The Solar TErrestrial RElations Observatory (STEREO) was originally designed for a two year heliocentric-orbit mission to study coronal mass ejections (CMEs), solar energetic particles (SEPs), and the solar wind. After over ten years of continuous science data collection, the twin NASA STEREO observatories have significantly advanced our understanding of Heliophysics. This mission was the first to image CMEs all the way from the Sun to Earth and has allowed us to observe solar activity over a full 360 degrees around the Sun.
STEREO has demonstrated the importance of a point of view beyond the Sun-Earth line to significantly improve CME arrival time estimates. It has also enhanced our understanding of CME structure and trajectories and the longitudinal distribution of SEPs. STEREO was also the first to use one launch vehicle to insert two spacecraft into opposing heliocentric orbits, undergo a 3.5 month long superior solar conjunction, implement unattended daily science operations on two deep space observatories, maintain 7 arcsec continuous pointing without gyros, and detect and attempt to recover a spacecraft after a 22 month long communications anomaly at a range of 2 AU. This paper discusses significant performance and science results after the first ten years of operations of the STEREO mission from its journey around the Sun. • 02.0102 Overview of the Spacecraft Design for the Psyche Mission Concept William Hart (NASA Jet Propulsion Laboratory), Peter Lord (SSL), Lindy Elkins Tanton (Arizona State University), Maria De Soria Santacruz Pich (Jet Propulsion Laboratory), Danielle Marsh (NASA Jet Propulsion Laboratory), Steve Snyder (Jet Propulsion Laboratory), Noah Warner (), Zachary Pirkl (San Jose State University) Presentation: William Hart - Sunday, March 4th, 04:55 PM - Jefferson. In January 2017, Psyche and a second mission concept were selected by NASA for flight as part of the 14th Discovery mission competition.
Assigned for an initial launch date in 2023, the Psyche team was given direction shortly after selection to research the possibility for earlier opportunities. Ultimately, the team was able to identify a launch opportunity in 2022 with a reduced flight time to its destination. This was accomplished in large part to crosscutting trades centered on the electrical power subsystem. These trades were facilitated through the Psyche mission's planned use of Solar Electric Propulsion (SEP), which enables substantial flexibility with respect to trajectory design. In combination with low-thrust trajectory analysis tools, the team was able to robustly converge to solutions with a higher fidelity and accuracy of results. These trades also took advantage of the 1300 series product line produced by Space Systems Loral (SSL), which enabled power growth while maintaining strong system-level heritage through its modular design that has been utilized on a large number of geostationary (GEO) communications satellites.
This paper presents an overview of the Psyche mission concept, and the unique architecture that enables the use of commercially developed electric propulsion and space power systems from Space Systems Loral (SSL) to provide flexibility in mission design. This paper then discusses the trades that allowed the Psyche team to meet a 2022 launch date. • 02.0103 The Ice, Cloud, and Land Elevation Satellite-2 – Overview, Science, and Applications Mark Seidleck (NASA - Goddard Space Flight Center) Presentation: Mark Seidleck - Sunday, March 4th, 05:20 PM - Jefferson. The Ice, Cloud, and Land Elevation Satellite – 2 (ICESat-2) was designed and built by NASA. It is planned to launch in September of 2018 with a designed mission life of 3 years to measure the elevation changes of the earth’s land and ice masses.
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This presentation describes the mission overview, spacecraft and instrument design and operation, anticipated science, and potential applications for the data. The discussion will also include a description of the technical issues encountered during development of the ATLAS instrument. • 02.0104 Juno at Jupiter: The Mission and Its Path to Unveiling Secrets of the History of the Solar System Stuart Stephens (Jet Propulsion Laboratory) Presentation: Stuart Stephens - Sunday, March 4th, 09:00 PM - Jefferson. This paper describes the Juno mission, focusing on its orbits at Jupiter, how the plan evolved, and science return so far. Previous papers described the history of the mission, its development, launch, and early cruise, spacecraft operations lessons learned, and pre-arrival plans for orbital operations. Juno is a NASA New Frontiers spacecraft currently in a near-polar highly elliptical 53-day orbit at the solar system’s largest planet. Since arriving at Jupiter in July 2016, it has used 9 science investigations to study its atmospheric composition and structure, magnetic and gravity fields, and polar and extended magnetosphere.
A radiation monitoring investigation contributes to our understanding of the Jupiter environment. Juno's primary science goal is to understand the origin and evolution of Jupiter, thereby shedding light on how the Earth and other planets formed. Its baseline science objectives will be satisfied with 32 orbits (minimum mission: 8 evenly spaced, or any 16), a spin-stabilized solar powered spacecraft, an electronics vault for radiation shielding, and a robust science payload including microwave receivers, X- and Ka-band radio science hardware, vector magnetometers, high- and low-energy charged particle detectors, radio and plasma wave antennas, UV and IR spectroscopic imagers, and a visible light public outreach camera. Science observations are made in a limited number of orientations, including one for gravity science (spin axis and main antenna pointing to Earth), and one for microwave atmospheric sounding (spin axis perpendicular to orbit plane to allow nadir pointing in spin plane). Prime science data are collected near closest approach (perijove), along with calibrations, occasional remote sensing, and continued magnetospheric observations in the outer parts of the orbits. The mission plan for Juno’s science orbits at Jupiter evolved following the 2005 proposal.
This was partly a result of design and operations choices, e.g., mission design (cruise and early orbital trajectory), orbit period (11, 14, then 53 days), perijove orientations (rationales for more than two), and DSN coverage (combinations of 34-m and 70-m antennas). In large part, the resulting choices were motivated by achieving or improving the science data return. The plan was also refined by flight system lessons learned, e.g., after safe modes or other anomalies.
Selected preliminary science results are summarized, including how they benefited from the choices made as the plan changed. Juno has begun to unveil Jupiter – peeling apart its interior by measuring its gravity and magnetic fields, using microwaves to probe the atmosphere down to 100s of km, exploring its polar and extended magnetosphere as never before, and imaging the poles for the first time. In the process, it is revealing secrets of the history of the Earth and the solar system. • 02.0105 Mission Concept for a Europa Lander Jennifer Dooley (Jet Propulsion Laboratory, California Institute of Technology) Presentation: Jennifer Dooley - Sunday, March 4th, 09:25 PM - Jefferson.
A NASA HQ-directed study team led by JPL with partners including APL, MSFC, GSFC, LaRC and Sandia National Laboratory has recently presented a mission concept for a Europa Lander that would search for bio-signatures and signs of life in the near-subsurface of the Jovian moon. The mission would follow the Europa Multiple-Flyby Mission Clipper planned for launch in June of 2022, which would provide reconnaissance imagery and other data to the Lander for use in selecting a scientifically compelling site and certifying it for engineering safety. The Europa Lander concept accommodates the Model Payload identified by the Europa Lander Science Definition Team (SDT) and documented in the Europa Lander Study 2016 Report released in February of 2017. The currently envisioned Europa Lander would launch on an SLS Block 1B as early as October of 2025 into a delta-VEGA trajectory, arriving in the Jovian system as early as July of 2030. The baseline design of the integrated flight system includes a dedicated Carrier and Relay Stage, a Deorbit Vehicle composed of a Deorbit Stage consisting of a solid rocket motor (SRM), an MSL-like sky-crane Descent Stage, and a Lander which accommodates the instrument suite.
The Lander would be powered by primary batteries over a 20-day surface mission. The science goals envisioned by the SDT require five samples taken from a depth of 10cm, a depth chosen to ensure minimal radiation processing of the potential biomarkers. Mission challenges include the large launch mass, unknown terrain topography, surface composition and materials properties, the high radiation environment, and complying with the stringent planetary protection requirements. The mission concept uses a strategy of early risk reduction and overlapping requirements to provide robustness to harsh and uncertain environments.
