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Robotic exploration of the solar system using small satellites (SmallSats) is gaining popularity because of SmallSats’ lower cost and faster development cycle compared to primary science missions. A potential obstacle for deployment of SmallSats in deep space is the limitation associated with the communications link imposed by SmallSats' frugal power and antenna size. These technical limitations constrain the range and throughput of SmallSats and may impact their navigation performance. To allow SmallSats to play a greater role in deep space science endeavors, we propose an enhanced communications architecture that will lessen the communications disparity between larger, more expensive missions and SmallSats. The objectives of the study are to define a communications architecture, in-flight and flight-to-ground, that is needed for planetary SmallSats, both standalone or in support of a larger mission; and to develop a guideline for primary spacecraft that would communicate with ride-along SmallSats. This work treats the space segment and the ground network as a whole to provide for improvements not only via investments in technology, but also via improvements in operational procedures. The study includes reviews of current communication concepts, technologies, and procedures, as well as an assessment of navigation needs, evaluation of the communication architecture performance for a range of destinations and applications, and identification of any gaps in capability. The study product is a set of recommendations to space agencies and the planetary SmallSat community. The main thrust of these recommendations is investment in technology, such as radios, antennas, protocols, and methodologies. The concepts for improving communications capability of deep space SmallSats are discussed in a series of three articles, Improving Small Satellite Communications in Deep Space. In this article, Part I, we discuss SmallSat direct-to-Earth links, as well as SmallSat communications equipment. In Part II, we discuss navigation topics, proximity links and networks, and the use of the communications channel for science observations. We present the ground network in Part III. Definitions and assumptions are provided here in Part I.

The Lunar Polar Hydrogen Mapper (LunaH-Map) mission will map the distribution of hydrogen around the lunar South Pole using a miniature neutron spectrometer. The mission builds upon a decade of lunar science, which has revealed both regional and more localized enrichments of water ice near the lunar poles. Localized enrichments are primarily within permanently shadowed regions (PSRs) and craters throughout the South Pole. The spatial extent of these regions is often below the resolution of previous neutron instruments that have flown on lunar missions. The neutron leakage spectrum from planetary surfaces is primarily sensitive to hydrogen abundance in the top meter of regolith, however, for neutron spectrometers with omnidirectional sensitivity, the spatial resolution is limited by the spacecraft orbital altitude above the surface. A low altitude measurement from a distance on the same scale of the PSRs could spatially isolate and constrain the hydrogen enrichments both within and around within those regions. A small spacecraft mission is ideally suited to acquire the low-altitude measurements required to localize hydrogen enrichments using neutron spectroscopy at the lunar South Pole. LunaH-Map will use a solid iodine ion propulsion system, X-Band radio communications through the NASA Deep Space Network, star tracker, Command & Data Handling System, and EPS systems from Blue Canyon Technologies, solar arrays from MMA Designs, LLC, mission design and navigation by KinetX. Spacecraft systems design, integration, qualification, test, and mission operations are performed by Arizona State University, AZ Space Technologies and Qwaltec.

ASTERIA (Arcsecond Space Telescope Enabling Research in Astrophysics) was a 6-unit CubeSat technology demonstration mission that deployed from the International Space Station 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 to mature technology both in space and on the ground. One of the technologies was a completely cloud-based ground system leveraging Amazon Web Services (AWS) Ground Station service. Announced in November 2018 and launched in May 2019, AWS Ground Station is a fully managed ground station service that aims to reduce the overhead associated with developing and maintaining ground system infrastructure throughout the mission lifecycle. AWS Ground Station makes available the suite of features required for any ground system in support of low-Earth orbit (LEO) and medium-Earth Orbit (MEO) satellite operations on-demand and without setting up or maintaining long-term contracts. Charges are incurred on a per-minute basis for antenna usage during scheduled tracks. Support is available for S-band uplink and downlink, along with X-band narrowband and wideband downlink. Missions that use the service may reserve tracks with any licensed AWS Ground Station antennas located across each service region and have direct access to any AWS services in support of mission operations. The cloud-based architecture built around the AWS Ground Station service greatly enhanced ASTERIA mission operations by enabling end-to-end pass automation, on-demand contact scheduling and contingency planning, along with more efficient data downlink through station availability and station-to-station handovers. It incorporated open-source software, particularly NASA's AMMOS Instrument Toolkit (AIT) and Open Mission Control Technologies (OpenMCT), along with the AWS application programming interfaces (API) to the Ground Station, Elastic Compute Cloud (EC2) and Simple Storage Service (S3) services. After showcasing operability in August 2019, the team continued using and improving this novel ground system architecture until the end of mission in December 2019. This paper describes the cloud-based ground system, how it was designed, tested, and evaluated with an in-orbit spacecraft, the operational capabilities that it enabled, along with lessons learned and recommendations for future missions.

