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NASA NTRS2019

NISAR flight feed assembly measurement campaign

NISAR (NASA ISRO SAR, National Aeronautics and Space Administration, Indian Space Research Organization, Synthetic Aperture Radar) is an Earth science project currently in its final development phase at NASA Jet Propulsion Laboratory (JPL) and at ISRO. Due for launch in 2022, it will assess how our planet changes overtime by measuring differences in the Earth’s solid surface due to factors like climate change, movement and melting of glaciers, earthquakes, land-slides, deforestation, agriculture and others. The enabling instrument for this mission is a dual band radar (L-Band and S-Band) that feeds a 12m deployable mesh reflector. This paper describes the antenna measurement campaign of the L-Band feed array developed at JPL. The SBand feed array is developed by ISRO and is not part of this paper. While Measurements of the Engineering Model have been published before, this paper focuses on the first measurements done with the Flight Model.

Harrell, Jefferson A.↗
NASA NTRS2020

NISAR flight feed passive antenna measurements

NISAR (NASA ISRO SAR, National Aeronautics and Space Administration, Indian Space Research Organization, Synthetic Aperture Radar) is an Earth science project currently in its final development phase at NASA Jet Propulsion Laboratory (JPL) and at ISRO. Due for launch in 2022 it will assess how our planet changes over time by measuring differences in the Earth’s solid surface due to factors like climate change, movement and melting of glaciers, earthquakes, land-slides, deforestation, agriculture and others. The enabling instrument for this mission is a dual band radar (L-band and S-band) that feeds a 12m deployable mesh reflector. This paper describes the measurement campaign of the L-band flight feed in its passive configuration. Further measurements will be done using the antenna with the active radar components but they are not part of this paper.

Harrell, Jefferson A.↗
NASA NTRS2021

The Evolution of Team-X: 25 Years of Concurrent Engineering Design Experience

Established in 1995 in response to NASA’s “Faster, Better, Cheaper” era, Team-X was born from a need to perform rapid space mission design for principal investigator-led competed proposals. The success and sustainability of Team-X over the 25 years that have followed is directly attributable to the Team-X business model and its evolution over time. While dozens of organizations and institutions have emulated the Team-X design process, there are nuances to the Team-X business model that are unique to JPL, and explain why it is different than other concurrent design teams. One of the key components of a business model is the customer segments that are served. Team-X was founded to conduct the Pre-Phase A work necessary to formulate a portfolio of multiple planetary mission concepts, but has since expanded to include the capability to conduct studies for Earth science, astrophysics, and heliophysics missions as well as Human Exploration and Operations missions and space technology development. Team-X delivers value to its clients both in terms of speed and cost. Team-X has also added value by creating teams to enable the development of Instrument and SmallSat Concepts. Value has further been enhanced through a revision of its process for reviews and the addition of pre-design architecting capabilities. Other aspects of the Team-X infrastructure, in addition to study process, have enabled it to succeed for over a quarter century. From our most important resource, the people, our tools, especially for cost estimating, as well as our increasing capable IT infrastructure have contributed to our capability to meet the demands of our clients. The Team-X business model and its evolution over time, position it well for success in the decades to come.

Murphy, Jonathan↗
NASA NTRS2022

NISAR Flight Feed Passive Antenna Measurements

NISAR (NASA ISRO SAR, National Aeronautics and Space Administration, Indian Space Research Organization, Synthetic Aperture Radar) is an Earth science project currently in its final development phase at NASA Jet Propulsion Laboratory (JPL) and at ISRO. Due for launch in 2022 it will assess how our planet changes over time by measuring differences in the Earth’s solid surface due to factors like climate change, movement and melting of glaciers, earthquakes, land-slides, deforestation, agriculture and others. The enabling instrument for this mission is a dual band radar (L-band and S-band) that feeds a 12m deployable mesh reflector. This paper describes the measurement campaign of the L-band flight feed in its passive configuration. Further measurements will be done using the antenna with the active radar components but they are not part of this paper.

