Showing papers on "Spacecraft propulsion published in 2008"

Space Applications of High-Power Microwaves

Abstract: Schemes have been suggested for transferring energy from Earth-to-space, space-to-Earth, and space-to-space using high-power microwave (HPM) beams. All use power beaming. Microwave beams have been studied for propelling spacecraft for launch to orbit, orbit raising, launch from orbit into interplanetary and interstellar space, and deployment of large space structures. The microwave thermal rocket, called the ldquomicrowave thermal thruster,rdquo is a reusable single-stage vehicle that uses an HPM beam to provide power to a heat-exchanger propulsion system, with double the specific impulse of conventional rockets. Orbital missions include orbit raising and space solar power. Microwave-propelled sails are a new class of spacecraft that promises to revolutionize future space probes. Experiments and simulations have verified that sails riding beams can be stable on the beam for conical sail shapes. Beam-driven sail flights have now demonstrated the basic features of the beam-driven propulsion. Beams can also carry angular momentum and communicate it to a sail to help control it in flight. An early mission for microwave space propulsion is dramatically shortening the time needed for sails to escape Earth's orbit. A number of missions for beam-driven sails have been quantified for high-velocity mapping of the outer solar system, Kuiper Belt, the Heliopause, and the penultimate interstellar precursor mission. For large HPM systems at fixed effective isotropic radiated power, minimum capital cost is achieved when the cost is equally divided between antenna gain and radiated power. This is a driver when considering design of power-beaming systems such as interstellar Beacons, which the Search for Extraterrestrial Intelligence is searching for. Much of the technical means for these applications are already in hand. Microwave and millimeter-wave array antennas are already in use for astronomy; sources at high frequencies are being developed for fusion and the military. Development of high-power arrays is needed. A synergistic way to develop a space power-beaming infrastructure is incremental buildup, addressing lower power applications first, and then upgrading.

Solar Sails: A Novel Approach to Interplanetary Travel

Abstract: Acknowledgements.- Preface.- Foreword. Part I: Space Engines: Past and Present.- A Historical Introduction to Space Propulsion.- The Rocket: How It Works in Space.- Rocket Problems and Limitations- Non-Rocket-In-Space Propulsion.- The Solar-Sail Reality: from the Oceans to Space.- Part II Space Mission by Sail.- Principles of Space Sailing.- What is a Space Sailcraft?- Sails vs. Rockets.- Exploring and Developing Space by Sailcraft.- Riding a Beam of Light.- Part III Construction of Sailcraft.- Designing a Solar Sail.- Building a Sailcraft.- Progress to Date.- Future Plans.- Part IV Breakthroughs in Space.- The IKAROS/JAXA Mission.- The NanoSail-D2/NASA Mission.- New Projects in Progress.- Part V Space Sailing: Some Technical Aspects.- Space Sources of Light.- Modeling Thrust via Electromagnetic Radiation Pressure and Diffraction- Sailcraft Trajectories.- Sails in Space Environment.

Permanent Magnet Helicon Source for Ion Propulsion

Abstract: Helicon sources have been proposed by at least two groups for generating ions for space propulsion: the Helicon Double Layer Thruster (HDLT) concept at the Australian National University and the Variable Specific Impulse Magnetohydrodynamic Rocket (VASIMR) concept at the Johnson Space Center in Houston. These sources normally require a large electromagnet and power supply to produce the magnetic field. At this stage of research, emphasis has been on the plasma density and ion current that can be produced, but not much on the weight, size, impulse, and gas efficiency of the thruster. This paper concerns the source itself and shows that great savings in size and weight can be obtained by using specially designed permanent magnets (PMs). This PM helicon design, originally developed for plasma processing of large substrates, is extended here for ion thrusters of both the HDLT and VASIMR types. Measured downstream densities are on the order of 1012 cm-3 , which should yield much higher ion currents than reported so far. The design principles have been checked experimentally, showing that the predictions of the theory and computations are reliable. The details of two new designs are given here to serve as examples to stimulate further research on the use of such sources as thrusters.