Early risk reduction efforts are aimed at maturing technologies associated with the sampling system, the intelligent landing system, high specific energy batteries, low mass and power motor controllers, and a thermal sterilization system. • 02.0106 Instrument Commissioning Timeline for NASA-ISRO Synthetic Aperture Radar (NISAR) Priyanka Sharma (Jet Propulsion Laboratory) Presentation: Priyanka Sharma - Sunday, March 4th, 09:50 PM - Jefferson. The NASA-ISRO Synthetic Aperture Radar (NISAR) is a joint collaboration between NASA and India’s national space agency, the Indian Space Research Organization (ISRO).
This Earth-orbiting Radar mission, which will be launched from Sriharikota (India) in December 2021, is designed to systematically and globally study solid Earth, ice masses, and ecosystems, all of which are sparsely sampled at present. The NISAR mission concept is derived from the DESDynI-Radar mission (Deformation, Ecosystem Structure and Dynamics of Ice), which was one of the four Tier 1 missions recommended by the National Research Council (NRC) in the 2007 Earth Science Decadal Survey.
NISAR will be launched on ISRO’s GSLV Mark-II launch vehicle, and will be carrying two Radar instruments: NASA’s L-SAR (L-band SAR, 24 cm wavelength) and ISRO’s S-SAR (S-band instrument, 12 cm wavelength). The Radars will employ a SweepSAR technique to attain a large swath (>240 km) for global data collection via a repeat orbit of 12 days in duration. The main objective of the NISAR mission is to enable studies of the causes and consequences of land surface change on Earth. Multiple scientific and applications disciplines will benefit from and utilize data from this mission, including solid Earth deformation, ecosystems, cryospheric studies, natural disasters and hazard assessment. The NISAR flight project entered Phase C (implementation phase) in August 2016. Following NISAR’s launch, the first 90 days will be dedicated to performing ‘in-orbit checkout’, or the ‘commissioning’ period, during which there will be a step-by-step buildup in capability to full observatory operations. Activities performed during commissioning will be aimed at demonstrating the full functionality of the Radar instruments (L-SAR and S-SAR), the reflector antenna, and the spacecraft and flight systems; characterizing and confirming their nominal performance within specifications.
All ground systems and infrastructure, including Ground Data Systems (GDS), Science Data Systems (SDS) and Mission Operations Systems (MOS), as well as the compatibility of all system interfaces, will be tested and validated. Calibration strategies for monitoring instrument calibration stability will be developed and tested by performing initial instrument calibrations. The commissioning phase will be divided into four sub-phases, starting with ‘Initial Checkout’, during which spacecraft and flight engineering systems will be powered on and calibrated. This will be followed by the ‘Reflector Boom Assembly (RBA) Deployments’ sub-phase. ‘Spacecraft checkout’ will follow the RBA deployment, during which the GPS sub-system will be turned on, and orbit-raising maneuvers will be executed to transfer the observatory from the initial injection orbit after launch to a science-like orbit (within 5-10 km of the NISAR Reference Science Orbit (RSO)). The final sub-phase of commissioning will be ‘instrument checkout’ during which the observatory will reach the RSO, both the L-SAR and S-SAR instruments will be powered on, and their performance will be characterized and calibrated.
The research described here was carried out in part at the Jet Propulsion Laboratory, California Institute of Technology, under a contract with the National Aeronautics and Space Administration. • 02.0108 The Surface Water and Ocean Topography Mission Guy Zohar (NASA Jet Propulsion Lab), Daniel Esteban Fernandez (NASA/JPL), Daniel Limonadi (JPL), Parag Vaze (Jet Propulsion Laboratory), Said Kaki (Jet Propulsion Laboratory) Presentation: Guy Zohar - Monday, March 5th, 04:30 PM - Jefferson. The SWOT mission is a partnership between two communities, physical oceanography and hydrology, to share high vertical accuracy and high spatial resolution topography data produced by the science payload, whose principal instrument is a Ka-band radar Interferometer. The SWOT mission will provide large-scale data sets of ocean sea-surface height resolving scales of 15km (in wavelength) and larger, allowing the characterization of ocean mesoscale and submesoscale circulation.
Present altimeter constellations can only resolve the ocean circulation at wavelengths larger than 200km. SWOT will address fundamental questions on the dynamics of ocean variability at wavelengths shorter than 200km, which encompasses mesoscale and submesoscale processes such as the formation, evolution, and dissipation of eddy variability (including narrow currents, fronts, and quasi-geostrophic turbulence) and their role in air-sea interaction. The SWOT mission will also provide measurements of water storage changes in terrestrial surface water bodies and will provide estimates of discharge in large (wider than 100m) rivers, globally. The SWOT measurements will directly measure the surface water (lakes, reservoirs, rivers, and wetlands) component of the water cycle. The core of the SWOT payload consists of a Ka-band Radar Interferometer (KaRIn) - a dual-antenna synthetic aperture radar specifically designed to make high precision height and backscatter measurements enabling the key oceanographic and hydrology data sets.
The SWOT payload also includes: a Nadir Altimeter (NA) system, a radiometer for tropospheric path delay corrections, and a precision orbit determination instrument suite consisting of Global Positioning System-Payload (GPSP), Doppler Orbitography and Radiopositioning Integrated by Satellite (DORIS) receivers, and a Laser Retro-reflector Assembly (LRA). SWOT is a partnership mission between NASA, the French Space Agency (CNES), and The Canadian Space Agency (CSA). The spacecraft bus and command & data ground stations are provided by CNES, while the launch vehicle and payload module are provided by NASA/JPL. The SWOT payload module instrument suite consists of the NASA/JPL provided KaRIn instrument, cross-track radiometer, GPS, and laser retroreflector, as well as the CNES provided DORIS and nadir pointing radar altimeter.
CNES also provides the Radio Frequency Unit (RFU) which is a key component of the KaRIn instrument. CSA contributes the Extended Interaction Klystron (EIK) which is one of the KaRIn subsystem. This paper describes the mission design, implementation of the Payload, and some of the interface and design challenges. • 02.0110 Europa Clipper Project Update Todd Bayer (NASA Jet Propulsion Lab), Maddalena Jackson (Jet Propulsion Laboratory) Presentation: Todd Bayer - Monday, March 5th, 04:55 PM - Jefferson. Europa, the fourth largest moon of Jupiter, is believed to be one of the best places in the solar system to look for extant life beyond Earth. The 2011 Planetary Decadal Survey, Vision and Voyages, states: “Because of this ocean’s potential suitability for life, Europa is one of the most important targets in all of planetary science”. Exploring Europa to investigate its habitability is the goal of the proposed Europa Clipper mission.
This exploration is intimately tied to understanding the three “ingredients” for life: liquid water, chemistry, and energy. The Europa Clipper mission would investigate these ingredients by comprehensively exploring Europa’s ice shell and liquid ocean interface, surface geology and surface composition to glean insight into the inner workings of this fascinating moon. In addition, a lander mission is seen as a possible future step, but current data about the Jovian radiation environment and about potential landing site hazards and potential safe landing zones is insufficient. Therefore an additional goal of the mission would be to characterize the radiation environment near Europa and investigate scientifically compelling sites for hazards, to inform a potential future landed mission.
