Within the Mars science community, there is growing interest in the role that small spacecraft missions can play in increasing the breadth and frequency of Mars exploration. Fueled by significant advances in miniaturization of science instruments and avionics, innovative concepts for delivery of small spacecraft to Mars, and emerging low-cost capabilities in the fast-growing NewSpace marketplace, small spacecraft missions appear capable of achieving compelling Mars science results at unprecedentedly low mission cost, far below the current NASA Discovery Program cost cap. We report on the development of a cost model for small spacecraft Mars orbiters, providing insight into the dependence of mission cost on key mission parameters, including science payload characteristics (e.g., mass, power), mission design parameters (e.g., ΔV), and mission risk categorization. These results can help guide NASA and others in establishing budget guidelines for a new class of low-cost, small spacecraft missions.

Mars Sample Return (MSR) continues to be a high priority in the planetary science community and a decades-long goal of international planetary exploration programs. Options for architectures and mission concepts are currently under study by NASA and ESA to find potential partnership opportunities to achieve MSR in the 2020s. The major elements of a potential MSR campaign have significant architectural flexibility and mission launch, arrival, and return options. The decision criteria often depend on mission design and functional allocations across many elements. This paper outlines the reference architecture and key trades among the campaign elements.

A Mars Sample Return (MSR) campaign would involve a series of three flight missions to acquire and cache Mars samples, retrieve those samples and launch them into Mars orbit, and then capture these samples and return them to Earth. Relay communications would be crucial for supporting this campaign, characterized by multiple critical events, complex surface operations, and an on-orbit Mars rendezvous. The existing Mars relay network offers significant capability, and efforts are underway to maximize the likelihood that one or more of these current assets will still be operational in the timeframe of an MSR campaign. In addition, the Earth Return Orbiter (ERO) element of a campaign could serve as a primary relay asset, if it can achieve a useful relay orbit by the time of arrival of the Sample Retrieval Lander mission. We describe key operational challenges of the MSR campaign that would drive the required relay capabilities, and characterize the performance of the existing relay orbiters as well as ERO itself in meeting those relay needs.

The interface between a spacecraft and its ground operations segment includes the flow of commands, configuration, and sequencing elements to the spacecraft, and the flow of telemetry and data products from the spacecraft. Creating and implementing a complete definition of this interface simplifies and standardizes mission operations, allowing easy sharing of operations personnel across missions. Early spacecraft featured a simple flight / ground interface (FGI) using hardware command decoding in the radio, driven by technological limitations of the time. Modern spacecraft use command and data handling (CDH) avionics on which flight software executes, which in turn controls and configures the mission, executes subsystem and instrument instructions, and implements critical fault protection actions. Deep space missions feature advanced operations software for running sequenced activities over a period of weeks, which allows them to function with only infrequent ground contact. This approach comes at the cost of increased complexity in the FGI, requiring expensive modifications to heritage flight software and ground systems. By hosting the interface in the radio instead of the CDH avionics, modern missions can approximate the FGI design simplicity of early spacecraft, with significant advantages for vendor competition, lowered costs, standardization of operations, and reduction of implementation risk.

Mission Timeline Control Scheme • Jettison timing controls amount of aerocapture ΔV • Guidance algorithm: numerical predictor-corrector • Only atmospheric accelerometer measurements required • Monte Carlo results show robustness to uncertainties in entry FPA, atmosphere, and burns Trajectory Modeling • End-to-end simulation models vehicle trajectory and uncertainties • 3-DOF integration of atmospheric equations of motion • Key outputs: • Deceleration • Heat Rates • Orbit uncertainties

Current and near-term Mars relay telecommunications services are provided by a set of NASA and ESA Mars science orbiters equipped with UHF relay communication payloads employing operationally simple low-gain antennas. These have been extremely successful in supporting a series of landed Mars mission, greatly increasing data return relative to direct-to-Earth lander links. Yet their relay services are fundamentally constrained by the short contact times available from the selected science orbits. Future Mars areostationary orbiters, flying in circular, equatorial, 1-sol orbits, offer the potential for continuous coverage of Mars landers and rovers, radically changing the relay support paradigm. Achieving high rates on the longer slant ranges to areostationary altitude will require steered, high-gain links. Both RF and optical options exist for achieving data rates in excess of 100 Mb/s. Several point designs offer a measure of potential user burden, in terms of mass, volume, power, and pointing requirements for user relay payloads, as a function of desired proximity link performance.

Current Mars science orbiters carry UHF proximity payloads to provide limited access and data services to landers and rovers on Mars surface. In the era of human spaceflight to Mars, very high rate and reliable relay services will be needed to serve a large number of supporting vehicles, habitats, and orbiters, as well as astronaut EVAs. These will likely be provided by a robust network of orbiting assets in very high orbits, such as areostationary orbits. In the decade leading to that era, telecommunications orbiters can be operated at areostationary orbit that can support a significant population of robotic precursor missions and build the network capabilities needed for the human spaceflight era. Telecommunications orbiters of modest size and cost, delivered by Solar Electric Propulsion to areostationary orbit, can provide continuous access at very high data rates to users on the surface and in Mars orbit.