Alpha Centauri is a triple star system with two Sun-like stars, α Cen A (V = 0.01) and B (V = 1.33), and a third fainter red dwarf star, Proxima Centauri. Most current transit missions cannot produce precision photometry of α Cen A and B as their detectors saturate for these very bright stars. The Arcsecond Space Telescope Enabling Research in Astrophysics (ASTERIA) was a technology demonstration mission that successfully demonstrated two key technologies necessary for precision photometry achieving line-of-sight fine-pointing stability of 0.5″ rms and focal plane temperature control of ±0.01 K over a period of 20 minutes. The payload consisted of a 6.7 cm aperture diameter refractive camera and used a scientific complementary metal-oxide semiconductor detector that enabled monitoring of the brightest stars without saturating. We obtained spatially unresolved (blended) observations of α Cen A and B during opportunistic science campaigns as part of ASTERIA’s extended mission. The resulting 1σ photometric precision for the blended α Cen A and B data is 250 ppm (parts per million) per 9 s exposure. We do not find evidence of transits in the blended data. We establish limits for transiting exoplanets around both α Cen A and B using transit signal injection and recovery tests. We find that ASTERIA is sensitive to planets with radii as small as 3.0 R ⊕ around α Cen A and 3.7 R ⊕ around α Cen B, corresponding to signals of ∼500 ppm (signal-to-noise ratio = 5.0) in the blended data, with periods ranging from 0.5 to 6 days.

This chapter aims to provide the reader with an overview of the main aspects of designing a telecommunication system for a CubeSat or a SmallSat. The main factors to consider when designing such a system (data generation, electrical power system constraints, ground station availabilities, range, pointing, and spectrum regulations) are presented. Then an overview of the iterative design process is given, and key analysis (coverage and link analysis) is discussed in details. Finally the main telecommunication components required for the system such as antennas and transceivers are presented. The chapter concludes with some considerations for optical telecommunication systems.

ASTERIA (Arcsecond Space Telescope Enabling Research In Astrophysics) is a 6U CubeSat space telescope (10 cm x 20 cm x 30 cm, 10 kg). ASTERIA's primary mission objective was demonstrating two key technologies for reducing systematic noise in photometric observations: high-precision pointing control and high-stabilty thermal control. ASTERIA demonstrated 0.5 arcsecond RMS pointing stability and $\pm$10 milliKelvin thermal control of its camera payload during its primary mission, a significant improvement in pointing and thermal performance compared to other spacecraft in ASTERIA's size and mass class. ASTERIA launched in August 2017 and deployed from the International Space Station (ISS) November 2017. During the prime mission (November 2017 -- February 2018) and the first extended mission that followed (March 2018 - May 2018), ASTERIA conducted opportunistic science observations which included collection of photometric data on 55 Cancri, a nearby exoplanetary system with a super-Earth transiting planet. The 55 Cancri data were reduced using a custom pipeline to correct CMOS detector column-dependent gain variations. A Markov Chain Monte Carlo (MCMC) approach was used to simultaneously detrend the photometry using a simple baseline model and fit a transit model. ASTERIA made a marginal detection of the known transiting exoplanet 55 Cancri e ($\sim2$~\Rearth), measuring a transit depth of $374\pm170$ ppm. This is the first detection of an exoplanet transit by a CubeSat. The successful detection of super-Earth 55 Cancri e demonstrates that small, inexpensive spacecraft can deliver high-precision photometric measurements., Comment: 23 pages, 9 figures. Accepted in AJ

During the Apollo era of lunar exploration, mysterious albedo patterns, called swirls, captured the imagination of the scientific community. A key aspect of this interest was due to the discovery that the swirls are associated with localized relatively strong remnant magnetic fields. Analysis of returned soil samples revealed that solar wind, galactic cosmic rays, and micrometeorite impacts change the albedo of surface soil grains, a process known as space weathering that reduces regolith reflectance over time. Thus it was natural to invoke local magnetic structures as shields that retard space weathering resulting in relatively less space weathering of the regolith in regions with the strongest localized magnetic fields. However, the origins of the magnetic anomalies and associated swirls remain enigmatic to this day. We designed the Swirl CubeSat to determine the nature of remnant lunar magnetic fields and investigate their role in moderating space weathering of the regolith and assess their suitability for radiation protection of surface assets. Swirl has one focused observational objective: characterize the magnetic field associated with the Reiner Gamma Swirl (RGS) at sub-kilometer spatial sampling, with 0.5 nTesla accuracy and 100 m spatial precision. In the Swirl mission concept, the Swirl spacecraft, a 6U CubeSat, would deploy as a secondary payload from a vehicle on a deep space trajectory. A series of maneuvers would then place the spacecraft in a low orbit that would then be modified to have a periapse of 5–10 km for thirty orbits passing over RGS. Prime Swirl observations were designed during these low-altitude passes and consist of high-resolution vector magnetic field measurements and monochrome navigation imaging.