The proposed Europa Clipper mission concept envisions sending a flight system, consisting of a spacecraft equipped with a payload of NASA-selected scientific instruments, to execute numerous flybys of Europa while in Jupiter orbit. A key challenge is that the flight system must survive and operate in the intense Jovian radiation environment, which is especially harsh at Europa. The innovative design of this multiple-flyby tour is an enabling feature of this mission: by minimizing the time spent in the radiation environment the spacecraft complexity and cost has been significantly reduced compared to previous mission concepts. The spacecraft would launch from Kennedy Space Center (KSC), Cape Canaveral, Florida, USA, on a NASA supplied launch vehicle, no earlier than 2022. The proposed mission would be formulated and implemented by a joint Jet Propulsion Laboratory (JPL) and Applied Physics Laboratory (APL) Project team. In January 2017 the proposed Clipper mission passed its System Requirements Review / Mission Definition Review (SRR/MDR) and in February NASA approved the mission for entry into Phase B, the Preliminary Design phase, with a planned launch in 2022. The Clipper name was officially sanctioned.
The Flight System Preliminary Design Review (PDR) is planned for October 2017 followed by Subsystem PDRs and then by the Project PDR in July 2018. Mass and power allocations have been made; a new tour trajectory has been adopted, key sizing design elements have been frozen (propulsion tanks, solar array, etc), and a down-selection to one launch vehicle by NASA is anticipated sometime this year. A decision to down-select to the NASA Space Launch System (SLS) would enable launch onto a direct-to-Jupiter trajectory, allowing significant system simplifications. This paper will describe the progress of the Europa Clipper Mission since January 2017, including maturation of the spacecraft design, requirements, system analyses, and mission trajectories. • 02.0111 Third Generation Commercial Solar Electric Propulsion: A Foundation for Space Exploration Missions Peter Lord (SSL) Presentation: Peter Lord - Monday, March 5th, 05:20 PM - Jefferson. Second-generation Solar Electric Propulsion (SEP) systems now in commercial production are a key step forward enabling cost-effective and mass-efficient in space transportation for exploration missions.
As designed with 5 kW class Hall Effect thrusters, they are suitable for Discovery class missions carrying significant sensors and/or targeting faraway objects. Performance evolution though a judicious methodology of 3x power/thrust scale up with existing SEP flight systems will provide for similar low-risk steps toward 15-kW and then 50-kW class thrusters including essential high power solar arrays and higher voltage power processing equipment. Larger payload missions supported by the latter capabilities directly benefit from US Government and commercial development of flexible solar arrays and very long life, third generation high-power Hall thruster technologies with high Isp performance. Fielding these technologies for exploration missions will benefit from commercial space product development approaches which are strongly focused on early product configuration decisions, modularity and scalability. Commercial practices require high reliability features for 15 year missions with rigorous qualification to an envelope of application environments and products designed for manufacturability. The benefits resulting from these uncompromising capabilities based approach include schedule-certain development, qualification, product insertion, and cost-effectiveness of recurring production. Valuable SEP technology enhancements in work at US Government labs should be transitioned to industry for commercial development completion to further benefit the cost effectiveness of these missions.
Anchoring these technologies in ongoing commercial production is vital to achieving both low cost and high reliability. In addition to a SEP roadmap covering enhanced commercial, smaller planetary, and large space transportation capabilities, two examples are highlighted of exploration missions based on commercial production SEP spacecraft optimized for deep space operations.
One illustrates how the Psyche Discovery missions SEP chassis design was extracted from SSL’s commercial product line as a FFP commercial item for the exploration of then main belt asteroid 16 Psyche. The other shows how a much larger Power Propulsion Element (PPE) or “space tug” in the 50 kW class that is capable of Cislunar and Mars missions and can be performed with replication of the same, modular commercial hardware and further enhanced in capability with an additional ~3x scale up of key SEP components. • 02. Impot Rapide 2010 Keygen Music. 0112 Cassini’s Grand Finale: A Mission Planning Retrospective Erick Sturm (Jet Propulsion Laboratory) Presentation: Erick Sturm - Monday, March 5th, 09:00 PM - Jefferson. In April 2017, Cassini began its Grand Finale, a series of 22½ orbits that take Cassini between Saturn and its rings.
At the completion of the Grand Finale in September 2017, Cassini will have flown through the inner fringes of Saturn’s D ring and upper reaches of its atmosphere. Significant work was done in the years leading up to the Grand Finale studying the region, assessing its hazards, and defining contingency plans to ensure the spacecraft’s health and safety and the return of its unique science data from that region. This paper will look back at the planned operational scenarios and contingencies for the unknown environment of Saturn’s proximal region and compare them with the actual, as-flown Grand Finale. The dust hazard contingency plan will be covered, along with how the project was able to leverage it to remove (rather than add) a dust hazard protection when the environment proved less hazardous than predicted. It will also discuss the atmosphere contingency plan and pop-down scenario and what will have happened in the final (and lowest) five orbits of the Grand Finale to which they both pertain.
The predicted versus actual propellant use and orbital maneuver strategy will be included. It will also describe the as-flown final plunge and compare it to the planned end-of-mission timeline. • 02.0113 InSight: Mars Geophysical Lander Tom Hoffman (Jet Propulsion Laboratory) Presentation: Tom Hoffman - Monday, March 5th, 09:25 PM - Jefferson.
The InSight Mission will uncover the geophysical characteristics of Mars and use comparative planetary geophysical techniques to better understand the formation and evolution of Mars and thus by extension other terrestrial planets. The InSight spacecraft has heritage from the 2001 Mars Lander which was used for the Phoenix mission and from more modern missions for the spacecraft avionics. The mission also carries several instruments and sensors designed to achieve the science mission objectives. International partners contributed several of these sensors.
This paper will describe the InSight mission and science objectives as well as some of the changes made to the mission when the launch date was postponed from 2016 to 2018. We present an update on the Sun Radio Interferometer Space Experiment (SunRISE) mission concept, which was selected in Fall 2017 for Step 2 of the NASA Small Explorer (SMEX) Mission of Opportunity (MoO) program. SunRISE is space-based sparse array, composed of formation flying SmallSats designed to localize the radio emission associated with coronal mass ejections (CMEs) from the Sun. Radio emission from CMEs is a direct tracer of the particle acceleration in the inner heliosphere and potential magnetic connections from the lower solar corona to the larger heliosphere.
Furthermore, CME radio emission is quite strong such that only a relatively small number of antennas is required, and a small mission would make a fundamental advancement. Indeed, the state-of-the-art for tracking CME radio emission is defined by single antennas (Wind/WAVES, Stereo/SWAVES) in which the tracking is accomplished by assuming a frequency-to-density mapping. We present the most recent updates on this mission concept, starting with an overview of the science requirements. It then focuses on the SunRISE concept of operations, which consists of six SmallSats placed in a geostationary graveyard orbit for 6 months to achieve the aforementioned science goals. Each of the mission phases are discussed in detail.
Finally, the paper dives into SunRISE’s science data processing process, including simulation results, demonstrating how the data will produce sufficient to achieve the science goals. • 02.0205 Venus Origins Explorer (VOX): A Proposed New Frontiers Mission Suzanne Smrekar (NASA Jet Propulsion Lab), Michael Lisano (Jet Propulsion Laboratory) Presentation: Suzanne Smrekar - Wednesday, March 7th, 08:55 AM - Madison. Of all known bodies in the galaxy, Venus remains the most Earth-like in terms of size, composition, surface age, and distance from the sun.
Although not currently habitable, Venus lies within the ‘Goldilocks zone’, and may have been habitable before Earth. What caused Venus to follow a divergent path to its present hostile environment, devoid of oceans, magnetic field, and plate tectonics that may have enabled Earth’s long-term habitability? Venus Origins Explorer (VOX) determines how Earth’s twin diverged, and enables breakthroughs in our understanding of terrestrial planet evolution and habitability in our own solar system — and others.
