The previous Decadal Survey included a Working Group report on "Explorers, Suborbital, and Other Platforms", included in the Decadal as "Appendix C".The

The positions, experiences, and viewpoints expressed below are those of

When the first CubeSats were launched nearly two decades ago, few people believed that the miniature satellites would likely prove to be a useful scientific tool. Skeptics abounded. However, the last decade has seen the highly successful implementation of space missions that make creative and innovative use of fast-advancing CubeSat and small satellite technology to carry out important science experiments and missions. Several projects now have used CubeSats to obtain first-of-their-kind observations and findings that have formed the basis for high-profile engineering and science publications, thereby establishing without doubt the scientific value and broad utility of CubeSats. In this paper, we describe recent achievements and lessons learned from a representative selection of successful CubeSat missions with a space weather focus. We conclude that these missions were successful in part because their limited resources promoted not only mission focus but also appropriate risk-taking for comparatively high science return. Quantitative analysis of refereed publications from these CubeSat missions and several larger missions reveals that mission outcome metrics compare favorably when publication number is normalized by mission cost or if expressed as a weighted net scientific impact of all mission publications.

Accurate knowledge of the full, three-dimensional electric field vector is of fundamental importance in understanding electrodynamics of a vast variety of space plasmas. However, heliophysics research still lacks access to the reliable parallel electric field measurements required to close many significant science questions. This uncertainty represents a significant barrier to progress in the field. The only way to close this major observational gap is a profound change in electric field instrument design. A new electric field instrument called Grotifer is now being designed to address the need for highly accurate three-dimensional electric field measurements while enabling lower cost missions and constellation missions in deep space. Grotifer (Giant rotifer) is a reference to the rotifer, also known as the “wheel animalcule.” Similarly, Grotifer consists of mounting detectors on two rotating plates, orthogonal to each other, on a non-rotating central body. The two rotating plates provide continuous high-accuracy three-dimensional measurements of both electric fields and magnetic fields. The Grotifer design leverages more than 50 years of expertise in delivering highly accurate spin plane electric field measurements, while overcoming inaccuracies generated by spin axis electric field measurements. Our current efforts focus on designing Grotifer as a SmallSat (27U CubeSat). That said, Grotifer could also become part of the payload on a much larger platform. In the future, one could imagine fleets of Grotifers studying electrodynamics at many points, facilitating differentiation between spatial and temporal dynamics. Plasma detectors could also be added to the rotating plates to cover the full phase space better than is done on spinning spacecraft, leading to more complete correlation studies of the fields and plasmas.

Earth and Space Science Open Archive PosterOpen AccessYou are viewing the latest version by default [v1]Meet Grotifer: a CubeSat that Will Provide Highly Accurate Three-Component Electric Field Measurements throughout the HeliosphereAuthorsSolèneLejosneiDDavidAuslanderiDJohnBonnelliDDavidKlumparJeremyMcCauleyRubinMeuchelForrestMozeriDDavidPankowJohnSampleiDLarrySpringerSee all authors Solène LejosneiDCorresponding Author• Submitting AuthorSpace Sciences Laboratory, University Of California, BerkeleyiDhttps://orcid.org/0000-0003-4238-8579view email addressThe email was not providedcopy email addressDavid AuslanderiDMechanical Engineering DepartmentUniversity of CaliforniaBerkeleyiDhttps://orcid.org/0000-0002-1377-0624view email addressThe email was not providedcopy email addressJohn BonnelliDSpace Sciences Laboratory, University Of California, BerkeleyiDhttps://orcid.org/0000-0002-0675-7907view email addressThe email was not providedcopy email addressDavid KlumparSpace Science and Engineering Laboratory, Department of Physics, Montana State University - Bozemanview email addressThe email was not providedcopy email addressJeremy McCauleySpace Sciences Laboratory, University Of California, Berkeleyview email addressThe email was not providedcopy email addressRubin MeuchelSpace Science and Engineering Laboratory, Department of Physics, Montana State University - Bozemanview email addressThe email was not providedcopy email addressForrest MozeriDSpace Sciences Laboratory, University Of California, BerkeleyiDhttps://orcid.org/0000-0002-2011-8140view email addressThe email was not providedcopy email addressDavid PankowSpace Sciences Laboratory, University Of California, Berkeleyview email addressThe email was not providedcopy email addressJohn SampleiDSpace Science and Engineering Laboratory, Department of Physics, Montana State University - BozemaniDhttps://orcid.org/0000-0002-9516-9292view email addressThe email was not providedcopy email addressLarry SpringerSpace Science and Engineering Laboratory, Department of Physics, Montana State University - Bozemanview email addressThe email was not providedcopy email address

