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Electric propulsion technology is widely used today, and multiple thrusters exist for primary electric propulsion application. NASA and the U.S. commercial market developed several thrusters suitable for primary electric propulsion on full scale spacecraft. The

The main goal of this paper is to investigate feasibility of using all-electric propulsion system for a mid-light business jet aircraft in the near future (20–30 years from now). The secondary goal is to assess the impact of using such system on operating costs and emission reduction. This paper presents calculations of business jet aircraft mission energy demands and compares them with batteries capabilities. Three different types of lithium batteries are investigated in terms of their energy densities projected for three different time frames. Mass of batteries that is required to provide demanded amount of energy to perform the mission is compared with the maximum mass of fuel that the baseline aircraft is able to take. On this basis, the feasibility of all-electric propulsion system is assessed. Additionally, in order to show the limitations of such system, maximum range is calculated for the mass of batteries that would potentially enable to perform the flight. Furthermore, CO2 and NOx emission of the baseline aircraft engines are compared with the amount of gaseous pollutants which are emitted by the power plant, when energy needed to recharge batteries is being produced. Finally, the potential fuel cost reduction is calculated based on the cost of electricity that would be used to recharge batteries.

At present, very few CubeSats have flown in space featuring propulsion systems. Of those that have, the literature is scattered, published in a variety of formats (conference proceedings, contractor websites, technical notes, and journal articles), and often not available for public release. This paper seeks to collect the relevant publically releasable information in one location. To date, only two missions have featured propulsion systems as part of the technology demonstration. The IMPACT mission from the Aerospace Corporation launched several electrospray thrusters from Massachusetts Institute of Technology, and BricSAT-P from the United States Naval Academy had four micro-Cathode Arc Thrusters from George Washington University. Other than these two missions, propulsion on CubeSats has been used only for attitude control and reaction wheel desaturation via cold gas propulsion systems. As the desired capability of CubeSats increases, and more complex missions are planned, propulsion is required to accomplish the science and engineering objectives. This survey includes propulsion systems that have been designed specifically for the CubeSat platform and systems that fit within CubeSat constraints but were developed for other platforms. Throughout the survey, discussion of flight heritage and results of the mission are included where publicly released information and data have been made available. Major categories of propulsion systems that are in this survey are solar sails, cold gas propulsion, electric propulsion, and chemical propulsion systems. Only systems that have been tested in a laboratory or with some flight history are included.

The growing competitiveness in the commercial space market has raised the interest in operating small spacecraft at very low altitudes. To make this feasible, the space industry has started developing propulsion options tailored specifically to these platforms. This paper presents a review of emerging micropropulsion technologies and evaluates their applicability to microsatellite missions in the altitude range 250–500 km. The results of the proposed analysis are demonstrated on two different remote sensing applications.

The advances in solar technology and electric propulsion now offer the promise of new, very capable space transport systems that will allow us to effectively explore the solar system. NASA has developed many concepts of space-powered spacecraft with power levels ranging from tens to hundreds of kilowatts for robotic and asteroid-driven missions and Mars. This paper describes two electrical/chemical propulsion concepts developed over the last 5 years and discusses how they could be used to humanize the solar system. A possible sustainable solution for the supply of spacecraft power would be to achieve and utilization nuclear fusion energy. The paper briefly presents some contributions to obtaining nuclear fusion energy as a viable alternative to current energies. For the energy of spacecraft of the future, the combination of photovoltaic energy (obtained from the stars) and that produced by a nuclear reactor on fusion is essential. NASA is developing a strategy to send a crew to Mars by 2030. To achieve this goal, NASA plans to develop the technology for long-haul flights including advanced transportation work and living systems. Among these technologies, Solar Electric Propulsion (PES) has been identified as very effective in moving large masses through interplanetary space. For decades it has been known that missions outside the low Earth's orbit can be made cost-effective by PSA, but yet such space missions have not yet been done because the manufacturing technology is not advanced enough. NASA's recent investments in solar power systems and propulsion systems have now matured so that the 50 kW PSA is already ready to be put on flight missions. It has been demonstrated analytically that these technologies can be resized to systems with the power of several hundred kilowatts.

A prediction of future directions and trends of the satellite propulsion market could be helpful in the formation of technology development agendas proposed by various global entities such as commercial companies, space agencies, or research institutions. Possible market evolutions are presented, in light of past and present technology development, and the authors' estimation of the upcoming development tracks of the various forms and subclasses of electric propulsion (EP). The history of EP is reviewed in the context of a technology life cycle, with the conclusion that EP is in its early majority pragmatist phase. Specific applications for in-space propulsion, including geostationary communication satellites, low Earth orbit mega constellations, CubeSats, interplanetary missions, and Earth observation satellites, are paired with the appropriate corresponding EP technologies. Rising developers and major influences are noted. It is predicted that EP technologies will likely fragment and diversify ...