At the time of the Decadal Survey (2011), the full capabilities of near-IR instruments to map global mineralogy from orbit and present-day radar techniques to detect active deformation were not fully appreciated. VOX leverages these methods to answer essential questions, many of which can only be answered with high resolution, global data. New Frontiers science objectives are met using orbital, global reconnaissance and in-situ noble gas measurements: 1.Atmospheric physics/chemistry: noble gases, their isotopes, and light stable isotopes to constrain atmospheric sources, escape processes, and volcanic outgassing; global search for volcanically outgassed water.
2.Past hydrological cycles: global tessera composition to determine the role of volatiles in crustal formation, determines if global, ‘catastrophic’ resurfacing occurred, and assess initial volatile sources and outgassing history. 3.Crustal physics/chemistry: determine global variations in crustal mineralogy/chemistry, tectonic framework and heat flow, whether catastrophic resurfacing occurred, what types of geologic processes are currently active and possible crustal recycling. 4.Crustal weathering: global mineralogy distinguishes between surface-atmosphere weathering reactions by quantifying the redox state and the chemical equilibrium of the near-surface atmosphere.
5.Atmospheric properties/winds: map cloud particle modes and their temporal variations, and track cloud-level winds in the polar vortices. 6.Surface-atmosphere interactions: mineralogy maps distinguish between models for surface-atmosphere chemical interactions; search for new and/or recent volcanism and outgassed water.
VOX consists of: 1) an Atmosphere Sampling Vehicle (ASV), and 2) an Orbiter that performs global reconnaissance using two instruments and a gravity science investigation. The ASV dips into the well-mixed atmosphere to deliver an atmospheric sample to the Venus Original Constituents Experiment (VOCE). Measurement of noble gases reveal the source and evolution of volatiles in the inner solar system. From orbit, Venus Emissivity Mapper (VEM) provides global surface mineralogy and searches for active and/or recent volcanism. Venus Interferometric Synthetic Aperture Radar (VISAR) generates long-awaited high-resolution imaging and digital elevation models, and possible active deformation locations with repeat-pass interferometry to reveal geologic evolution. Ka-band tracking provides the gravity field resolution needed to estimate global elastic thickness. VOX is the logical next mission to Venus because it delivers: 1) top priority atmosphere, surface, and interior science objectives; 2) key global data for comparative planetology; 3) high-resolution global topography, composition, and imaging to optimize future landed missions; 4) opportunities for revolutionary discoveries including active geologic processes, with a 3-year long orbital mission and proven implementation; and 5) 44 Tb of data to fuel next generation scientists.
• 02.0206 Observatory Design for the Imaging X-Ray Polarimeter Explorer (IXPE) Mission William Deininger (Ball Aerospace) Presentation: William Deininger - Wednesday, March 7th, 09:20 AM - Madison. The Imaging X-Ray Polarimeter Explorer (IXPE) Mission is an international collaboration lead by NASA MSFC as the PI institution and includes Ball and CU/LASP, as well as the Italian Space Agency with IAPS/INAF and INFN as major international partners. The goal of IXPE is to expand understanding of high-energy astrophysical processes and sources, in support of NASA’s first science objective in Astrophysics: “Discover how the universe works.” Polarization uniquely probes physical anisotropies—ordered magnetic fields, aspheric matter distributions, or general relativistic (GR) coupling to black-hole spin—that are not otherwise measurable. MSFC provides the x-ray telescopes and SOC along with mission management and systems engineering. Ball is responsible for the spacecraft, payload I&T, spacecraft I&T, system I&T, launch and operations. The MOC is located at CU/LASP. IAPS/INAF and INFN provide the unique polarization-sensitive detectors, detector units and payload computer.
The IXPE Observatory consists of spacecraft and payload modules built up in parallel to form the Observatory during system integration and test. IXPE’s payload is a set of three identical, imaging, X-ray polarimeter systems mounted on a common optical bench and co-aligned with the pointing axis of the spacecraft.
Each system, operating independently, comprises a 4-m-focal length Mirror Module Assembly that focuses X-rays onto a polarization-sensitive imaging detector separated by the deployable boom. Each Detector Unit (DU) contains its own electronics, which communicate with the payload computer that in turn interfaces with the spacecraft.
Each DU has a multi-function filter wheel assembly for in-flight calibration checks and source flux attenuation. The IXPE Observatory is based on the Ball Configurable Platform (BCP)-100 spacecraft architecture. The modular design allows for concurrent payload and spacecraft development with a well-defined, clean interface that reduces technical and schedule risk.
IXPE is leveraging the flexibility of the BCP-100 architecture to accommodate the IXPE science payload. The IXPE payload is mounted on the spacecraft top deck.
The IXPE Observatory is designed to launch on a Pegasus XL or larger launch vehicle. IXPE is designed as a 2-year mission with launch in November 2020. IXPE launches to a cirular LEO orbit at an altitude of 540 km and an inclination of 0 degrees. The payload uses a single science operational mode capturing the x-ray data from the targets. The mission design follows a simple observing paradigm: pointed viewing of known x-ray sources (with known locations in the sky) over multiple orbits (not necessarily consecutive orbits) until the observation is complete. The Observatory communicates with the ASI-contributed Malindi ground station via S-band link.
The IXPE Project completed its Phase A activities in July 2016 with the submission of the Concept Study Report to the NASA Explorers Program Office as a Small Explorer (SMEX) Mission. NASA considered three SMEX mission concepts for flight and selected the IXPE Project as the winner in January 2017. The IXPE Project is working towards SRR in September 2017 and PDR in February 2018. This paper summarizes the IXPE mission science objectives and describes the Observatory implementation concept including the payload and spacecraft elements.
• 02.0208 Analysis of CMBR Using a Nano-satellite Dhananjay Sahoo (Manipal Institute of Technology), Avish Gupta (Manipal University) Presentation: Dhananjay Sahoo - Wednesday, March 7th, 09:45 AM - Madison. We propose a distributed close-range survey of hundreds of asteroids representing many asteroid families, spectral types and sizes. This can be implemented by a fleet of nanospacecraft (e.g., 3 to 6-unit CubeSats) equipped with miniature imaging and spectral instruments (from near ultraviolet to near infrared). To enable the necessary large delta-v, each spacecraft carries a single electric sail tether which taps the momentum from the solar wind.
Data are stored in a flash memory during the mission and downlinked at an Earth flyby. This keeps deep-space network telemetry costs down, despite the large number of individual spacecraft. To navigate without the use of the deep-space network, optical navigation is required to track stars, planets and asteroids. The proposed mission architecture is scalable both scientifically and financially.
A fleet of 50 spacecraft will be able to obtain images and spectral data from 200 to 300 near-Earth objects and main belt asteroids. It allows study of those asteroid families and spectroscopic types for which currently no close-range observations are available. This paper presents science objectives, overall science traceability matrix, example targets and technical challenges associated with the mission. We open for a discussion the preliminary requirements and mission design. • 02.03 Ian Clark (Jet Propulsion Laboratory) & Ian Clark (Jet Propulsion Laboratory) & Ian Clark (Jet Propulsion Laboratory) & Ian Clark (Jet Propulsion Laboratory) & Ian Clark (Jet Propulsion Laboratory) & Ian Clark (Jet Propulsion Laboratory) & Ian Clark (Jet Propulsion Laboratory) & Ian Clark (Jet Propulsion Laboratory) & Ian Clark (Jet Propulsion Laboratory) & Ian Clark (Jet Propulsion Laboratory) & Ian Clark (Jet Propulsion Laboratory) & Ian Clark (Jet Propulsion Laboratory) & Ian Clark (Jet Propulsion Laboratory).
