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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 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.

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.

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.

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.

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.

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.

International audience; PRECISE focuses on the research and development of a MEMS-based monopropellant micro chemical propulsion system for highly accurate attitude control of satellites. The availability of such propulsion systems forms the basis for defining new mission concepts such as formation flying and rendezvous manoeuvres. These concepts require propulsion systems for precise attitude and orbit control manoeuvrability. Application-oriented aspects are addressed by two end-users who are planning a formation flying mission for which the propulsion system is crucial. Basic research is conducted aiming at improving crucial MEMS technologies required for the propulsion system. Research and development also focuses on the efficiency and reliability of critical system components. System analysis tools are enhanced to complement the development stages. Finally, the propulsion system will be tested in a simulated space vacuum environment. These experiments will deliver data for the validation of the numerical models

Since September 2001, NASA's In-Space Propulsion Technology (ISPT) program has been developing technologies for lowering the cost of planetary science missions. Recently completed is the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost. Two other cost saving technologies nearing completion are the NEXT ion thruster and the Aerocapture technology project. Under development are several technologies for low-cost sample return missions. These include a low-cost Hall-effect thruster (HIVHAC) which will be completed in 2011, light-weight propellant tanks, and a Multi-Mission Earth Entry Vehicle (MMEEV). This paper will discuss the status of the technology development, the cost savings or performance benefits, and applicability of these in-space propulsion technologies to NASA's future Discovery, and New Frontiers missions, as well as their relevance for sample return missions.

The idea for using the International Space Station (ISS) as a platform for exploration has matured in the past few years and the concept continues to gain momentum. ISS provides a robust infrastructure which can be used to test systems and capabilities needed for missions to the Moon, Mars, asteroids and other potential destinations. International cooperation is a critical enabler and ISS has already demonstrated successful management of a large multi-national technical endeavor. Systems and resources needed for expeditions can be aggregated and thoroughly tested at ISS before departure thus providing wide operational flexibility and the best assurance of mission success. A small part of ISS called an Exploration Platform (ISS-EP) can be placed in cislunar space providing immediate benefits and flexibility for future exploration missions. We will show how ISS and the ISS-EP can be used to reduce risk and improve the operational flexibility for missions beyond low Earth orbit. Life support systems and other technologies developed for ISS can be evolved and adapted to the ISS-EP and other exploration spacecrafts. New technology, such as electric propulsion and advanced life support systems can be tested and proven at ISS as part of an incremental development program. Commercial companies who are introducing transportation and other services will benefit with opportunities to contribute to the mission since ISS will serve as a focal point for the commercialization of low earth orbit services. Finally, we will show how the use of ISS provides immediate benefits to the scientific community because its capabilities are available today and certain critical aspects of exploration missions can be simulated.

The past and the recent status of nuclear propulsion (NP) for application to space mission is presented. The case for using NP in manned space missions is made based on fundamental physics and on the necessity to ensure safe radiation doses to future astronauts. In fact, the presence of solar and galactic-cosmic radiation poses substantial risks to crews traveling for months in a row to destinations such as asteroids and Mars. Since passive or active shields would be massive to protect against the more energetic part of the radiation energy spectrum, the only alternative is to reduce dose by traveling faster. Hence the importance of propulsion systems with much higher specific impulse than that of current chemical systems, and thus the use of nuclear propulsion. Nuclear-thermal and nuclear-electric propulsions are then discussed in view of their potential application to missions now in the preliminary planning stage by space agencies and industries and being considered by the ISECG international panel. In this context, recent ideas for future use of the ISS that may require NP are also presented.

During the summer of 2010, NASA's Office of Chief Technologist assembled 15 civil service teams to support the creation of a NASA integrated technology roadmap. The Aero-Space Technology Area Roadmap is an integrated set of technology area roadmaps recommending the overall technology investment strategy and prioritization for NASA's technology programs. The integrated set of roadmaps will provide technology paths needed to meet NASA's strategic goals. The roadmaps have been reviewed by senior NASA management and the National Research Council. With the exception of electric propulsion systems used for commercial communications satellite station-keeping and a handful of deep space science missions, almost all of the rocket engines in use today are chemical rockets; that is, they obtain the energy needed to generate thrust by combining reactive chemicals to create a hot gas that is expanded to produce thrust. A significant limitation of chemical propulsion is that it has a relatively low specific impulse. Numerous concepts for advanced propulsion technologies with significantly higher values of specific impulse have been developed over the past 50 years. Advanced in-space propulsion technologies will enable much more effective exploration of our solar system, near and far, and will permit mission designers to plan missions to “fly anytime, anywhere, and complete a host of science objectives at the destinations” with greater reliability and safety. With a wide range of possible missions and candidate propulsion technologies with very diverse characteristics, the question of which technologies are ‘best’ for future missions is a difficult one. A portfolio of technologies to allow optimum propulsion solutions for a diverse set of missions and destinations are described in the roadmap and herein.

