Solar Electric Propulsion Research Articles

Optimized Low-Thrust Orbit Transfer for Space Tugs

Introduction Recent advances in low-cost, commercial launch systems predict rapid insertion of relatively large payloads into Low Earth Orbit (LEO). At the same time, extensive work performed on radiation hardening of solar cells—specifically GaAs, InP and InGaP/GaAs cells—has permitted repeated transfers through the Van Allen belt with minimal loss of conversion efficiency. 1 The combined impact of these two technological advances suggests that a market for an in-space transportation system capable of ferrying payloads from LEO to GEO and beyond is not only possible, but desirable. The space tug configuration demonstrates a viable candidate for an Earth-Moon cargo vehicle system satisfying NASA’s Vision for Space Exploration. Establishment of permanent outposts on the Moon will require the transfer of large quantities of cargo to the Moon’s surface at low cost. A high power, reusable, solar electric propulsion system optimized for both transfer mass and transfer time is capable of delivering high-tonnage cargo per year to the surface of the Moon. The cost saving from both the use of smaller launch vehicles and the reuse of the power and propulsion subsystems makes such a mission system an attractive choice for the Lunar Exploration Program. In this note, the authors apply an optimization technique previously developed by Burton and Wassgren 2 to a reusable electrically-propelled Earth-Moon “space tug” transfer vehicle proposed by Spores et al. 3 Both transfer mass and mission time are optimized, and the results show that significant reductions in the round trip transfer time can be achieved. In addition, either the initial launch mass can be significantly reduced, while keeping the yearly delivered payload constant, or the yearly delivered cargo can be significantly increased by appropriately increasing the available power.

On the design of low‐thrust transfers through L1 for the ESMO mission

AbstractThe following paper summarises the mission analysis studies performed for the phase A of the solar electric propulsion option of the European Student Moon Orbiter (ESMO) mission. ESMO is scheduled to be launched in 2011, as an auxiliary payload on board an Ariane 5. Hence the lack of influence on the launch date. Two methods were devised to assess efficiently wide launch windows for the Earth‐Moon transfer, they are presented below. (© 2008 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)

Earth-moon Trajectory Optimization Using Solar Electric Propulsion

The optimization of the Earth-moon trajectory using solar electric propulsion is presented. A feasible method is proposed to optimize the transfer trajectory starting from a low Earth circular orbit (500 km altitude) to a low lunar circular orbit (200 km altitude). Due to the use of low-thrust solar electric propulsion, the entire transfer trajectory consists of hundreds or even thousands of orbital revolutions around the Earth and the moon. The Earth-orbit ascending (from low Earth orbit to high Earth orbit) and lunar descending (from high lunar orbit to low lunar orbit) trajectories in the presence of J2 perturbations and shadowing effect are computed by an analytic orbital averaging technique. A direct/indirect method is used to optimize the control steering for the trans-lunar trajectory segment, a segment from a high Earth orbit to a high lunar orbit, with a fixed thrust-coast-thrust engine sequence. For the trans-lunar trajectory segment, the equations of motion are expressed in the inertial coordinates about the Earth and the moon using a set of nonsingular equinoctial elements inclusive of the gravitational forces of the sun, the Earth, and the moon. By way of the analytic orbital averaging technique and the direct/indirect method, the Earth-moon transfer problem is converted to a parameter optimization problem, and the entire transfer trajectory is formulated and optimized in the form of a single nonlinear optimization problem with a small number of variables and constraints. Finally, an example of an Earth-moon transfer trajectory using solar electric propulsion is demonstrated.

The Deep Space 1 Mission: Flight Validation of New Technology

On September 22 2001 NASA’s Deep Space 1 (DS1) spacecraft passed 2171 km from the nucleus of comet 19P/Borrelly. DS1, the first mission in NASA’s New Millennium Program (NMP), was also the first mission to use solar electric propulsion (SEP) as primary propulsion. In addition, it was NASA’s first spacecraft with both an imaging system and a plasma mass spectrometer to encounter a comet The primary focus of the NMP is to identify and flight validate new, advanced technologies that hold great promise for revolutionizing observations from earth orbit or in deep space. Such technologies, despite being technically mature, have not been demonstrated in space and hence are perceived to present a fairly high risk to missions that propose to use them for the first time. In order to help reduce the costs and risks to future missions that use these technologies, NMP oversees a series of dedicated “technology validation” missions into deep space and around Earth to validate these technologies and prove that they perform as intended in space. Since DS1 there have been four completed follow-on NMP validation activities (DS2, Earth Observing I, Space Technology 5 (ST5), and ST 6 Autonomous Sciencecraft Experiment) and four more are under development (ST 6 Inertial Stellar Compass, ST7 Disturbance Reduction System, ST8 and five concepts in Phase A for ST9). NMP projects can be either stand alone missions or payloads on other earth orbiting or deep space missions. The DS1 payload contained a set of technologies that were successfully validated in the first phase of the mission. The principal, or project defining technology validated on DS1 was SEP. The SEP primary propulsion system on DS1 ionized xenon gas in a thrust