Rocket and Spacecraft Propulsion

Abstract: History and principles of rocket propulsion -- The thermal rocket engine -- Liquid propellant rocket engines -- Solid propellant rocket motors -- Launch vehicle dynamics -- Electric propulsion -- Nuclear propulsion -- Advanced thermal rockets.The revised edition of this practical, hands-on book discusses the range of launch vehicles in use today throughout the world, and includes the very latest details of some of the advanced propulsion systems currently being developed. The author covers the fundamentals of the subject, from the basic principles of rocket propulsion and vehicle dynamics through the theory and practice of liquid and solid propellant motors, to new and future developments. The revised edition will stick to the same principle of providing a serious exposition of the principles and practice of rocket propulsion, but from the point of view of the user and enquirer who is not an engineering specialist. Most chapters will remain substantially the same as the first edition; they will be updated where necessary and errata corrected. The main revisions will be to the chapter on electric propulsion where there have been significant new developments both in engine types and in practical applications. This is now seen as the key to planetary exploration by robotic probes and should therefore be reflected. Nuclear propulsion has emerged from the doldrums and is now seen as a definite possibility for outer solar system robotic exploration; and as enabling technology for a human mars expedition. A new chapter on nuclear thermal propulsion has been added to reflect this revival of interest.

Dawn ion propulsion system – initial checkout after launch

Abstract: The first 80 days after launch of the Dawn mission were dedicated to the checkout of the spacecraft with a major emphasis on the ion propulsion system. All three ion thrusters, all three thruster-gimbal assemblies, both power processor units, both digital interface and control units, and the entire xenon feed system were completely checked out and every component was found to be in good health. Direct thrust measurements agreed well with preflight expected values for all three thrusters over the entire throttle range. Thruster electrical operating parameters and power processor units efficiencies also agreed well with preflight expected values based on acceptance test data. Two of the three ion thrusters were fully checked out within 30 days after launch. Checkout of all three thrusters was completed 64 days after launch. Deterministic thrusting with the IPS began on December 17, 2007.

Options for the human exploration of Mars using Solar Electric propulsion

Abstract: Solar Electric propulsion (SEP) is examined as a candidate transportation option for human missions to Mars. Focus is given to an Earth-escape staging concept. This concept uses a SEP system to transfer from low earth orbit (LEO) to a high-energy elliptical parking orbit (HEEPO) and a chemical propulsion system to transfer from the HEEPO to a hyperbolic escape trajectory. LEO to Earth escape performance of these combined transportation systems is comparable to that of a nuclear thermal rocket (NTR). As a result, a mass efficient non-nuclear transportation architecture with fast, 180 day, Earth-to-Mars piloted transit times is enabled.

Geosynchronous-Earth-Orbit Communication Satellite Deliveries with Integrated Electric Propulsion

Abstract: DOI: 10.2514/1.35322 The use of electric propulsion was evaluated for transfer of communication satellites from geosynchronous transfer orbits to geosynchronous earth orbits. Recent communication satellite designs, normal launch vehicle delivery orbits, and integrated electric propulsion subsystems (with input powers less than nominal satellite power levels) were assumed to minimize required changes to present and near-term launchers and spacecraft. The capture fraction of recent communication satellites that could have been delivered was evaluated versus launcher delivery capability, launch site, and in-space propulsion characteristics. Electric propulsion significantly increases the capturefractionoflauncherswithgeosynchronoustransfermassdeliverycapabilitieslessthan(dependentonlaunch site) about 4500 to 5500 kg. Insertion times at given launch sites were found to be accurately specified by the satellite power-to-mass ratiosandtheassumedelectricpropulsionspecific-impulse/efficiencycharacteristics. Insertiontimes less than 100 days were found for satellites with high power-to-mass ratios that used high thrust-to-power electric propulsion options. The influence of power for electric propulsion beyond that used for satellite payloads and housekeeping was also assessed, and insertion times less than a month will require powers significantly higher than presently installed.

Analysis of Displaced Solar Sail Orbits with Passive Control

Abstract: T HEuse of solar radiation pressure wasfirst proposed by a Soviet pioneer of astronautics, Tsiolkovski, and the technology was greatly developed by NASA for a proposed comet Halley rendezvous mission in the 1970s [1,2]. Recently, many space applications of solar sails are proposed because solar sails enable some special missions which would be impossible for any conventional space propulsion. Such missions include displaced solar orbits, geocentric halo orbits, Mercury sun-synchronous polar orbit, artificial Lagrange points, and so on. Leipold and Wagner investigated the Mercury sun-synchronous polar orbit using solar sail propulsion to explore the inner solar system [3]. West investigated the new artificial Lagrange points created by solar sails to provide early warning of solar plasma storms before they reach the Earth [4]. McInnes and Simmons have done much work on the dynamics and control of solar sails on different exotic trajectories [5,6]. The stability of solar sails on displaced solar orbitswith passive control is investigated in [7], and the results show that the sails are stable if the sail pitch angle is fixed with respect to a rotating frame. The passive stability can be realized by designing the configuration of the sail, which is investigated in [8]. Passive control is a good option for the solar sail because its large and complex structure may introduce some difficulties for active control. In this Note, the global stability of the solar sail with passive control is investigated by considering the dynamics in an inertial frame. It is found that the sail is stable with any initial values, and the sail will oscillate in the vicinity of a nominal orbit that is uniquely determined by the angular momentum of the sail. The amplitudes of the oscillations are determined by the initial values of the radius and angular velocity.