Distributed propulsion is one of the revolutionary candidates for future aircraft propulsion. In this journal article, the potential role of distributed propulsion technology in future aviation is investigated. Following a historical journey that revisits distributed propulsion technology in unmanned air vehicles and military aircraft, features of this specific technology are highlighted in synergy with an electric aircraft concept and a first-of-a-kind comparison to commercial aircraft employing distributed propulsion arrangements. In light of propulsion–airframe integration and complementary technologies such as boundary layer ingestion, thrust vectoring and circulation control, transpired opportunities and challenges are addressed in addition to a number of identified research directions proposed for future aircraft. The motivation behind enhanced means of communication between engineers, researchers and scientists has stimulated a novel proposed definition for the distributed propulsion technology in aviation and is presented herein.

The status and the main results achieved by Project 242 are presented. Project 242 is a programme (funded by ASI—1999/2002 from an idea of Carlo Rubbia) that studied a new concept of space propulsion motor by using direct conversion of the kinetic energy of fission fragments into increasing of enthalpy of a propellant gas. Project 242 studied the application of this propulsion system to a manned mission to Mars. Preliminary results were very satisfactory and it has been observed that a propulsion system with these characteristics could make the mission feasible. Results for other unmanned missions to the outer solar system are also presented.

One of the main challenge in the design of an active removal system for space debris is the high Δ V required both to approach space debris lying in different orbits and to de-orbit/re-orbit them. Indeed if the system does not target a number of objects during its lifetime the cost of the removal will be far too high to be considered as the basis of an economically viable business case. Using a classical chemical propulsion (CP) system, the Δ V is limited by the mass of propellant that the system can carry. This limitation is greatly reduced if electrical propulsion is considered. Electrical propulsion (EP) systems are indeed characterized by low propellant mass requirements, however this comes at the cost of higher electrical power and, typically, higher complexity and mass of the power supply system. Because of this, the use of EP systems has been, therefore, primarily limited to station keeping maneuvers. However in the recent past, the success of missions using EP as primary propulsion (e.g. GOCE, SMART-1, Artemis, Deep Spcae1, Hayabusa) makes this technology a suitable candidate for providing propulsion for an active debris removal system. This study case will provide the analysis of the possible application of electrical propulsion systems in such a context, presenting a number of possible mission profiles. This paper will start with the description of possible mission concepts and the assessment of the EP technology, comparing near-term propulsion options, that best fits the mission. A more detailed analysis follows with the relevant trade-off to define the characteristics of the final system and its size in terms of mass and power required. A survey of available space qualified EP systems will be performed with the selection of the best candidates to be used and/or developed for an active debris removal system. The results of a similar analysis performed for a classical CP system are then presented and the two options are compared in terms of total cost of the mission. The output of this study case shows that the EP is a suitable solution in a very demanding Δ V system such as active debris removal satellites and that the current technology is sufficiently mature to be used in the very near future.

In this paper we propose a pole-sitter spacecraft hybridising solar electric propulsion (SEP) and solar sailing. The intriguing concept of a hybrid propulsion spacecraft is attractive: by combining the two forms of propulsion, the drawbacks of the two systems cancel each other, potentially enabling propellant mass saving, increased reliability, versatility and lifetime over the two independent systems. This almost completely unexplored concept will be applied to the continuous monitoring of the Earth's polar regions through a pole-sitter, i.e. a spacecraft that is stationary above one pole of the Earth. The continuous, hemispherical, real-time view of the pole will enable a wide range of new applications for Earth observation and telecommunications. In this paper, families of 1-year-periodic, minimum-propellant orbits are found, for different values of the sail lightness number and distance from the pole. The optimal control problem is solved using a pseudo-spectral method. The process gives a reference control to maintain these orbits. In addition, for stability issues, a feedback control is designed to guarantee station-keeping in the presence of injection errors, sail degradation and temporary SEP failure. Results show that propellant mass can be saved using a medium-sized solar sail. Finally, it is shown that the feedback control is able to maintain the spacecraft on-track with only minimal additional effort from the SEP thruster.