The paper considers the problem of equipping a small spacecraft (SS) of the AIST-2 series with electric propulsion for orbital plane maneuvers, such as station keeping and attitude control. Solution of this problem will increase the satellite's operational service life on a low-Earth orbit and, in future, allow attitude control of a satellite constellation.

A spacecraft propulsion is any method used to accelerate spacecraft and artificial satellites. There are several different methods, each with advantages and disadvantages, spacecraft propulsion being an active area of research. However, most current spacecraft are propelled by forcing a gas exits through the rear of the vehicle at high speed through supersonic nozzle of a rocket engine. All spacecraft are using chemical rockets (fuel or solid fuel) to launch, although some (such as Pegasus missiles and Space Ship One) are using air jet engines in the first step. Most satellites have simple chemical thrusters (often missiles mono) or missiles resistojet to maintain orbit. Soviet bloc satellites have used electric propulsion for decades and the new Western geo-orbital spacecraft begin using electric propulsion for orbit maintenance of north-south. There is a need increasingly of more new propulsion systems, modern, technology-based.

The paper is devoted to the design features of the prospective Russian INTERHELIO-PROBE spacecraft using, depending on the configuration version, an electric or chemical propulsion system as a sustainer. The scientific goal of the mission is the study of near-solar space from close distances (60–70 solar radii). The paper presents the description of several versions of the spacecraft options depending on the installed propulsion system, as well as the main characteristics of the flight profile depending on the engine type.

Aspects of the design of small spacecraft with electric propulsion power plants for investigating minor bodies in the Solar System are examined. The results of design and ballistic analysis of transfer into an orbit of terrestrial asteroids using electric propulsion thrusters are given. The possible concept design of the spacecraft is determined and the structure of a small spacecraft with an electric propulsion power plant is presented. Parameters of the electric propulsion power plant of a small spacecraft for a flight to the minor bodies of the Solar System are estimated.

Outer planet exploration missions require significant propulsive capability, particularly to achieve orbit insertion. Missions to explore the moons of outer planets place even more demanding requirements on propulsion systems, since they involve multiple large Δ V maneuvers. Hybrid rockets present a favorable alternative to conventional propulsion systems for many of these missions. They typically enjoy higher specific impulse than solids, can be throttled, stopped/restarted, and have more flexibility in their packaging configuration. Hybrids are more compact and easier to throttle than liquids and have similar performance levels. In order to investigate the suitability of these propulsion systems for exploration missions, this paper presents novel hybrid motor designs for two interplanetary missions. Hybrid propulsion systems for missions to Europa and Uranus are presented and compared to conventional in-space propulsion systems. The hybrid motor design for each of these missions is optimized across a range of parameters, including propellant selection, O / F ratio, nozzle area ratio, and chamber pressure. Details of the design process are described in order to provide guidance for researchers wishing to evaluate hybrid rocket motor designs for other missions and applications.

NASA is actively funding research into advanced, unconventional aircraft and engine architectures to achieve drastic reductions in vehicle fuel burn, noise, and emissions. One such concept is being explored by The Boeing Company, the General Electric Company, Virginia Polytechnic Institute and State University, and the Georgia Institute of Technology under the Subsonic Ultra Green Aircraft Research Project. A major cornerstone of this research is evaluating the potential performance benefits that can be attributed to using hybrid-electric propulsion. Hybrid-electric propulsion in this context involves a non-Brayton power generation or storage source, such as a battery or a fuel cell that can be used to provide additional propulsive energy to a conventional Brayton-cycle-powered turbofan engine. This research constructs an integrated Numerical Propulsion System Simulation hybrid-electric propulsion model capable of predicting hybrid-electric engine performance throughout the operational envelope. The syste...

Spacecraft propulsion system is a kind of rocket engine that has been developed from the end of 1950s for attitude control and orbit maintenance of satellite. Since the spacecraft propulsion system has to be used for a relatively long time, therefore, stability of propellant and life of thruster could be very important factor for propulsion system design. Recently, green propellant propulsion and all electrical propulsion system have became very important issue, and we also need a development according to well organized plan. In this paper, we will introduce the development status, key technologies and development prospect of spacecraft propulsion system.