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 cislunar space station, which is a strategic gateway towards permanent settlement of the Moon and Mars. A station in deep space will require structures with large surfaces, very high volume to mass ratio and high-packing efficiency. The transportation of bulky payloads to beyond low-earth orbits pose formidable cost and logistical challenges. This requires a paradigm shift in research towards methods to build and assemble 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 blocks 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.

In the last decade, great progress was made in the development of small satellites and CubeSats. Several small spacecraft were designed, fabricated, launched, and successfully operated in low Earth orbit. While more companies and space agencies are becoming interested in CubeSats and small satellite, also the mission goals for these spacecraft are gradually changing: these small spacecraft are starting to be considered for deep-space, interplanetary exploration. Given the limited size, mass and power capabilities of these small platforms, one of the most interesting problems to address is how to develop a communication system to allow small satellites to communicate from very far distance in the solar system. This paper aims to address this problem by proposing cooperative communication approaches in which multiple CubeSats communicate cooperatively together to improve the link performance with respect to the case of a single satellite transmitting. Three approaches are proposed: a beam-forming approach, a coding approach, and a network approach. The approaches are applied to the specific case of the SOLARA/SARA concept: a proposed constellation of CubeSats at the Lunar Lagrangian point L1 which aims to perform radio astronomy at very low frequencies (30 kHz to 3 MHz). The paper describes the development of the approaches, the simulation and a graphical user interface developed in Matlab which allows to perform trade-offs across multiple constellation׳s configurations.

The emergence of small satellites and CubeSats for interplanetary exploration will mean hundreds if not thousands of spacecraft exploring every corner of the solar-system. Current methods for communication and tracking of deep space probes use ground based systems such as the Deep Space Network (DSN). However, the increased communication demand will require radically new methods to ease communication congestion. Networks of communication relay satellites located at strategic locations such as geostationary orbit and Lagrange points are potential solutions. Instead of one large communication relay satellite, we could have scores of small satellites that utilize phase arrays to effectively operate as one large satellite. Excess payload capacity on rockets can be used to warehouse more small satellites in the communication network. The advantage of this network is that even if one or a few of the satellites are damaged or destroyed, the network still operates but with degraded performance. The satellite network would operate in a distributed architecture and some satellites maybe dynamically repurposed to split and communicate with multiple targets at once. The potential for this alternate communication architecture is significant, but this requires development of satellite formation flying and networking technologies. Our research has found neural-network control approaches such as the Artificial Neural Tissue can be effectively used to control multirobot/multi-spacecraft systems and can produce human competitive controllers. We have been developing a laboratory experiment platform called Athena to develop critical spacecraft control algorithms and cognitive communication methods. We briefly report on the development of the platform and our plans to gain insight into communication phase arrays for space., Comment: 4 pages, 10 figures, IEEE Cognitive Communications for Aerospace Applications Workshop

Stationary orbits around Mars, which are known as areostationary orbits, have similar characteristics as Earths geostationary orbits. An areostationary Relay Satellite is expected to be used for future communication network for the exploration of Mars. An areostationary Relay Satellite at 17,000 km above Mars surface is always in the same place in the sky on Mars, receiving data from the rovers/landers and low Mars orbiters and transmitting it to an Earth station. The areostationary satellite enables almost all time communication, except for certain durations when Mars is in between the areostationary Relay Satellite and the Earth. In this paper we define a code division multiple access (CDMA) communication system between various assets such as rovers/landers on Mars surface, low Mars orbiters such as MRO, CubeSats in vicinity of Mars, and an areostationary relay satellite at 17,000 km above Mars surface. We assume data rates between 50 Kbps and 1 Mbps operating at 8.40 GHz frequency. We define the transmit power, antenna size, and data rates for the assets communicating with the areostationary relay satellite. Since the areostationary relay satellite simultaneously receives data from many users we have chosen the code division multiple access (CDMA) as a multiple access scheme, which does not require time/frequency coordination among users. Further-more we compute the link budgets for various communication links, we define the modulation format, the PN spreading and chip rates. We provide a new theoretical analysis for multiuser interference, and simulate the performance of coded CDMA communication system. For coded system we use standard CCSDS LDPC codes proposed by NASA for space applications. Using CDMA, the communication structure of all assets on Mars operates at one single RF frequency. However the initial states of spreading PN codes assigned to users are different such that if all users communicate simultaneously with the areostationary relay satellite over a given band will have low cross correlation that reduces the multiuser interference.