The QED engine system: Direct‐electric fusion‐powered rocket propulsion systems

Abstract: Practical ground‐to‐orbit and inter‐orbital space flights both require propulsion systems of large flight‐path‐averaged specific impulse (Isp) and engine system thrust‐to‐mass‐ratio (F/me=[F]) for useful payload and structure fractions in single‐stage vehicles (Hunter 1966). Current rocket and air‐breathing engine technologies lead to enormous vehicles and small payloads; a natural result of the limited specific energy available from chemical reactions. While nuclear energy far exceeds these specific energy limits (Bussard and DeLauer 1958), the inherent high‐Isp advantages of fission propulsion concepts for space and air‐breathing flight (Bussard and DeLauer 1965) are negated for manned systems by the massive radiation shielding required by their high radiation output (Bussard 1971). However, there are well‐known radiation‐free nuclear fusion reactions (Gross 1984) between isotopes of selected light elements (such as H+11B, D+3He) that yield only energetic charged particles, whose energy can be converted directly into electricity by confining electric fields (Moir and Barr 1973,1983). New confinement concepts using magnetic‐electric‐potentials (Bussard 1989a) or inertial‐collisional‐compression (ICC) (Bussard 1990) have been found that offer the prospect of clean, compact fusion systems with very high output and low mass. Their radiation‐free d.c. electrical output can power unique new electron‐beam‐driven thrust systems of extremely high performance. Parametric design studies show that such charged‐particle electric‐discharge engines (‘‘QED’’ engines) might yield rocket propulsion systems with performance in the ranges of 2<[F]<6 and 1500

Laser-powered multi-newton thrust space engine with variable specific impulse

Abstract: Recently we became interested in applying previous work with liquid fueled laser powered minithrusters for spacecraft orientation to the conceptual design of a multi-newton thruster based on the same principles. Solid-fuel configurations (such as the fuel tapes used in the Photonic Associates microthruster) are not amenable to the range of mass delivery rates (g/s to g/s) necessary for such an engine. We will discuss problems for this design which have been solved, including identifying a practical method of delivering liquid fuel to the laser focus, avoiding splashing of liquid fuels under pulsed laser illumination, and avoiding optics clouding due to ablation backstreaming on optical surfaces from the laser-fuel interaction region. We have already shown that Isp = 680 seconds can be achieved by a viscous liquid fuel based on glycidyl azide polymer and an IR-dye laser absorber. The final problem is mass: we will discuss a notional engine design which fits within a 10-kg "dry mass" budget. This engine, 80kg mass with fuel, is designed to fit within a 180-kg spacecraft, and use 3kW of prime power to deliver a Δv of 17.5 km/s to the spacecraft in sixteen months. Its specific impulse will be adjustable over the range 200

Optimal trajectories towards near-earth-objects using solar electric propulsion (sep) and gravity assisted maneuver

Abstract: The solar electric propulsion could be the best option for the transports of the future due to its high specific impulse when compared to the chemical propulsion. Electric propellants are being extensively used to assist the propulsion of terrestrial satellites for the maneuvers of orbit correction and as primary propulsion in missions toward other bodies of the solar system. In this work the optimization of interplanetary missions using solar electric propulsion (SEP) and Gravity Assisted Maneuver to reduce the costs of the mission, is considered. The high specific impulse of electric propulsion makes a Gravity Assisted Maneuver 1 year after departure convenient. Missions for several Near Earth Asteroids will be considered. The analysis suggests criteria for the definition of initial solutions demanded for the process of optimization of trajectories. Trajectories to the asteroid 2002TC70 are analyzed. Direct trajectories, trajectories with 1 gravity assisted at the Earth and with 2 gravity assisted with the Earth and either Mars are presented. Shall be analyzed missions with thrusters PPS1350 and the Phall 1 for performance comparison. An indirect optimization method will be used in the simulations.