Harrell, Jefferson A.↗
NASA NTRS2022

Evolution of the NISAR Feed Antenna Design due to Environmental Requirements

NISAR (NASA ISRO SAR, National Aeronautics and Space Administration, Indian Space Research Organization, Synthetic Aperture Radar) is an Earth science project currently in its final development phases at NASA Jet Propulsion Laboratory (JPL) and at ISRO. Due for launch in 2023 it will assess how our planet changes over time by measuring differences in the Earth’s solid surface due to factors like climate change, movement and melting of glaciers, earthquakes, land-slides, deforestation, agriculture and others. The enabling instrument for this mission is a dual band radar (L-Band and S-Band) that feeds a 12m deployable mesh reflector. Given the extreme accuracy needed to perform these measurements from Low Earth Orbit (LEO) some of the most stringent requirements were pushed onto the antenna subsystem to control phase and gain stability over time. In this paper we describe how environmental requirements affected the antenna design and what techniques were used to meet these requirements.

Vacchione, Joseph D.↗
NASA NTRS2017

From Prototype to Flight: Qualifying a Ka-Band Parabolic Deployable Antenna (KaPDA) for CubeSats

CubeSats have experienced a number of exciting technological advancements in the past several years. However, until recently, there has been very limited development in the area of high gain CubeSat antennas, which are critical for both high data rate communications and radar science. A Ka-band high gain antenna would provide a 10,000 times increase in data communication rates over an X-band patch antenna and a 100 times increase over state-of-the-art S-band parabolic antennas. Because of this, three years ago the Jet Propulsion Laboratory (JPL) initiated a research and technology development effort to advance CubeSat communication capabilities, with one of the key thrusts being the Ka-band parabolic deployable antenna (KaPDA). This antenna started with the ambitious goal of fitting a 42 dB, 0.5 meter, 35 Ghz antenna in a 1.5U canister. This paper discusses the process of taking the antenna from a first prototype to the flight design, how the design successfully met its goals, and lessons learned. A prototype antenna was constructed in early 2015, and then upgraded to an engineering model at the end of 2016. KaPDA will be flying on the RainCube mission, and earth science CubeSat. KaPDA is the second deployable parabolic antenna to fly on a CubeSat, and the first of its kind to operate at Ka-band enabling a number of opportunities for high rate deep space antenna communications and radar science.

Sauder, Jonathan↗
NASA NTRS2015

Multiplying Mars Lander Opportunities with Marsdrop Microlanders

From canyons to glaciers, from geology to astrobiology, the amount of exciting surface science awaiting us at Mars greatly outstrips available mission opportunities. Based on the thrice -flown Aerospace Corporation Earth Reentry Breakup Recorder (REBR), we present a method for accurate landing of small instrument payloads on Mars, utilizing excess cruise -stage mass on larger missions. One to a few such microlanders might add 1-5% to the cost of a primary mission with inconsequential risk. Using the REBR and JPL Deep Space 2 starting points for a passively stable entry vehicle provides a low mass and low ballistic coefficient, enabling subsonic d employment of a steerable parawing glider, capable of 10+ km of guided flight at a 3:1 glide ratio. Originally developed for the Gemini human space program, the parawing is attractive for a volume -limited microprobe, minimizing descent velocity, and providing sufficient remaining volume for a useful scientific payload. The ability to steer the parawing during descent opens unique opportunities, including terrain- relative navigation for landing within tens of meters of one of several specified targets within a given uncertainty ellipse. In addition to scientific value, some Mars human exploration Strategic Knowledge Gaps could be addressed with deployment of focused instruments at multiple locations.

Staehle, Robert L.↗
NASA NTRS2019

Requirements Development and Management on the Psyche Project

In January 2017, Psyche was one of two mission concepts selected by NASA for flight as part of the 14th Discovery mission competition. The project has been staffing up and maturing the spacecraft, instrument and mission system baseline designs on the path towards a 2022 launch. During much of 2018, the Project has been executing the lifecycle stage called Phase B, “Preliminary Design and Technology Completion,” one key element of which is the development and management of requirements at various levels. In the case of the Psyche project, this process has been particularly unique for several reasons. The project utilizes a Solar Electric Propulsion (SEP) Chassis from Space Systems Loral (SSL), a high volume manufacturer of commercial geostationary (GEO) telecom spacecraft based on the 1300 satellite bus. While SSL has an extensive, well-vetted set of requirements based on their very successful Earth-orbiting product line, translating that heritage to a deep space science mission required special care. In addition to the differences associated with the deep space environment and longer communication times, new interfaces had to be incorporated. While a substantial portion of the Flight System consists of the SEP Chassis, there were several new interfaces within various subsystems between SSL components and those provided by JPL and other contractors. Managing these interfaces through requirements at a relatively higher level than normally seen on internal or external builds proved challenging. Finally, the Psyche spacecraft plans to host the flight terminal of the Deep Space Optical Communications (DSOC) technology demonstration, which is itself a separate project with its own requirements that must be flowed down and managed. This paper will present an overview of the requirement development and management process for the Psyche project. It will discuss in detail the various challenges summarized above, the methods and decisions chosen to address them, and evaluate their overall effectiveness at this stage in the project.