NASA's Space Technology Mission Directorate has been recently developing critical technologies for high-power solar electric propulsion (SEP), including large deployable solar array structures and high-power electric propulsion components. An ion propulsion system based on these developments has been considered for many SEP technology demonstration missions, including the Asteroid Redirect Robotic Mission (ARRM) concept. These studies and the high-power SEP technology developments have generated excitement within NASA about the use of the ARRM ion propulsion system design for other types of potential missions. One application of interest is for Mars missions, especially with the types of orbiters now under consideration for flights in the early 2020's to replace the aging Mars Reconnaissance Orbiter. High-power SEP can deliver large payloads to Mars with many additional capabilities, including large orbital plane changes and round-trip missions, compared to chemically-propelled spacecraft. Another application for high-power SEP is for exo-planet observation missions, where a large starshade spacecraft would need to be repositioned with respect to its companion telescope relatively frequently and rapidly. SEP is an enabling technology for the ambitious science goals of these types of missions. This paper will discuss the benefits of high-power SEP for these concepts based on the STMD technologies now under development.

Comet C/2013 A1 (Siding Spring) will have a close encounter with Mars on 2014 October 19. We model the dynamical evolution of dust grains from the time of their ejection from the comet nucleus to the close encounter with Mars, and determine the flux at Mars. Constraints on the ejection velocity from Hubble Space Telescope observations indicate that the bulk of the grains will likely miss Mars, although it is possible that a few percent of the grains with higher velocities will reach Mars, peaking approximately 90-100 minutes after the close approach of the nucleus, and consisting mostly of millimeter-radius grains ejected from the comet nucleus at a heliocentric distance of approximately 9 AU or larger. At higher velocities, younger grains from submillimeter to several millimeters can also reach Mars, although an even smaller fraction of grains is expected have these velocities, with negligible effect on the peak timing. Using NEOWISE observations of the comet, we can estimate that the maximum fluence will be of the order of 10(-7) grains m(-2). We include a detailed analysis of how the expected fluence depends on the grain density, ejection velocity, and size frequency distribution, to account for current model uncertainties and in preparation of possible refined model values in the near future.

Mars Sample Return (MSR) missions could benefit from the high specific impulse of Solar Electric Propulsion (SEP) to achieve lower launch masses than with chemical propulsion. SEP presents formulation challenges due to the coupled nature of launch vehicle performance, propulsion system, power system, and mission timeline. This paper describes a SEP orbiter-sizing tool, which models spacecraft mass & timeline in conjunction with low thrust round-trip Earth-Mars trajectories, and presents selected concept designs. A variety of system designs are possible for SEP MSR orbiters, with large dry mass allocations, similar round-trip durations to chemical orbiters, and reduced design variability between opportunities.

Since the selection of the proposed Mars 2020 mission as a Rover with the capability of sample collection and caching, there has been renewed interest in subsequent mission concepts to return Mars samples to Earth. The general architecture for this series of missions is outlined in the Planetary Science Decadal Survey of 2011. The role of the Sample Return Orbiter (SRO) in The 2011 Decadal Survey MSR architecture was to collect an orbiting sample (OS) from low Mars orbit and deliver it to Earth's surface. The architecture focused on chemical propulsion orbiters with ballistic and aerobraking trajectories that were dedicated entirely to the capture of orbiting samples and returning them to the surface of the Earth. Recent concepts have explored the use of Solar Electric Propulsion (SEP) to Mars and for the return to Earth. SEP could enable significant mission flexibility which includes: lower launch mass or increased mass delivery capability to Mars orbit and return to Earth; longer launch periods for both launch and Earth return; consistency of design across launch opportunities; access to both high and low Mars orbit altitudes; increased on-orbit ΔV budgets for orbit changes and sample rendezvous; and greater control over Earth arrival speed and geometry. With this flexibility come opportunities to: save launch cost; add functions such as remote sensing observations, secondary payload deployment, and relay telecommunications; and choose between direct return of Mars samples to the Earth's biosphere or capturing them to a stable long-term orbit around the Earth. This paper compares the previous SRO chemical-ballistic concepts with the recent SEP orbiter concepts. We will show the potential benefits gained by the inherent flexibility of SEP as applied to launch mass, launch periods, Earth return opportunities, on-orbit ΔV and other architectural drivers.

The JPL Rapid Mission Architecture (RMA) capability is a novel collaborative team-based approach to generate new mission architectures, explore broad trade space options, and conduct architecture-level analyses. RMA studies address feasibility and identify best candidates to proceed to further detailed design studies. Development of RMA first began at JPL in 2007 and has evolved to address the need for rapid, effective early mission architectural development and trade space exploration as a precursor to traditional point design evaluations. The RMA approach integrates a small team of architecture-level experts (typically 6-10 people) to generate and explore a wide-ranging trade space of mission architectures driven by the mission science (or technology) objectives. Group brainstorming and trade space analyses are conducted at a higher level of assessment across multiple mission architectures and systems to enable rapid assessment of a set of diverse, innovative concepts. This paper describes the overall JPL RMA team, process, and high-level approach. Some illustrative results from previous JPL RMA studies are discussed.