This study presents a novel hazard avoidance guidance method using a dynamic safety margin index, to enhance robustness in the highly uncertain environments of planet and small body landing. As future planetary landing and sample return missions will seek to land in areas with high scientific value which may be located in hazardous terrains, onboard hazard avoidance capability is indispensable. Moreover, the dynamics environment of planet or small body landing is very uncertain due to many sources of perturbations, and the accuracy of lander state estimation in real time is limited. To cope with the impact of state uncertainty on hazard avoidance performance, this study introduces a dynamic safety margin index that changes with the lander state uncertainty, and derives the hazard avoidance guidance law based on evaluation of this index.
The safety margin index takes into account the state uncertainty and control constraints of the lander, and quantitatively describes the safety state of the lander with respect to the hazards around. The index is then used to derive the guidance law that makes the system globally stable and guides the lander to the desired final landing state without collision with any hazard. The impact of the lander state uncertainty on trajectory safety is considered and quantified in safety margin index evaluation and in the development of the guidance law, so the proposed algorithm is adaptive to the lander state uncertainty exhibited in planetary landing practice. The hazard avoidance performance with limited control ability is also improved as the control constraints are considered.
No offline trajectory generation is required, so the method is appropriate for real-time hazard avoidance following online hazard detection. The behavior and performance of the proposed guidance method is investigated via a set of numerical simulations, and the results show that the hazard avoidance performance with state uncertainty and control constraints is improved using the proposed method, advantageous to practical implementation in the uncertain dynamics environment of planetary landing. • 02.0302 A Scanning LIDAR System for Active Hazard Detection and Avoidance during Landing on Europa Eric Schindhelm (Ball Aerospace & Technologies Corporation), Reuben Rohrschneider (Ball Aerospace), Shane Roark (Ball Aerospace and Technologies Corp), Carl Weimer (Ball Aerospace and Technologies Corp) Presentation: Eric Schindhelm - Thursday, March 8th, 08:55 AM - Madison. Active hazard detection and avoidance will be required for landing on Europa due to a lack of a priori knowledge of surface features similar in size to the lander. A light detection and ranging (LIDAR) instrument can provide both long distance (8 km) ranging and close-range (500 m) imaging to enable real-time hazard detection during landing operations. An example space-qualified LIDAR instrument is the Vision Navigation Sensor (VNS) on the Sensor Test for Orion Relative-navigation Risk Mitigation (STORRM) mission in 2011 on STS-134. The VNS consists of a single box housing a laser, transmit and receive optics, focal plane assembly, electronic assemblies, and mechanical components.
The instrument operates in dual mode to change the field of illumination for near or far targets. On STS-134 the VNS successfully acquired imaging range data as the Shuttle docked with ISS. The VNS was subsequently installed on ISS in February, 2017, as part of NASA’s Raven technology demonstration and is operating with more integrated processing algorithms.
Raven’s flash LIDAR observes vehicles as they approach and depart ISS, performing calculations onboard to test autonomous rendezvous capability. We will present potential modifications to the VNS system that could address the unique challenges posed by the Europa environment and landing operations, while reducing payload size, weight and power. One key technology for making the flash LIDAR applicable to Europa landing is the addition of electronically steerable laser projection optics to provide flexibility in operations and to optimize the use of the limited number of photons available. This principle is important in any application where mass and power are limited commodities, and can be performed with either an acousto-optic beam modulator (AOBD) or spatial light modulator (SLM). The addition of Electronically Steerable Flash LIDAR (ESFL) capability to the VNS enables the system to offer the functionality of both scanning and flash LIDARs simultaneously, without any mechanisms. The ESFL also adds in the capability to distribute the LIDAR power into multiple simultaneous beamlets which can further optimize mission efficiency.
A brass board ESFL system with an AOBD was developed and successfully demonstrated on airborne platforms to advance the path to space flight. A bread board ESFL system utilizing an SLM has been demonstrated in the lab, providing another viable option for improving the target acquisition range while still providing the imaging capability at close range. We will present these different systems within the context of the landing Concept of Operations as laid out in the Europa Lander Science Definition Team Report. • 02.0305 Lunar and Mars Ascent and Entry Crew Abort Modes for the Hercules Single-Stage Reusable Vehicle David Komar (NASA - Langley Research Center) Presentation: David Komar - Thursday, March 8th, 09:20 AM - Madison. The Hercules concept is a multi-functional, modular, operationally flexible, single-stage, reusable vehicle designed to operate between low Mars orbit (LMO) and a Mars surface base utilizing oxygen and methane propellants manufactured at the Mars base from Martian resources. Its primary function is cargo and crew transport between LMO and the Mars surface base.
The Hercules vehicle supports an alternative human spaceflight strategy with the goal of affordably establishing a permanent and self-sustaining settlement on Mars in the next half century, as a prelude to colonization, with NASA playing a major role. This strategy, referred to here as Base-First, briefly postpones early human landings on Mars until key technologies and systems are demonstrated and matured, and a significant amount of infrastructure is established on Mars to safely support humans. The Hercules vehicle is designed to maximize crew safety by providing full coverage crew abort capability during both Mars ascent and Mars entry, either through abort-to-surface or abort-to-orbit. This paper will outline each of the abort modes and discuss the Hercules vehicle design along with the base and orbital infrastructure required to enable the full coverage abort capability. For each abort mode, trajectory simulations are flown that illustrate the requisite design capabilities and highlight the sensitivity to key design variables.
Although the Hercules vehicle is designed for Mars operations, this paper also discusses an alternative strategic approach where initial crewed flights are performed in support of a human Lunar campaign that serve to demonstrate and mature the Hercules vehicle nominal and abort operations. For Lunar missions, the Hercules vehicle operates between the Lunar surface and a Deep Space Gateway (DSG) orbiting the moon, with propellant resupplied from Earth to the vehicle at the DSG.
Abort modes for Lunar ascent and descent are discussed highlighting the abort trajectory analysis and the vehicle modifications specific for Lunar operations. • 02.0307 The Conceptual Design of a Novel Simple and Small-sized Mars Lander Ryohei Takahashi (The University of Tokyo), Ryo Sakagami (The University of Tokyo), Akifumi Wachi (University of Tokyo), Shinichi Nakasuka (The University of Tokyo) Presentation: Ryohei Takahashi - Thursday, March 8th, 09:45 AM - Madison.
This paper presents a simulation tool for planetary landing operations near a plume source on the bottom of a tiger stripe canyon on the South Polar Terrain of Saturn's moon Enceladus, reports on its development status, and presents results from the landing simulation. Enceladus is a promising hot spot for astrobiology in the solar system. A concept studied at the Institute for Space Technology and Space Applications (ISTA) of Bundeswehr University Munich in the context of the DLR funded Enceladus Explorer project (EnEx) aimed to place a lander near one of the plume sources on the bottom of a “tiger stripe” canyon on the south pole of Saturn's moon Enceladus. A spacecraft landing near one of the plume sources and deploying a melting probe to sample relatively shallow liquid water in the ice under that plume source would be able to look for signatures of life before they are degraded by exposure to the vacuum of space. The lander would have to meet very challenging landing accuracy and reliability requirements on an exceptionally challenging terrain.
To perform this challenging landing, a sophisticated landing Guidance, Navigation, and Control (GN&C) system would be necessary. For the lander to land within the narrow canyon floor with the required accuracy, a terrain relative navigation (TRN) function can use sensors such as optical and thermal cameras, LIDAR, etc. To navigate relative to detected terrain features.