The aerospace industry is recently expressing a growing interest in green, safe and non-toxic propellants for the propulsion systems of the new generation of space vehicles, which is especially true in the case of Cubesat micro-propulsion systems. Demanding requirements are associated to the future missions and challenges offered by this class of spacecraft, where the availability of a propulsion system might open new possibilities for a wide range of applications including orbital maintenance and transfer, formation flying and attitude control. To accomplish these requirements, Delft University of Technology is currently developing two different concepts of water-propelled micro-thrusters based on MEMS technologies: a free molecular micro-resistojet operating with sublimating solid water (ice) at low plenum gas pressure of less than 600 Pa, and a more conventional micro-resistojet operating with liquid water heated and vaporized by means of a custom designed silicon heating chamber. In this status review paper, the current design and future expected developments of the two micro-propulsion concepts is presented and discussed, together with an initial analysis of the expected performance and potential operational issues. Results of numerical simulations conducted to optimize the design of the heating and expansion slots, as well as a detailed description of the manufacturing steps for the conventional micro-resistojet concept, are presented. Some intended steps for future research activities, including options for thrust intensity and direction control, are briefly introduced.

Liquid hydrogen (LH2) and liquid oxygen (LO2) cryogenic propellants can dramatically enhance NASA’s ability to explore the solar system due to their superior specific impulse ( I sp ) capability. Although these cryogenic propellants can be challenging to manage and store, they allow significant mass advantages over traditional hypergolic propulsion systems and are therefore enabling for many planetary science missions. New cryogenic storage techniques such as subcooling and the use of advanced insulation and low thermal conductivity support structures will allow for the long term storage and use of cryogenic propellants for solar system exploration and hence allow NASA to deliver more payloads to targets of interest, launch on smaller and less expensive launch vehicles, or both. These new cryogenic storage technologies were implemented in a design study for the Titan Orbiter Polar Surveyor (TOPS) mission, with LH2 and LO2 as propellants, and the resulting spacecraft design was able to achieve a 43% launch mass reduction over a TOPS mission, that utilized a traditional hypergolic propulsion system with mono-methyl hydrazine (MMH) and nitrogen tetroxide (NTO) propellants. This paper describes the cryogenic propellant storage design for the TOPS mission and demonstrates how these cryogenic propellants are stored passively for a decade-long Titan mission that requires the cryogenics propellants to be stored for 8.5 years.

While electric propulsion for warships has existed for nearly a century, it has only been since the end of the Cold War that modern integrated power systems have been developed and implemented on U.S. Navy warships. The principal enablers have been the products of research and development for rotating machines (generators and propulsion motors), power electronics (power conversion and motor drives), energy storage, and controls. The U.S. Navy has implemented this advanced technology incrementally. Notably, DDG 1000 with its integrated propulsion system and CVN 78 with its electromagnetic aircraft launch system will soon join the fleet and mark another important advance to the electric warship. In the future, the integration of electric weapons such as railguns, high power radars, and lasers will result in the final achievement of the electric warship.

It is nowadays clear that future space access and activity sustainability is greatly endangered by the large amount of space debris populating the near-earth region. In the last few years, a number of different active debris removal (ADR) methods have been proposed and each of them may represent a valuable solution for space debris belonging to specific classes or types or orbiting in particular space regions. Regardless of the method identified as the most suitable, an ADR mission scenario can be thought as composed of different phases in which a deorbiting platform is in charge of approaching a target debris, bringing it to a lower altitude orbit and, in the case of a multiple-target mission, releasing it and chasing a second one. Considering the high total impulse typical of these kinds of missions, electric propulsion (EP) plays a key role in reducing the propellant mass consumption required for each maneuver and thus increasing the mass available to deorbit a relevant number of debris per mission. A low-power and low-cost EP system based on the Alta HT-100 Hall effect thruster is here considered with the aim of highlighting the advantages offered by this system for these kinds of missions.

This article offers a summary of past efforts in the development of Nuclear Thermal Propulsion systems for space transportation. First, the generic principle of thermal propulsion is outlined: a propellant is directly heated by a power source prior to being expanded which creates a thrusting force on the rocket. This enables deriving a motivation for the use of Nuclear Thermal Propulsion (NTP) relying on nuclear power sources. Then, a summary of major families of NTP systems is established on the basis of a literature survey. These families are distinguished by the nature of their power source, the most important being systems with radioisotope, fission, and fusion cores. Concepts proposing to harness the annihilation of matter and anti-matter are only touched briefly due to their limited maturity. For each family, an overview of physical fundamentals, technical concepts, and – if available – tested engines' propulsion parameters is given.