Powered Flight of Electron Cyclotron Resonance Ion Engines on Hayabusa Explorer

The electron cyclotron resonance ion engine has long life and high reliability because of electrodeless plasma generation in both the ion generator and the neutralizer. Four μ10s, each generating a thrust of 8 mN, specific impulse of 3200 s, and consuming 350 W of electric power, propelled the Hayabusa asteroid explorer launched on May 2003. After vacuum exposure and several baking runs to reduce residual gas, the ion engine system established continuous acceleration. Electric propelled delta-V Earth gravity assist, a new orbit change scheme that uses electric propulsion with a high specific impulse was applied to change from a terrestrial orbit to an asteroid-based orbit. In 2005, Hayabusa, using solar electric propulsion, managed to successfully cover the solar distance between 0.86 and 1.7 AU. It rendezvoused with, landed on, and lifted off from the asteroid Itokawa. During the 2-year flight, the ion engine system generated a delta-V of 1400 m/s while consuming 22 kg of xenon propellant and operating for 25,800 h.

Thermal Model of the Hollow Cathode Using Numerically Simulated Plasma Fluxes

The first results from a hollow cathode thermal model are presented. The thermal model uses the spatially distributed plasma fluxes, calculated by a two-dimensional axisymmetric model of the plasma inside the emitter region, as the heat source to predict the hollow cathode and insert temperatures. The computed insert temperature profiles are compared with measured values for a cathode similar to the 0.635-cm diameter discharge cathode used in the NASA solar electric propulsion technology applications readiness ion engine. The insert temperatures can be used to predict cathode life from barium depletion. The present results and related experimental studies yield three important conclusions. First, the emitter in the NASA solar electric propulsion technology applications readiness hollow cathode does not operate in the emission-limited regime. The thermionic electron current is about 20 A higher than the discharge current and requires significant reverse electron flux from the plasma to satisfy current continuity. Second, the high-plasma density near the centerline of the cathode results in power deposition on the orifice plate that is more than twice the emitter power deposition. Third, despite a higher heat load to the orifice plate, the operating temperature where it would be measured by a thermocouple is approximately 100°C lower than the emitter. The lower orifice plate temperature is due to poor thermal contact between the emitter and the cathode tube and higher than anticipated radiative losses from the external surface of the heater.

Concept study on Deep Space Orbit Transfer Vehicle

In this paper, the concept of Orbit Transfer Vehicle for Deep Space Exploration (Deep Space OTV) is proposed, and its effectiveness and feasibility are discussed. Basic concept is the separation of two functions required for the deep space exploration, the transportation to the destination, and the exploration at the destination. Deep Space OTV is a spacecraft specialized for the transportation to the deep space destination. It is an expendable spacecraft propelled by solar electric propulsion. The payload of Deep Space OTV is Explorer, which is a spacecraft specialized for the exploration at the deep space destination. The effectiveness of the concept is discussed qualitatively, focused on the merits of the separations of two functions. The feasibility of Deep Space OTV is discussed based on the conceptual design of the spacecraft and its applicability to deep space missions. Several deep space missions are modeled and the payload capacity of Deep Space OTV is estimated. The missions include Asteroid rendezvous, Mars orbiter, Lunar lander, and so on.

Simulation study of solar wind push on a charged wire: basis of solar wind electric sail propulsion

Abstract. One possibility for propellantless propulsion in space is to use the momentum flux of the solar wind. A way to set up a solar wind sail is to have a set of thin long wires which are kept at high positive potential by an onboard electron gun so that the wires repel and deflect incident solar wind protons. The efficiency of this so-called electric sail depends on how large force a given solar wind exerts on a wire segment and how large electron current the wire segment draws from the solar wind plasma when kept at a given potential. We use 1-D and 2-D electrostatic plasma simulations to calculate the force and present a semitheoretical formula which captures the simulation results. We find that under average solar wind conditions at 1 AU the force per unit length is (5±1×10−8 N/m for 15 kV potential and that the electron current is accurately given by the well-known orbital motion limited (OML) theory cylindrical Langmuir probe formula. Although the force may appear small, an analysis shows that because of the very low weight of a thin wire per unit length, quite high final speeds (over 50 km/s) could be achieved by an electric sailing spacecraft using today's flight-proved components. It is possible that artificial electron heating of the plasma in the interaction region could increase the propulsive effect even further.