To ensure a safe landing in the hazardous terrain, a hazard detection and avoidance (HDA) function must be able to assess if the originally planned landing site is safe, and if not to then autonomously command a re-targeting to another safer spot. The guidance function must then calculate a viable trajectory and thrust arc to the newly chosen landing site. To validate that the landing satisfies the accuracy and reliability requirements we are developing a tool in Matlab to simulate the operation of the autonomous landing GN&C system.
In this paper we describe the tool and we present its development status. We then demonstrate that the accuracy and safety requirements are met for landing on an adequately realistic “tiger stripe” canyon-like simulated terrain.
Based on the simulation results we also propose possible modifications to the landing GN&C system. • 02.0309 Overview and Reconstruction of the ASPIRE Project's SR01 and SR02 Supersonic Flight Tests Clara O'farrell (Jet Propulsion Laboratory), Chris Karlgaard (Analytical Mechanics Associates, Inc.), Soumyo Dutta (NASA Langley Research Center), Eric Queen (NASA), Ian Clark (Jet Propulsion Laboratory) Presentation: Clara O'farrell - Thursday, March 8th, 10:35 AM - Madison. The Advanced Supersonic Parachute Inflation Research and Experiments (ASPIRE) project is a risk-reduction activity for NASA's upcoming Mars2020 mission which is investigating the supersonic deployment, inflation, and aerodynamics of two candidate Disk-Gap-Band (DGB) designs. The two parachutes under investigation are a built-to-print version of the DGB used by the Mars Science Laboratory and a full-scale strengthened version of this parachute that has the same geometry but differs in materials and construction.
Starting in the fall of 2017, the parachutes will be tested at deployment conditions representative of flight at Mars by sounding rockets launched out of NASA's Wallops Flight Facility (WFF). Game Sound Museum Famicom Edition Flaco. The first flight test (SR01) of the built-to-print parachute will take place in September of 2017, followed by the first test of the strengthened parachute during flight SR02 in early November of 2017. During both tests, sounding rockets will deliver a payload containing the parachute pack, the deployment mortar, and the ASPIRE instrumentation suite up to altitudes of approximately 45 km at supersonic speeds. The instrumentation suite includes a GLN-MAC IMU, a GPS unit, a C-band transponder for radar tracking, three load pins at the parachute triple bridles, and three high-speed/high-resolution cameras trained on the canopy during inflation.
In addition, the atmospheric conditions at the time of flight will be characterized by means of meteorological balloons carrying radiosondes and meteorological sounding rockets carrying inflatable ROBIN spheres. These data will allow the reconstruction of the test conditions, parachute loads, and parachute aerodynamic performance in flight. In addition, the imagery from the three on-board cameras will allow the reconstruction of the three-dimensional geometry of the canopy during inflation. This paper will describe the first two sounding rocket flight tests, SR01 and SR02.
It will provide an overview of flight operations, the data acquired during testing, the techniques used for post-flight reconstruction, and the reconstructed performance of the test vehicle and parachute system for each flight. • 02.0310 Concept of Autonomous GNC Aided Operation System for Future China Manned Lunar Pinpoint Landing Xiuqiang Jiang (Nanjing University of Aeronautics and Astronautics), Shuang Li (Nanjing University of Aeronautics and Astronautics) Presentation: Xiuqiang Jiang - Thursday, March 8th, 11:00 AM - Madison. Manual control is widely used in manned space missions including Apollo manned lunar landing. With the advance of technology, China indeed has some advantages although its progress lags behind the United States and the Soviet Union in this field. On the one hand, China’s manned spaceflight rendezvous and docking mission successfully verified the TV-guidance based manual rendezvous and docking technologies.
On the other hand, China’s unmanned lunar soft landing mission has successfully verified the autonomous Guidance Navigation and Control (GNC) technologies for powered descent, hazard avoidance and safe landing. In order to achieve safely pinpoint landing close to the scheduled lunar facilities, it is necessary to develop an autonomous GNC aided operation system for future manned lunar pinpoint landing.
This system makes the autonomous GNC mode and the manual mode mutually cooperated and backuped, which is conductive to enhance operation efficiency, accuracy, reliability and safety. This paper will report a conceptual design of the autonomous GNC aided operation system for future manned lunar pinpoint landing based on China’s current technological accumulations. First, the integrated navigation system solves and outputs the real-time absolute/relative flight state parameters of the lunar module, and the outputs of each sensors and navigation modes are also displayed. Second, the terrain recognition system real-time outputs the three-dimensional terrain and television images of local lunar surface, the corresponding hazardous/safe zone distribution map and fuel consumption map for hazard avoidance, the optimal landing site and candidate landing sites.
Third, the guidance and control system real-time outputs the difference between reference trajectory and real trajectory, the deviation between predicted and theoretical landing sites, corresponding hazard avoidance route to reach selected landing sites, state of thrusters and remaining fuel. Then, a human-machine interface is utilized for astronauts to check, confirm or revise the GNC modes according to the above maps, images, parameters, flight trajectory, and the remaining fuel. The guidance mode and destination can be modified at any time throughout the descent and landing process, and the corresponding guidance law will be automatically implemented.
The astronauts can modify the selected landing site or directly manually point out the desired landing position on the television image, and the system performs the solution and guidance control accordingly. Additionally, astronauts can also switch the operation mode to completely manual control mode at any time according to the actual situation. • 02.0311 Overview of the Mars 2020 Parachute Risk Reduction Plan Christopher Tanner (Jet Propulsion Laboratory), Ian Clark (Jet Propulsion Laboratory) Presentation: Christopher Tanner - Thursday, March 8th, 11:25 AM - Madison. In 2012, the Mars Science Laboratory (MSL) landed safely on the surface of Mars using a supersonic Disk-Gap-Band (DBG) parachute, which was qualified for flight via a subsonic wind tunnel test program. Results of the Low-Density Supersonic Decelerators (LDSD) program have called into question the methodology and principles that form the foundation of the MSL subsonic test program. LDSD discovered that quasi-static subsonic proof loading a parachute via ground testing may not provide canopy stresses that sufficiently bound the stresses experienced during a rapid supersonic inflation at Mars. Additionally, the single supersonic flight of the MSL parachute cannot be solely relied upon to demonstrate sufficiency as the peak parachute inflation load at Mars for the Mars 2020 mission cannot be guaranteed to be less than or equal to that of MSL.
To place the Mars 2020 parachute system into a position of acceptable risk, a series of risk reduction steps were initiated starting in 2015. First, two parachute assemblies have been pursued in parallel: a Build-to-Print (BTP) MSL parachute, designed and manufactured by Pioneer Aerospace Corporation, which maintains the heritage of the successful MSL parachute, and a strengthened parachute, designed and manufactured by Airborne Systems North America, which uses higher strength materials throughout the parachute assembly but maintains the same overall size and general configuration as the MSL parachute to minimize changes to the flight vehicle and to maintain aerodynamic similarity. Second, the workmanship of each parachute system was tested using a subsonic wind tunnel test program nearly identical to that of MSL. Finally, full-scale parachutes from each vendor will experience at least one supersonic inflation at Mars-relevant Mach numbers and atmospheric densities at Earth via a supersonic sounding rocket test campaign. This paper presents high-level details regarding the risk reduction strategy, the two candidate parachute configurations, the ground test program, and the supersonic flight test program. • 02.0316 Sustaining PICA TPS for Future NASA Robotic Science Missions Including NF-4 and Discovery Mairead Stackpoole (NASA Ames Research Center), Ethiraj Venkatapathy (NASA ARC) Presentation: Mairead Stackpoole - Thursday, March 8th, 11:50 AM - Madison. Phenolic Impregnated Carbon Ablator (PICA), invented in the mid 1990’s, is a low-density ablative thermal protection material proven capable of meeting sample return mission needs from the moon, asteroids, comets and other “unrestricted class V destinations” as well as for Mars.