This paper presents an optimal control strategy suitable for investigating the mission performance of a spacecraft equipped with a NASAʼs Evolutionary Xenon Thruster. Using an accurate mathematical model of the propulsion system, based on both experimental data and a semi-empirical schematization of the power generation system, the minimum-time rendezvous problem with various celestial bodies is investigated through an indirect approach. Unlike a conventional approach, in which the thruster power is varied with continuity within a given range, here the solar electric propulsion system is described using a finite number of operation points, characterized by different pairs of thruster input power and propellant mass flow rate. The model includes the possibility of obtaining coasting phases along the spacecraft trajectory. The optimal control strategy is found for a three-dimensional problem in which the actual shape and relative orientation of the initial and final orbits are taken into account. The paper presents the simulation results for a number of minimum-time trajectories towards inner planets (Mercury, Venus, and Mars) and three Near Earth Asteroids (NEAs) of particular interest for the scientific community: 1998 KY-26, 99942 Apophis, and 2007 VK-184.

In the early to mid-2000s, NASA made substantial progress in the development of solar sail propulsion systems. Solar sail propulsion uses the solar radiation pressure exerted by the momentum transfer of reflected photons to generate a net force on a spacecraft. To date, solar sail propulsion systems were designed for large robotic spacecraft. Recently, however, NASA has been investigating the application of solar sails for small satellite propulsion. The NanoSail-D is a subscale solar sail system designed for possible small spacecraft applications. The NanoSail-D mission flew on board the ill-fated Falcon Rocket launched August 2, 2008, and due to the failure of that rocket, never achieved orbit. The NanoSail-D flight spare is ready for flight and a suitable launch arrangement is being actively pursued. This paper will present an introduction solar sail propulsion systems and an overview of the NanoSail-D spacecraft.

Current plans call for the first human missions to Mars to be launched perhaps as early as 2035. The recently completed “Mars Design Reference Architecture 5.0” study defines a conceptual mission architecture and identifies enabling technologies. NASA is beginning long range development on key technologies needed for these missions because it will take many years for them to reach maturity. The ISS and the lunar outpost will be used as test beds for these technologies to reduce risk and prepare for human exploration of Mars. NASA’s Exploration Technology Development Program is maturing technologies and demonstrating operational scenarios for lunar exploration that are extensible to future human missions to Mars. These include entry, descent, and landing systems for large payloads; fission surface power systems; liquid oxygen–liquid methane propulsion systems; cryogenic fluid management; closed-loop life support; small pressurized rovers for surface mobility; in-situ resource utilization; radiation shielding; and optical communications. Advanced technologies will enable more affordable and sustainable Mars exploration.

The invention of gravity-propelled interplanetary space travel (also known as “gravity-assist trajectories”) in the early 1960s broke the high-energy barrier of classical space travel based on reaction propulsion, and made possible the exploration of the entire solar system with instrumented spacecraft. In this concept, a free-fall spacecraft is launched from a launch planet P1 to a nearby planet P2 such that its gravitational field (superimposed on the gravitational field of the Sun) catapults the vehicle to another planet P3, which in turn is used to repeat the process. Thus, through a series of planetary encounters, a gravity-propelled trajectory P1-P2-P3-P4-…-PN is generated. This paper describes how the invention was conceived and how the difficult mathematical problem of computing the trajectories was solved in order to numerically investigate and use the invention in actual missions. The crucial roles played by the UCLA Computing Facility and the Departments of Mathematics and Physics are also described.

Studies over the last decade have shown radioisotope-based nuclear electric propulsion to be enhancing and, in some cases, enabling for many potential robotic science missions. Also known as radioisotope electric propulsion (REP), the technology offers the performance advantages of traditional reactor-powered electric propulsion (i.e., high specific impulse propulsion at large distances from the Sun), but with much smaller, affordable spacecraft. Future use of REP requires development of radioisotope power sources with system specific powers well above that of current systems. The US Department of Energy and NASA have developed an advanced Stirling radioisotope generator (ASRG) engineering unit, which was subjected to rigorous flight qualification-level tests in 2008, and began extended lifetime testing later that year. This advancement, along with recent work on small ion thrusters and life extension technology for Hall thrusters, could enable missions using REP sometime during the next decade.

著者人数：15名, Accepted: 2009-05-26, 資料番号: SA1000722000

Liquid-propellant rocket engines are widely used all over the world, thanks to their high performances, in particular high thrust-to-weight ratio. The present paper presents a general panorama of liquid propulsion as a contribution of the IAF Advanced Propulsion Prospective Group. After a brief history of its past development in the different parts of the world, the current status of liquid propulsion, the currently observed trends, the possible areas of future improvement and a summarized road map of future developments are presented. The road map includes a summary of the liquid propulsion status presented in the “Year in review 2007” of Aerospace America. Although liquid propulsion is often seen as a mature technology with few areas of potential improvement, the requirements of an active commercial market and a renewed interest for space exploration has led to the development of a family of new engines, with more design margins, simpler to use and to produce associated with a wide variety of thrust and life requirements.