Interplanetary1 CubeSats and small satellites have potential to provide means to explore space and to perform science in a more affordable way. As the goals for these spacecraft become more ambitious in space exploration, the communication systems currently implemented will need to be improved to support those missions. One of the bottlenecks is the antennas' size, due to the close relation between antenna gain and dimensions. Hence, a possible solution is to develop inflatable antennas which can be packaged efficiently, occupying a small amount of space, and they can provide, once deployed, large dish dimension and correspondent gain. A prototype of a 1 m inflatable antenna for X-Band has been developed in a joint effort between JPL and ASU. After initial photogrammetry tests and radiation tests, it was discovered that the design was not able to meet the required gain. As a result, a new design, based on a spherical inflatable membrane, is proposed. This new design will allow reaching a more stable inflatable surface, hence improving the electromagnetic performance. This paper will detail the principle challenges in developing this new antenna focusing on: design, EM analysis, fabrication and tests.

This letter a. rectifies misrepresentations of the classical model b. identifies methodological shortcomings in the authors's approach and c. compares performance of the authors' "calibration weighting" with equal weights and global weights of the classical model on the large expert judgement data set made available in (Cooke and Goossens 2008)

Summary Expert elicitation is increasingly applied to different research areas. Multiple approaches have been implemented, but the development of methods to quantify experts' biases and calibration represents a challenge. As a result, the integration of multiple and often conflicting opinions can be demanding, owing to the complexity of properly weighting experts' contributions. We propose an approach to address this problem when probability densities for seed calibration variables are not available. The methodology generates an expert score that is employed to aggregate multiple-expert assessments. The approach has been experimentally applied to engineering design risk analysis. Results indicate that the approach improves the quality of the estimations. The weighted aggregations of experts' estimates based on the experts' scores achieve better results than the corresponding aggregations based on experts' opinions equally weighted.

CubeSats and small satellites have potential to provide means to explore space and to perform science in a more affordable way. As the goals for these spacecraft become more ambitious in space exploration, moving from Low Earth Orbit (LEO) to Geostationary Earth Orbit (GEO) or further, the communication systems currently implemented will not be able to support those missions. One of the bottlenecks in small spacecraft communication systems is represented by antennas' size, due to the close relation between antenna gain and dimensions. Current antennas for CubeSats are mostly dipole or patch antennas with limited gain. Deployable (not inflatable) antennas for CubeSats are currently being investigated, but these solutions are affected by the challenge of packaging the whole deployable structure in a small spacecraft. The work that we propose represents the first attempt to develop an inflatable antenna for CubeSats. Inflatable structures and antennas can be packaged efficiently occupying a small amount of space, and they can provide, once deployed, large dish dimension and correspondent gain. Inflatable antennas have been previously tested in space (Inflatable Antenna Experiment, STS-77). However they have never been developed for small spacecraft such as CubeSats, where the packaging efficiency, the deployment, and the inflation represent a challenge. Our study explores for the first time the possibility of developing such antenna in a way compatible with CubeSat dimensions and constraints. The research provides answers on the possible dimensions for an inflatable antenna for small satellites, on the gain and resolution that can be achieved, and on the deployment and inflation mechanism compatible with CubeSat. Future work in the development of the antenna will include the test of the antenna in flight during a specific technical demonstration mission. The article is structured as follows: context and motivation for Cubesat inflatable antenna are described; then a study to design the antenna which achieves the required performance metrics, while respecting the constraints imposed by CubeSat structure, is presented.