Its low density and efficient performance characteristics have proven effective for use from Discovery to Flagship class missions. It is important that NASA maintain this TPS material capability and ensure its availability for future NASA use.
The rayon based carbon precursor raw material used in PICA preform manufacturing required replacement and requalification at least twice in the past 25 years and a third substitution is now needed. The carbon precursor replacement challenge is twofold – the first involves finding a long-term replacement for the current rayon and the second is to assess its future availability periodically to ensure it is sustainable and be alerted if additional replacement efforts need to be initiated. Rayon is no longer a viable process in the US and Europe due to environmental concerns. In the early 80’s rayon producers began investigating a new method of producing a cellulosic fiber through a more environmentally responsible process. This cellulosic fiber, lyocell, is a viable replacement precursor for PICA fiberform. This presentation reviews current SMD-PSD funded PICA sustainability activities in ensuring a rayon replacement for the long term is identified and in establishing that the capability of the new PICA derived from an alternative precursor is in family with previous versions of the so called “heritage” PICA.
• 02.0317 A Momentum-Based Method for Predicting the Peak Opening Load of Supersonic Parachutes David Way (NASA - Langley Research Center) Presentation: David Way - Thursday, March 8th, 04:30 PM - Madison. In this paper, a new empirical indicator for predicting the peak opening loads of supersonic parachutes is presented. The proposed indicator is proportional to twice the free-stream dynamic pressure and the projected area of the parachute, which is equivalent to estimating the opening load as a percentage of the free-stream momentum flux through the projected area at the moment of peak inflation. The form of this expression is motivated by a classical control volume analysis of the aerodynamic forces acting on a parachute during inflation, under the simplifying assumptions of quasi-static and one-dimensional flow.
For parachute geometries and flight conditions typical of Mars Entry, Descent, and Landing systems, the largest contribution to the total drag is shown to be a momentum flux term that is associated with the entrainment of atmosphere within the inflating parachute volume. Using this new method, empirical constants are calculated from existing flight reconstruction data and are shown to have a smaller standard deviation than similar constants determined using the customary indicator form, which is based on the steady-state subsonic drag and proportional to the parachute reference area. These empirical constants are also compared to an analytic estimate, derived from the control volume analysis, and shown to have excellent agreement across a wide range of Mach numbers and dynamic pressures for several parachute geometries. While opening loads estimated using both methods produce similar results at low supersonic Mach numbers typical of past inflations, the proposed method predicts notably larger loads at higher Mach numbers, those above Mach 2.0, due to the omission of any Mach Efficiency Factor.
Several current Mars EDL projects have adopted this new indicator. • 02.0318 An Analytical Mars Entry Guidance Method for Higher Deployment Altitude considering Uncertainties Zeduan Zhao (), Pingyuan Cui () Presentation: Zeduan Zhao - Thursday, March 8th, 04:55 PM - Madison.
The atmospheric entry phase is challenging and vital for the whole Mars entry, descent and landing (EDL) process. In the present EDL scheme for Mars landing missions, the altitude and accuracy of parachute deployment have greatly decided the final landing accuracy and safety. The Entry Terminal Point Controller (ETPC) adapted from the Apollo guidance law has helped the Mars Science Laboratory mission to land the Curiosity rover on Mars surface successfully with greatly improved landing accuracy and altitude compared with the former landing missions. ETPC is an analytical guidance law with prediction and correction capabilities. Based on the measured acceleration and altitude rate, ETPC predicts the terminal range error and gives the commanded bank angle. In this research, the authors try to develop an analytical entry guidance method based on ETPC, by considering improving the deployment altitude and eliminating the tracking error caused by the parameter uncertainties. According to the controllability of the entry vehicle, the original range control phase is further divided into two smaller phases: range control phase and altitude/range control phase.
The trigger for the altitude/range control phase is carefully designed to make a compromise between the final range accuracy and the altitude performance. The ETPC is improved in terms of two aspects. Firstly, the state is extended to consider the parameter uncertainties in the ballistic coefficient and atmospheric density models, which will cause terminal range and altitude errors. By taking uncertain parameters as augmented state variables, the final range or altitude error caused by the effect of the parameter deviation can be calculated analytically. Secondly, the influence coefficients in the altitude/range phase are redesigned according to the new guidance objective. Different from MSL’s ETPC where the guidance objective is only the terminal range, the new objective adopts the formulation of a weighted sum of the terminal range and deployment altitude.
The weights can be tested and designed according to the mission requirements. The energy is taken as the independent variable for its mono-increasing/mono-decreasing property. To evaluate the performance of the proposed guidance method, Monte Carlo tests are designed and carried out in the MATLAB environment. The results indicate that when the altitude control is included in the improved ETPC, the terminal 3σ altitude distribution will decrease and the average deployment altitude increases.
However, as the entry vehicle has to use its limited control ability to eliminate the altitude and range tracking errors simultaneously, the 3σ range distribution increases compared to the original ETPC. By modulating the trigger for the altitude/range control phase, as well as the weights for the final range and altitude in the performance index, mission designers will get compromised altitude and range performance, which might increase the possibility for the final successful landing. • 02.0319 Divert Capability Analysis and Corresponding Guidance Design for Mars Landing Tong Qin (Beijing Institute of Technology) Presentation: Tong Qin - Thursday, March 8th, 05:20 PM - Madison. Defined as the largest horizontal distance that the lander can cover during the powered descent phase, the divert capability determines whether the lander can reach the distant landing target safely from the parachute jettison point.This paper furthers the research on divert capability in three aspects.
First, previously unrevealed engineering constraints on divert capability are revealed, including initial powered descent conditions, path constraint, thrust bounds, and residual fuel amount. The initial states are closely related to the divert capability since the lander can divert further along the initial horizontal velocity and shorter against it. In powered descent phase, the larger velocity component is along the vertical direction. Thus, as an average for this sensitivity analysis, the lander is assumed to be moving vertically at parachute jettison, and only the effect of initial altitude and vertical velocity on DC is analyzed. The path constraint requires the descent trajectory constructed by the guidance law stays above Mars surface.
The engine thrust, representing the maneuverability, varies from a minimum value greater than zeros to its ultimate value corresponding to the physical process in which the engine cannot be shut down once started up. The residual fuel is expressed by the propellant mass fraction which is the ratio of the fuel mass to the total mass. Next, considering all these four factors, the methods to obtain the divert capability analysis using convex programming guidance(CPG)law, first order guidance (FPG)law, second order guidance (SPG)law, are developed. Using CPG, the former three constraints are considered in the optimization problem.
Therefore, if a solution is found for this optimization problem, it will satisfy all constraints.The largest horizontal distance allowed by the residual fuel is taken as the divert capability. For the FPG and SPG, the acceleration profile is only guaranteed to satisfy the terminal boundary condition.
Even if the fuel cost is within the total amount, it should be checked whether other constraints are violated. Thus an iterative searching algorithm is developed to obtain the divert capability while satisfying all constraints. Then, considering the performance trade between divert capability and fuel consumption, the flight time selection in polynomial guidance laws is analyzed.
Two methods for selecting flight time are proposed. One is to obtain the maximum divert capability with a limited residual fuel amount, and anther is to minimize the fuel consumption while satisfying the divert capability requirement. • 02.0320 Space Mission Hibernation Mode Design: Lessons Learned from Rosetta and Other Missions Arthur Chmielewski (Jet Propulsion Laboratory), John West () Presentation: Arthur Chmielewski - Thursday, March 8th, 09:00 PM - Madison.
