbit (LEO) are severely affected by the build-up of orbital debris. A key practice, to be compliant with IADC (Inter-Agency Space Debris Coordination Committee) mitigation guidelines, is the removal of space systems that interfere with the LEO region not later than 25 years after the End of Mission. It is important to note that the current guidelines are not generally legally binding, even if different Space Agencies are now looking at the compliance for their missions. If the guidelines will change in law, it will be mandatory to have a postmission disposal strategy for all satellites, including micro and smaller classes. A potential increased number of these satellites is confirmed by different projections, in particular in the commercial sector. Micro and smaller spacecraft are, in general, not provided with propulsion capabilities to achieve a controlled re-entry, so they need different de-orbit disposal methods. When considering the utility of different debris mitigation methods, it is useful to understand which spacecraft subsystems are most likely to fail and how this may affect the operation of a de-orbit system. This also helps the consideration of which components are the most relevant or should be redundant depending on the satellite mass class. This work is based on a sample of LEO and MEO satellites launched between January 2000 and December 2014 with mass lower than 1000 kg. Failure analysis of satellite subsystems is performed by means of the Kaplan-Meier survival analysis; the parametric fits are conducted with Weibull distributions. The study is carried out by using the satellite database SpaceTrak™ which provides anomalies, failures, and trends information for spacecraft subsystems and launch vehicles. The database identifies five states for each satellite subsystem: three degraded states, one fully operational state, and one failed state (complete failure). The results obtained can guide the identification of the

SMART-1: the first spacecraft of the future

Science.gov (United States)

2003-09-01

gather high-value scientific and technological data. Another innovation lies in the industrial policy applied to this mission. SMART-1 is a good example of an ESA mission in which a comparatively small company such as the Swedish Space Corporation (SSC) has been selected as prime contractor. “The experience of SSC in highly successful projects at national level was a key factor in the decision, as was ESA's goal of fostering a balanced industrial landscape in Europe,” says Niels Jensen of ESA’s Directorate of Industrial Matters and Technology Programmes. The magic of ion engines Solar-electric propulsion, one of the main technologies to be tested by SMART-1, is a new technique that uses 'ion engines'. These work by expelling a continuous beam of charged particles --ions-- at the back of the engine, which produces a thrust in the opposite direction and therefore pushes the spacecraft forward. The energy to feed the engine comes from the solar panels, hence the name 'solar-electric propulsion'. Engineers have been working on ion engines for decades, but only recently have obstacles such as the lack of power availability from a spacecraft’s solar panels been overcome. Recent missions have been using ion thrusters mainly for attitude control and orbit station keeping. In the recent case of ESA’s telecommunication satellite Artemis, the onboard availability of ion thrusters was actually what allowed the mission to be rescued. Having been left by the launcher on an unplanned orbit, Artemis was slowly - but safely - brought up to its final working orbit by the power of its ion engines, initially designed for orbit maintenance only. Starting with SMART-1, the first European spacecraft to use an ion engine as its main propulsion system, the amazing advantages of this method can now be fully exploited. Ion engines are very efficient: they deliver about ten times as much impulse per kilogram of propellant used. This gives a substantial reduction in the mass of the fuel

Operational Philosophy Concerning Manned Spacecraft Cabin Leaks

Science.gov (United States)

DeSimpelaere, Edward

2011-01-01

The last thirty years have seen the Space Shuttle as the prime United States spacecraft for manned spaceflight missions. Many lessons have been learned about spacecraft design and operation throughout these years. Over the next few decades, a large increase of manned spaceflight in the commercial sector is expected. This will result in the exposure of commercial crews and passengers to many of the same risks crews of the Space Shuttle have encountered. One of the more dire situations that can be encountered is the loss of pressure in the habitable volume of the spacecraft during on orbit operations. This is referred to as a cabin leak. This paper seeks to establish a general cabin leak response philosophy with the intent of educating future spacecraft designers and operators. After establishing a relative definition for a cabin leak, the paper covers general descriptions of detection equipment, detection methods, and general operational methods for management of a cabin leak. Subsequently, all these items are addressed from the perspective of the Space Shuttle Program, as this will be of the most value to future spacecraft due to similar operating profiles. Emphasis here is placed upon why and how these methods and philosophies have evolved to meet the Space Shuttle s needs. This includes the core ideas of: considerations of maintaining higher cabin pressures vs. lower cabin pressures, the pros and cons of a system designed to feed the leak with gas from pressurized tanks vs. using pressure suits to protect against lower cabin pressures, timeline and consumables constraints, re-entry considerations with leaks of unknown origin, and the impact the International Space Station (ISS) has had to the standard Space Shuttle cabin leak response philosophy. This last item in itself includes: procedural management differences, hardware considerations, additional capabilities due to the presence of the ISS and its resource, and ISS docking/undocking considerations with a