Elkins-Tanton, Linda T.↗
NASA NTRS2005

An Overview of Wind-Driven Rovers for Planetary Exploration

The use of in-situ propulsion is considered enabling technology for long duration planetary surface missions. Most studies have focused on stored energy from chemicals extracted from the soil or the use of soil chemicals to produce photovoltaic arrays. An older form of in-situ propulsion is the use of wind power. Recent studies have shown potential for wind driven craft for exploration of Mars, Titan and Venus. The power of the wind, used for centuries to power wind mills and sailing ships, is now being applied to modern land craft. Efforts are now underway to use the wind to push exploration vehicles on other planets and moons in extended survey missions. Tumbleweed rovers are emerging as a new type of wind-driven science platform concept. Recent investigations by the National Aeronautics and Space Administration (NASA) and Jet Propulsion Laboratory (JPL) indicate that these light-weight, mostly spherical or quasi-spherical devices have potential for long distance surface exploration missions. As a power boat has unique capabilities, but relies on stored energy (fuel) to move the vessel, the Tumbleweed, like the sailing ships of the early explorers on earth, uses an unlimited resource the wind to move around the surface of Mars. This has the potential to reduce the major mass drivers of robotic rovers as well as the power generation and storage systems. Jacques Blamont of JPL and the University of Paris conceived the first documented Mars wind-blown ball in 1977, shortly after the Viking landers discovered that Mars has a thin CO2 atmosphere with relatively strong winds. In 1995, Jack Jones, et al, of JPL conceived of a large wind-blown inflated ball for Mars that could also be driven and steered by means of a motorized mass hanging beneath the rolling axis of the ball. A team at NASA Langley Research Center started a biomimetic Tumbleweed design study in 1998. Wind tunnel and CFD analysis were applied to a variety of concepts to optimize the aerodynamic characteristics of the Tumbleweed Rovers. Bare structures, structures carrying sails and a tumbleweed plant (of the Salsola genus) were tested in Langley's wind tunnels. Thomas Estier of the Swiss Federal Institute of Technology developed a memory metal collapsible structure, the Windball. Numerous other researchers have also suggested spherical rovers.

Hajos, Gregory A.↗
NASA NTRS2003

DSMS investment in support of satellite constellations and formation flying

Over the years, NASA has supported unmanned space missions, beyond earth orbit, through a Deep Space Mission System (DSMS) that is developed and operated by the Jet Propulsion Laboratory (JPL) and subcontractors. The DSMS capabilities have been incrementally upgraded since its establishment in the late '50s and are delivered primarily through three Deep Space Communications Complexes (DSCC 's) near Goldstone, California, Madrid, Spain, and Canberra, Australia and from facilities at JPL. Traditionally, mission support (tracking, command, telemetry, etc) is assigned on an individual-mission basis, between each mission and a ground-based asset, independent of other missions. As NASA, and its international partners, move toward flying fullconstellations and precision formations, the DSMS is developing plans and technologies to provide the requisite support. The key activities under way are: (1) integrated communications architecture for Mars exploration, including relays on science orbiters and dedicated relay satellites to provide continuous coverage for orbiters, landers and rovers. JPL is developing an architecture, as well as protocols and equipment, required for the cost-effective operations of such an infrastructure. (2) Internet-type protocols that will allow for efficient operations across the deep-space distances, accounting for and accommodating the long round-trip-light-time. JPL is working with the CCSDS to convert these protocols to an international standard and will deploy such protocol, the CCSDS File Delivery Protocol (CFDP), on the Mars Reconnaissance Orbiter (MRO) and on the Deep Impact (01) missions. (3) Techniques to perform cross-navigation between spacecrafi that fly in a loose formation. Typical cases are cross-navigation between missions that approach Mars and missionsthat are at Mars, or the determination of a baseline for missions that fly in an earth-lead- lag configuration. (4) Techniques and devices that allow the precise metrology and controllability of tightformations for precision constellation missions. In this paper we discuss the four classes of constellatiodformation support with emphasis of DSMS current status (technology and implementation) and plans in the first three areas.

