Combining Electric and Sail Propulsion for Interplanetary Sample Return

Solar sails have long been known to be an attractive method of propulsion in the inner solar system if the areal density of the overall spacecraft (S/C) could be reduced to ∼10 g/m2. It has also long been recognized that the figure (precise shape) of useful solar sails needs to be reasonably good, so that the reflected light goes mostly in the desired direction. If one could make large reflective surfaces with reasonable figure at an areal density of ∼10 g/m2, then several other attractive options emerge. One is to use such sails as solar concentrators for solar-electric propulsion. Current flight solar arrays have a specific output of ∼100W/kg at 1 Astronomical Unit (AU) from the sun, and near-term advances promise to significantly increase this figure. A S/C with an areal density of 10 g/m2 could accelerate up to 29 km/s per year as a solar sail at 1 AU. Using the same sail as a concentrator at 30 AU, the same spacecraft could have up to ∼45 W of electric power per kg of total S/C mass available for electric propulsion (EP). With an EP system that is 50% power-efficient, exhausting 10% of the initial S/C mass per year as propellant, the exhaust velocity is ∼119 km/s and the acceleration is ∼12 km/s per year. This hybrid thus opens attractive options for missions to the outer solar system, including sample-return missions. If solar-thermal propulsion were perfected, it would offer an attractive intermediate between solar sailing in the inner solar system and solar electric propulsion for the outer solar system. In the example above, both the solar sail and solar electric systems don't have a specific impulse that is near-optimal for the mission. Solar thermal propulsion, with an exhaust velocity of the order of 10 km/s, is better matched to many solar system exploration missions. This paper derives the basic relationships between these three propulsion options and gives examples of missions that might be enabled by such hybrids.

Solar electric propulsion by a solar power sail for small spacecraft missions to the outer solar system