Mission Analysis and Orbit Control of Interferometric Wheel Formation Flying

Science.gov (United States)

Fourcade, J.

Flying satellite in formation requires maintaining the specific relative geometry of the spacecraft with high precision. This requirement raises new problem of orbit control. This paper presents the results of the mission analysis of a low Earth observation system, the interferometric wheel, patented by CNES. This wheel is made up of three receiving spacecraft, which follow an emitting Earth observation radar satellite. The first part of this paper presents trades off which were performed to choose orbital elements of the formation flying which fulfils all constraints. The second part presents orbit positioning strategies including reconfiguration of the wheel to change its size. The last part describes the station keeping of the formation. Two kinds of constraints are imposed by the interferometric system : a constraint on the distance between the wheel and the radar satellite, and constraints on the distance between the wheel satellites. The first constraint is fulfilled with a classical chemical station keeping strategy. The second one is fulfilled using pure passive actuators. Due to the high stability of the relative eccentricity of the formation, only the relative semi major axis had to be controlled. Differential drag due to differential attitude motion was used to control relative altitude. An autonomous orbit controller was developed and tested. The final accuracy is a relative station keeping better than few meters for a wheel size of one kilometer.

Shared control on lunar spacecraft teleoperation rendezvous operations with large time delay

Science.gov (United States)

Ya-kun, Zhang; Hai-yang, Li; Rui-xue, Huang; Jiang-hui, Liu

2017-08-01

Teleoperation could be used in space on-orbit serving missions, such as object deorbits, spacecraft approaches, and automatic rendezvous and docking back-up systems. Teleoperation rendezvous and docking in lunar orbit may encounter bottlenecks for the inherent time delay in the communication link and the limited measurement accuracy of sensors. Moreover, human intervention is unsuitable in view of the partial communication coverage problem. To solve these problems, a shared control strategy for teleoperation rendezvous and docking is detailed. The control authority in lunar orbital maneuvers that involves two spacecraft as rendezvous and docking in the final phase was discussed in this paper. The predictive display model based on the relative dynamic equations is established to overcome the influence of the large time delay in communication link. We discuss and attempt to prove via consistent, ground-based simulations the relative merits of fully autonomous control mode (i.e., onboard computer-based), fully manual control (i.e., human-driven at the ground station) and shared control mode. The simulation experiments were conducted on the nine-degrees-of-freedom teleoperation rendezvous and docking simulation platform. Simulation results indicated that the shared control methods can overcome the influence of time delay effects. In addition, the docking success probability of shared control method was enhanced compared with automatic and manual modes.

Precise Orbit Determination of QZS-1

Science.gov (United States)

Hugentobler, U.; Steigenberger, P.; Rodriguez-Solano, C.; Hauschild, A.

2011-12-01

QZS-1, the first satellite of the Japanese Quasi Zenith Satellite System (QZSS) was launched in September 2010. Transmission of the standard codes started in December 2010 and the satellite was declared healthy in June 2011. Five stations of the COoperative Network for GIOVE Observation (CONGO) were upgraded to provide QZSS tracking capability. These five stations provide the basis for the precise orbit determination (POD) of the QZS-1 spacecraft. The stability and consistency of different orbital arc lengths is analyzed based on orbit fit residuals, day boundary discontinuities, and Satellite Laser Ranging residuals. As QZS-1 simultaneously transmits navigation signals on three frequencies in the L1, L2, and L5 band, different ionosphere-free linear combinations can be formed. The differences of the orbits computed from these different observables (ionosphere-free linear combination of L1/L2 and L1/L5) as well as the stability of the differential code biases estimated within the POD are studied. Finally, results of the attitude determination based on the navigation signal transmission from two different antennas onboard QZS-1 are presented.