space operations↗
NASA NTRS2000

Nozomi Cis-Lunar Phase Orbit Determination

Japan's Institute of Space and Astronautical Science (ISAS) launched Nozomi, its first mission to the planet Mars using the newly developed M-V launch vehicle on July 3, 1998. Scientific objectives of the mission are to study the structure and dynamics of the Martian upper atmosphere and its interaction with the solar wind. Nozomi is a cooperative mission between ISAS and the National Aeronautics and Space Administration (NASA). The NASA contribution includes navigation and tracking services provided by the Jet Propulsion Laboratory (JPL). The spacecraft also serves as an engineering demonstration of basic technology for planetary exploration. One of the new technologies was a unique trajectory, developed by ISAS, which used solar gravitational perturbations at the weak stability boundary as an aid to achieve an Earth-Mars transfer orbit. This trajectory saves approximately 120 m/s of Delta V compared to direct hyperbolic insertion and is considered an enabling technology for the mission. Nozomi was the first spacecraft to employ this trajectory and provided on-orbit validation of the technique. The trajectory was achieved by initially placing the spacecraft in a highly elliptical cis-lunar phasing orbit. Six maneuvers were performed during this period to correct injection errors and target an outbound lunar swingby in September 1998. The gravity assist from the lunar swingby raised apogee to the vicinity of the weak stability boundary. After three more targeting maneuvers, Nozomi performed an inbound lunar swingby followed immediately by a powered Earth swingby in late December 1998. A 420 m/s Trans Mars Insertion (TMI) burn at the final Earth periapsis was intended to place the spacecraft on a heliocentric trajectory leading to Mars orbit insertion in October 1999. Orbit determination for Nozomi is performed in parallel by both ISAS and the Multi-Mission Navigation (MMNAV) group at JPL. This was an advantage for the mission because each group would generate solutions based on data collected from their respective tracking networks. Spacecraft events, such as sequence uplinks and maneuvers, were generally scheduled during passes at the Usuda tracking station in Japan. As a result, maneuver design and reconstruction was derived from MMNAV solutions based on JPL tracking data obtained immediately prior to or following maneuvers. Data was also exchanged between ISAS and MMNAV so orbit determination could be performed on joint data sets in support of critical targeting late in the cis-lunar phase. In this paper, information regarding the MMNAV orbit determination effort for the first six months of the mission is presented. The spacecraft trajectory is characterized first, followed by a discussion of the orbit determination estimation procedure and models. Results from selected orbit solutions are presented and compared against reconstructed trajectories. One area of emphasis in this paper is orbit determination in the vicinity of the weak stability boundary. Precise navigation was necessary to target the second lunar swingby and the powered Earth swingby. Delivery accuracy of 150 m was required for these critical encounters, but a number of factors contributed to the general degradation of orbit determination accuracy. This included the fact that the spacecraft was at apogee, at a range of 1.7 million km and moving at less than I km/sec perpendicular to the line of sight. Nozomi was also close to zero degrees declination where there are known limitations on orbit determination performance. Finally, S-band tracking data was acquired through the Nozomi backup low gain antenna. This antenna is offset from the axis of this spin stabilized spacecraft and superimposed large signatures in the Doppler and range data. These difficulties were overcome by combining long data arcs, spanning several maneuvers, with a high fidelity solar pressure model. The model included a physically accurate representation of the spacecraft structure and a high time resolution orientation model. Observation modeling included the removal of the spin induced Doppler bias, spin signature and per pass correction of range calibration errors applied for data leading up to critical events. As a result, all orbit determination goals were met. A second area of emphasis in this paper is the JPL tracking and orbit determination effort in support of the TMI maneuver. TMI occurred out of contact with ground stations and the JPL Goldstone tracking complex had the first pass following the bum. As a result, MMNAV had the responsibility to make a rapid assessment of the maneuver performance. MMNAV made the determination that a 100 m/s under bum had occurred and promptly informed ISAS via voice lines. ISAS immediately began preparations for a correction maneuver (TMIc), which had to be performed during the next Usuda pass. The near real time assessment by MMNAV provided accurate antenna frequency and pointing updates for the spacecraft acquisition at Usuda and the close coordination between the two agencies enabled the design and successful execution of the TMc maneuver. Propellant consumption during the correction burn dictated that the mission be redesigned. ISAS developed a new plan which adds 3 full solar orbits, two Earth swingbys and one lunar swingby with arrival at Mars in January 2004. The final Mars orbit will still enable the mission to achieve all of its science objectives.