We compare a variety of mission scenarios to assess the strengths and weaknesses of options for Mars exploration. The mission design space is modeled along two dimensions: trajectory architectures and propulsion system technologies. We examine direct, semi-direct, stop-over, semi-cycler, and cycler architectures, and we include electric propulsion, nuclear thermal rockets, methane and oxygen production on Mars, Mars water excavation, aerocapture, and reusable propulsion systems in our technology assessment. The mission sensitivity to crew size, vehicle masses, and crew travel time is also examined. Many different combinations of technologies and architectures are applied to the same Mars mission to determine which combinations provide the greatest potential reduction in the injected mass to LEO. We approximate the technology readiness level of a mission to rank development risk, but omit development cost and time calculations in our assessment. It is found that Earth–Mars semi-cyclers and cyclers require the least injected mass to LEO of any architecture and that the discovery of accessible water on Mars has the most dramatic effect on the evolution of Mars exploration.

India's low power GEOSAT communication satellite to be launched in 2008 will be using electrical propulsion system as an experimental payload augmenting the chemical propulsion system. India's low power communication typically can provide 700 W of power in the beginning of life. Hence, a power and mass optimized electrical propulsion system (EPS) is designed to provide inclination and eccentricity control during North South Station keeping operations. The EPS is based on four 350 W rated Hall effect thrusters (HET) designed, developed and qualified as per the required standards of GEOSAT mission. This paper presents the overview of the Electrical Propulsion System and its elements along with the results of the qualification tests and intended flight operations.

Innovative interplanetary deep space missions, like a main belt asteroid sample return (SR) mission, require ever larger velocity increments ( Δ V 's) and thus ever more demanding propulsion capabilities. Providing much larger exhaust velocities than chemical high-thrust systems, electric low-thrust space-propulsion systems can significantly enhance or even enable such high-energy missions. In 1995, a European-Russian Joint Study Group (JSG) presented a study report on “Advanced Interplanetary Missions Using Nuclear-Electric Propulsion” (NEP). One of the investigated reference missions was an SR from the main belt asteroid (19) Fortuna. The envisaged nuclear power plant, Topaz-25, however, could not be realized and also the worldwide developments in space reactor hardware stalled. In this paper, we investigate, whether such a mission is also feasible using a solar electric propulsion (SEP) system and compare our SEP results to corresponding NEP results.

Human exploration and development of space is being pursued by spacefaring nations to explore, use, and enable the development of space and expand the human experience there. The goals include: increasing human knowledge of nature’s processes using the space environment; exploring and settling the solar system; achieving routine space travel; and enriching life on Earth through living and working in space. A crucial aspect of future space missions is the development of infrastructure to optimize safety, productivity, and costs. A major component of mission execution is operations management. NASA’s International Space Station is providing extensive experience in both infrastructure and operations. In view of this, a vigorously organized approach is needed to implement successful space-, planet-, and ground-based research and operations that entails wise and efficient use of technical and human resources. Many revolutionary technologies being pursued by researchers and technologists may be vital in making space missions safe, reliable, cost-effective, and productive. These include: ionic polymer–metal composite technology; solid-state lasers; time-domain sensors and communication systems; high-temperature superconductivity; nanotechnology; variable specific impulse magneto plasma rocket; fuzzy logic; wavelet technology; and neural networks. An overview of some of these will be presented, along with their application to space missions. © 2008 Elsevier Ltd. All rights reserved.

Future space exploration activities very likely will incorporate a large number of human missions to various locations on the Moon surface and other destinations near the Earth–Moon system. A wide range of space infrastructure elements will be required in low Earth orbit and waypoints to these destinations. It is reasonable to start assembly and mission logistics from today's frequently used near-Earth orbits, such as the orbit of the International Space Station (ISS). This allows for using launch systems and other infrastructure elements currently available in order to reduce costs. Besides new transfer, ascent and re-entry vehicles as well as planetary surface installations, space stations beyond low Earth orbit, namely in the Earth–Moon system, can be crucial elements for forthcoming exploration missions. The lunar Lagrange points have unique advantages in this context and offer promising options for the humans' next steps into space. As shown in this paper, a lunar space station (LSS) is not too farfetched; and it is definitely not science fiction but rather the next logical step for solar system exploration. Like a stepping-stone, it could enable the immediate start of sustained development of outer space with technologies available today. The proposed lunar exploration architecture targets at an LSS at the lunar Lagrange point 1 to be used as a gateway for mid-term lunar surface exploration missions or servicing missions to other destinations. For study purposes the challenging constraint of only using existing and tailored infrastructure elements and technology of European/Russian heritage is assumed. The results manifest the feasibility of such a space station between 2015 and 2020. Presented are station configuration and modules, transfer vehicles for assembly, crew and cargo transport, an enhanced life support system and a logistics concept. The concept outlines enhancements of the current transportation and station infrastructure and shows that the ISS as a transportation node can beneficially support lunar scenarios.

