The Double Asteroid Redirection Test (DART) spacecraft was developed to provide the first measurement for orbital deflection of an asteroid upon intentional impact. The NEXT ion engine is part of the mission, on its maiden voyage. As part of the pre-launch risk reduction, erosion characteristics of the extraction grid system were evaluated using laser measurements of sputtered molybdenum atoms over the envelope of potential throttle conditions for the mission. Erosion rate dependence on propellant flow rate as well as relative density and directionality of molybdenum sputter from grid center to edge were measured. Sputtered atoms were found to have average radial velocity directed toward the engine perimeter and increasing with radial distance. The relative contribution of source and facility background gas and other sources of accelerator grid current was examined as well as the influence of several engine operating parameters. Facility background gas was found to influence engine operation more than a wall-mounted pressure gauge and typical assumptions about ingestion would indicate. Far-field flux was estimated over the full angular range based on the near-field relative density and velocity results and relying on quartz crystal microbalance data at one location to fix absolute numbers everywhere. The results substantially deepen knowledge and understanding of the complex grid erosion process of the engine and its lifetime, as grid failure via erosion is the normal life limiter. Study results are also relevant to thruster–spacecraft integration issues such as molybdenum deposition rate on solar cells and other spacecraft surfaces.

Dynetics, Inc., 1000 Explorer Blvd NW, Huntsville, AL 35806, jimmy.allen@dynetics.com Southern Research, 757 Tom Martin Dr. Birmingham, AL 35211, mjohns@southernresearch.org, mpatterson@southernresearch.org NASA Marshall Space Flight Center, Huntsville, AL 35812, michael.houts@nasa.gov Argonne National Laboratory, 9700 Cass Ave., Argonne, IL 60439, fheidet@anl.gov, mjarrett@anl.gov Idaho National Laboratory, 1955 N Fremont Ave, Idaho Falls, ID 83415, nicholas.smith2@inl.gov The University of Michigan, 500 S. State Street, Ann Arbor, MI 48109, jefoster@umich.edu,

In support of the Double Asteroid Redirection Test (DART) mission, laboratory measurements were made on the NEXT ion engine, which will be used for the spacecraft's in-space propulsion [1]. This study revisits a small range of mission-specific 2.7A throttle levels to understand the effect of in-flight flow rate variability, investigate intermediate throttle conditions, and improve measurement methodology. This paper specifically examines the far-field plume divergence and backflow ion flux distribution of the NEXT, while a companion paper examines the charge state distributions.

Risk mitigation tests have been conducted by the NASA Glenn Research Center and The Aerospace Corporation in support of the DART Mission. The tests focused on NEXT performance characterizations intended to ensure its operations and characteristics are compatible with the DART mission operations, and to assist in the definition of the propulsion system. Tests were performed at the Aerospace Corporation and they involved: flow sensitivity-analyses, steady-state performance characterizations, and measurements of thruster erosion. The tests also involved defining, demonstrating, verifying, and evaluating the start-up sequences and a beam current regulation algorithm consistent with DART mission requirements. It was found that NEXT thruster operations are compatible with the proposed relaxation of flow control ranges for ignition and for steady-state operation.

Electric propulsion (EP) is an important technology for NASA. It has played a major role on three missions, namely Deep Space 1, Dawn and Space Technology 7, and it is planned for use on many more. The ion propulsion system for the ongoing Dawn mission has achieved several notable accomplishments, including providing a total velocity change (delta-V) of over 11 km/s to the spacecraft. As a result of these successes, solar electric propulsion (SEP) is now broadly recognized as an essential technology for both robotic and human exploration. NASA is currently conducting many projects focused on research and development of EP for a variety of applications. All three of NASA's mission directorates that deal directly with space exploration are actively engaged in supporting work in this area. This paper describes these projects in more detail, including the specific engineering activities being conducted at NASA's main centers for EP technology development, namely Glenn Research Center (GRC) and the Jet Propulsion Laboratory (JPL).

The Annular Ion Engine concept represents an evolutionary development in gridded ion thruster technology with the potential for delivering revolutionary capabilities.