Advanced Communication Technology Satellite (ACTS) Multibeam Antenna On-Orbit Performance

Science.gov (United States)

1995-01-01

The NASA Lewis Research Center's Advanced Communication Technology Satellite (ACTS) was launched in September 1993. ACTS introduced several new technologies, including a multibeam antenna (MBA) operating at extremely short wavelengths never before used in communications. This antenna, which has both fixed and rapidly reconfigurable high-energy spot beams (150 miles in diameter), serves users equipped with small antenna terminals. Extensive structural and thermal analyses have been performed for simulating the ACTS MBA on-orbit performance. The results show that the reflector surfaces (mainly the front subreflector), antenna support assembly, and metallic surfaces on the spacecraft body will be distorted because of the thermal effects of varying solar heating, which degrade the ACTS MBA performance. Since ACTS was launched, a number of evaluations have been performed to assess MBA performance in the space environment. For example, the on-orbit performance measurements found systematic environmental disturbances to the MBA beam pointing. These disturbances were found to be imposed by the attitude control system, antenna and spacecraft mechanical alignments, and on-orbit thermal effects. As a result, the MBA may not always exactly cover the intended service area. In addition, the on-orbit measurements showed that antenna pointing accuracy is the performance parameter most sensitive to thermal distortions on the front subreflector surface and antenna support assemblies. Several compensation approaches were tested and evaluated to restore on-orbit pointing stability. A combination of autotrack (75 percent of the time) and Earth sensor control (25 percent of the time) was found to be the best way to compensate for antenna pointing error during orbit. This approach greatly minimizes the effects of thermal distortions on antenna beam pointing.

Temperature Condition and Spherical Shell Shape Variation of Space Gauge-Alignment Spacecraft

Directory of Open Access Journals (Sweden)

V. S. Zarubin

2016-01-01

Full Text Available A high precision spherical shell is one of the geometrical shape embodiments of a gaugealignment spacecraft to determine and control a radar channel energy potential of the ground-based complex for the traffic control of space objects. Passive relays of signals and some types of smallsized instrumentation standard reflectors used for radar gauge and alignment have the same shape. Orbits of the considered spacecraft can be either circular with a height of about 1000 km, including those close to the polar, or elliptical with an apogee of up to 2200 km.In case there is no thermal control system in spacecrafts of these types the solar radiation is a major factor to define the thermal state of a spherical shell in the illuminated orbit area. With the shell in fixed position with respect to direction towards the Sun an arising uneven temperature distribution over its surface leads to variation of the spherically ideal shell shape, which may affect the functional characteristics of the spacecraft. The shell rotation about an axis perpendicular to the direction towards the Sun may reduce an unevenness degree of the temperature distribution.The uneven temperature distribution over the spherical shell surface in conditions of the lowEarth space and this unevenness impact on the shell shape variation against its spherical shape can be quantively estimated by the appropriate methods of mathematical modeling using modification of a previously developed mathematical model to describe steady temperature state of such shell on the low-Earth orbit. The paper considers the shell made from a polymeric composite material. Its original spherical shape is defined by rather low internal pressure. It is assumed that equipment in the shell, if any, is quite small-sized. This allows us to ignore its impact on the radiative transfer in the shell cavity. Along with defining the steady temperature distribution over the shell surface at its fixed orientation with respect to

The Laser Ranging Experiment of the Lunar Reconnaissance Orbiter: Five Years of Operations and Data Analysis

Science.gov (United States)

Mao, Dandan; McGarry, Jan F.; Mazarico, Erwan; Neumann, Gregory A.; Sun, Xiaoli; Torrence, Mark H.; Zagwodzki, Thomas W.; Rowlands, David D.; Hoffman, Evan D.; Horvath, Julie E.;

Electric sail elliptic displaced orbits with advanced thrust model

Science.gov (United States)

Niccolai, Lorenzo; Quarta, Alessandro A.; Mengali, Giovanni

2017-09-01

This paper analyzes the performance of an Electric Solar Wind Sail for generating and maintaining an elliptic, heliocentric, displaced non-Keplerian orbit. In this sense, this paper extends and completes recent studies regarding the performances of an Electric Solar Wind Sail that covers a circular, heliocentric, displaced orbit of given characteristics. The paper presents the general equations that describe the elliptic orbit maintenance in terms of both spacecraft attitude and performance requirements, when a refined thrust model (recently proposed for the preliminary mission design) is taken into account. In particular, the paper also discusses some practical applications on particular mission scenarios in which an analytic solution of the governing equations has been found.