Accomplishment and prospect of isas electric propulsion

The paper investigates the utility of how nuclear thermal rocket (NTR) upper stage propulsion could substantially reduce the interplanetary transit time while greatly enhancing the economics of performing the manned Mars mission. au ECCV GTO IMLEO Isp kg kg/kw km km/sec kw LEO MEV mw NTR sec !N Nomenclature astronomical unit Earth capture crew vehicle geosynchronous transfer orbit initial mass in low Earth orbit specific impulse kilogram kilograms per kilowatt kilometer kilometers per second kilowatt low Earth orbit Mars excursion vehicle megawatt nuclear thermal rocket seconds delta V 1. Background Initial responses to the launch cost issue have been to suppose the development of heavy lift launch vehicles that are assumed to have fundamental launch cost savings. In general, most studies have isolated launch vehicle systems from the interplanetary vehicle assembled in low Earth orbit. This paper investigates a potential synergism between applying nuclear thermal rockets as both upper stages of the launch vehicle and as departure propulsion for a Mars vehicle departure stage. II. InterPlanetary Vehicle Description The mission to Mars is a propulsion technology dominated mission. The velocity change requirements (llV or delta V) are very large and very sensitive to trip time constraints. Due to this, there has been a heavy emphasis on high specific impulse (Isp) propulsion systems for performing the manned missions to Mars. The fundamental problem of high Isp systems is that they are generally thermally limited by material temperature limits. This issue is eliminated by electric propulsion systems that use electro-magnetic forces for accelerating a working fluid, but are thrust limited due to practical limitations in power supply mass. More exotic propulsion systems tend to use plasma working fluids that can only be contained by such methods as magnetic nozzles. The Space Exploration Initiative (SET) has a long range goal to perform a manned mission to Mars. Preliminary design goals of an interplanetary vehicle for this mission stress extremely high propulsion requirements to minimize transit times while allowing significant stay times on Mars. Related to this key trade between propulsion performance and trip time is the space transportation economics of 20 or more Earth launches required of the current launch vehicle architecture to place adequate vehicle mass in low Earth orbit to perform a single manned mission to Mars. The large initial mass in low Earth orbit (IMLEO) to perform the manned Mars mission is recognized as a dominant mission cost driver using the current launch infrastructure. IMLEO has been used as a preliminary figure of merit to evaluate various design approaches. * Senior Advanced Propulsion Engineer, Member AIAA Most analysis to date has focused on nuclear thermal rockets (Isp ... 1000 sec) and megawatt class electric propulsion systems (Isp ... 5000+ sec). The vehicle strategies have been to split the mission between a one way only cargo mission to Mars and a fast transfer manned vehicle. Mars rendezvous is essential for mission success, and techniques to slow at Earth return have used Earth capture crew vehicles (ECCV) with very high excess velocities (9+ km/sec) that are ballistically braked. Analyzed here is a very simple single vehicle. It uses an NTR driven, high thrust, Earth departure stage that is dropped once the vehicle has entered a fast Mars transfer ellipse. Once the vehicle has exited the deep Earth gravity well, it relies solely on a 5000 second Isp electric propulsion system for capture at Mars, departure from Mars, and final mission capture at Mars with zero excess velocity (a conservative assumption) The payload to Mars is broken into two elements. The ECCV is assumed to weigh 40,000 kg for a crew of four, and is the crew's interplanetary travel habitation module. The Mars excursion vehicle (MEV) is the element used to drop the crew from Mars orbit to the planet surface, be the crew's planetary habitation module, carry all science equipment, and carry the ascent module for returning to Mars orbit. The MEV has a baseline weight of 84,000 kg for a nominal 30 day planetary stay time and is scaled with the squareroot of normalized stay time. The MEV is not returned to Earth. Consumables are estimated at 10 kg per person per mission day. optimized for minimum l1V by adjusting the up leg perihelion and aphelion while maintaining the constraints of Earth rendezvous and travel time. Figure 2 (with the radius in millions of kilometers) shows a transfer ellipse for the up leg for the 320 day travel time, 80 day Mars stay time, 0 degree mean Mars true anomaly case. Figure 3 shows the down leg of that mission. The mission l1V requirement of 29.1 krn/sec is divided into 4.2 krn/sec for Earth departure (in heliocentric space), 6.0 krn/sec for Mars capture, 7.4 krn/sec for Mars departure, and 11.5 krn/sec for Earth capture. Transfer Ellipse Up. III. Interplanetary Mission Modeling The problem of analyzing the interplanetary mission can be driven by very exacting issues such as actual planetary configurations for particular mission opportunities, the requirements for travel time, and the Mars stay time. The issue that needs near term resolution, however, is the general efficacy of various propulsion and/or vehicle configuration options for the Mars mission in general. To answer these type of questions, it was necessary to generalize the Mars transfer problem. To do so, simplifying assumptions were that the Mars and Earth orbits were coplanar, and that the Earth orbit was circular. The nominal planetary configurations for Mars, Earth, and Venus are shown in Figure 1. The radius units are in astronomical units (Earth radius =1 au) and the angular units are the true anomaly of the orbital positions in degrees (Mars perihelion = 0 degrees). tojMJoC:: ~ •• ~: ••• . -