Characterizing high-energy-density propellants for space propulsion applications

Abstract: A technique for computationally determining the thermophysical properties of high-energy-density matter propellants is presented. High-energy-density matter compounds are of interest in the liquid rocket engine industry due to their high-density and high-energy content relative to existing industry-standard propellants. To accurately model rocket engine performance, cost, and weight in a conceptual design environment, several thermodynamic and physical properties are required over a range of temperatures and pressures. The approach presented here combines quantum mechanical and molecular dynamic calculations and group additivity methods. A method for improving the force field model coefficients used in the molecular dynamics simulations is included. This approach is used to determine thermophysical properties for two high-energy-density matter compounds of interest: quadricyclane and 2-azido-N, N-dimethylethanamine. The modified force field approach provides results that more accurately match experimental data than the unmodified approach. Launch vehicle and lunar lander case studies are presented to quantify the system-level impact of employing quadricyclane and 2-azido-N, N-dimethylethanamine rather than industry-standard propellants.In both cases, the use of high-energy-density matter propellants provides reductions in vehicle mass compared with industry-standard propellants. The results demonstrate that high-energy-density matter propellants can be an attractive technology for future launch vehicle and lunar lander applications.

A concept of ferroelectric microparticle propulsion thruster

Abstract: A space propulsion concept using charged ferroelectric microparticles as a propellant is suggested. The measured ferroelectric plasma source thrust, produced mainly by microparticles emission, reaches ∼9×10−4N. The obtained trajectories of microparticles demonstrate that the majority of the microparticles are positively charged, which permits further improvement of the thruster.

Simplest AB‐thermonuclear space propulsion and electric generator

Abstract: Purpose – This paper aims to suggest a new thermonuclear space propulsion and electric generator for aerospace.Design/methodology/approach – Methods of thermonuclear physics are used for research.Findings – The paper applies, develops and researches mini‐sized Micro‐AB thermonuclear reactors for space propulsion and space power systems. These small engines directly convert the high‐speed charged particles produced in the thermonuclear reactor into vehicle thrust or vehicle electricity with maximum efficiency. The simplest AB‐thermonuclear propulsion offered allows spaceships to reach speeds of 20,000‐50,000 km/s (1/6 of light speed) for fuel ratio 0.1 and produces a huge amount of useful electric energy. The offered propulsion system permits flight to any planet of the solar system in a short time and to the nearest non‐Sun stars by E‐being or intellectual robots during a single human life period.Research limitations/implications – Technical limitations may be apparent.Originality/value – The theory of th...

Cost-Efficient Storage of Cryogens

Abstract: NASA's cryogenic infrastructure that supports launch vehicle operations and propulsion testing is reaching an age where major refurbishment will soon be required. Key elements of this infrastructure are the large double-walled cryogenic storage tanks used for both space vehicle launch operations and rocket propulsion testing at the various NASA field centers. Perlite powder has historically been the insulation material of choice for these large storage tank applications. New bulk-fill insulation materials, including glass bubbles and aerogel beads, have been shown to provide improved thermal and mechanical performance. A research testing program was conducted to investigate the thermal performance benefits as well as to identify operational considerations and associated risks associated with the application of these new materials in large cryogenic storage tanks. The program was divided into three main areas: material testing (thermal conductivity and physical characterization), tank demonstration testing (liquid nitrogen and liquid hydrogen), and system studies (thermal modeling, economic analysis, and insulation changeout). The results of this research work show that more energy-efficient insulation solutions are possible for large-scale cryogenic storage tanks worldwide and summarize the operational requirements that should be considered for these applications.

VASIMR ® VX-200: High power electric propulsion for space transportation beyond LEO