Capture orbits around asteroids by hitting zero-velocity curves

Science.gov (United States)

Wang, Wei; Yang, Hongwei; Zhang, Wei; Ma, Guangfu

2017-12-01

The problem of capturing a spacecraft from a heliocentric orbit into a high parking orbit around binary asteroids is investigated in the current study. To reduce the braking Δ V, a new capture strategy takes advantage of the three-body gravity of the binary asteroid to lower the inertial energy before applying the Δ V. The framework of the circular restricted three-body problem (CR3BP) is employed for the binary asteroid system. The proposed capture strategy is based on the mechanism by which inertial energy can be decreased sharply near zero-velocity curves (ZVCs). The strategy has two steps, namely, hitting the target ZVC and raising the periapsis by a small Δ V at the apoapsis. By hitting the target ZVC, the positive inertial energy decreases and becomes negative. Using a small Δ V, the spacecraft inserts into a bounded orbit around the asteroid. In addition, a rotating mass dipole model is employed for elongated asteroids, which leads to dynamics similar to that of the CR3BP. With this approach, the proposed capture strategy can be applied to elongated asteroids. Numerical simulations validate that the proposed capture strategy is applicable for the binary asteroid 90 Antiope and the elongated asteroid 216 Kleopatra.

Spacecraft Line-of-Sight Stabilization Using LWIR Earth Signature

Science.gov (United States)

Quadrelli, Marco B.; Piazzolla, Sabino

2012-01-01

The objective of this study is to investigate the potential of using the bright and near-uniform Earth infrared (or wavelength infrared, LWIR) signature as a stable reference for accurate (micro-rad or less) inertial pointing and tracking on-board an space vehicle, including the determination of the fundamental limits of applicability of the proposed method for space missions. We demonstrate sub-micro radian level pointing accuracy under a representative set of disturbances experienced by the spacecraft in orbit.

The BepiColombo MORE gravimetry and rotation experiments with the ORBIT14 software

Science.gov (United States)

Cicalò, S.; Schettino, G.; Di Ruzza, S.; Alessi, E. M.; Tommei, G.; Milani, A.

2016-04-01

The BepiColombo mission to Mercury is an ESA/JAXA cornerstone mission, consisting of two spacecraft in orbit around Mercury addressing several scientific issues. One spacecraft is the Mercury Planetary Orbiter, with full instrumentation to perform radio science experiments. Very precise radio tracking from Earth, on-board accelerometer and optical measurements will provide large data sets. From these it will be possible to study the global gravity field of Mercury and its tidal variations, its rotation state and the orbit of its centre of mass. With the gravity field and rotation state, it is possible to constrain the internal structure of the planet. With the orbit of Mercury, it is possible to constrain relativistic theories of gravitation. In order to assess that all the scientific goals are achievable with the required level of accuracy, full cycle numerical simulations of the radio science experiment have been performed. Simulated tracking, accelerometer and optical camera data have been generated, and a long list of variables including the spacecraft initial conditions, the accelerometer calibrations and the gravity field coefficients have been determined by a least-squares fit. The simulation results are encouraging: the experiments are feasible at the required level of accuracy provided that some critical terms in the accelerometer error are moderated. We will show that BepiColombo will be able to provide at least an order of magnitude improvement in the knowledge of Love number k2, libration amplitudes and obliquity, along with a gravity field determination up to degree 25 with a signal-to-noise ratio of 10.

Low-cost autonomous orbit control about Mars: Initial simulation results

Science.gov (United States)

Dawson, S. D.; Early, L. W.; Potterveld, C. W.; Königsmann, H. J.