Venus is Earth’s closest planetary neighbour and both bodies are of similar size and mass. As a consequence, Venus is often described as Earth’s sister planet. But the two worlds have followed very different evolutionary paths, with Earth having benign surface conditions, whereas Venus has a surface temperature of 464 °C and a surface pressure of 92 bar. These inhospitable surface conditions may partially explain why there has been such a dearth of space missions to Venus in recent years.The oxygen isotope composition of Venus is currently unknown. However, this single measurement (Delta ^{17}text{O}) would have first order implications for our understanding of how large terrestrial planets are built. Recent isotopic studies indicate that the Solar System is bimodal in composition, divided into a carbonaceous chondrite (CC) group and a non-carbonaceous (NC) group. The CC group probably originated in the outer Solar System and the NC group in the inner Solar System. Venus comprises 41% by mass of the inner Solar System compared to 50% for Earth and only 5% for Mars. Models for building large terrestrial planets, such as Earth and Venus, would be significantly improved by a determination of the Delta ^{17}text{O} composition of a returned sample from Venus. This measurement would help constrain the extent of early inner Solar System isotopic homogenisation and help to identify whether the feeding zones of the terrestrial planets were narrow or wide.Determining the Delta ^{17}text{O} composition of Venus would also have significant implications for our understanding of how the Moon formed. Recent lunar formation models invoke a high energy impact between the proto-Earth and an inner Solar System-derived impactor body, Theia. The close isotopic similarity between the Earth and Moon is explained by these models as being a consequence of high-temperature, post-impact mixing. However, if Earth and Venus proved to be isotopic clones with respect to Delta ^{17}text{O}, this would favour the classic, lower energy, giant impact scenario.We review the surface geology of Venus with the aim of identifying potential terrains that could be targeted by a robotic sample return mission. While the potentially ancient tessera terrains would be of great scientific interest, the need to minimise the influence of venusian weathering favours the sampling of young basaltic plains. In terms of a nominal sample mass, 10 g would be sufficient to undertake a full range of geochemical, isotopic and dating studies. However, it is important that additional material is collected as a legacy sample. As a consequence, a returned sample mass of at least 100 g should be recovered.Two scenarios for robotic sample return missions from Venus are presented, based on previous mission proposals. The most cost effective approach involves a “Grab and Go” strategy, either using a lander and separate orbiter, or possibly just a stand-alone lander. Sample return could also be achieved as part of a more ambitious, extended mission to study the venusian atmosphere. In both scenarios it is critical to obtain a surface atmospheric sample to define the extent of atmosphere-lithosphere oxygen isotopic disequilibrium. Surface sampling would be carried out by multiple techniques (drill, scoop, “vacuum-cleaner” device) to ensure success. Surface operations would take no longer than one hour.Analysis of returned samples would provide a firm basis for assessing similarities and differences between the evolution of Venus, Earth, Mars and smaller bodies such as Vesta. The Solar System provides an important case study in how two almost identical bodies, Earth and Venus, could have had such a divergent evolution. Finally, Venus, with its runaway greenhouse atmosphere, may provide data relevant to the understanding of similar less extreme processes on Earth. Venus is Earth’s planetary twin and deserves to be better studied and understood. In a wider context, analysis of returned samples from Venus would provide data relevant to the study of exoplanetary systems.

The electron cyclotron resonance ion engine, “μ10,” has a long life and high reliability because of electrodeless plasma generation in both the ion generator and the neutralizer. Four μ10, each generating a thrust of 8 mN, specific impulse of 3,200 seconds, and consuming 350 W of electric power, propel the “HAYABUSA” asteroid explorer that was launched on May 2003. After vacuum exposure and several runs of baking to reduce residual gas, the ion engine system established continuous acceleration. 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 distance between 0.86 AU and 1.7 AU in the solar system, as well as rendezvous with, land on, and lift off from the asteroid Itokawa. During the 3-year flight, the ion engine system generated a delta-V of 1,400 m/s while consuming 22 kg of xenon propellant and operating for 25,900 hours.

Project Lyra: Sending a spacecraft to 1I/’Oumuamua (former A/2017 U1), the interstellar asteroid

The limitations of conventional chemical propulsion for missions to the outer planets, comets, and asteroids were illustrated in Table 5.2 of the previous chapter. An alternative approach to the use of swingby manoeuvres is to exploit the very high specific impulses that can be achieved through the use of electric propulsion, eg. exhaust velocities the order of 30 km/s instead of about 3 km/s associated with chemical propellants. The source of the electrical energy can be either solar (solar electric propulsion SEP) or nuclear (nuclear electric propulsion NEP). As a result of the diminishing power of the Sun with distance, SEP is not useful beyond about 5 AU from the Sun. Nuclear energy generation is of course independent of the Sun but has unfortunately become less attractive in recent years due to an influential Earth-bound environmental lobby.

Development of the mathematical basis for the generation of multiple target low-thrust trajectories. The unique trajectory shaping capabilities of electric propulsion make it ideally suited for modifying the path of the spacecraft to effect close encounters of selected asteroids with only slight penalties in performance. Each intermediate encounter with an asteroid involves the optimization of four parameters, the encounter time and the three components of the relative velocity at encounter. A second type of mission amenable to this analysis is that of asteroid or comet sample return missions. For missions of this type, both launch date and division of departure and return flight times are capable of being optimized. Use is made of the traditional calculus of variations to derive the necessary conditions that must be satisfied for both types of missions. In addition, examples are presented of both a Comet Encke rendezvous mission with multiple encounters of the asteroids Aschera and Alekto and also a sample return mission from the asteroid Flora.