This paper describes the mission concept for "RadPC-Lunar", a technology demonstration of a novel computer architecture that can recover from faults caused by ionizing radiation. RadPC-Lunar is a payload that was selected by NASA in 2019 to go to the Moon through its Commercial Lunar Payload Services (CLPS) project as part of the Artemis lunar program. RadPC-Lunar will travel to the lunar surface in 2022-23 onboard a commercial lander and will spend a minimum of 7-days in the Mare Crisium. This demonstration will serve two important purposes in support of future lunar missions. First, it will demonstrate a key technology for computationally intense autonomous lunar activities such as in situ resource utilization, robotic surface operations, and entry/descent/landing maneuverers while providing increased reliability over the state-of-theart in space computers. Second, it will provide a characterization of the radiation effects environment of the lunar surface by tracking upsets within the computing fabric and correlating them to data from on-board dosimeters. The payload can also provide a unique set of measurements on the ionizing radiation environment as it passes through the Earth's magnetosphere during transit to the Moon.This paper describes the overall mission concept of RadPC-Lunar in addition to the details of the design-of-experiments and the types of data that will be collected. This paper will be of interest to engineers and scientists studying the lunar transit and lunar surface radiation environment and those working with radiation tolerant avionics. The timing of the presentation will allow the RadPC-Lunar team to solicit feedback from the aerospace community that can influence its design-of-experiments prior to completion in order to maximize the return of the mission.

FIREBIRD-II (Focused Investigations of Relativistic Electron Burst Intensity, Range, and Dynamics) is a National Science Foundation CubeSat mission exploring relativistic electron microbursts. The mission consists of two identically instrumented CubeSats that were launched into a near polar orbit on January 31, 2015. Each spacecraft has two solid state detectors that return high cadence (10's of ms) measurements of the electron population. Both units operated continuously for almost 5 years and one unit continues to operate and return high quality data over 6 years after launch.

This is a COSPAR roadmap to advance the frontiers of science through innovation and international collaboration using small satellites. The world of small satellites is evolving quickly and an opportunity exists to leverage these developments to make scientific progress. In particular, the increasing availability of low-cost launch and commercially available hardware provides an opportunity to reduce the overall cost of science missions. This in turn should increase flight rates and encourage scientists to propose more innovative concepts, leading to scientific breakthroughs. Moreover, new computer technologies and methods are changing the way data are acquired, managed, and processed. The large data sets enabled by small satellites will require a new paradigm for scientific data analysis. In this roadmap we provide several examples of long-term scientific visions that could be enabled by the small satellite revolution. For the purpose of this report, the term “small satellite” is somewhat arbitrarily defined as a spacecraft with an upper mass limit in the range of a few hundred kilograms. The mass limit is less important than the processes used to build and launch these satellites. The goal of this roadmap is to encourage the space science community to leverage developments in the small satellite industry in order to increase flight rates, and change the way small science satellites are built and managed. Five recommendations are made; one each to the science community, to space industry, to space agencies, to policy makers, and finally, to COSPAR.

We present the observation of a spatially large microburst with multiple bounces made simultaneously by the Focused Investigation of Relativistic Electron Bursts: Intensity, Range, and Dynamics II (FIREBIRD-II) CubeSats on 2 February 2015. This is the first observation of a microburst with a subsequent decay made by two coorbiting but spatially separated spacecraft. From these unique measurements, we place estimates on the lower bounds of the spatial scales as well as quantify the electron bounce periods. The microburst's lower bound latitudinal scale size was 29 +/- 1 km and the longitudinal scale size was 51 +/- 1 km in low Earth orbit. We mapped these scale sizes to the magnetic equator and found that the radial and azimuthal scale sizes were at least 500 +/- 10 km and 530 +/- 10 km, respectively. These lower bound equatorial scale sizes are similar to whistler mode chorus wave source scale sizes, which supports the hypothesis that microbursts are a product of electron scattering by chorus waves. Lastly, we estimated the bounce periods for 200- to 800-keV electrons and found good agreement with four common magnetic field models. Plain Language Summary Microbursts are a subsecond impulsive increase of electron precipitation from the outer Van Allen radiation belt into the atmosphere, believed to be an important loss process of radiation belt electrons. Here we present an observation of a microburst observed simultaneously by the twin Focused Investigation of Relativistic Electron Bursts: Intensity, Range, and Dynamics II CubeSats. This unique observation allowed us to calculate the microburst's spatial scale size and electron bounce periods. The spatial scale size in low Earth orbit was found to be a few tens of kilometers in size, one of the largest reported in literature. We then magnetically mapped this scale size to the region near the microburst's generation region and found it to be around 500 km, similar to the spatial scale size of the waves that are believed to be responsible for microburst generation. This observation shows an example of how large microbursts can be, and it sheds light on its scattering mechanism.