Ryne, Mark↗
NASA NTRS2018

Spacecraft Contamination Control Challenges for Space Missions with Organic Compound Detection Capabilities and for Potential Sample Return

Contamination control is critical to missions with organic compound detection capabilities, and for missions targeting acquisition of samples for potential return to Earth. Significant challenges are being addressed in the design of current flight projects and conceptual mission studies at JPL. These challenges extend to both orbiting spacecraft, as well as landed missions, for future missions to Mars and Europa, and potential missions to Titan and Enceladus. Contamination control during all phases of a mission, from preliminary design through operation, is fundamental to ensure that organic compounds of terrestrial origin are controlled to ensure successful completion of science objectives. This presentation summarizes contamination control challenges specific to landed missions (which include sample acquisition, encapsulation, caching, potential sample return, and UV instruments), and orbiting missions (modeling interactions between the spacecraft, and local exospheres and plumes).

Soares, Carlos↗
NASA NTRS2022

Mars 2020 Perseverance Rover SHERLOC Instrument Isolation System

The NASA Jet Propulsion Laboratory (JPL) successfully landed the Mars 2020 Perseverance Rover at Jezero Crater on the Martian surface. Perseverance’s main mission objective is to cache Martian rock and regolith samples in hermetically sealed tubes to be brought back to Earth by future missions. To achieve this goal the rover is equipped with a 2-meter-long Robotic Arm (RA) which manipulates the Turret to interact with the surface. The Turret is comprised of science instruments and tools that enable surface sampling and science. The backbone of the Turret hardware is the Rotary Percussive Coring Drill. The science instruments, which are directly mounted to the structural housing of this drill, must be able to withstand not only the launch and entry, descent and landing (EDL) loads of the vehicle but also the dynamic percussive environment, induced by the drill throughout surface operations. This paper discusses the technical hardware design, development and testing of a vibration isolation system to protect the Scanning Habitable Environments with Raman and Luminescence for Organics and Chemicals (SHERLOC) Instrument from these dynamic environments. The hardware design aspect of this problem was limited by both size and mass constraints. The Turret is tightly configured and the physical space in which hardware can reside is restricted due to multiple axes of surface interaction points. Another driving requirement is the extreme range of thermal non-operational environment from -135C to +90C. This, in addition to SHERLOC being a heavier instrument than previously integrated on a turret, made the use of isolation systems employed on previous rover missions problematic. The final design incorporates custom wire mesh springs preloaded within titanium hexapod struts equipped with flexure ends. Custom versions of commercially available wire mesh springs were designed and tested. These custom springs differed in dimension, density and stiffness from the off the shelf options. The newly designed springs went through various stages of flight acceptance testing. Detailed physics-based modeling to understand what loads the isolation system would experience was used to define the test conditions for the springs and entire isolation subsystem. The testing program developed around this was extensive to fully characterize the isolation performance over cleanliness levels, temperature and life. Another major aspect of this work was developing a method of replicating the drill’s percussive environment for performance and life testing of Turret mounted hardware. Attaching the instrument to an operating drill for testing was not a viable option. A method of using a Highly Accelerated Life Test (HALT) table was developed and applied across the project as the most representative and tunable option to physically simulate this new environment. This method, along with launch environment random vibration testing, was used to build an entire dynamic qualification and flight hardware test program. The results of the design, modeling, testing and analysis of the SHERLOC Isolation System is described throughout the following paper.