1999-11-01

Interest in studying the possibility of extraterrestrial life has led to the re-emergence of the Red Planet as a major target of planetary exploration. Currently proposed missions in the post-2000 period are routinely calling for rendezvous with ascent craft, long-term orbiting of, and sample-return from Mars. Such missions would benefit greatly from autonomous orbit control as a means to reduce operations costs and enable contact with Mars ground stations out of view of the Earth. This paper present results from initial simulations of autonomously controlled orbits around Mars, and points out possible uses of the technology and areas of routine Mars operations where such cost-conscious and robust autonomy could prove most effective. These simulations have validated the approach and control philosophies used in the development of this autonomous orbit controller. Future work will refine the controller, accounting for systematic and random errors in the navigation of the spacecraft from the sensor suite, and will produce prototype flight code for inclusion on future missions. A modified version of Microcosm's commercially available High Precision Orbit Propagator (HPOP) was used in the preparation of these results due to its high accuracy and speed of operation. Control laws were developed to allow an autonomously controlled spacecraft to continuously control to a pre-defined orbit about Mars with near-optimal propellant usage. The control laws were implemented as an adjunct to HPOP. The GSFC-produced 50 × 50 field model of the Martian gravitational potential was used in all simulations. The Martian atmospheric drag was modeled using an exponentially decaying atmosphere based on data from the Mars-GRAM NASA Ames model. It is hoped that the simple atmosphere model that was implemented can be significantly improved in the future so as to approach the fidelity of the Mars-GRAM model in its predictions of atmospheric density at orbital altitudes. Such additional work

High-Resolution Mid-IR Imaging of Jupiter's Great Red Spot: Comparing Cassini, VLT and Subaru Observations

Science.gov (United States)

Fletcher, Leigh N.; Orton, G. S.; Yanamandra-Fisher, P.; Irwin, P. G. J.; Baines, K. H.; Edkins, E.; Line, M. R.; Mousis, O.; Parrish, P. D.; Vanzi, L.; Fuse, T.; Fujoyoshi, T.

2008-09-01

In the eight years since the Cassini fly-by of Jupiter, the spatial resolution of ground-based observations of Jupiter's giant anticyclonic storm systems (the Great Red Spot, Oval BA and others) using 8m-class telescopes has surpassed the resolution of the Cassini/CIRS maps. We present a time-series of mid-IR imaging of the Great Red Spot (GRS) and its environs from the VISIR instrument on the Very Large Telescope (UT3/Melipal) and the COMICS instrument on the Subaru telescope (Hawaii). The NEMESIS optimal-estimation retrieval algorithm (Irwin et al., 2008) is used to analyse both the 7-25 micron filtered imaging from 2005-2008 and Cassini/CIRS 7-16 micron data from 2000. We demonstrate the ability to map temperatures in the 100-400 mbar range, NH3, aerosol opacity and the para-H2 fraction from the filtered imaging. Furthermore, the Cassini/CIRS spectra are used to map the PH3 mole fraction around the GRS. The thermal field, gaseous composition and aerosol distribution are used as diagnostics for the atmospheric motion associated with the GRS. Changes in the atmospheric state in response to close encounters with Oval BA and other vortices will be assessed. These results will be discussed in light of their implications for the planning of the Europa-Jupiter System Mission.

Using The Global Positioning System For Earth Orbiter and Deep Space Network

Science.gov (United States)

Lichten, Stephen M.; Haines, Bruce J.; Young, Lawrence E.; Dunn, Charles; Srinivasan, Jeff; Sweeney, Dennis; Nandi, Sumita; Spitzmesser, Don

1994-01-01

The Global Positioning System (GPS) can play a major role in supporting orbit and trajectory determination for spacecraft in a wide range of applications, including low-Earth, high-earth, and even deep space (interplanetary) tracking.

SSS-A spacecraft and experiment description.

Science.gov (United States)

Longanecker, G. W.; Hoffman, R. A.

1973-01-01

The scientific objectives of the Explorer-45 mission are discussed. The primary objective is the study of the ring current responsible for the main phase of magnetic storms. Closely associated with this objective is the determination of the relationship between magnetic storms, substorms, and the acceleration of charged particles in the magnetosphere. Further objectives are the measurement of a wide range of proton, electron and alpha-particle energies, and studies of wave-particle interactions responsible for particle transport and loss in the inner magnetosphere. The orbital parameters, the spacecraft itself, and some of its unique features, such as the data handling system, which is programmable from the ground, are described.