One of the key components for future space exploration, manned or unmanned, is the availability of propulsion systems beyond the state of the art. The rapid development in conventional propulsion systems since the middle of the 20th century has already reached the limits of chemical propulsion technology. To enhance mission radius, shorten the transit time and also extend the lifetime of a spacecraft more efficient, but still powerful propulsion system must be developed. Apart from the propulsion system a major weight contribution arises from the required energy source. Envisioning rapid development of future high average power laser systems and especially the ICAN project we review the prospect of advanced space propulsion based on laser systems.

Although the harsh space environment imposes many severe challenges to space pioneers, space exploration is a realistic and profitable goal for long-term humanity survival. One of the viable and promising options to overcome the harsh environment of space is nuclear propulsion. Particularly, the Nuclear Thermal Rocket (NTR) is a leading candidate for near-term human missions to Mars and beyond due to its relatively high thrust and efficiency. Traditional NTR designs use typically high power reactors with fast or epithermal neutron spectrums to simplify core design and to maximize thrust. In parallel there are a series of new NTR designs with lower thrust and higher efficiency, designed to enhance mission versatility and safety through the use of redundant engines (when used in a clustered engine arrangement) for future commercialization. This paper proposes a new NTR design of the second design philosophy, Korea Advanced NUclear Thermal Engine Rocket (KANUTER), for future space applications. The KANUTER consists of an Extremely High Temperature Gas cooled Reactor (EHTGR) utilizing hydrogen propellant, a propulsion system, and an optional electricity generation system to provide propulsion as well as electricity generation. The innovatively small engine has the characteristics of high efficiency, being compact and lightweight, and bimodal capability. The notable characteristics result from the moderated EHTGR design, uniquely utilizing the integrated fuel element with an ultra heat-resistant carbide fuel, an efficient metal hydride moderator, protectively cooling channels and an individual pressure tube in an all-in-one package. The EHTGR can be bimodally operated in a propulsion mode of 100 MWth and an electricity generation mode of 100 kWth, equipped with a dynamic energy conversion system. To investigate the design features of the new reactor and to estimate referential engine performance, a preliminary design study in terms of neutronics and thermohydraulics was carried out. The result indicates that the innovative design has great potential for high propellant efficiency and thrust-to-weight of engine ratio, compared with the existing NTR designs. However, the build-up of fission products in fuel has a significant impact on the bimodal operation of the moderated reactor such as xenon-induced dead time. This issue can be overcome by building in excess reactivity and control margin for the reactor design.

Although it is clear that electric propulsion can deliver more payload mass when compared to conventional chemical propulsion, near-term transfers to geostationary-equatorial orbit will likely use both propulsion modes in order to reduce the transfer time. Determining the best starting orbit for the subsequent electric-propulsion phase is the key to computing time-constrained maximum-payload transfers to geostationary orbit. Numerical optimization methods are used to determine the unique optimal starting orbit for a given low-thrust velocity increment ΔV to be performed by the electric-propulsion stage. This approach yields a purely analytical algorithm that can determine spacecraft mass requirements for a desired electric propulsion system and desired transfer time. Numerical examples are presented in order to demonstrate how this tool can be used to rapidly perform trade studies for transfers that use chemical and electric propulsion stages.

Electric propulsion is a spacecraft propulsion technology offering higher exhaust velocities and higher thrust efficiency than classical chemical rockets. Such technology allows for high total impulse missions with reduced propellant mass consumption, but it requires a power generation system that becomes heavier and heavier with thrust magnitude. As a result, the design of an electric propulsion mission cannot ignore the spacecraft mass breakdown because it might happen that power generation system, propellant, and other spacecraft subsystems result in a too low or even null payload mass fraction. The paper presents a semianalytic optimization of the design of an electric powered spacecraft identifying those missions more suited for a given thruster and/or power generation system technology. The relations presented, moreover, can also be used to identify the best thruster for delivering a given payload mass into a target orbit in a prescribed time. A basic spacecraft mass breakdown and analytic closed form approximations are used to identify spacecraft optimal configurations intended as the best combinations of power level, thruster characteristics, and thrust duration resulting in the maximum payload mass fraction achievable for a given electric propulsion mission.

Technological aspects and market-wise applications of electric-powered UAVs are emerging. UAVs proved to be a candidate to be an indispensable tool in our life. In this chapter, applications and future trends of UAVs are discussed. Electric propulsion of UAVs seems to be a key enabler for the most of the applications. It can be concluded that the more autonomous and lighter the UAV is, the more electric power is preferred.