Spacecraft is complex systems that involve different subsystems and multiple relationships among them. For these reasons, the design of a spacecraft is an evolutionary process that starts from requirements and evolves over time across different design phases. During this process, a lot of changes can happen. They can affect mass and power at component, subsystem, and system levels. Each spacecraft has to respect the overall constraints in terms of mass and power: for this reason, it is important to be sure that the design does not exceed these limitations. Current practice in the system model primarily deals with this problem by allocating margins on individual components and on individual subsystems. However, a statistical characterization of the fluctuations in mass and power of the overall system (i.e., the spacecraft) is missing. This lack of an adequate statistical characterization would result in a risky spacecraft design that might not fit the mission constraints and requirements, or in a conservative design that might not fully utilize the available resources. Due to the complexity of the problem and due to the different expertise and knowledge required to develop a complete risk model for a spacecraft design, this research is focused on risk estimation for a specific spacecraft subsystem, the communication subsystem. The current research aims to be a “proof of concept” of a risk-based design optimization approach, which can then be further expanded to the design of other subsystems as well as to the whole spacecraft. The objective of this paper is to develop a mathematical approach to quantify the likelihood that the major design drivers of mass and power of a space communication system would meet the spacecraft and mission requirements and constraints through the mission design lifecycle. Using this approach the communication system designers will be able to evaluate and compare different communication architectures in a risk tradeoff prospective. The results described in the presentation include a baseline communication system design tool, and a statistical characterization of the design risks through a combination of historical mission data and expert opinion contributions. An application example of the communication system of a university spacecraft is presented.

In this paper a communication system for CubeSats in formation that operate in the vicinity of the Moon is proposed. A CDMA system for the fleet of CubeSats in the vicinity of the moon to communicate with the Earth station is considered. This is an extension of our previous proposed CDMA system for a concept Constellation of CubeSats. In this paper the Doppler effects on CDMA communication system performance for a constellation of CubeSats around the Moon will be investigated. As an example we have estimated the maximum Doppler and Doppler rate profile of a SOLARA/SARA CubeSat scenario. We investigate the Effects of Doppler Shift/Rate on the CDMA system performance as a result of the CubeSat constellation orbiting in a halo orbit around Earth-Moon Lagrange Point L1. A detailed analysis and simulation of system in presence of Doppler frequency and unknown carrier phase will be performed. First we define the CDMA system for uplink and downlink between the vicinity of the Moon and Earth station. Link budgets will be provided both for uplink and downlink. Bandwidth limitations imposed by the spectral standards will be investigated for modulation formats. All system simulations are done using Simulink Matlab platform. For highly efficient nonlinear power amplifiers, we prefer to use a filtered offset QPSK with phase modulation, which is a CCSDS standard for constant envelope signaling. This allows us to use a nonlinear amplifier at CubeSat to operate at saturation point for the highest efficiency. Filtered offset QPSK with phase modulation is much more bandwidth efficient scheme. We demonstrate that this modulation format satisfies the two international spectral standards. For estimated and specified Doppler frequencies and Doppler rates we design frequency-tracking loops to track the Doppler frequency and Doppler rate in the presence of data with filtered offset QPSK with phase modulation. For carrier phase offsets a well-designed tracking loop is derived with specified loop bandwidth for the same modulation format. We use the CCSDS standard LDPC codes for space applications to meet the link budget margins.

CubeSats1 and small satellites have potential to provide means to explore space and to perform science in a more affordable way. As the goals for these spacecraft become more ambitious in space exploration, moving from Low Earth Orbit (LEO) to Geostationary Earth Orbit (GEO) or further, the communication systems currently implemented will need to be improved to support those missions. One of the bottlenecks is the antennas' size, due to the close relation between antenna gain and dimensions. Hence, a possible solution is to develop inflatable antennas which can be packaged efficiently, occupying a small amount of space, and they can provide, once deployed, large dish dimension and correspondent gain. A prototype of a 1 m inflatable antenna for X-Band has been developed in a joint effort between JPL and ASU. This paper will detail the principle challenges in developing the antenna technology focusing on: design, EM analysis, fabrication and tests.

Cathode Anode Satellite Thruster for Orbital Reposition (CASTOR) is an orbital manoeuvre and transfer micro-satellite bus developed at MIT Space System Laboratory. The technical objective of the mission is achieving 1 km/s of delta-V over a 1 year mission in Low Earth Orbit (LEO). This will be accomplished using a novel electric propulsion system, the Diverging Cusped Field Thruster (DCFT), which enables high efficiency orbital changes of the ESPA-ring class satellite. CASTOR is capable of improving rapid access to space capabilities by providing an orbital transfer platform with a very high performance to mass ratio, thus greatly reducing launch costs and allowing for highly efficient orbital manoeuvre. Furthermore, CASTOR is highly scalable and modular, allowing it to be adapted to a wide range of scales and applications. CASTOR is developed as part of the University Nanosatellite Program (UNP) funded by Air Force Research Laboratory (AFRL). In order to accomplish CASTOR mission objective, a highly optimized, scalable, light weight, and low cost communication system needed to be developed. These constraints imply the development of trade studies to select the final communication system architecture able to maximize the amount of data transmitted, while guaranteeing reliability, redundancy and limited mass, power consumption, and cost. A special attention is also required to guarantee a reliable communication system in cases of tumbling, or in case of strong Doppler shift which is inevitable due to the high delta-V capabilities of the vehicle. In order to accomplish all the mission requirements, different features have been introduced in the design of the communication system for this mission. Specifically, customized patch antennas have been realized, and a customized communication protocol has been designed and implemented. The communication subsystem has been validated through an intense testing campaign which included software tests in the laboratory, hardware tests in anechoic chamber, and in flight tests through a balloon experiment. The article presents an overview of CASTOR mission, a presentation of the trade studies analysis and of the final communication architecture selected, a description of the customized antenna developed, of the customized protocol designed, and a presentation of the results of the tests performed.