The Annular Ion Engine (AIE) concept represents an evolutionary development in gridded ion thruster technology with the potential for delivering revolutionary capabilities. It has this potential because the AIE concept: (a) enables scaling of ion thruster technology to high power at specific impulse (Isp) values of interest for near-term mission applications, ≤ 5000 sec; and (b) it enables an increase in both thrust density and thrust-topower (F/P) ratio exceeding conventional ion thrusters and other electric propulsion (EP) technology options, thereby yielding the highest performance over a broad range in Isp. The AIE concept represents a natural progression of gridded ion thruster technology beyond the capabilities embodied by NASA’s Evolutionary Xenon Thruster (NEXT) [1]. The AIE would be appropriate for: (a) applications which require power levels exceeding NEXT’s capabilities (up to about 14 kW [2]), with scalability potentially to 100’s of kW; and/or (b) applications which require F/P conditions exceeding NEXT’s capabilities. This publication synopsizes the technology development status of the AIE and a potential forward-path for further concept maturation. 1. CONCEPT DESCRIPTION All concepts disclosed in the publication are covered under U.S. Patent #8,468,794 (“Electric Propulsion Apparatus”, June 25, 2013), U.S. Patent #9,279,368 (“Multi-Thruster Propulsion Apparatus”, March 29, 2016), and PatentsPending filed under both U.S. and International Patent Applications. Assignee: United State Government. The AIE concept consists of an annular discharge chamber, with a set of annular ion optics, potentially configured with a centrally-mounted neutralizer cathode assembly. This concept is illustrated in Fig. 1. Fig. 2 shows a secondgeneration (‘GEN2’) full-scale AIE under test with beam extraction [3]. Figure 1. AIE Concept in Cross-Section and Frontal View 2. CONCEPT ADVANTAGES The AIE has significant potential advantages over conventional ion thrusters, and other EP thruster concepts [4]. Attributes and technical descriptions of the AIE are listed in Tab. 1, and include: very high thrust density; extensibility to high power; improved efficiencies and F/P; enhanced life time; and improved packaging over State-of-the-Art (SOA) EP thrusters. Figure 2. Full-Scale GEN2 AIE in Test; 65 cm outside Beam Diameter https://ntrs.nasa.gov/search.jsp?R=20170000387 2020-01-05T12:51:35+00:00Z

Gridded ion engines have the highest efficiency and total impulse of any mature electric propulsion technology, and have been successfully implemented for primary propulsion in both geocentric and heliocentric environments with excellent ground/in-space correlation of performance. However, they have not been optimized to maximize thrust-to-power, an important parameter for Earth orbit transfer applications. This publication discusses technology development work intended to maximize this parameter. These activities include investigating the capabilities of a non-conventional design approach, the annular engine, which has the potential of exceeding the thrust-to-power of other EP technologies. This publication discusses the status of this work, including the fabrication and initial tests of a large-area annular engine. This work is being conducted in collaboration among NASA Glenn Research Center, The Aerospace Corporation, and the University of Michigan.

The annular ion engine concept consists of a cylindrical ion thruster with a centrally located stalk that provides increased anode area. The increased electron collection area allows for increased discharge plasma current, and thus high power operation. The central stalk also provides support for the ion optics, allowing for larger beam area and scale up potential. Discharge performance of a 42 cm annular ion engine has previously been studied. Here, the discharge uniformity of a 65 cm annular ion engine during simulated beam extraction is assessed using three diagnostics: Faraday probes, Langmuir probes, and a fast camera. The percent uniformity was found to vary from 89.8% to 97.4% for discharge powers under 1.5 kW. The data presented herein demonstrates the feasibility of scalability of the annular ion engine. PhD candidate, Nuclear Engineering and Radiological Sciences, 2355 Bonisteel Blvd., AIAA student member Professor, Nuclear Engineering and Radiological Sciences, 2355 Bonisteel Blvd., AIAA member Senior Technologist, Power and In-Space propulsion Division, 2100 Brookpark Rd./MS 301-3, AIAA senior member Research Engineer, Propulsion and Propellants Branch, 21000 Brookpark Rd./MS 301-3, AIAA member Member of Technical Sta↵, Space Materials Laboratory, P.O. Box 92957 M2-341, AIAA member Senior Scientist, Space Materials Laboratory, P.O. Box 92957 M2-341, AIAA senior member

Gridded ion thrusters provide excellent thruster performance and have successfully been implemented on both geocentric and heliocentric missions. While ion thrusters have a substantial number of attractive technological attributes, they are often classified as inherently low thrust density devices. This manuscript details an ongoing collaborative effort among the NASA Glenn Research Center, the University of Michigan, and The Aerospace Corporation investigating ion engine design modifications for high thrust-density/high thrust-topower operation. Measurements were performed at The Aerospace Corporation in a 2.4-m diameter × 9.8-m long cryopumped vacuum chamber on an engineering model NEXT engine with a reduced interelectrode gap. The perveance, discharge losses, and far-field current density were characterized at operating conditions consistent with high thrust-to-power operation. The total voltage needed to achieve a given beam current was reduced by a factor of 10% with the reduced grid-gap optics, which is in close agreement with the Child Langmuir equation. The thrust loss correction factor ranged from 0.961 to 0.979 and was consistently higher than the predicted values. The discharge losses decreased with increasing beam current, with a minimum value of 150 W/A at a beam current of 5.50 A.