Developing a capability for the assembly of large space structures has the potential to increase the capabilities and performance of future space missions and spacecraft while reducing their cost. One such application is a megawatt-class solar electric propulsion (SEP) tug, representing a critical transportation ability for the NASA lunar, Mars, and solar system exploration missions. A series of robotic assembly experiments were recently completed at Langley Research Center (LaRC) that demonstrate most of the assembly steps for the SEP tug concept. The assembly experiments used a core set of robotic capabilities: long-reach manipulation and dexterous manipulation. This paper describes cross-cutting capabilities and technologies for in-space assembly (ISA), applies the ISA approach to a SEP tug, describes the design and development of two assembly demonstration concepts, and summarizes results of two sets of assembly experiments that validate the SEP tug assembly steps.

The National Aeronautics and Space Administration continues to develop and refine various transportation options to successfully field a human Mars campaign. One of these transportation options is the Hybrid Transportation System which utilizes both solar electric propulsion and chemical propulsion. The Hybrid propulsion system utilizes chemical propulsion to perform high thrust maneuvers, where the delta-V is most optimal when ap- plied to save time and to leverage the Oberth effect. It then utilizes solar electric propulsion to augment the chemical burns throughout the interplanetary trajectory. This eliminates the need for the development of two separate vehicles for crew and cargo missions. Previous studies considered single point designs of the architecture, with fixed payload mass and propulsion system performance parameters. As the architecture matures, it is inevitable that the payload mass and the performance of the propulsion system will change. It is desirable to understand how these changes will impact the in-space transportation system's mass and power requirements. This study presents an in-depth sensitivity analysis of the Hybrid crew transportation system to payload mass growth and solar electric propulsion performance. This analysis is used to identify the breakpoints of the current architecture and to inform future architecture and campaign design decisions.

Human missions to Mars present several major challenges that must be overcome, including delivering multiple large mass and volume elements, keeping the crew safe and productive, meeting cost constraints, and ensuring a sustainable campaign. Traditional methods for executing human Mars missions minimize or eliminate in-space assembly, which provides a narrow range of options for addressing these challenges and limits the types of missions that can be performed. This paper discusses recent work to evaluate how the inclusion of in-space assembly in space mission architectural concepts could provide novel solutions to address these challenges by increasing operational flexibility, robustness, risk reduction, crew health and safety, and sustainability. A hierarchical framework is presented to characterize assembly strategies, assembly tasks, and the required capabilities to assemble mission systems in space. The framework is used to identify general mission system design considerations and assembly system characteristics by assembly strategy. These general approaches are then applied to identify potential in-space assembly applications to address each challenge. Through this process, several focus areas were identified where applications of in-space assembly could affect multiple challenges. Each focus area was developed to identify functions, potential assembly solutions and operations, key architectural trades, and potential considerations and implications of implementation. This paper helps to identify key areas to investigate were potentially significant gains in addressing the challenges with human missions to Mars may be realized, and creates a foundation on which to further develop and analyze in-space assembly concepts and assembly-based architectures.

This paper describes a case study for applying innovative architectures related to electrified propulsion for aircraft. Electric and hybrid electric propulsion for aircraft has gained widespread and significant attention over the past decade. The driver for industry interest has principally been the need to reduce emissions of combustion engine exhaust products and noise, but increasingly studies revealed potential for overall improvement in energy efficiency and mission flexibility of new aircraft types. In this work, a conceptual new type for a skydiver lift mission aircraft is examined. The opportunities which electric hybridisation offers for this role is analysed in comparison with conventional legacy type propulsion systems. For a conventional commercial skydiving mission, an all-electric propulsion system is shown as viable, and a hybrid-electric system is shown to reduce aircraft fuel costs and CO2 emissions whilst maintaining conventional aero-engine operational benefits. The new paradigm for aircraft development which hybrid electric propulsion enables has highlighted significant issues with aircraft certification practices as they exist today. The advancement of aircraft design and production to harness the value of new propulsion systems may require adaption and development of certification standards to cater for these new technologies.

The National Aeronautics and Space Administration (NASA) Solar Electric Propulsion Technology Demonstration Mission (SEP TDM), in conjunction with PC Krause and Associates, has created a Simulink-based power architecture model for a 50 kilo-Watt (kW) solar electric propulsion system. NASA has extended this model to investigate 150 kW solar electric propulsion systems. Increasing the power capability to 150 kW is an intermediate step to the anticipated power requirements for Mars and other deep space applications. The high-power solar electric propulsion capability has been identified as a critical part of NASA’s future beyond-low-Earth-orbit for human-crewed exploration missions. This paper presents four versions of a 150 kW architecture, simulation results, and a discussion of their merits.