Abstract: () The Variable Specific Impulse Magnetoplasma Rocket (VASIMR™) is a high power magnetoplasma rocket, capable of Isp/thrust modulation at constant power. The plasma is produced by a helicon discharge. The bulk of the energy is added by ion cyclotron resonance heating (ICRH.) Axial momentum is obtained by adiabatic expansion of the plasma in a magnetic nozzle. Thrust/specific impulse ratio control in the VASIMR™ is primarily achieved by the partitioning of the RF power to the helicon and ICRH systems, with the proper adjustment of the propellant flow. Ion dynamics in the exhaust were studied using probes, gridded energy analyzers (RPA's), microwave interferometry and optical techniques. This paper will summarize results from high power ICRH experiments performed on the VX-100 using argon plasma during 2007. An overview of the way forward will be touched on briefly, with some emphasis on the fact that VASIMR™ is now being developed by private enterprise. The opportunities and challenges of this situation will be reviewed. VASIMR™ was originally designed to serve as the sustainer engines for the manned Mars mission, for which it remains one of the leading candidate systems. A number of other uses of the VASIMR™ have been studied in some detail, including robotic missions to the outer planets, ISSO reboost, station keeping and sustained maneuvering of robotic craft in Earth orbit, and lunar cargo hauling. We will explore the latter in this paper. A number of studies have illustrated the cost and mass efficiency of solar-electric propulsion as an alternative to chemical propulsion for hauling cargo from low Earth orbit to low lunar orbit; recent studies considered the Hall thruster in this application. Here, we present the results of a payload vs. specific impulse trade study for a six-month Earth-Moon transit time, and compare the technical and economic features of the VASIMR™ technology to other electric thrusters for this application.

Ultra-compact, high-performance aerovortical rocket thruster

Abstract: An ultra-compact aerovortical swirl-enhanced combustion (ASC) system features an aerovortical swirl generator for use in rocket thrusters utilizing hypergolic or non-hypergolic propellants. The ACS thruster can be sized for diameters ranging from about 0.5 to about 2.0 inches, and producing thrust levels of approximately 5 lb f to about 250 lb f . A plurality of helicoid flow channels in the swirl generator introduces swirl into a flow stream of a first propellant within ultra-compact sized rocket thrusters. The ASC system also includes injectors for introducing a second liquid propellant into the swirling flowfield to promote rapid and efficient atomization, mixing and vigorous combustion, which, results in major improvements in combustion and propulsion performance over current rocket thrusters, but in much shorter combustor systems. Hence, the ultra-compact ASC system is a substantial improvement in small bipropellant chemical propulsion thrusters, which can be utilized in-space satellite, spacecraft maneuvering and attitude/orbit control.

Solar Sail Technology for Nanosatellites

Abstract: Solar sailing is an attractive means of spacecraft propulsion because it extracts momentum from electromagnetic radiation supplied by the Sun. This allows a solar sail spacecraft to accomplish new classes of missions that would otherwise require a prohibitive amount of propellant. Unfortunately, solar radiation pressure is meager, requiring a large area-tomass ratio to produce significant accelerations. A solar sailing nanosatellite achieves this with a dramatically reduced mass allowing a smaller sail to generate significant accelerations. This sacrifice of slightly less capability for an immense gain in reduced cost and complexity is the motivation behind Stanfords solar sailing nanosatellite project, SailSat.

An analytical optimization method for electric propulsion orbit transfer vehicles

Abstract: Due to electric propulsion’s inherent propellant mass savings over chemical propulsion, electric propulsion orbit transfer vehicles (EPOTVs) are a highly efficient mode of orbit transfer. When selecting an electric propulsion device (ion, MPD, or arcjet) and propellant for a particular mission, it is preferable to use quick, analytical system optimization methods instead of time intensive numerical integration methods. It is also of interest to determine each thruster’s optimal operating characteristics for a specific mission. Analytical expressions are derived which determine the optimal specific impulse (Isp) for each type of electric thruster to maximize payload fraction for a desired thrusting time. These expressions take into account the variation of thruster efficiency with specific impulse. Verification of the method is made with representative electric propulsion values on a LEO‐to‐GEO mission. Application of the method to specific missions is discussed.

Mission Assessment of the Faraday Accelerator with Radio-frequency Assisted Discharge (FARAD)

Abstract: Pulsed inductive thrusters have typically been considered for future, high-power, missions requiring nuclear electric propulsion. These high-power systems, while promising equivalent or improved performance over state-of-the-art propulsion systems, presently have no planned missions for which they are well suited. The ability to efficiently operate an inductive thruster at lower energy and power levels may provide inductive thrusters near term applicability and mission pull. The Faraday Accelerator with Radio-frequency Assisted Discharge concept demonstrated potential for a high-efficiency, low-energy pulsed inductive thruster. The added benefits of energy recapture and/or pulse compression are shown to enhance the performance of the pulsed inductive propulsion system, yielding a system that con compete with and potentially outperform current state-of-the-art electric propulsion technologies. These enhancements lead to mission-level benefits associated with the use of a pulsed inductive thruster. Analyses of low-power near to mid-term missions and higher power far-term missions are undertaken to compare the performance of pulsed inductive thrusters with that delivered by state-of-the-art and development-level electric propulsion systems.

