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Viewgraphs on orbital debris environmental monitor (ODEM) are presented. Topics covered include: Long Duration Exposure Facility (LDEF); interplanetary dust experiment; orbital debris clouds; mapping and modeling of orbital debris clouds; and solar maximum mission spacecraft.
In the past 25 years, the majority of interplanetary spacecraft have been powered by nuclear sources. However, as the emphasis on smaller, low cost missions gains momentum, the majority of missions now being planned will use photovoltaic solar arrays. This will present challenges to the solar array builders, inasmuch as planetary requirements usually differ from earth orbital requirements. In addition, these requirements often differ greatly, depending on the specific mission; for example, inner planets vs. outer planets, orbiters vs. flybys, spacecraft vs. landers, and so on. Also, the likelihood of electric propulsion missions will influence the requirements placed on solar array developers. The paper will discuss representative requirements for a range of planetary missions now in the planning stages. Insofar as inner planets are concerned, a Mercury orbiter is being studied with many special requirements. Solar arrays would be exposed to high temperatures and a potentially high radiation environment, and will need to be increasingly pointed off sun as the vehicle approaches Mercury. Identification and development of cell materials and arrays at high incidence angles will be critical to the design. Missions to the outer solar system that have been studied include a Galilean orbiter and a flight to the Kuiper belt. While onboard power requirements would be small (as low as 10 watts), the solar intensity will require relatively large array areas. As a result, such missions will demand extremely compact packaging and low mass structures to conform to launch vehicle constraints. In turn, the large are, low mass designs will impact allowable spacecraft loads. Inflatable array structures, with and without concentration, and multiband gap cells will be considered if available. In general, the highest efficiency cell technologies operable under low intensity, low temperature conditions are needed. Solar arrays will power missions requiring as little as approximately 100 watts, up to several kilowatts (at Earth) in the case of solar electric propulsion missions. Thus, mass and stowage volume minimization will be required over a range of array sizes. Concentrator designs, inflatable structures, and the combination of solar arrays with the telecommunications system have been proposed. Performance, launch vehicle constraints, an cost will be the principal parameters in the design trade space. Other special applications will also be discussed, including requirements relating to planetary landers and probes. In those cases, issues relating to shock loads on landing, operability in (possibly dusty) atmospheres, and extreme temperature cycles must be considered, in addition to performance, stowed volume, and costs.
The recent discovery of possible fossilized life forms within a Martian meteorite has rekindled interest in human exploration of the planet. While the global scientific community must carefully study this physical evidence as well as await the results of 4 space probes to Mars in the coming years, others are looking ahead to the next logical step of sending humans to the planet. Such a feat, while a tremendous scientific undertaking, is nearly within the grasp of current technology. The travel distance involved and the harsh Martian environment itself impose great challenges to human physiology as well as to spacecraft design. Mission planners face challenges from the long-duration microgravity environment, complex life-support systems, interplanetary radiation, and the psychological stressors involved in such an endeavor.
In the past 25 years, the majority of interplanetary spacecraft have been powered by nuclear sources. However, as the emphasis on smaller, low cost missions gains momentum, more deep space missions now being planned have baselined photovoltaic solar arrays due to the low power requirements (usually significantly less than 100 W) needed for engineering and science payloads. This will present challenges to the solar array builders, inasmuch as planetary requirements usually differ from earth orbital requirements. In addition, these requirements often differ greatly, depending on the specific mission; for example, inner planets vs. outer planets, orbiters vs. flybys, spacecraft vs. landers, and so on. Also, the likelihood of electric propulsion missions will influence the requirements placed on solar array developers. This paper will discuss representative requirements for a range of planetary and deep space science missions now in the planning stages. We have divided the requirements into three categories: Inner planets and the sun; outer planets (greater than 3 AU); and Mars, cometary, and asteroid landers and probes. Requirements for Mercury and Ganymede landers will be covered in the Inner and Outer Planets sections with their respective orbiters. We will also discuss special requirements associated with solar electric propulsion (SEP). New technology developments will be needed to meet the demanding environments presented by these future applications as many of the technologies envisioned have not yet been demonstrated. In addition, new technologies that will be needed reside not only in the photovoltaic solar array, but also in other spacecraft systems that are key to operating the spacecraft reliably with the photovoltaics.