For more than a decade, the Aeronautics and Astronautics Department at MIT has offered undergraduate students the opportunity of conceiving, developing, implementing and operating new spacecraft's missions. During a three term class, junior and senior students experience all the challenges of a true engineering team project: design, analysis, testing, technical documentation development, team management, and leadership. Leadership instruction is an important part of the curricula; through the development of leadership skills, students learn to manage themselves and each other in a more effective way, increasing the overall productivity of the team. Also, a strong leadership education is a key factor in improving the abilities of future engineers to be effective team members and leaders in the companies and agencies in which they will work. However, too often leadership instruction is presented in an abstract way, which does not provide students with suggestions for immediate applicability. As a consequence, students underestimate the potential that leadership education can have on the development of their projects. To counteract that effect, a new approach for teaching “practical” leadership has been developed. This approach is composed of a set of activities developed to improve students' leadership skills in the context of a project. Specifically, this approach has been implemented in the MIT satellite development class. In that class, students experienced the challenges of building two satellites: CASTOR and Exoplanet. These two missions are real space projects which will be launched in the next two years, and which involve cooperation with different entities (MIT, NASA, and Draper). Hence, the MIT faculty was interested in developing leadership activities to improve the productivity of the teams in a short time. In fact, one of the key aspects of the approach proposed is that it can be quickly implemented in a single semester, requiring no more than 4 h of activity. Data collected show that the approach improved the ability of students to interact productively with each other. This suggests that the activity can also be used in different contexts where a rapid and effective way of improving leadership and team membership is required. The article presents an overview of MIT satellite development class and of the two missions used as test cases, a detailed description of the leadership approach implemented, and of the results obtained.

The inflatable antenna for CubeSat is a 1 meter antenna reflector designed with one side reflective Mylar, another side clear Mylar with a patch antenna at the focus. The development of this technology responds to the increasing need for more capable communication systems to allow CubeSats to operate autonomously in interplanetary missions. An initial version of the antenna for the S-Band was developed and tested in both anechoic chamber and vacuum chamber. Recent developments in transceivers and amplifiers for CubeSat at X-band motivated the extension from the S-Band to the X-Band. This paper describes the process of extending the design of the antenna to the X-Band focusing on patch antenna redesign, new manufacturing challenges and initial results of experimental tests.

In this paper a communication system for CubeSats in formation to operate in the vicinity of the Lunar Lagrangian L1 is proposed. CubeSats will collect lunar scientific data and will perform surface observations. An improved low complexity CDMA system for CubeSats for communications between the Lunar L1 and Earth station is considered. The complexity of a coded CDMA transmitter is lower than the complexity of the CDMA receiver with decoder therefore for downlink communications it makes sense to use encoders such as space standard LDPC code followed by a spread spectrum transmitter for CDMA systems for CubeSats. For the uplink an uncoded CDMA system is chosen since the uplink transmit power is expected to be high enough to support the use of uncoded CDMA system. The uncoded CDMA yields receivers for CubeSats that have low complexity implementation. For uplink since there would be no multipath the use of orthogonal spreading codes is more appropriate. The choice of orthogonal codes would reduce the multiuser interference at CubeSats. For the downlink, based on the available bandwidth, and the data rates, a reasonable processing gain could be obtained. Thus the multiuser interference degradation due to the other CubeSats could be made small at the Earth station. In this paper we analyzed and simulated the proposed improved CDMA system for a concept Constellation of CubeSats. All system simulations are done using Simulink Matlab platform. For highly efficient nonlinear power amplifiers we use a filtered offset QPSK with phase modulation which is a CCSDS standard for constant envelope signaling. This allows a nonlinear amplifier at CubeSat to operate at saturation point for the highest efficiency. We compare the difference in performance with our current pulse shaped BPSK (using half-sine wave). Certainly if there is no bandwidth limitation (due to the spectral standard masking) we could as well use unfiltered rectangular pulses which would produce constant envelope signaling. But it seems that rectangular pulses will not satisfy the bandwidth limitation imposed by the spectral standard. Filtered offset QPSK with phase modulation is much more bandwidth efficient scheme.