Nuclear electric ion propulsion is considered for three sample deep space missions starting from a 500 km low Earth orbit encompassing the transfer of a 100 MT payload into a 1500 km orbit around Mars, the rendezvous of a 10 MT payload with the Jovian moon Europa and the rendezvous of a similar payload with Saturn's moon Titan. Near term ion engine and space nuclear reactor technology are assumed. It is shown that nuclear electric ion propulsion offers more than twice the payload for the Mars mission relative to the case when a nuclear thermal rocket is used for the trans-Mars injection maneuver at Earth, and about the same payload advantage relative to the case when solar electric propulsion is used for the Mars heliocentric transfer. For missions to the outer planets nuclear electric ion propulsion increases the payload mass fraction by a factor of two or more compared with high thrust systems that utilize gravity assist trajectories.

Nuclear and radioisotope powered electric thrusters are being developed as primary in-space propulsion systems for potential future robotic and piloted space missions. Possible applications for high power nuclear electric propulsion include orbit raising and maneuvering of large space platforms, lunar and Mars cargo transport, asteroid rendezvous and sample return, and robotic and piloted planetary missions, while lower power radioisotope electric propulsion could significantly enhance or enable some future robotic deep space science missions. This paper provides an overview of recent U.S. high power electric thruster research programs, describing the operating principles, challenges, and status of each technology. Mission analysis is presented that compares the benefits and performance of each thruster type for high priority NASA missions. The status of space nuclear power systems for high power electric propulsion is presented. The paper concludes with a discussion of power and thruster development strategies for future radioisotope electric propulsion systems

Supporting NASA's Science Mission Directorate, the In-Space Propulsion Technology (ISPT) Program is developing solar sail propulsion (SSP) for use in robotic science and exploration of the solar system. SSP will provide longer on-station operation, increased scientific payload mass fraction, and access to previously inaccessible orbits for multiple potential science missions. Two different 20-m solar sail systems were produced and successfully completed functional vacuum testing last year in NASA Glenn's Space Power Facility at Plum Brook Station, Ohio. The sails were designed and developed by ATK Space Systems and L’Garde, respectively. These sail systems consist of a central structure with four deployable booms that support the sails. These sail designs are robust enough for deployments in a one atmosphere, one gravity environment, and are scalable to much larger solar sails—perhaps as much as 150 m on a side. In addition, computation modeling and analytical simulations have been performed to assess the scalability of the technology to the large sizes ( > 150 m) required for first generation solar sail missions. Life and space environmental effects testing of sail and component materials are also nearly complete. This paper will summarize recent technology advancements in solar sails and their successful ambient and vacuum testing.

The Dawn project is progressing toward its 2007 launch on a mission to orbit main belt asteroids (1) Ceres and (4) Vesta. Designed to provide insights into important questions about the evolution of the solar system, Dawn will spend more than 0.5 years in orbit about each of these bodies. This challenging mission is enabled by an ion propulsion system. In contrast to missions that use conventional chemical propulsion, the use of this system creates a strong coupling of allowable flight system mass and available power, thereby requiring different methods of managing these and other technical resources. Now that the project is nearing launch, the refinement of resource estimates allows the identification of excess margin, which is being applied in novel ways to benefit the project both in development and in operations.

The present investigation points out the potential of continuously propelled spacecraft for piloted Mars missions and compares them to impulsive propulsion (chemical and nuclear thermal) and ballistic trajectories. Although the results are related to piloted Mars missions, the stated issues raised hold true for a broad range of space missions. It is demonstrated that the use of impulsive propulsion leads to inflexible missions and may result in long total mission durations. Meanwhile, the use of continuous electric propulsion not only guarantees short total mission durations of Mars missions with moderate masses but also results in highly flexible missions. These criteria can be met with a continuous electric propulsion system that provides a thrust level of 100 N and 3000 s of specific impulse. Great potential lies in electric hybrid thrusters. The high-power, two-stage hybrid plasma thruster TIHTUS is currently being developed at the Institute of Space Systems (IRS). Its technology including preliminary laboratory testing results are presented.