The atmospheric degradation of piperazine was investigated using an indoor smog chamber. Experiments were carried out in the presence of nitrogen oxides (NOx), ozone or nitric acid. Piperazine reacted rapidly under all evaluated conditions: irradiated in the presence of NOx and with ozone and nitric acid in the dark. Gas phase products from the oxidation of piperazine were identified by infrared spectroscopy, DNPH cartridges followed by HPLC analysis, and by sampling chamber gas through Tenax sorbent material followed by analysis using thermal desorption GC-ITMS (gas chromatography ion trap mass spectrometry).Eight compounds were positively identified, with a further nine compounds tentatively identified using GC-MS based on molecular weight and mass spectra. Ammonia formation was observed from piperazine oxidation, and its formation was from the subsequent reactions of photooxidation products of piperazine rather than directly from the reaction of piperazine. The nitrosamine and nitramine expected from piperazine, N-nitrosopiperazine, and N-nitropiperazine, were both identified and confirmed using (NO)-N-15, with a tentative maximum yield of nitrosamine of less than 5% observed.Aerosol yields, relative to total piperazine reacted not including that which absorbed to the walls, were considerably high but were not able to be quantified absolutely due to unusual behaviour of the scanning mobility particle sizer instrument to aerosol containing amines. The reaction of piperazine with gas phase nitric acid gave rise to immediate formation of aerosol. (C) 2015 Elsevier Ltd. All rights reserved.

This presentation describes results from the end-of-test performance characterization of NASA's Evolutionary Xenon Thruster (NEXT) Long-Duration Test (LDT). Sub-component performance as well as overall thruster performance is presented and compared to results over the course of the test. Overall wear of critical thruster components is also described, and an update on the first failure mode of the thruster is provided.

Since 2001, the In-Space Propulsion Technology (ISPT) program has been developing and delivering in-space propulsion technologies for NASA's Science Mission Directorate (SMD). These in-space propulsion technologies are applicable, and potentially enabling for future NASA Discovery, New Frontiers, Flagship and sample return missions currently under consideration. The ISPT program is currently developing technology in three areas that include Propulsion System Technologies, Entry Vehicle Technologies, and Systems Mission Analysis. ISPT's propulsion technologies include: 1) the 0.6-7 kW NASA's Evolutionary Xenon Thruster (NEXT) gridded ion propulsion system; 2) a 0.3-3.9kW Hall-effect electric propulsion (HEP) system for low cost and sample return missions; 3) the Xenon Flow Control Module (XFCM); 4) ultra-lightweight propellant tank technologies (ULTT); and 5) propulsion technologies for a Mars Ascent Vehicle (MAV). The HEP system is composed of the High Voltage Hall Accelerator (HiVHAc) thruster, a power processing unit (PPU), and the XFCM. NEXT and the HiVHAc are throttle-able electric propulsion systems for planetary science missions. The XFCM and ULTT are two component technologies which being developed with nearer-term flight infusion in mind. Several of the ISPT technologies are related to sample return missions needs like: MAV propulsion and electric propulsion. And finally, one focus of the SystemsMission Analysis area is developing tools that aid the application or operation of these technologies on wide variety of mission concepts. This paper provides a brief overview of the ISPT program, describing the development status and technology infusion readiness.

Ion thruster technology offers the highest performance and efficiency of any mature electric propulsion thruster.It has by far the highest demonstrated total impulse of any technology option, demonstrated at input power levels appropriate for primary propulsion.It has also been successfully implemented for primary propulsion in both geocentric and heliocentric environments, with excellent ground/in-space correlation of both its performance and life.Based on these attributes there is compelling reasoning to continue the development of this technology: it is a leading candidate for high power applications; and it provides risk reduction for as -yet unproven alternatives.As such it is important that the operational limitations of ion thruster technology be critically examinedand in particular for its application to primary propulsionits capabilities relative to thrust density and thrust-to-power ratio be understood.This publication briefly addresses some of the considerations relative to achieving high thrust density and maximizing thrust-to-power ratio with ion thruster technology, and discusses the status of development work in this area being executed under a collaborative effort among

AbstractThis paper provides a brief overview of the history of electric propulsion (EP) at the National Aeronautics and Space Administration (NASA) Glenn Research Center (GRC). Since the inception of EP technology, GRC has played a pivotal role in developing the technology and transitioning it to flight. Research and development efforts by GRC’s employees, grantees, and contractors brought operational EP systems to commercial satellites in the early 1990s and to NASA and other government agency missions starting in the late 1990s. GRC’s early construction of unique EP test infrastructure and NASA’s and GRC’s sustained investments in EP research and development over several decades were strategic in the creation of technologies that continue to greatly benefit the competitiveness and capabilities of U. S. space systems and missions.