In1 the past, great progress has been made in the development of small satellites and CubeSats, mainly for low Earth orbit. Currently, a new technological trend is the development of technologies and strategies for potential interplanetary applications of small platforms (CubeSats/Satellites). Given the limited size, mass and power capabilities of these small platforms, one of the most interesting problems is how to allow small satellites to communicate from very far distance in the solar system. This paper aims to review and possibly combine two solutions for the problem: the use of inflatable antenna reflectors and the arrays across multiple spacecraft. An overview of the inflatable antenna technology, its development and tests, its applicability in terms of frequencies and sizes, and the advantages with respect to other technologies is described. An overview of cooperative communication techniques across small platforms is presented and the main challenges of arraying antennas on different spacecraft are underlined. Finally, the two solutions are combined to provide a first order quantification of the advantages in terms of EIRP, data rate and range.

12Spacecraft are complex systems that involve different subsystems and multiple relationships among them. The design of a spacecraft is an evolutionary process that starts from requirements and evolves across different design phases. During this process, a lot of changes can happen and affect mass and power at component level, at subsystem level, and even at system level. Each spacecraft has to respect overall constraints in terms of mass and power: for this reason, it's important to be sure that the design does not exceed these limitations. Current practice in system modeling primarily deals with this problem allocating margins. However, a statistical characterization of the fluctuations in mass and power in the early design phases is missing. This lack of statistical characterization would result either in a risky spacecraft design that might not fit the mission constraints and requirements, or in a conservative design that might not fully utilize the available resources. The authors have previously developed a mathematical approach ( [1] [2] [3] [4]) to quantify the likelihood that the major design drivers of mass and power of a space communication system would meet the spacecraft and mission requirements and constraints through the mission design lifecycle. The methodology uses a combination of statistics based on data and probabilities assessed by experts in the field. Additionally, the methodology is innovative in the approach developed to elicit experts' opinions [5] which allows the analyst to identify biases or mis-calibrations which can affect the final estimation. The work presented in this paper extends the approach developed to the entire spacecraft design, focusing on small spacecraft (less than 50 Kg mass) due to the recent interest in the space community in such platforms. The spacecraft is analyzed in its subsystems and related components. A database is constructed to compute data statistics. An interview process to elicit expert opinions in all the areas of spacecraft design is developed and the results are presented. A test case made of a small satellite mission (CASTOR) is presented and the results are described.

CubeSats and small satellites have potential to provide means to explore space and to perform science in a more affordable way. As the goals for these spacecraft become more ambitious in space exploration, moving from Low Earth Orbit (LEO) to Geostationary Earth Orbit (GEO) or further, the communication systems currently implemented will not be able to support those missions. One of the bottlenecks is the antennas’ size, due to the close relation between antenna gain and dimensions. Current antennas for CubeSats are mostly dipole or patch antennas with limited gain. Deployable (not inflatable) antennas for CubeSats are currently being investigated, but these solutions are affected by the challenge of packaging the whole deployable structure in a small spacecraft. The work that we propose represents the first attempt to develop an inflatable antenna for CubeSats. Inflatable structures and antennas can be packaged efficiently, occupying a small amount of space, and they can provide, once deployed, large dish dimension and correspondent gain. Inflatable antennas have been previously tested in space (Inflatable Antenna Experiment, STS-77). However they have never been developed for small spacecraft such as CubeSats, where the packaging efficiency, the deployment, and the inflation represent a challenge. Previous works developed by the authors described trade-off analysis, preliminary design and radiation model for the antenna. The research presented in this paper is focused specifically on implementation and testing. Details of the antenna’s fabrication and related issues are illustrated as well as the mechanism to fold and deploy the antenna in space. Finally, results of the experimental tests (vacuum chamber and anechoic chamber) are described. Future work in the development of the antenna will include the improvement of the fabrication process and the design of a 3U CubeSat mission to be proposed as a technical demonstration.