This software provides capabilities for autonomous cross-cueing and coordinated observations between multiple orbital and landed assets. Previous work has been done in re-tasking a single Earth orbiter or a Mars rover in response to that craft detecting a science event. This work enables multiple spacecraft to communicate (over a network designed for deep-space communications) and autonomously coordinate the characterization of such a science event. This work investigates a new paradigm of space science campaigns where opportunistic science observations are autonomously coordinated among multiple spacecraft. In this paradigm, opportunistic science detections can be cued by multiple assets where a second asset is requested to take additional observations characterizing the identified surface feature or event. To support this new paradigm, an autonomous science system for multiple spacecraft assets was integrated with the Interplanetary Network DTN (Delay Tolerant Network) to provide communication between spacecraft assets. This technology enables new mission concepts that are not feasible with current technology. The ability to rapidly coordinate activities across spacecraft without requiring ground in the loop enables rapid reaction to dynamic events across platforms, such as a survey instrument followed by a targeted high resolution instrument, as well as regular simultaneous observations.
The Mars Science Laboratory spacecraft, carrying the Curiosity rover to Mars, hit the top of the Martian atmosphere just 200 meters from where it had been predicted more than six days earlier, and 2.6 million kilometers away. This un-expected level of accuracy was achieved by a combination of factors including: spacecraft performance, tracking data processing, dynamical modeling choices, and navigation filter setup. This paper will describe our best understanding of what were the factors that contributed to this excellent interplanetary trajectory prediction performance. The accurate interplanetary navigation contributed to the very precise landing performance, and to the overall success of the mission.
The Parker Solar Probe and the incoming Solar Orbiter mission will provide measurements of solar energetic particle (SEP) events at close heliocentric distances from the Sun. Up to present, the largest data set of SEP events in the inner heliosphere are the observations by the two Helios spacecraft. Aims.We re-visit a sample of 15 solar relativistic electron events measured by the Helios mission with the goal of better characterizing the injection histories of solar energetic particles and their interplanetary transport conditions at heliocentric distances <1AU. Methods. The measurements provided by the E6 instrument on board Helios provide us with the electron directional distributionsin eight di erent sectors that we use to infer the detailed evolution of the electron pitch-angle distributions. The results of a Monte Carlo interplanetary transport model, combined with a full inversion procedure, were used to fit the observed directional intensities in the 300–800 keV nominal energy channel. Unlike previous studies, we have considered both the energy and angular responses of the detector. This method allowed us to infer the electron release time profile at the source and determine the electron interplanetary transport conditions. Results. We discuss the duration of the release time profiles and the values of the radial mean free path, and compare them with thevalues reported previously in the literature using earlier approaches. Five of the events show short injection histories (<30 min) at the Sun and ten events show long-lasting (>30 min) injections. The values of mean free path range from 0.02AU to 0.27AU.Conclusions. The inferred injection histories match with the radio and soft X-ray emissions found in literature.We find no dependence of the radial mean free path on the radial distance. In addition, we find no apparent relation between the strength of interplanetary scattering and the size of the solar particle release.