A Model of Titan-like Chemistry to Connect Experiments and Cassini Observations

Science.gov (United States)

Raymond, Alexander W.; Sciamma-O’Brien, Ella; Salama, Farid; Mazur, Eric

2018-02-01

A numerical model is presented for interpreting the chemical pathways that lead to the experimental mass spectra acquired in the Titan Haze Simulation (THS) laboratory experiments and for comparing the electron density and temperature of the THS plasma to observations made at Titan by the Cassini spacecraft. The THS plasma is a pulsed glow-discharge experiment designed to simulate the reaction of N2/CH4-dominated gas in Titan's upper atmosphere. The transient, one-dimensional model of THS chemistry tracks the evolution of more than 120 species in the direction of the plasma flow. As the minor species C2H2 and C2H4 are added to the N2/CH4-based mixture, the model correctly predicts the emergence of reaction products with up to five carbon atoms in relative abundances that agree well with measured mass spectra. Chemical growth in Titan's upper atmosphere transpires through ion–neutral and neutral–neutral chemistry, and the main reactions involving a series of known atmospheric species are retrieved from the calculation. The model indicates that the electron density and chemistry are steady during more than 99% of the 300 μs long discharge pulse. The model also suggests that the THS ionization fraction and electron temperature are comparable to those measured in Titan's upper atmosphere. These findings reaffirm that the THS plasma is a controlled analog environment for studying the first and intermediate steps of chemistry in Titan's upper atmosphere.

Validation of Galileo orbits using SLR with a focus on satellites launched into incorrect orbital planes

Science.gov (United States)

Sośnica, Krzysztof; Prange, Lars; Kaźmierski, Kamil; Bury, Grzegorz; Drożdżewski, Mateusz; Zajdel, Radosław; Hadas, Tomasz

2018-02-01

The space segment of the European Global Navigation Satellite System (GNSS) Galileo consists of In-Orbit Validation (IOV) and Full Operational Capability (FOC) spacecraft. The first pair of FOC satellites was launched into an incorrect, highly eccentric orbital plane with a lower than nominal inclination angle. All Galileo satellites are equipped with satellite laser ranging (SLR) retroreflectors which allow, for example, for the assessment of the orbit quality or for the SLR-GNSS co-location in space. The number of SLR observations to Galileo satellites has been continuously increasing thanks to a series of intensive campaigns devoted to SLR tracking of GNSS satellites initiated by the International Laser Ranging Service. This paper assesses systematic effects and quality of Galileo orbits using SLR data with a main focus on Galileo satellites launched into incorrect orbits. We compare the SLR observations with respect to microwave-based Galileo orbits generated by the Center for Orbit Determination in Europe (CODE) in the framework of the International GNSS Service Multi-GNSS Experiment for the period 2014.0-2016.5. We analyze the SLR signature effect, which is characterized by the dependency of SLR residuals with respect to various incidence angles of laser beams for stations equipped with single-photon and multi-photon detectors. Surprisingly, the CODE orbit quality of satellites in the incorrect orbital planes is not worse than that of nominal FOC and IOV orbits. The RMS of SLR residuals is even lower by 5.0 and 1.5 mm for satellites in the incorrect orbital planes than for FOC and IOV satellites, respectively. The mean SLR offsets equal -44.9, -35.0, and -22.4 mm for IOV, FOC, and satellites in the incorrect orbital plane. Finally, we found that the empirical orbit models, which were originally designed for precise orbit determination of GNSS satellites in circular orbits, provide fully appropriate results also for highly eccentric orbits with variable linear

Validation of double Langmuir probe in-orbit performance onboard a nano-satellite

Science.gov (United States)

Tejumola, Taiwo Raphael; Zarate Segura, Guillermo Wenceslao; Kim, Sangkyun; Khan, Arifur; Cho, Mengu