The institute of aerospace engineering at TU Dresden is developing a highly miniaturized field emission electric propulsion (FEEP) system for Nano-satellites, called CubeSats. The presented propulsion system NanoFEEP is composed of the actual field emission thrusters and a carbon nanotubes silicon chip used as a cold electron source for neutralizing the electric charging behavior of the spacecraft. The NanoFEEP thrusters use Gallium as metal propellant and provide a continuous controllable thrust of up to 20 micro-newton. This thrust level is sufficient to control the attitude of the miniature CubeSat satellites and even perform orbit maneuvers. Two satellites missions will demonstrate these capabilities in the near future. One mission will show the feasibility of formation flying of CubeSats. The other mission will demonstrate the possibility of space debris avoidance with our propulsion system by deorbiting the satellite after mission completed.

The L1-L2 and L2-L1 missions using electric propulsion are studied. The possibility and the necessity of the and gravity-assisted maneuvers are considered. Total flight time was chosen as the optimization criterion for the variational problem of the low thrust spacecraft transfer. As the result of simulation, control programs, corresponding trajectories, and a set of output parameters were determined for certain values of acceleration and jet stream velocity of the propulsion system.

Hybrid propulsion is currently favored for a Mars Ascent Vehicle (MAV) concept from a thermal performance and Gross Lift Off Mass standpoint. However, it is at a relatively low level of maturity compared to conventional propulsion options. Technology development efforts are currently underway to bring hybrid propulsion to a technology readiness level that would enable its infusion into potential Mars Sample Return. A new propellant combination is being considered for this design that has excellent low temperature behavior. Preliminary results of two ground test campaigns are currently underway to characterize this propellant combination. Hotfire testing is being carried out in parallel at Parabilis Space Technologies and Space Propulsion Group. In addition to the new propellant combination, several other technologies are being pursued for a potential hybrid MAV: hypergolic ignition and Liquid Injection Thrust Vector Control. Both of these technologies have been applied in other rocket applications, e.g. liquid propulsion commonly uses hypergolic propellants and missiles, such as the Minuteman II, have used LITVC in the past. Hypergolic ignition, when oxidizer and fuel combust upon contact, is highly desirable for multiple starts required by the MAV concept. Therefore, testing at Penn State and Purdue is being completed in this area. An updated hybrid propulsion system design for a Mars Ascent Vehicle concept based on JPL's current understanding of potential Mars Sample Return requirements will be presented, leveraging the advances in technology development as well as updated understanding of how requirements may evolve.

The Jet Propulsion Laboratory is investing in a suite of core flight system technologies to enable CubeSats to conduct missions in deep space. These will be demonstrated on currently funded missions, such as INSPIRE, MarCO, and Lunar Flashlight, which will be among the first CubeSat missions to leave Earth's orbit and explore deep space, Mars, and the Moon, respectively. Other concepts may consider using these technologies to explore Venus, asteroids, Europa, Titan, and other areas of the solar system. These missions and concepts can be enabled by the development of miniaturized yet performant command and data handling, power, software, and communications systems specifically designed for deep space applications. JPL is pushing the state of the art in small subsystems to augment NASA's history of exploration. While the CubeSat/SmallSat component market has grown significantly to benefit LEO applications, only a few vendors are actively developing avionics and instrument interface electronics capable of meeting the stringent environmental, reliability, and performance requirements of deep space missions. These electronics and systems need to be specifically designed to handle harsh radiation and thermal environments as well as extended mission durations, where a CubeSat may begin its science observations after a multi-year cruise. Deep space missions also require additional technologies, such as radio transponders for interplanetary navigation. This paper first summarizes the systems-level developments of the enabling technologies of the JPL avionics bus, looking at maturing hardware and as well as future evolutions of technologies. Second, the paper discusses potential science instruments and applications that could be accommodated by these unique flight systems, either within a CubeSat or SmallSat form factor. Finally, the paper pairs technologies and instruments and showcases potential science missions enabled by this novel capability.

Recently missions like Hayabusa and Dawn have shown the relevance and benefits of low-thrust spacecraft concerning the exploration of our solar system. In general, the efficiency of low-thrust propulsion is one means of improving mission payload mass. At the same time, gravity-assist maneuvers can serve as mission enablers, as they have the capability to provide “free energy.” A combination of both, gravity-assist and low-thrust propulsion, has the potential to generally improve mission performance, i.e. planning and optimization of gravity-assist sequences for low-thrust missions is a desirable asset. Currently no established methods exist to include the gravity-assist partners as optimization variable for low-thrust missions. The present paper explains how gravity-assists are planned and optimized, including the gravity-assist partners, for high-thrust missions and discusses the possibility to transfer the established method, based on the Tisserand Criterion, to low-thrust missions. It is shown how the Tisserand Criterion needs to be adapted using a correction term for the low-thrust situation. It is explained why this necessary correction term excludes an a priori evaluation of sequences and therefore their planning and an alternate approach is proposed. Preliminary results of this method, by application of a Differential Evolution optimization algorithm, are presented and discussed, showing that the method is valid but can be improved. Two constraints on the search space are briefly presented for that aim.