Presolar grains condensed in the outflows of evolved red giant stars and the ejecta of supernovae (SNe) and novae. These grains have greatly anomalous isotopic compositions compared to solar system material, reflecting their stellar origins [1]. They have been identified in primitive meteorites, interplanetary dust particles (IDPs), Antarctic micrometeorites, and comet Wild 2 samples returned by NASA’s Stardust mission. Presolar silicates are one of the most abundant presolar phases and their concentrations extend up to 1.5% in primitive IDPs believed to derive from comets [2]. These grains are highly susceptible to alteration and destruction by secondary processing in the interstellar medium, nebula, and asteroid or comet parent body. Presolarsilicate abundance variations between primitive meteorites and chemical and mineralogical studies provide indications of the extent of secondary hydrothermal alteration [3, 4]. The abundance of presolar SiC grains is generally consistent among chondrites, but lower abundances in some meteorites have been attributed to thermal alteration [5, 6]. Presolar grain abundance variations attributed to localized alteration have also been reported within different regions of a chondrite [e.g., 7].Carbonaceous chondrites from the CO and CR groups have the highest abundances of presolar silicates among meteorites, attesting to their primitive nature. The CO3 chondrite Dominion Range (DOM) 08006 has the highest presolar O-rich grain abundance of ~260 ppm [8, 9]. MillerRange (MIL) 090019 is classified as a CO3.1 chondrite and has affinities to Acfer 094, DOM 08004/6 and Allan Hills (ALH) 77307. These chondrites have high presolar silicate abundances and contain high abundances of various types of refractory inclusions. We previously conducted detailed studies of CAIs in MIL 090019 [10, 11]. Here we evaluate its presolar grain inventory to assess the degree of parent body alteration and compare to other chondrites
Dawn is a low-thrust interplanetary spacecraft currently orbiting the dwarf planet Ceres. After successfully completing its Vesta and Ceres prime missions, Dawn is now in its extended mission, continuously exploring Ceres. After losing the second reaction wheel assembly at the end of the Vesta mission, the feasibility of the Ceres mission was at risk due to the potentially significant increase in hydrazine consumption from using only the reaction control system thrusters. This paper summarizes the intense, collaborative efforts undertaken by the project to conserve hydrazine prior to Vesta departure, during cruise to Ceres, and throughout the Ceres mission. The project’s efforts in minimizing the number of turns are discussed. A special emphasis is given to describing various efforts taken by the Attitude Control Subsystem team in an attempt to reduce hydrazine consumption. These efforts include: changing default slew rates, control gain tuning, changing nadir pointing strategy with Ahead Cross Nadir pointing, hybrid control implementation using hydrazine based thrusters and two remaining reaction wheels, optimal science turn location analysis, and reaction wheel angular momentum management strategies for reducing the number of momentum unloadings. In addition to the Attitude Control Subsystem team’s efforts, the Mission Design and Navigation team’s efforts taken in designing hydrazine friendly orbit transfers are discussed. The trade study performed collaboratively by the Attitude Control Subsystem team and the Science Operations Support Team in choosing hydrazine friendly off-nadir targets is also discussed. Various simulation results are presented and compared against actual flight data obtained. Lastly, remaining hydrazine status at the end of the prime Ceres mission and the hydrazine management planning for the extended mission are discussed.
Excitement is building for the first launch of NASA’s Space Launch System (SLS), a unique exploration asset for the agency’s Artemis lunar program as well as for a new generation of science missions. SLS is designed for an array of missions beyond Earth’s orbit. The flexible system, which can be configured for Orion, cargo or Orion with co-manifested payload missions, offers high escape velocities to send more mass to deep space destinations. When configured with an 8.4 m-diameter fairing, SLS offers unmatched payload volume for human exploration and science missions. The initial Block 1 variant will insert at least 26 metric tons (t) to trans-lunar injection (TLI) and the more powerful Block 1B vehicle will launch 34-37 t to TLI using a new-development upper stage. Much of the initial SLS Block 1 vehicle is complete, including the upper stage and payload section, the core stage, engines and the solid rocket boosters. The first mission, Artemis I, launching from modernized and upgraded facilities at Kennedy Space Center (KSC), will be an uncrewed test flight of SLS, Orion and ground processing, with a primary objective of testing Orion’s heat shield at lunar re-entry velocity. Artemis I will have accommodations for 13 6U CubeSat payloads. These CubeSat missions will be deployed along the upper stage disposal trajectory after Orion separates from the vehicle. A rare opportunity for CubeSats to be deployed beyond low Earth orbit (LEO), Artemis I CubeSat missions range from searching for hydrogen and other volatiles on the lunar South Pole to studying the acceleration mechanisms of solar and interplanetary particles from a heliocentric trajectory. With manufacturing of the initial vehicle complete, fabrication and procurement is progressing for the second flight of SLS and Orion, Artemis II. Also an SLS Block 1 and Orion flight launching from KSC, Artemis II will mark the return of American astronauts to deep space with a lunar flyby-free return trajectory mission. With the Artemis III flight, NASA has the goal to land the first woman and the next man on the Moon. Infrastructure beyond SLS will be required for this effort, including elements of the lunar Gateway as well as lunar rovers, landers and additional commercially supplied launch services. SLS, as the only vehicle with the capability to lift 26 t of mass to TLI in its initial Block 1 variant, will remain a key component of this new-era exploration program. Future variants – Block 1B and Block 2 –will lift 34-45 t to TLI. This paper will discuss the status of testing and integration for the Artemis I vehicle, manufacturing progress for the second vehicle and the manifest outlook for primary, co-manifested and secondary payloads in the current deep space exploration environment.