Jupiter Europa Orbiter (JEO), the proposed NASA element of the proposed joint NASA-ESA Europa Jupiter System Mission (EJSM), could launch in February 2020 and conceivably arrive at Jupiter in December of 2025. The concept is to perform a multi-year study of Europa and the Jupiter system, including 30 months of Jupiter system science and a comprehensive Europa orbit phase of 9 months. This paper provides an overview of the JEO concept and describes the Europa Science phase orbit design and the related science priorities, model pay-load and operations scenarios needed to conduct the Europa Science phase. This overview is for planning and discussion purposes only.

The Europa Jupiter System Mission (EJSM) has been prioritized as the next Outer Planets Flagship Mission that would be devoted to exploring the emergence of habitable worlds around gas giants. This joint NASA and ESA endeavor would focus on the Galilean moons Europa and Ganymede but would also investigate Io, Callisto, and the Jupiter system as a whole. The NASA-contributed Jupiter Europa Orbiter (JEO) and the ESA-contributed Jupiter Ganymede Orbiter (JGO) would be launched on separate launch vehicles in 2020. Here we focus on JEO. After 2-3 years of performing science in the Jovian system, JEO would orbit Europa and would operate in a high radiation environment. The life-limiting radiation environment complicates hardware and software performance as well as operations strategy. These challenges would require the Europa science goals to be met in an efficient duration of 9 months. To maximize the science while in orbit at Europa, JEO needs to develop strategies to make the system easily operable and robust to radiation degradation. In addition to radiation tolerant hardware and software designs along with robust margins, the project has already been considering operability features in its baseline design for JEO. Operability is the combination of aspects of a system that make it simple and inexpensive to operate, robust to changing system behavior, responsive to modified goals, and adaptive to deviations in expected environments or operating conditions. For JEO, being robust and flexible to meet science goals in the face of a harsh environment is paramount. The mission must be designed to have multiple means of meeting critical science goals; accommodate operation in the face of radiation-based noise, degradation and failure; and be flexible to changing science goals based upon discoveries. As a result, JEO has kept science goals and operability in mind for all system design processes and trade studies. The most stringent and driving operational requirements and con- - straints for the JEO concept are encountered during Europa Science orbit phase. Some of the key operability issues incorporated from the earliest concept studies include: (1) Make the flight and ground systems operable and survivable for a high intensity, rapid turn-around operations environment in Europa orbit in the presence of radiation based anomalies. (2) Use modern system engineering methods to model the system behavior as early as possible to balance mission scope with system capability, complexity, risk, and cost. Systems design based on behavior models allows for accommodation of changing behaviors in operating the systems. (3) Use lessons learned from previous applicable missions to guide design philosophy and trade studies specifically for operability issues. In this paper, we discuss the key flight system design trades, operations scenarios and lessons learned developed in recent JEO mission studies.

In December 2007, NASA’s Planetary Science Division announced its intent to conduct Phase-2 studies for the next Outer Planets Flagship Mission (OPFM). One of the several tasks identified in the OPFM Statement of Work from NASA to the Jet Propulsion Laboratory (JPL) was to perform a Mission Operations Lessons Learned Study (referred to here on as the LL Study) with special focus on Phase E cost drivers and operations. The intent of the LL Study was to safely lower Phase E operations costs from those traditional to this class of mission. Consequently JPL requested its OPFM partner, the Johns Hopkins University Applied Physics Laboratory (JHU/APL), to lead the LL Study.

Missions to explore Europa have been imagined ever since the Voyager mission first suggested that Europa was geologically very young. Subsequently, Galileo supplied fascinating new insights into that satellite's secrets. The Jupiter Europa Orbiter (JEO) would be the NASA-led portion of the Europa Jupiter System Mission (EJSM), an international mission with orbiters developed by NASA, ESA and possibly JAXA. JEO would address a very important subset of the complete EJSM science objectives and is designed to function alone or in conjunction with ESA's Jupiter Ganymede Orbiter (JGO).

In 2007 a JPL Rapid Mission Architecture (RMA) analysis team identified and evaluated a broad set of mission architecture options for a suite of scientific exploration objectives targeting the Saturnian moon Enceladus. Primary science objectives were largely focused on examination of the driving mechanisms and extent of interactions by the plumes of Enceladus recently discovered by Cassini mission science teams. Investigation of the architectural trade space spanned a wide range of options, from high-energy flybys of Enceladus as a re-instrumented expansion on the Cassini mission, to more complex, multi-element combinations of Enceladus orbiters carrying multiple variants of in-situ deployable systems. Trajectory design emerged as a critical element of the mission concepts, enabling challenging missions on Atlas V and Delta IV-Heavy class launch vehicles. Various Enceladus Flagship-class mission concepts identified were analyzed and compared against several first-order figures of merit, including mass, cost, risk, mission timeline, and associated science value with respect to accomplishment of the full set of science objectives. Results are presented for these comparative analyses and the characterization of the explored trade space.