2018-03-01

Many plasma measurement systems have been proposed and used onboard different satellites to characterize space plasma. Most of these systems employed the technique of Langmuir probes either using the single or double probes methods. Recent growth of lean satellites has positioned it on advantage to be used for space science missions using Langmuir probes because of its simplicity and convenience. However, single Langmuir probes are not appropriate to be used on lean satellites because of their limited conducting area which leads to spacecraft charging and drift of the instrument's electrical ground during measurement. Double Langmuir probes technique can overcome this limitation, as a measurement reference in relation to the spacecraft is not required. A double Langmuir probe measurement system was designed and developed at Kyushu Institute of Technology for HORYU-IV satellite, which is a 10 kg, 30 cm cubic class lean satellite launched into Low Earth Orbit on 17th February 2016. This paper presents the on-orbit performance and validation of the double Langmuir probe measurement using actual on-orbit measured data and computer simulations.

CASSINI E/J/S/SW MIMI INCA SENSOR UNCALIBRATED DATA V1.1

Data.gov (United States)

National Aeronautics and Space Administration — The Cassini Magnetospheric Imaging Instrument(MIMI) Imaging Neutral Camera (INCA) uncalibrated data set includes all data collected from the MIMI Data Processing...

CASSINI E/J/S/SW MIMI INCA SENSOR UNCALIBRATED DATA V1.0

Data.gov (United States)

National Aeronautics and Space Administration — The Cassini Magnetospheric Imaging Instrument(MIMI) Imaging Neutral Camera (INCA) uncalibrated data set includes all data collected from the MIMI Data Processing...

Cassini CAPS Identification of Pickup Ion Compositions at Rhea

Science.gov (United States)

Desai, R. T.; Taylor, S. A.; Regoli, L. H.; Coates, A. J.; Nordheim, T. A.; Cordiner, M. A.; Teolis, B. D.; Thomsen, M. F.; Johnson, R. E.; Jones, G. H.; Cowee, M. M.; Waite, J. H.

2018-02-01

Saturn's largest icy moon, Rhea, hosts a tenuous surface-sputtered exosphere composed primarily of molecular oxygen and carbon dioxide. In this Letter, we examine Cassini Plasma Spectrometer velocity space distributions near Rhea and confirm that Cassini detected nongyrotropic fluxes of outflowing CO2+ during both the R1 and R1.5 encounters. Accounting for this nongyrotropy, we show that these possess comparable along-track densities of ˜2 × 10-3 cm-3. Negatively charged pickup ions, also detected during R1, are surprisingly shown as consistent with mass 26 ± 3 u which we suggest are carbon-based compounds, such as CN-, C2H-, C2-, or HCO-, sputtered from carbonaceous material on the moon's surface. The negative ions are calculated to possess along-track densities of ˜5 × 10-4 cm-3 and are suggested to derive from exogenic compounds, a finding consistent with the existence of Rhea's dynamic CO2 exosphere and surprisingly low O2 sputtering yields. These pickup ions provide important context for understanding the exospheric and surface ice composition of Rhea and of other icy moons which exhibit similar characteristics.

Programs To Optimize Spacecraft And Aircraft Trajectories

Science.gov (United States)

Brauer, G. L.; Petersen, F. M.; Cornick, D.E.; Stevenson, R.; Olson, D. W.

1994-01-01

POST/6D POST is set of two computer programs providing ability to target and optimize trajectories of powered or unpowered spacecraft or aircraft operating at or near rotating planet. POST treats point-mass, three-degree-of-freedom case. 6D POST treats more-general rigid-body, six-degree-of-freedom (with point masses) case. Used to solve variety of performance, guidance, and flight-control problems for atmospheric and orbital vehicles. Applications include computation of performance or capability of vehicle in ascent, or orbit, and during entry into atmosphere, simulation and analysis of guidance and flight-control systems, dispersion-type analyses and analyses of loads, general-purpose six-degree-of-freedom simulation of controlled and uncontrolled vehicles, and validation of performance in six degrees of freedom. Written in FORTRAN 77 and C language. Two machine versions available: one for SUN-series computers running SunOS(TM) (LAR-14871) and one for Silicon Graphics IRIS computers running IRIX(TM) operating system (LAR-14869).