One very cost-effective route into orbit is as an Auxiliary Microsatellite on the ASAP-5 structure. All planned flights of the ASAP-5 structure are into GTO. The Advanced Microsatellite Mission study carried out within the ESA General Studies Programme addressed the feasibility of a deep space mission starting from this point. One of the major configuration drivers in the ASAP-5 launch envelope is the available volume. Low-density nondeployable equipment (tanks, star sensor baffles, etc.) becomes design drivers. Optimisation for power and volume has led to an Attitude Control System architecture which is gyro-based (using the latest high-performance MEMS devices) making use of opportunistic calibration by minimally baffled star sensors. Solar electric propulsion appears the leading candidate for missions requiring large Δ V . There has historically been a quest for ever higher values of I sp for SEP thrusters, but it is shown that a full mission-level trade-off taking into account the restricted power availability on this class of spacecraft often leads to a lower value of I sp in order to minimise operations costs and improve the timeliness of scientific observations. It is concluded that useful deep space missions can be performed by a spacecraft complying with the ASAP-5 constraints.

It is generally acknowledged that new challenging scientific goals combined with the variety and complexity of newly proposed instruments are increasingly dictating the system and mission design of space missions in general, and of planetary exploration missions in particular. On the other hand, economical restrictions in public budgets, decreasing political support for scientific space research and technological hurdles are neutralizing and constraining the scientific momentum. New low-cost planetary missions, which despite the above mentioned constraints can guarantee an excellent scientific return, are requiring a new approach in system engineering and management, by applying innovative and creative spirit but keeping the cost awareness high during the complete development process. The present paper discusses a collection of the new challenges surfacing in the various fields of system design for future planetary missions.

Development is underway on a unique high-power solar concentrator array called stretched lens array (SLA) for direct drive electric propulsion. These SLA performance attributes closely match the critical needs of solar electric propulsion (SEP) systems, which may be used for “space tugs” to fuel-efficiently transport cargo from low Earth orbit (LEO) to low lunar orbit (LLO), in support of NASA's robotic and human exploration missions. Later SEP systems may similarly transport cargo from the Earth–Moon neighborhood to the Mars neighborhood. This paper will describe the SLA SEP technology, discuss ground tests already completed, and present plans for future ground tests and future flight tests of SLA SEP systems.

MAUS (Moltiplicatore Avanzato Ultracompatto Spaziale) is a nuclear reactor concept design capable to ensure a reliable, long-lasting, low-mass, compact energy supply needed for advanced, future space missions. The exploration of the solar system and the space beyond requires the development of nuclear energy generators for supplying electricity to space-bases, spacecrafts, probes or satellites, as well as for propelling ships in long space missions. For propulsion, the MAUS nuclear reactor could be used to power electric ion drive engines. An ion engine is able to build up to very high velocities, far greater than chemical propulsion systems, but has high power and long service requirements. The MAUS concept is described, together with the ion propulsion engine and together with the reference thermoionic process used to convert the thermal power into electricity. The design work has been performed at the Nuclear Engineering and Energy Conversion Department of the University of Rome “La Sapienza” starting from 1992 on an issue submitted by the Italian Space Agency (ASI), in cooperation with the research laboratories of ENEA.

A study was conducted that shows how a single radioisotope electric propulsion (REP) spacecraft design could be used for various missions throughout the solar system. This spacecraft design is based on an REP feasibility design from a study performed by NASA Glenn Research Center and the Johns Hopkins University Applied Physics Laboratory. The study also identifies technologies that need development to enable these missions. The mission baseline for the REP feasibility design study is a Trojan asteroid orbiter. This mission sends an REP spacecraft to Jupiter's leading Lagrange point where it would orbit and examine several Trojan asteroids. The spacecraft design from the REP feasibility study would also be applicable to missions to the Centaurs, and through some change of payload configuration could accommodate a comet sample-return mission. Missions to small bodies throughout the outer solar system are also within reach of this spacecraft design. This set of missions, utilizing the common REP spacecraft design, is examined, and required design modifications for specific missions are outlined.

SMART-1 is the first of the small missions for advanced research and technology as part of ESA's science programme “Cosmic vision”. It was successfully launched on September 27, 2003 and is presently traveling towards its destination, the Moon. The main objective of the mission, to demonstrate solar electric primary propulsion for future Cornerstones (such as Bepi-Colombo), has already been achieved. At the time of writing the electric propulsion system has been working already for more than 3400 h and has provided a Delta-V to the spacecraft of more than 2500 m/s. The other technology objectives are also being fulfilled by the verification of the proper functioning of such on-board experiments like the X-Ka band transponder, the X-ray spectrometer, the near IR spectrometer, the laser link, etc. The scientific objectives are related to lunar science and will be fulfilled once the spacecraft enters its operational lunar orbit, currently expected for January 2005. SMART-1 lunar science investigations will include studies of the chemical composition of the Moon, of geophysical processes, environment and high-resolution studies in preparation for future steps of lunar exploration. SMART-1 has been an innovative mission in many aspects and we are now drawing some preliminary conclusions about the lessons to be learnt. The paper describes the spacecraft and the technology elements with particular emphasis to the technology nature of the mission. The on-board avionics employs many novel designs for spacecraft, including a serial CAN bus for data communication, autonomous star trackers and extensive use of auto-code generation for implementing the attitude control system and the failure, detection, isolation and recovery (FDIR). Finally, the orbital operation phase currently ongoing, including the routine electric propulsion operations and the instrument commissioning, is providing a wealth of data and lesson-learnt useful for future autonomous planetary missions.