Interplanetary Cubesats would enable low-cost missions for high-quality scientific and exploration programs. In particular cubeSats in formation have been proposed to operate in the vicinity of the Lunar Lagrangian L1 to collect lunar scientific data and to perform surface observation. In this paper we present a low complexity CDMA system for CubeSats (M small spacecraft) for communications between the Lunar L1 and Earth station. It is well known that the complexity of a CDMA transmitter is much lower than the complexity of the CDMA receiver. Moreover, the complexity of a channel encoder is always much lower than the complexity of the channel decoder. So for downlink communications it makes sense to use encoders for modern codes such as Turbo and LDPC followed by a spread spectrum transmitter for CDMA systems for CubeSats. Here we used an LDPC coded CDMA with BPSK modulation with rectangular and half-sine pulse shaping. Except for the PN generator seed numbers, the communication structure of all CubeSats would be identical and operating at one single RF frequency. For the uplink we may choose an uncoded CDMA system since the uplink transmit power is expected to be high enough to support the use of uncoded CDMA system. In addition since there would be no multipath for the uplink (broadcast channel) the use of orthogonal spreading codes such as Walsh codes is appropriate. The choice of orthogonal codes would reduce the multiuser interference. However due to some limitation (bandwidth, data rates, and M) we may be forced to use nonorthogonal PN codes. In addition, one of the spreading codes will not carry any data, which acts as an unmodulated pilot to reduce the complexity of synchronization. The proposed uncoded CDMA yields receivers for CubeSats that have low complexity implementation. Each component of CubeSats could easily extract its own received data with almost no interference from other users in case of orthogonal spreading codes. For the downlink, depending on the available bandwidth, and the data rates, a large processing gain could be obtained if the N is not large. Thus the multiuser interference degradation due to the other CubeSats could be made small at the Earth station. If N is large, and the bandwidth and data rates do not allow large processing gains then the multiuser interference could be high. In such cases we could use a simple parallel interference cancellation method with two stages that dramatically improves the system performance for the downlink. In this paper we accurately analyzed and simulated the proposed CDMA system for a concept Constellation of 20 CubeSats (M=20). All system simulations are done using Simulink platform.

The design of a spacecraft is an evolutionary process that starts from requirements and evolves over time across different design phases. During this process, a lot of changes can happen. They can affect mass and power at component level, at subsystem level, and even at system level. Each spacecraft has to meet overall constraints in terms of mass and power: for this reason, it's important to be sure that the design does not exceed these limitations. Current practice in system modeling deals with this problem by allocating margins on single components and on each of the subsystems. However, a statistical characterization of these fluctuations in mass and power is missing, and the consequence is a design that either is too risky and does not fit the mission constraints, or is too conservative and generates an inefficient utilization of resources.

CubeSats and small satellites are becoming a way to explore space and to perform science more affordably. As the goals for these spacecraft become more ambitious in terms of physical distance (moving from Low Earth Orbit (LEO) to Geostationary Earth Orbit (GEO) or further), and of the amount of data to relay back to Earth (from Kbits to Mbits), the communication systems currently implemented will not be able to fully support those missions. Two of the possible strategies to solve this issue are the following: 1. Increasing the time available to communicate to the ground by using existing satellite networks as communication relays. 2. Equipping the satellite with a more powerful antenna compatible with the volume and mass constraints imposed by CubeSats and small satellites.

We propose to simulate the geostationary orbit (GEO) with a “belt” of satellites, angularly and regularly spaced, orbiting in the same lower equatorial orbit, in such a number that a ground station does not need to change antenna pointing, acting therefore as if it were connected to a geostationary satellite. The equivalent GEO is accomplished by using a “distributed” geostationary satellite (D–GEO). The belt can be used, cooperatively, by any ground station located in a large equatorial area around the globe, thus minimizing system costs. The complexity of a D–GEO system may be worth its costs because it multiplies what is naturally a unique physical resource, such as the geostationary orbit. We set the main parameters for comparing D–GEO and GEO satellite systems, serving approximately the same global area centered at the equator, and apply them to few examples.

MIT is developing an Earth-based hopper prototype for autonomous planetary exploration. This paper describes the initial development of the hopper prototype and its application as a testbed for lunar hopper technologies. The prototype hopper consists of a gravity-canceling system, composed of battery-powered electric ducted fans, and an integrated lunar hopper system, which operates using cold gas propulsion. The gravitycanceling system is designed to allow the lunar hopper system to operate in an environment which is dynamically similar to the surface of the Moon, although it also allows for recovery and safe touchdown if the cold gas system experiences a failure. The environment provided by the electric ducted fans applies the fan thrust against Earth’s gravity to obtain a desired upward acceleration, allowing the cold gas jets to operate as if they were in a fractionalgravity environment. The key design features of the integrated gravity-canceling and lunar hopping systems are presented in this paper.