Secondary spacecraft are being put in orbit at the rate of hundreds per year, with each year smashing the record set by the previous. These 'piggyback' missions (which reach orbit using the unused mass/volume in a launch vehicle) come from every source: universities, government labs, large contractors, startups and even art collectives (?!?!). There are Beepsats, journal-worthy science missions, operational sigint missions, radio Amateur missions, technology demonstrations and venture-capital-backed remote sensing missions.
For today, secondary spacecraft are almost exclusively low-Earth-orbiting missions, though even that limitation is changing. Despite the breadth of mission sponsors, mission developers, mission types and orbital domains, our terminology is unchanged and contradictory: cubesat (or CubeSat), picosat, smallsat. More importantly, decisionmakers hear a phrase like 'half of all CubeSats fail' and assume that there's something inherently flawed in the CubeSat concept -- instead of recognizing that half of 'all' CubeSats fail because hobbyists make up more than half of all CubeSats, and their ad hoc practices fail two-thirds of the time. In this paper, we propose two taxonomies for secondary spacecraft: the first, based on the type of mission developer (e.g., Hobbyist, Crafter, Industrialist) and the second based on the type of mission flown (e.g., constellation, experiment, short operational, long operational and extreme operational). We will contrast these taxonomies against the existing ones (such as NASA's Class A/B/C/etc). Finally, we will use these new taxonomies to parse our existing secondary spacecraft database to provide meaningful examinations of mission types, mission success rates and opportunities for further study.
• 02.0403 Modular Inflatable Space Structures Jekan Thangavelautham (University of Arizona), Aman Chandra (Arizona State University) Presentation: Jekan Thangavelautham - Wednesday, March 7th, 09:25 PM - Lake/Canyon. There is a growing need to develop a human focused exploration program and support infrastructure, including relay sites in deep space. One of the first targets will be cis-lunar space, which is a strategic gateway towards permanent settlement of the Moon and Mars. Developing a station in deep space will require space structures to have large surfaces, with very high volume to mass ratio and high-packing efficiency. The prohibitive costs of transporting bulky, heavy payloads to beyond low-earth orbits pose a formidable challenge. This requires a paradigm shift in research towards developing methods to build and assemble increasingly low-mass, large and complex structures in space instead of transporting them from Earth and deploying them on-site.
On-site additive construction and assembly methods hold promise, but they face a major challenge of still having to transport large and heavy robotic equipment required to perform complex construction. This paper presents an alternate and more feasible pathway in developing small structural units that can be quickly shaped and assembled with limited external support. Inflatable structures hold that promise as they are low-mass, can be quickly reshaped, inflated and rigidized into desired modular units that are assembled into large, complex structures. Our present work extends the inflatables concept to study modular, inflatable building blocks that can be assembled into pre-determined geometries. The inflatable building blocks would be assembled into communication relays, science instrument antennas, structures to hold solar panels and large reflectors. Our efforts aim to identify common desired structural design traits in these modular units to enable them to be multi-functional building blocks that can be assembled into more complex functional blocks.
Our design methodology focuses on simplicity of deployment mechanisms and high-scalability over varying sizes. Finally, this paper will provide preliminary feasibility of the modular inflatable building block concept and analyze the applications of this technology towards assembly of large structures in deep space.
• 02.05 Richard Volpe (Jet Propulsion Laboratory) & Richard Volpe (Jet Propulsion Laboratory) & Richard Volpe (Jet Propulsion Laboratory) & Richard Volpe (Jet Propulsion Laboratory) & Richard Volpe (Jet Propulsion Laboratory) & Richard Volpe (Jet Propulsion Laboratory) & Richard Volpe (Jet Propulsion Laboratory) & Richard Volpe (Jet Propulsion Laboratory) & Richard Volpe (Jet Propulsion Laboratory) & Richard Volpe (Jet Propulsion Laboratory) & Richard Volpe (Jet Propulsion Laboratory) & Richard Volpe (Jet Propulsion Laboratory) & Richard Volpe (Jet Propulsion Laboratory). This paper presents the validation of the end-to-end BiBlade comet surface sampling chain, including Touch-and-Go (TAG) activities.
The sampling chain consists of the BiBlade sampler, a structural compliant interface to a would be spacecraft, a 3 degree of freedom planar robotic arm, a sample volume measurement device, and two sample vaults. The end-to-end sample chain is integrated into a purpose built full-scale dynamic spacecraft emulator, enabling to be performed the full sequence from surface approach to collected sample storage. Touchdown, sample triggering, sample acquisition, and ascent from surface are a critical set of sequences dependent on spacecraft dynamics and the sampler interaction with a comet surface. Of special importance are the rotational and linear inertia's provided by the 2200kg spacecraft emulator during TAG operations. The full-scale dynamic spacecraft emulator is intended to validate these complex interactions between the sampler, comet surface simulants, and spacecraft mass. The full-scale dynamic Touch-and-Go spacecraft emulator consists of a 2200kg mass on a frictionless, planar air bearing system, fourteen horizontal air thrusters providing 3 degree of freedom motion, onboard computing for motion control, and an off-board motion capture system providing localization.
This robotic test-bed can approach a simulated comet surface in a controlled fashion to validate the BiBlade TAG performance in nominal and off-nominal approach velocity scenarios. Following TAG, sample transfer, volume verification, and storage is performed. • 02.0502 Development and Validation of a Fiberscope Sample Imaging System for In-situ Sample Assessment Risaku Toda (Jet Propulsion Laboratory, California Institute of Technology), Scott Moreland (Jet Propulsion Laboratory), Mircea Badescu (Jet Propulsion Laboratory), Paul Backes (Jet Propulsion Laboratory), Vladimir Arutyunov (Jet Propulsion Laboratory), Jacob Tims (JPL), Valerie Scott (), Harish Manohara (Jet Propulsion Laboratory) Presentation: Risaku Toda - Thursday, March 8th, 05:20 PM - Jefferson. In sample return missions, in-situ sample assessment tool could play critical role to significantly improve the outcome of primary mission objectives. NASA's Apollo lunar landing missions (Apollo 11 through 17) involved human astronauts available on-site to visually confirm the lunar sample collection. However, most robotic sample return missions that have been launched since (Russian Luna missions, NASA’s Genesis, Stardust and JAXA’s Hayabusa missions) have lacked in-situ sample verification mechanisms. NASA’s OSIRIS-REx mission, launched in 2016, incorporates a sample mass measurement mechanism based on the spacecraft inertia change before and after the asteroid sample collection.
We are developing a in-situ sample assessment tool called Fiberscope Sample Imaging (FiSI) system for a notional comet surface sample return mission. This tool would image acquired sample just after the comet sample acquisition and the sample volume would be analyzed from the images. If sample quantity is deemed insufficient, the sample capture maneuver would be re-attempted until the baseline sample quantity is positively confirmed. During the IEEE Aerospace 2016 conference, we presented the FiSI system based on nine fiberscope imagers. Each fiberscope contains 30,000 picture elements and has 55 degrees field-of-view.
The nine fiberscopes capture nine images of acquired comet sample from nine viewpoints through the narrow opening of a sampler head. The wide-swath coverage of nine images are analyzed to estimate acquired comet sample quantity. This IEEE Aerospace 2018 paper will describe further development and validation of this FiSI system. The illumination scheme is modified to enhance imaging robustness for dusty sample objects. The illumination is provided by ten bundles of illumination fibers that are separated from the imaging fiberscopes. Two LED illumination source provide switchable illumination that can be adjusted for different sample shapes and brightness.