Application of Solar Electric Propulsion to a Comet Surface Sample Return Mission

Current NSTAR (planned for the Discovery Mission: Dawn) and NASA's Evolutionary Xenon Thruster based propulsion systems were compared for a comet surface sample return mission to Tempe1 1. Mission and systems analyses were conducted over a range of array power for each propulsion system with an array of 12 kW EOL at 1 AU chosen for a baseline. Engine configurations investigated for NSTAR included 4 operational engines with 1 spare and 5 operational engines with 1 spare. The NEXT configuration investigated included 2 operational engines plus 1 spare, with performance estimated for high thrust and high Isp throttling modes. Figures of merit for this comparison include Solar Electric Propulsion dry mass, average engine throughput, and net non-propulsion payload returned to Earth flyby.

Assessment of Plasma Interactions and Flight Status of the HAYABUSA Asteroid Explorer Propelled by Microwave Discharge Ion Engines

Microwave discharge ion engines “ <tex xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">$mu 10$</tex> ” are dedicated to the main propulsion on the HAYABUSA asteroid explorer. In a development program, various tests and assessments were conducted on the ion engines and the spacecraft. They include endurance tests, an electromagnetic interference susceptibility test, an interference test between the plasma and communication microwave, a beam exhaust test on the spacecraft, assessments on the plasma interference with a solar array, and so on. The spacecraft was launched in deep space by the M-V rocket in May 2003. After vacuum exposure and several runs of baking for reduction of residual gas, the ion engine system established continuous acceleration of the spacecraft toward the asteroid ITOKAWA. The spacecraft passed through a perihelion of 0.86 astronomical unit (AU) in February 2004 and an aphelion of 1.7 AU in February 2005, becoming the first solar electric propulsion system to travel this far toward and away from the Sun. The HAYABUSA succeeded in rendezvousing with the target asteroid in September 2005.

Solar Electric and Chemical Propulsion for a Titan Mission

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Study of space transportation for space solar power systems

Space solar power systems (SSPSs) have the potential to provide abundant quantities of electric power for use on the Earth. One of the hurdles to them is the transportation of SSPSs to the operational geostationary Earth orbit (GEO). The objectives of this study are to examine the transportation of SSPSs, and to give a reference transportation scenario. This study presumes that the SSPSs have a mass of 10,000 tons each and are constructed at a rate of one per year. Reusable launch vehicles (RLVs) are assumed for the transportation to low Earth orbit (LEO), and reusable orbit transfer vehicles (OTVs) propelled by a solar electric propulsion system for the transportation from LEO to GEO. The payload element delivered to LEO by each launch is individually transferred by each OTV transportation service to GEO, where the elements are assembled into a whole SSPS. The OTV round-trip time is assumed to be a year. With these operations and reasonable estimations for the OTV subsystems, the OTV payload ratio was obtained. This, with an SSPS element mass, gave the total mass that has to be launched by RLVs. The result indicated that about 300 times of launch are required per year.

High-voltage array ground test for direct-drive solar electric propulsion

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.

Stretched lens array squarerigger (SLASR): A unique high-power solar array for exploration missions

Over the past two years, development has begun on a unique high-power solar concentrator array called stretched lens array squarerigger (SLASR). SLASR marries the high electrical performance and low mass of the ENTECH stretched lens array (SLA) flexible blanket concentrator with the high structural performance and low mass of the ATK space squarerigger (SR) platform. SLASR offers unprecedented performance metrics for a space solar array of any type, including: • high power capacity (100 kW to multi-MW Class) • compact launch stowage (80–120 kW/m 3 ) • state of the art areal power density (300–400 W/m 2 ) • outstanding array-level specific power (300–1000 W/kg) • high-voltage operation at low insulation mass penalty (300–600 V) • excellent manufacturability using proven processes (cells and lenses at MWs/year) • improved radiation hardness at low shielding mass penalty (enabling solar electric propulsion space tugs through the radiation belts) • exceptional cost effectiveness (50–75 percent less than high-efficiency planar arrays) This paper describes SLASR technology, discusses SLASR developments to date, and outlines plans for SLASR technology maturation. The advantages of SLASR over other high-power solar array technologies are presented, including a comparison of key performance parameters for large-scale solar electric propulsion space tug missions.