Human and robotic missions to Mercury and Saturn are presented and analyzed with a range of propulsion options. Historical studies of space exploration, planetary spacecraft, and astronomy, in-situ resource utilization (ISRU), and industrialization all point to the vastness of natural resources in the solar system. Advanced propulsion benefitted from these resources in many ways. While advanced propulsion systems were proposed in these historical studies, further investigation of nuclear options using high power nuclear thermal and nuclear pulse propulsion as well as advanced chemical propulsion can significantly enhance these scenarios. Updated analyses based on these historical visions are presented. Nuclear thermal propulsion and ISRU enhanced chemical propulsion landers are assessed for Mercury missions. At Saturn, nuclear pulse propulsion with alternate propellant feed systems and Saturn moon exploration with chemical propulsion and nuclear electric propulsion options are discussed. Issues with using in-situ resource utilization on Mercury missions are discussed. At Saturn, the best locations for exploration and the use of the moons Titan and Enceladus as central locations for Saturn moon exploration is assessed.

The International Space Station is the first space human outpost and over the last 15 years, it has represented a peculiar environment where science, technology and human innovation converge together in a unique microgravity and space research laboratory. With the International Space Station entering the second part of its life and its operations running steadily at nominal pace, the global space community is starting planning how the human exploration could move further, beyond Low-Earth-Orbit. According to the Global Exploration Roadmap, the Moon represents the next feasible path-way for advances in human exploration towards the nal goal, Mars. Based on the experience of the ISS, one of the most widespread ideas is to develop a Cislunar Station in preparation of long duration missions in a deep space environment. Cislunar space is de ned as the area of deep space under the influence of Earth-Moon system, including a set of special orbits, e.g. Earth-Moon Libration points and Lunar Retrograde Orbit. This habitat represents a suitable environment for demonstrating and testing technologies and capabilities in deep space. In order to achieve this goal, there are several crucial systems and technologies, in particular related to transportation and launch systems. The Orion Multi-Purpose Crew Vehicle is a reusable transportation capsule designed to provide crew transportation in deep space missions, whereas NASA is developing the Space Launch System, the most powerful rocket ever built, which could provide the necessary heavy-lift launch capability to support the same kind of missions. These innovations would allow quite-fast transfers from Earth to the Cislunar Station and vice versa, both for manned and unmanned missions. However, taking into account the whole Concept of Operations for both the growth and sustainability of the Cislunar Space Station, the Lunar Space Tug can be considered as an additional, new and fundamental element for the mission architecture. The Lunar Space Tug represents an alternative to the SLS scenario, especially for what concerns all unmanned or logistic missions (e.g. cargo transfer, on orbit assembly, samples return), from Low Earth Orbit to Cislunar space. The paper focuses on the mission analysis and conceptual design of the Lunar Space Tug to support the growth and sustainment of the Cislunar Station. Particular attention is dedicated to the analysis of the propulsion subsystem effects of the Lunar Space Tug design. Main results are presented and discussed, and main conclusions are drawn.

The aim of aircraft engines development is the propulsion which is characterized by high power-to-mass ratio. Therefore, the alternative solutions that provide the required power by the low weight propulsion are sought after. The main advantage of these solutions is improvement of environmental and economic properties. This paper presents the overview of solutions and studies conducted for the unmanned aerial vehicles propulsion. For the purposes of studies a test bench was prepared. Its enables the comparison of the propulsion operating parameters taking into account changes in the values of thrust and propulsion power. The summary includes a proposal to improve the environmental indicators of propulsion systems for unmanned aerial vehicles.

In parallel with the automotive industry, hybrid-electric propulsion is becoming a viable alternative propulsion technology for the aviation sector and reveals potential advantages including fuel savings, lower pollution, and reduced noise emission. Hybrid-electric propulsion systems can take advantage of the synergy between two technologies by utilizing both internal combustion engines and electric motors together, each operating at their respective optimum conditions. However, there can also be disadvantages to hybrid propulsion. We are conducting an analysis of hybrid-electric propulsion for aircraft, which is looking at modelling systems over a range of aircraft scale, from small UAVs to inter-city airliners. To support the theoretical models, a mid-scale hybrid-electric propulsion system for a single-seat manned aircraft is being designed, built, and tested to generate data for validation and development of the simulation models. This paper draws parallels between the synergy of hybrid-electric propulsion for automotive and aviation applications, and presents an innovative theoretical approach integrating several desktop PC software packages to analyse and optimize hybrid-electric technology for aircraft. Our findings to date indicate that hybrid-electric propulsion can have a significant impact in the small- and mid-scale sectors, but only a minor impact in the large-scale sector assuming battery energy densities predicted for the next decade. Fuel savings of up to 50 and 10 % have been calculated for a microlight aircraft and inter-city airliner, respectively, over the mission profiles considered.