We present initial dual spacecraft observations that for the first time both constrain the spatial scale size and provide spectral properties at medium energies of electron microbursts. We explore individual microburst events that occurred on 2 February 2015 using simultaneous observations made by the twin CubeSats which comprise the National Science Foundation (NSF) Focused Investigations of Relativistic Electron Bursts: Intensity, Range, and Dynamics (FIREBIRD II). During these microburst events, the two identically instrumented FIREBIRD II CubeSats were separated by as little as 11 km while traversing electron precipitation regions in low‐Earth orbit. These coincident microburst events map to size scales >120 km at the equator. Given the prevalence of coincident and noncoincident events we conclude that this is of the same order of magnitude as that of the spatial scale size of electron microburst, an unknown property that is critical for quantifying their overall role in radiation belt dynamics. Finally, we present measurements of electron microbursts showing that precipitation often occurs simultaneously across a broad energy range spanning 200 keV to 1 MeV, a new form of empirical evidence that provides additional insights into the physics of microburst generation mechanisms.

Nanosatellites, a class of small satellites, are becoming increasingly popular because of their small form factor and many other attractive features. In the process of qualifying nanosatellites for space readiness, their thermal behavior can be investigated in a laboratory setup using a thermal vacuum system to mimic orbital conditions. For these reasons, a thermal vacuum system suited specifically for nanosatellites was desired for performing thermal vacuum testing. Analytical calculations and laboratory testing were performed as part of the design of this thermal vacuum system. A set of simultaneous equations was solved using the LU Decomposition method to find the radiosities of several surfaces in an enclosure. The radiosities along with their respective view factors were then used to solve for the heat power required to heat and cool the thermal shroud under steady state conditions at the most extreme operating conditions expected. The analysis was performed on a system of three concentric cylinders of varying heights: the outer being the vacuum chamber wall, the middle the thermal shroud inside the chamber, and the inner the satellite. Under the most extreme operating conditions expected, the thermal shroud was cooled to −40°C and the satellite heated to 80°C during satellite cooling and the reverse during satellite heating. All surfaces in the enclosure were assumed to be diffuse, grey, opaque and isothermal. The thermal shroud was separated into two surfaces: the cylindrical shroud body and the shroud top disc. From the analytical results, the expected heating power for the shroud body was found to be 704.0 Watts, and 229.8 Watts for the shroud top. During cooling, where the temperatures were reversed, the expected heat power for the shroud body was calculated as −685.5 Watts, and −220.9 Watts for the shroud top. An experimental setup was tested under similar conditions as a comparison and as a method to validate the thermal shroud design and the analytical calculations. The shroud body and top heaters were selected to output 750 Watts and 230 Watts, respectively, and were driven at their maximum output, with the satellite held at −40°C. The shroud reached 80°C with no difficulty, indicating that the analytical calculations had correctly predicted the required heat power and that the design of the thermal shroud was capable of supporting testing under the most extreme conditions expected.

Ionospheric ions apparently accelerated transversely to the geomagnetic field in the topside ionosphere are regularly detected by the soft particle spectrometers on the ISIS satellites. Such gyro-accelerated ions are observed in association with precipitating auroral electrons. A detailed study of their relationship with such electrons and with field-aligned currents, together with simultaneous measurements of the local plasma composition and density, reveals the specific conditions present in the topside ionosphere during the generation of such transversely accelerated ions (TAI). A proposed mechanism for generation of TAI involving acceleration by electrostatic ion cyclotron waves is consistent with the present observations.

This chapter contains sections titled: Introduction Data Set Statistical Results Discussion Summary

Due to their small size and other attractive features, nanosatellites are becoming popular in space applications. Experimental investigation of the thermal behavior of such a satellite can be conducted in a laboratory setup using a thermal vacuum chamber to mimic the conditions of outer space. A small, cost effective thermal vacuum system was desired for performing thermal vacuum testing on nanosatellites. Numerical calculations and laboratory testing were performed as part of the design of this thermal vacuum system. A numerical method using the finite element method was employed to determine the amount of heat flux needed to be applied at the bottom plate of a satellite to achieve a certain rate of temperature increase in the plate. The numerical analysis was performed on a 40.5 kg satellite structure to predict the heat rate per unit area through its bottom surface when it was cycled in the temperature range of −40°C to +80°C with a rate of temperature change from 1°C/min to 5°C/min. A time dependent increase in temperature on the bottom wall was used as a boundary condition. The rest of the satellite walls were assumed to be insulated. Contact resistances between the components of the satellite structure were neglected. Temperature and heat flux distributions on various walls of the satellite were computed and reported in the study. From the numerical results, a maximum heat flux rate of 3,332 W/m2 was calculated on the bottom plate for a temperature increase rate of 1.5°C/min of the plate. A similar experimental setup was tested under similar conditions as a comparison and as a method to validate the thermal system design. Experimental results indicated a heat flux rate of 17,094 W/m2 through a test satellite. The difference between the numerical and experimental results is attributed to geometric differences between the numerical satellite model and the experimental test structure.

This chapter contains sections titled: Introduction Instrument Description Critical Design Elements Test and Calibration Results Modular Concept