Thermal analysis for folded solar array of spacecraft in orbit

International Nuclear Information System (INIS)

Yang, W.H.; Cheng, H.E.; Cai, A.

2004-01-01

The combined radiation-conduction heat transfer in folded solar array was considered as a three-dimensional anisotropic conduction without inner heat source. The three-dimensional equivalent conductivity in cell plate were obtained. The especially discrete equation coefficients of the nodes on the surfaces of adjacent cell plates were deduced by utilizing the simplified radiation network among the two adjacent cell plate surfaces and the deep cold space. All the thermal influence factors on the temperature response of the folded solar array were considered carefully. SIP method was used to solve the discrete equation. By comparing the calculation results under three cases, the temperature response and the maximum average difference of the folded solar array was obtained during the period of throw-radome of the launch vehicle and spread of the folded solar array. The obtained result is a valuable reference for the selection of the launch time of the spacecraft

Orbital Debris and NASA's Measurement Program

Science.gov (United States)

Africano, J. L.; Stansbery, E. G.

2002-05-01

Since the launch of Sputnik in 1957, the number of manmade objects in orbit around the Earth has dramatically increased. The United States Space Surveillance Network (SSN) tracks and maintains orbits on over nine thousand objects down to a limiting diameter of about ten centimeters. Unfortunately, active spacecraft are only a small percentage ( ~ 7%) of this population. The rest of the population is orbital debris or ``space junk" consisting of expended rocket bodies, dead payloads, bits and pieces from satellite launches, and fragments from satellite breakups. The number of these smaller orbital debris objects increases rapidly with decreasing size. It is estimated that there are at least 130,000 orbital debris objects between one and ten centimeters in diameter. Most objects smaller than 10 centimeters go untracked! As the orbital debris population grows, the risk to other orbiting objects, most importantly manned space vehicles, of a collision with a piece of debris also grows. The kinetic energy of a solid 1 cm aluminum sphere traveling at an orbital velocity of 10 km/sec is equivalent to a 400 lb. safe traveling at 60 mph. Fortunately, the volume of space in which the orbiting population resides is large, collisions are infrequent, but they do occur. The Space Shuttle often returns to earth with its windshield pocked with small pits or craters caused by collisions with very small, sub-millimeter-size pieces of debris (paint flakes, particles from solid rocket exhaust, etc.), and micrometeoroids. To get a more complete picture of the orbital-debris environment, NASA has been using both radar and optical techniques to monitor the orbital debris environment. This paper gives an overview of the orbital debris environment and NASA's measurement program.

Temperature variations in Titan's upper atmosphere: Impact on Cassini/Huygens

Directory of Open Access Journals (Sweden)

B. Kazeminejad

2005-06-01

Full Text Available Temperature variations of Titan's upper atmosphere due to the plasma interaction of the satellite with Saturn's magnetosphere and Titan's high altitude monomer haze particles can imply an offset of up to ±30K from currently estimated model profiles. We incorporated these temperature uncertainties as an offset into the recently published Vervack et al. (2004 (Icarus, Vol. 170, 91-112 engineering model and derive extreme case (i.e. minimum and maximum profiles temperature, pressure, and density profiles. We simulated the Huygens probe hypersonic entry trajectory and obtain, as expected, deviations of the probe trajectory for the extreme atmosphere models compared to the simulation based on the nominal one. These deviations are very similar to the ones obtained with the standard Yelle et al. (1997 (ESA SP-1177 profiles. We could confirm that the difference in aerodynamic drag is of an order of magnitude that can be measured by the probe science accelerometer. They represent an important means for the reconstruction of Titan's upper atmospheric properties. Furthermore, we simulated a Cassini low Titan flyby trajectory. No major trajectory deviations were found. The atmospheric torques due to aerodynamic drag, however, are twice as high for our high temperature profile as the ones obtained with the Yelle maximum profile and more than 5 times higher than the worst case estimations from the Cassini project. We propose to use the Cassini atmospheric torque measurements during its low flybys to derive the atmospheric drag and to reconstruct Titan's upper atmosphere density, pressure, and temperature. The results could then be compared to the reconstructed profiles obtained from Huygens probe measurements. This would help to validate the probe measurements and decrease the error bars.