Rocket based combined cycle propulsion unit having external rocket thrusters

Abstract: The present disclosure generally pertains to rocket based combined cycle (RBCC) propulsion units. In one exemplary embodiment, at least one rocket thruster is integrated with a jet engine but is external to the flow path of the jet engine, forming an altitude compensating plug nozzle. Since the rocket thruster is external to such flow path, the rocket flow from the rocket thruster interacts with the jet flow from the jet engine aft of the nozzle of the jet engine. Such interaction occurs without a significant performance penalty in the operation of the jet engine. In fact, it is possible that the interaction of the rocket flow with the jet flow may actually improve the efficiency of the jet engine under some conditions. Moreover, having the rocket thrusters positioned external to the flow path of the jet engine helps to avoid many of the problems plaguing conventional RBCC propulsion units.

A Spherical Torus Nuclear Fusion Reactor Space Propulsion Vehicle Concept for Fast Interplanetary Travel

Abstract: A conceptual vehicle design enabling fast outer solar system travel was produced predicated on a small aspect ratio spherical torus nuclear fusion reactor. Initial requirements were for a human mission to Saturn with a>5% payload mass fraction and a one way trip time of less than one year. Analysis revealed that the vehicle could deliver a 108 mt crew habitat payload to Saturn rendezvous in 235 days, with an initial mass in low Earth orbit of 2,941 mt. Engineering conceptual design, analysis, and assessment was performed on all major systems including payload, central truss, nuclear reactor (including diverter and fuel injector), power conversion (including turbine, compressor, alternator, radiator, recuperator, and conditioning), magnetic nozzle, neutral beam injector, tankage, start/re-start reactor and battery, refrigeration, communications, reaction control, and in-space operations. Detailed assessment was done on reactor operations, including plasma characteristics, power balance, and component design.

Status Update Report for the Peregrine 100km Sounding Rocket Project

Abstract: The Peregrine Sounding Rocket Program is a joint basic research program of NASA Ames Research Center, NASA Wallops, Stanford University and the Space Propulsion Group, Inc. (SPG). The goal is to determine the applicability of liquifying hybrid technology to a small launch system. The approach is to design, build, test and y a stable, efficient liquefying fuel hybrid rocket vehicle to an altitude of 100 km. The program was kicked o in October of 2006 and has seen considerable progress in the subsequent 18 months. Two virtually identical vehicles will be constructed and own out of the NASA Sounding Rocket Facility at Wallops Island. This paper presents the current status of the project as of June 2008. For background on the project, the reader is referred to last year's paper.

Fit for mission - design tailoring aspects of throttleable ducted rocket propulsion systems

Abstract: () A series of successful flight tests of Meteor (Fig. 1) with a development standard Boron loaded Throttleable Ducted Rocket (TDR) has demonstrated the capability and maturity of this superior propulsion system. Meteor is the medium to long range air to air missile, selected by six European nations (UK, Fr, It, Sp Swe, GE) for air superiority of their fighter aircraft Gripen, Rafale and EF2000 Typhoon. The paper summarizes the basic function of the TDR, the technical status and the major design drivers. Based on the existing technology from Meteor and multiple additional technology programmes, TDR propulsion systems for alternative missions are outlined with respect to the discussed design driving parameters. The Boron loaded TDR is a most promising candidate for future missile propulsion, providing a superior kinematic performance and mission flexibility.

NASA's In-Space Propulsion Technology Project Overview and Mission Applicability

Abstract: The In-Space Propulsion Technology Project, funded by NASA's Science Mission Directorate (SMD), is continuing to invest in propulsion technologies that will enable or enhance NASA robotic science missions. This paper provides development status, near-term mission benefits, applicability, and availability of in-space propulsion technologies in the areas of aerocapture, electric propulsion, and advanced chemical thrusters. Aerocapture investments have 1) improved models for: guidance, navigation, and control of blunt body rigid aeroshells, 2) atmospheric models for Earth, Titan, Mars and Venus, and 3) models for aerothermal effects. Investments in electric propulsion technologies have focused on completing the NEXT ion propulsion system, a 0.6-7kW throttle-able gridded ion system. The primary chemical propulsion investment is on a high-temperature storable bi-propellant rocket engine providing higher performance for lower cost. Development status of mid-term technology, the low-cost HiVHAC Hall thruster is also presented. In-space propulsion technologies are applicable, and potentially enabling for flagship destinations currently under evaluation, as well as having broad applicability to future Discovery and New Frontiers mission solicitations.