Chemical and Solar-Electric-Propulsion Systems Analyses for Mars Sample Return Missions

Conceptual in-space transfer stages, including those using solar electric propulsion and chemical propulsion with aerobraking or aerocapture assist at Mars, were evaluated. Roundtrip Mars sample return mission vehicles were analyzed to determine how technology selections influence payload delivery capability. Results show how specific engine, thruster, propellant, capture mode, trip time, and launch-vehicle technology choices would contribute to increasing payload delivered to Mars or decreasing the size of the required launch vehicles. Low-thrust trajectory analyses for solar-electric transfer were generated with the SEPTOP code.

SMART-1 after lunar capture: First results and perspectives

SMART-1 is a technology demonstration mission for deep space solar electrical propulsion and technologies for the future. SMART-1 is Europe’s first lunar mission and will contribute to developing an international program of lunar exploration. The spacecraft was launched on 27th September 2003, as an auxiliary passenger to GTO on Ariane 5, to reach the Moon after a 15-month cruise, with lunar capture on 15th November 2004, just a week before the International Lunar Conference in Udaipur. SMART-1 carries seven experiments, including three remote sensing instruments used during the mission’s nominal six months and one year extension in lunar science orbit. These instruments will contribute to key planetary scientific questions, related to theories of lunar origin and evolution, the global and local crustal composition, the search for cold traps at the lunar poles and the mapping of potential lunar resources

Advanced Microsatellite Mission—deep space applications and constraints

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.

Fuel-Optimal, Power-Limited Rendezvous with Variable Thruster Efficiency

The problem of minimum-fuel, time-fixed, three-dimensional rendezvous for a solar electric propulsion spacecraft is discussed. The problem is solved via an indirect approach. The formulation takes into account both a variable bounded specific impulse and a variable thruster efficiency and permits us to manage solutions with coast arcs. The thruster efficiency is assumed to vary with the specific impulse through a polynomial approximation. The optimal specific impulse control law is found to depend on the instantaneous values of the primer vector modulus, the spacecraft mass, the mass costate, and the thruster model. Optimal interplanetary trajectories toward Mars are discussed. It is shown that the inclusion of a variable efficiency thruster model has important effects on fuel consumption. In particular, the classic constant efficiency thruster model overestimates the final spacecraft mass.

APIES: A mission for the exploration of the main asteroid belt using a swarm of microsatellites

APIES (Asteroid Population Investigation and Exploration Swarm) is a mission developed by EADS Astrium in response to a European Space Agency (ESA) Call for Ideas for “swarm” missions, based on the utilization of a large number of spacecraft working cooperatively to achieve the mission objectives. The APIES baseline concept is centred on a “swarm” of 19 BElt Explorer (BEE) identical microsatellites, weighing less than 45 kg each, including their scientific payload, visiting over 100 Main Belt asteroids in multiple flybys. The BEEs are carried to the asteroid belt by a Hub and Interplanetary VEhicle (HIVE), a conventional spacecraft launched with a Soyuz–Fregat rocket, using solar electric propulsion for the transfer to the asteroid belt and acting as communication hub and control centre for the mission after the swarm deployment. Using the latest advances in systems miniaturization, propulsion, onboard autonomy and communications, the APIES mission can achieve its ambitious goal within the framework of a standard ESA mission, representing a novel mission concept example, whose feasibility is essentially linked to the use of microsatellite technology, enabling the achievement of science objectives unattainable with conventional spacecraft.

BepiColombo: Payload and mission updates

BepiColombo is an interdisciplinary mission to explore the planet Mercury through a partnership between ESA and Japan’s Aerospace Exploration Agency (JAXA). From their dedicated orbits two spacecraft, the Mercury Planetary Orbiter (MPO) and the Mercury Magnetospheric Orbiter (MMO), will be studying the planet and its environment. They will be launched together on a single Soyuz-Fregat. The launch is foreseen for May 2012 with arrival in January 2017. Solar electric propulsion will be used for the journey to Mercury. The MMO provided by JAXA focuses on investigating the wave and particle environment of the planet from an eccentric orbit. The MPO, a three-axis-stabilized and nadir-pointing spacecraft in a low-eccentricity polar orbit, is dedicated to the characterization of Mercury. It will provide high accuracy measurements of the planet’s interior structure, a characterization of Mercury’s exosphere and a full coverage of the planet surface at high resolution, whereby surface morphology will be correlated to surface composition. Major effort was put into optimizing the scientific return by defining the payload complement such that individual measurements can be interrelated and complement each other. In November 2004, the BepiColombo scientific payload has been officially approved. The MPO payload comprises 11 instruments/instrument packages; the MMO payload consists of 5 instruments/instrument packages. Together, the scientific payload of both spacecraft will provide the detailed information necessary to understand Mercury and its magnetospheric environment and to find clues to the origin and evolution of a planet close to its parent star. An overview is given of the status of BepiColombo will with special emphasis on the MPO.