The future space mission scenario of scientific and commercial/application missions impose very challenging requirements on the design and performance of space mechanisms and the tribology of their moving surfaces in contact. Taking as example the scenarios and preliminary design concepts of possible future ESA science missions, critical technological advancements of their mechanisms are being identified. The challenges of competitive mechanical product development for commercial/application missions and the role of equipment suppliers in the rapid development cycle are highlighted in selected examples of mechanism-related space equipment.

This paper focuses on the technical challenges that arise when electric propulsion is used on a small spacecraft such as SMART-1. The choice of electric propulsion influences not only the attitude control system and the power system, but also the thermal control as well as the spacecraft structure. A description is given on how the design of the attitude control system uses the possibility to control the alignment of the thrust vector in order to reduce the momentum build-up. An outline is made of the philosophy of power generation and distribution and shows how the thermal interfaces to highly dissipating units have been solved. Areas unique for electric propulsion are the added value of a thrust vector orientation mechanism and the special consideration given to the electromagnetic compatibility. SMART-1 is equipped with a thruster gimbal mechanism providing a 10° cone in which the thrust vector can be pointed. Concerning the electromagnetic compatibility, a discussion on how to evaluate the available test results is given keeping in mind that one of the main objectives of the SMART-1 mission is to assess the impact of electric propulsion on the scientific instruments and on other spacecraft systems. Finally, the assembly, integration and test of the spacecraft is described. Compared to traditional propulsion systems, electric propulsion puts different requirements on the integration sequence and limits the possibilities to verify the correct function of the thruster since it needs high quality vacuum in order to operate. Prime contractor for SMART-1 is the Swedish Space Corporation (SSC). The electric propulsion subsystem is procured directly by ESA from SNECMA, France and is delivered to SSC as a customer furnished item. The conclusion of this paper is that electric propulsion is possible on a small spacecraft, which opens up possibilities for a new range of missions for which a large velocity increment is needed. The paper will also present SMART-1 and show how the problems related to the accommodation of electric propulsion have been solved during design and planning of the project.

Modern state of development and application of electric propulsion (EP) in Russia is analyzed in this report. The characteristics of fields of advisable EPS application are presented. The power-to-weight and economic effectiveness of applying EP systems of different types for solving perspective tasks, including spacecraft transportation to operating orbits, its station keeping in the orbit and in the structure of satellite constellations, operation and stabilization and taking away from the orbit after the end of the lifetime are studied in this paper.

NASA's Integrated Space Transportation Plan (ISTP) calls for investments in Space Shuttle safety upgrades, second generation Reusable Launch Vehicle (RLV) advanced development and third generation RLV and in-space research and technology. NASA's third generation launch systems are to be fully reusable and operation by 2025. The goals for third generation launch systems are to reduce cost by a factor of 100 and improve safety by a factor of 10,000 over current systems. The Advanced Space Transportation Program Office (ASTP) at NASA's Marshall Space Flight Center in Huntsville, AL has the agency lead to develop third generation space transportation technologies. The Hypersonics Investment Area, part of ASTP, is developing the third generation launch vehicle technologies in two main areas, propulsion and airframes. The program's major investment is in hypersonic airbreathing propulsion since it offers the greatest potential for meeting the third generation launch vehicles. The program will mature the technologies in three key propulsion areas, scramjets, rocket-based combined cycle and turbine-based combination cycle. Ground and flight propulsion tests are being planned for the propulsion technologies. Airframe technologies will be matured primarily through ground testing. This paper describes NASA's activities in hypersonics. Current programs, accomplishments, future plans and technologies that are being pursued by the Hypersonics Investment Area under the Advanced Space Transportation Program Office will be discussed.