Today, research performed for new aircraft propulsion concepts is driven by the requirement of achieving significant emission reductions to meet the environmental objectives of future air traffic. A current trend visible in the aviation industry shows the attempt to reduce inflight emissions as well as overall energy consumption of conventional combustion engines through electrification via electrical energy sources. With the growing interest in novel hybrid electric propulsion concepts, in the same way, a demand for conceptual design and performance simulation methods in combination with analysis capabilities rises. Based on primarily introduced and discussed hybrid electric propulsion systems, this paper presents a set of aircraft propulsion system simulation (APSS) methods which allow for an integrated simulation and consistent analysis. Particularly, methods for the modelling of electric motor and battery systems as implemented in APSS are described in detail.

This paper offers a user-centric consolidation and comparison of the full range of government and commercial solar electric propulsion options available in the near term for primary propulsion on deep-space science missions of the class commonly proposed to NASA’s Discovery program. Unlike previous papers, this work does not emphasize feasibility from a mission-analysis perspective. Rather, it emphasizes requirements uniquely imposed by competitively reviewed cost-capped mission proposals, for which system-level flight heritage and cost credibility can trump sheer performance and mission capture. It describes criteria that mission architects and review boards can use to select and evaluate electric propulsion systems, provides descriptions of viable government and commercial electric propulsion system options, describes the modifications needed to adapt commercial electric propulsion systems to deep space, and discusses appropriate methods for costing commercial-based electric propulsion systems. It concl...

Currently, toxic and carcinogenic hydrazine propellants are commonly used in spacecraft propulsion. These propellants impose distinctive environmental challenges and consequential hazardous conditions. With an increasing level of future space activities and applications, the significance of greener space propulsion becomes even more pronounced. In this article, a selected number of promising green space propellants are reviewed and investigated for various space missions. In-depth system studies in relation to the aforementioned propulsion architectures further unveil possible approaches for advanced green propulsion systems of the future.

This paper presents a mission analysis comparison of human missions to asteroids using two distinct architectures. The objective is to determine if either architecture can reduce launch mass with respect to the other, while not sacrificing other performance metrics such as mission duration. One architecture relies on chemical propulsion, the traditional workhorse of space exploration. The second combines chemical and electric propulsion into a hybrid architecture that attempts to utilize the strengths of each, namely the short flight times of chemical propulsion and the propellant efficiency of electric propulsion. The architectures are thoroughly detailed, and accessibility of the known asteroid population is determined for both. The most accessible asteroids are discussed in detail. Aspects such as mission abort scenarios and vehicle reusability are also discussed. Ultimately, it is determined that launch mass can be greatly reduced with the hybrid architecture, without a notable increase in mission duration. This demonstrates that significant performance improvements can be introduced to the next step of human space exploration with realistic electric propulsion system capabilities. This leads to immediate cost savings for human exploration and simultaneously opens a path of technology development that leads to technologies enabling access to even further destinations in the future.

On 5 August 2012, the Mars Science Laboratory mission successfully landed Curiosity, the largest interplanetary rover ever built, on the surface of Mars. The entry, descent, and landing phase of this mission was by far the most complex landing ever attempted on a planetary body. The descent-stage propulsion subsystem was explicitly developed for the guided-entry and sky-crane maneuvers that enabled precise landing of Curiosity. Development of the descent-stage propulsion system required a number of new propulsion hardware developments, incorporating technologies not normally found in spacecraft propulsion subsystems. The in-flight performance of the descent-stage propulsion system was consistent with expectations, with the exception of higher than expected delivered thrust from the Mars lander engines. The descent-stage propulsion system completed its mission with significant capability margin with respect to thruster life, maximum delivered thrust, and available propellant at touchdown.

The research and curiosity about outer space had been always constant. Looking for others planets, ways, civilizations wherever the exploration of the space will be a thing which the human desire. The challenge here for several years was the obtaining energy sufficient for the application of the missions. So, presently the major objective in the missions is offer more autonomy to the spacecrafts and consequently to lower the cost of the missions. Solar Sails have long been envisaged as an enabling technology because is a promising low-cost option for space exploration for it uses for propulsion an abundant resource in space: solar radiation. In this paper a simple model of solar sail is shown and studied your performance of Control System.