Topics include: Measuring Image Navigation and Registration Performance at the 3-Sigma Level Using Platinum Quality Landmarks; Flight Dynamics Performances of the MetOp A Satellite during the First Months of Operations; Visual Navigation - SARE Mission; Determining a Method of Enabling and Disabling the Integral Torque in the SDO Science and Inertial Mode Controllers; Guaranteeing Pointing Performance of the SDO Sun-Pointing Controllers in Light of Nonlinear Effects; SDO Delta H Mode Design and Analysis; Observing Mode Attitude Controller for the Lunar Reconnaissance Orbiter; Broken-Plane Maneuver Applications for Earth to Mars Trajectories; ExoMars Mission Analysis and Design - Launch, Cruise and Arrival Analyses; Mars Reconnaissance Orbiter Aerobraking Daily Operations and Collision Avoidance; Mars Reconnaissance Orbiter Interplanetary Cruise Navigation; Motion Parameters Determination of the SC and Phobos in the Project Phobos-Grunt; GRAS NRT Precise Orbit Determination: Operational Experience; Orbit Determination of LEO Satellites for a Single Pass through a Radar: Comparison of Methods; Orbit Determination System for Low Earth Orbit Satellites; Precise Orbit Determination for ALOS; Anti-Collision Function Design and Performances of the CNES Formation Flying Experiment on the PRISMA Mission; CNES Approaching Guidance Experiment within FFIORD; Maneuver Recovery Analysis for the Magnetospheric Multiscale Mission; SIMBOL-X: A Formation Flying Mission on HEO for Exploring the Universe; Spaceborne Autonomous and Ground Based Relative Orbit Control for the TerraSAR-X/TanDEM-X Formation; First In-Orbit Experience of TerraSAR-X Flight Dynamics Operations; Automated Target Planning for FUSE Using the SOVA Algorithm; Space Technology 5 Post-Launch Ground Attitude Estimation Experience; Standardizing Navigation Data: A Status Update; and A Study into the Method of Precise Orbit Determination of a HEO Orbiter by GPS and Accelerometer.
All nuclear space acceleration system and basic spacecraft consisting of biconic earth entry module, mission module, vehicle control, experiment laboratories, and excursion module
Human exploration of space requires an understanding of the risks to which crews will be exposed during such missions, and the mitigation of those risks to the fullest extent practical. This becomes a greater imperative as we prepare for interplanetary expeditions involving long periods in weightlessness in transit to and then from the destination (a planet, such as Mars, or perhaps a point in space, such as the Lagrangian point L2), and exposure to the unique environment of the destination itself. We need to know, more definitively, what the risks are to human health, safety, and performance, and how to prevent or counteract them throughout all phases of a long duration mission. The Johnson Space Center's Space and Life Sciences Directorate and the National Space Biomedical Research Institute (NSBRI) have implemented an effort to identify the most critical risks confronting humans on such mission and the types of research and technology efforts required to mitigate and otherwise reduce the probability and severity of those risks. This paper describes the "Critical Path Roadmap Project" to define, assess and prioritize the risks and present the results of the assessment with an emphasis on the research and technology priorities to meet the challenge of long duration human spaceflight mission.
Galileo is now nearing the one-year mark along its three-year direct Earth-to-Jupiter transfer trajectory, having received it final interplanetary gravity assist from Earth on December 8, 1992.
The technical and economic feasibility of using electromagnetically launched EML payloads propelled from the Earth's surface to LEO, GEO, lunar orbit, or to interplanetary space was assessed. Analyses of the designs of rail accelerators and coaxial magnetic accelerators show that each is capable of launching to space payloads of 800 KG or more. A hybrid launcher in which EML is used for the first 2 KM/sec followed by chemical rocket stages was also tested. A cost estimates study shows that one to two EML launches per day are needed to break even, compared to a four-stage rocket. Development models are discussed for: (1) Earth orbital missions; (2) lunar base supply mission; (3) solar system escape mission; (4) Earth escape missions; (5) suborbital missions; (6) electromagnetic boost missions; and (7) space-based missions. Safety factors, environmental impacts, and EML systems analysis are discussed. Alternate systems examined include electrothermal thrustors, an EML rocket gun; an EML theta gun, and Soviet electromagnetic accelerators.