Through the application of advanced technologies and mission concepts, architectures for missions beyond Earth orbit have been dramatically simplified. These concepts enable a stepping stone approach to science driven; technology enabled human and robotic exploration. Numbers and masses of vehicles required are greatly reduced, yet the pursuit of a broader range of science objectives is enabled. The scope of human missions considered range from the assembly and maintenance of large aperture telescopes for emplacement at the Sun-Earth libration point L2, to human missions to asteroids, the moon and Mars. The vehicle designs are developed for proof of concept, to validate mission approaches and understand the value of new technologies. The stepping stone approach employs an incremental buildup of capabilities, which allows for future decision points on exploration objectives. It enables testing of technologies to achieve greater reliability and understanding of costs for the next steps in exploration.

A key feature of future deep-space science missions will be the need for significantly greater on-board propulsion capability. To meet this need, ion propulsion based on the technology that flew on NASA's Deep Space 1 spacecraft has now entered the mainstream of propulsion options available for deep-space missions. The next most likely science mission to use ion propulsion is the comet nucleus sample return (CNSR) mission. CNSR has recently been identified by the Solar System Exploration Subcommittee as the highest priority new mission for NASA's Exploration of the Solar System theme. Ion propulsion for CNSR enables the use of a smaller, less expensive launch vehicle, and significantly shortens the overall trip time. A trade study for CNSR was performed to identify engine and system technology improvements, which provide the greatest mission benefits for the lowest additional risk. This trade study indicated that the maximum specific impulse of the ion engine should be increased from 3100 to 3800 s and that the maximum engine input power should be increased from 2.3 to 3.2 kW . Simultaneously the engine total propellant throughput capability must be increased from the 80-kg NSTAR design point to approximately 180 kg . A focused technology program to make these advances is underway.

SMART-1 is the first of the Small Missions for Advanced Research in Technology of the ESA Horizons 2000 Science Plan. The main mission objective of SMART-1 is to demonstrate innovative and key technologies for scientific deep-space missions. One of the key technologies is the solar electric propulsion used as primary propulsion. The electric propulsion will be using 1400 W to transfer the 350 kg spacecraft from an Ariane 5 standard GTO to an elliptic Moon polar orbit, 10000×300 km . The total mission time is 24 months including a maximum of 18 months transfer time. The spacecraft development entered the detailed design and implementation phase in October 1999, under the responsibility of the Swedish Space Cooperation as prime contractor, and the flight acceptance is targeted for the 3rd quarter of 2002. The committed total life cost budget is 84 million Euro. Apart from the in-orbit demonstration of electric propulsion as primary propulsion, SMART-1 is implementing many other enabling technologies for deep-space missions such as • deep-space transponder with communication in X-band for uplink and in X- and Ka-band for downlink, • highly integrated and radiation tolerant microprocessors, FPGAs and memories, • high on-board autonomy driven by ground communication only once per 4 days, • maximum available power to electric propulsion by centralised and software-controlled power management, • minimum propellant consumption through innovative angular momentum management. In addition, the spacecraft avionics design is tailored to the low-cost philosophy by enabling flexible integration of commercial off the shelf (COTS) equipment. The scientific instruments, five in total, support characterisation of the electric propulsion thrust environment during the long transfer phase and detailed imaging and spectroscopy of the lunar surface in visible, infrared and X-ray during the Moon orbiting phase. Several of the instruments implement new enabling technologies, e.g. swept charge devices for the X-ray spectrometer and quarts gratings for the near-infrared reflectance spectrometer. The paper summarises the baseline mission design, the projected scientific and technology output, and describe the spacecraft bus design. The main part of the paper highlights and elaborate on a number of spacecraft bus technologies; the on-board autonomy, the avionics design, and the X/Ka-band transponder.

The Comet Nucleus Sample and Return (CNSR) mission has been of great interest for many years. The mission objective is very clear and it provides the clues to the solar system and the origin of life. The paper presents the feasible CNSR scenario taking the advantage of Solar Electric Propulsion to the comet Wilson-Harrington (WH). The plan is found feasible, and the launch window is in 2005 and the spacecraft can return the sample back to the Earth in only 8 years. Another window that is open in 2009–2010 is also provided. They are making use of the ion engine propulsion accumulating the orbital energy at Earth distance, followed by the Earth gravity assist. This is the unique point enabling even the smaller spacecraft to accomplish this high energy mission. This strategy accelerates the spacecraft orbital energy up to 36– 40 km 2 / s 2 in C3 that is adequate for the flight to the comet WH. The idea shown here also includes: (1) the rendezvous point to be at Earth distance so that the electric propulsion is utilized at maximum and (2) the target comet to be extinct so that the sampling and in situ measurement scenario robust. The trajectory synthesis developed here and tactical target selection make this impressive mission to be envisioned even by smaller missions. This is not an endorsed mission but ISAS may consider collaborating with the foreign space organization on the mission.