Krafchak, Todd↗
NASA NTRS1995

Some Recent Advances and Applications of Infrared Detector Technology at the Jet Propulsion Laboratory

The Jet Propulsion Laboratory has a long history in the development and application of infrared detector technology for the exploration of our planetary system. We will describe some recent applications of IR detector technology at JPL for ground based astronomy, the Earth Observing System, the Cassini Mission to Saturn, and the other proposed planetary missions. The infrared detector technologies to be discussed include single element, multiplexed linear and staring arrays, using a wide variety of infrared sensing materials including InSb, InGaAs, HgCdTe (PV and PC) and SiAs.

IR↗
NASA NTRS2017

Spacecraft Contamination Control Challenges for Space Missions with Organic Compound Detection Capabilities and for Potential Sample Return

Spacecraft contamination control is critical to space exploration missions with organic compound detection capabilities, and also for missions targeting acquisition of samples for potential return to Earth. Significant challenges are being addressed and resolved in the design of current flight projects and conceptual mission studies at JPL. These challenges extend to both orbiting spacecraft, as well as landed missions, for future missions to Mars and Europa, and potential missions to Titan and Enceladus. Contamination control during all phases of a mission, from preliminary design through operation, is fundamental to ensure that organic compounds of terrestrial origin are controlled to ensure successful completion of science objectives. This paper examines contamination control challenges specific to landed missions (which include sample acquisition, encapsulation, caching, potential sample return, and UV instruments), and orbiting missions (modeling interactions between the spacecraft and local exospheres and plumes).

Sylvia, Margarite↗
NASA NTRS1989

Galileo Science Writers' Briefing

This NASA Kennedy video production presents Part 1 of a press conference held at JPL on August 8, 1989. The briefing in its entirety covers the Galileo Project's mission design from launch to completion in 1997 and is moderated by JPL Public Information Mgr. Robert Macmillan. Part 1 of the 3 part video series includes presentations by Richard J. Spehalski (Galileo Project Manager) and Clayne M. Yeates (Acting Science Mission Design Manager). Mr. Spehalski's presentation includes actual footage of spacecraft preparations at Kennedy Space Center and slides of mission timelines. Dr. Yeates discusses the Galileo mission in chronological order and includes slides of the interplanetary trajectory, encounter geometry, propellant margins vs. launch date, and planned earth images.

Source record↗
NASA NTRS2020

Satellite-To-Satellite Imaging in Support of LEO Optical Navigation, Using the ASTERIA Cubesat

The Arcsecond Space Telescope Enabling Research in Astrophysics (ASTERIA) was a 6-unit CubeSat technology demonstration mission that was built at NASA’s Jet Propulsion Laboratory (JPL) and deployed from the International Space Station (ISS) on November 20th, 2017. After successfully completing its 90-day primary mission that demonstrated arcsecond-level line-of-sight pointing and focal plane thermal stability for exoplanet detection, it entered an extended mission performing onboard software demonstrations alongside science until end of mission in December 2019. At the end of its lifetime it was being used as a demonstration platform for several experiments, including low earth orbit (LEO) optical navigation operations. With its visible light astrometric camera and stable attitude control system, the ASTERIA spacecraft showed itself to be a capable platform for the imaging of geosynchronous satellites from LEO. This paper will describe the imagery attained in flight and also the image processing algorithms that were developed to render that imagery into navigation quality data. These algorithms dealt with hot pixel filtering, noise modeling, attitude registration, star signal rejection and satellite signal identification. Brightness prediction algorithms used for target selection will also be discussed.

Fesq, Lorraine↗
NASA NTRS2003

Nonlinear Dynamics of the Superfluid Transition: What may We learn on orbit?

Linear response (specifically, Fourier's Law) in He-4 has been observed to fail in heat flow experiments near the superfluid transition. A detailed analysis of the data suggests that the hydrostatic pressure gradient across the helium column limits the divergence of the correlation length in our earth-based experiments. This is consistent with other observations, such as the surprising lack of mutual friction and hysteresis near the superfluid transition, and a 'rounding' of the transition that appears to be independent of heat flux in the low heat flux limit. I will discuss these unusual results from earth-based measurements, and will show predictions for the very different results that may result when we make our measurements on orbit as part of the M1 Mission of the Low- Temperature, Microgravity Physics Facility. This work has been funded by the Fundamental Physics Discipline within the Physical Sciences Research Office of NASA, and is conducted by the DYNAMX (UNM) and CQ (Caltech) Groups, with assistance from the Low Temperature Science and Quantum Sensors Group at JPL.

Duncan, Rob↗