NASA’s Mars Study Capability Team continues the agency’s efforts to study and refine the nation’s plan to field a sustainable human Mars campaign. One of the primary open issues left unanswered during previous analysis cycles was the uncertainty of the impact to the vehicle performance requirement to deliver crew and cargo to the same landing site across multiple mission opportunities. The Mars Study Capability Team has recently developed an integrated trajectory optimization and system closure model to solve the complex interplanetary trajectory optimization using both low-thrust and high-thrust maneuvers. This paper demonstrates the capability of this new integrated trajectory design and optimization method as it applies to the landing site accessibility problem for the Hybrid transportation architecture. The results showed that the current vehicle is capable of reaching up to +20 degree latitude and down to -20 degree latitude in every mission opportunity from 2033 to 2052. However, reaching latitudes beyond +/- 20 required more propellant than the spacecraft is currently designed to carry for many of the mission opportunities.
The Charles Stark Draper Laboratory, Inc. is partnering with the NASA Marshall Space Flight Center (MSFC) Engineering Directorate's Avionics Design Division and Flight Mechanics & Analysis Division to develop and test a prototype small, low-weight, low-power, radiation-hardened, fault-tolerant mini-star tracker (fig. 1). The project is expected to enable Draper Laboratory and its small business partner, L-1 Standards and Technologies, Inc., to develop a new guidance, navigation, and control sensor product for the growing small sat technology market. The project also addresses MSFC's need for sophisticated small sat technologies to support a variety of science missions in Earth orbit and beyond. The prototype star tracker will be tested on the night sky on MSFC's Automated Lunar and Meteor Observatory (ALAMO) telescope. The specific goal of the project is to address the need for a compact, low size, weight, and power, yet radiation hardened and fault tolerant star tracker system that can be used as a stand-alone attitude determination system or incorporated into a complete attitude determination and control system for emerging interplanetary and operational CubeSat and small sat missions.
The quadrennium 1975-1978 was a period of great advance for solar-wind studies, a period that combined exploration of new regions with increased maturity in established fields of study. The Helios, Pioneer, and Voyager spacecraft have been exploring the inner and outer regions of the solar wind. There has been a rebirth of the study of possible relations between solar variability and Earth's climate and weather, stimulated largely by Eddy's investigation of the Maunder Minimum; the solar wind may well prove to be a significant link in solar-terrestrial relations. Unique coronal data from the SKYLAB 1973-1974 mission, in combination with satellite and ground-based observations, provided the basis for identification of coronal holes as the main source of highspeed solar wind. The interplanetary medium has continued to serve as a laboratory for the study of plasma processes that cannot yet be studied in terrestrial laboratories, providing insights of potential importance both for controlled fusion research and for astrophysics. It is ironic that such a productive period, the legacy of many past space missions, was also a time of severely limited opportunity for new space investigations; the outlook for the future is equally austere. Especially regrettable is the dearth of career opportunities for young scientists in this field; comparison of the bibliography of this report with that of its predecessor 4 years ago shows few new names. Despite such problems, research has continued with enthusiasm and much has been learned. The present report will survey selected topics related to the origin, expansion, and acceleration of the solar wind and the plasma physics of the interplanetary medium. Companion reports deal with a number of closely related topics, including the heliocentric distance and latitude variation of the solar wind and its fluctuations topology of the interplanetary magnetic field morphology of solar-wind streams and shocks, sunweather studies, and interplanetary manifestations of type-3 bursts. Of the subjects that fall within the scope of this report, the study of the relationship between coronal holes and solar-wind streams, and the associated revision of our ideas about solar wind acceleration and heating, have had the most impact; hence I review these topics in considerable detail. In addition, I discuss the topics of hydromagnetic waves and turbulence, and interplanetary electrons, as items of particular importance during the past quadrennium.