Miniaturization perspectives of electrostatic propulsion for small spacecraft platforms

Today, our questions and hypotheses about the Solar System's origin have surpassed our ability to deliver scientific instruments to deep space. The moons of the outer planets, the Trojan and Centaur minor planets, the trans-Neptunian objects (TNO), and distant Kuiper Belt objects (KBO) hold a wealth of information about the primordial conditions that led to the formation of our Solar System. Robotic missions to these objects are needed to make the discoveries, but the lack of deep-space propulsion is impeding this science. Radioisotope electric propulsion (REP) will revolutionize the way we do deep-space planetary science with robotic vehicles, giving them unprecedented mobility. Radioisotope electric generators and lightweight ion thrusters are being developed today which will make possible REP systems with specific power in the range of 5 to 10 W/kg. Studies have shown that this specific power range is sufficient to perform fast rendezvous missions from Earth to the outer Solar System and fast sample return missions. This whitepaper discusses how mobility provided by REP opens up entirely new science opportunities for robotic missions to distant primitive bodies. We also give an overview of REP technology developments and the required next steps to realize REP.

Solar power sail is a deep space probe to be powered by hybrid propulsion of solar photon acceleration and ion engines to explore outer planetary region of the Solar System without relying on nuclear technology. The Japan Aerospace Exploration Agency (JAXA) launched the world’s first deep space solar sail demonstration spacecraft “IKAROS” (Interplanetary Kite-craft Accelerated by Radiation Of the Sun) on May 21, 2010. IKAROS succeeded in deploying a 20m-span solar sail on June 9 and demonstrated several key technologies for solar sail utilizing the deep space flight environment. JAXA is currently studying an outer solar system exploration mission using the demonstrated solar power sail technology. The mission plans to fly for Jupiter, where the spacecraft drops a tiny Jovian probe and performs a swing-by for a Trojan asteroid. Current scenario consists of the rendezvous with one of the Trojan asteroids that are at the Lagrange points L4/L5 associated with Sun-Jupiter system. About as large as 50m sail should be deployed for this mission according to preliminary mission analysis and related research is intensively being carried out in JAXA. JAXA plans to initiate the project in a few years and looks at the launch around 2020.

The mission targets of the Chinese asteroid exploration are an earth co-orbiting celestial body 2016HO3 and the main-belt comet in the asteroid belt. Compared with Lunar Exploration and Mars Exploration missions, asteroid exploration mission is special. The light intensity varies widely during the mission, resulting in a large change in the output power range of the solar array, which seriously affects the power system design and the load design during the mission; the spacecraft uses a high-power electric propulsion system; the power system has strict weight requirements, the existing power supply system design cannot meet the requirements. Aiming at the design constraints of the large variation of solar wing light intensity, the large difference in power requirements between the electric propulsion and the platform, and the lightweight design of the spacecraft, a dual-bus power supply system is proposed. High-voltage bus power the electric propulsion system. The solar cell array directly converges to form a high-voltage bus, which supplies power for the electric propulsion system. The low-voltage bus adopts a semi-adjustable bus, and the solar cell array adopts a circular flexible solar wing. The main detector and return cabin are equipped with lithium-ion battery pack and lithium fluorocarbon battery pack respectively. Aiming at the requirement of long-term on-orbit operation of electric propulsion system, a solar wing output power measurement system is designed. Product test results show that the power system design can meet the working requirements of electric propulsion and platform loads. The power system design can provide reference for the other deep space exploration power system.

The significant increase of available electric power for the propulsion unit onboard satellites makes high and ultra-high specific impulse thrusters attractive for different applications and missions. Especially thrusters of the radio frequency type show a great potentiality for high Isp applications. Two different rf-thrusters are described: The RIT15LP engine optimized for a moderate input power consumption and the RIT15S designed for a high specific impulse. The main, but only difference between these engines is the extraction grid system. During the development of RIT15LP, principle improvements of rfthrusters were invented. The given description of RIT15LP is focused on these performance improvements whereas the advantages of the rf-principle for (ultra-) high specific impulse applications were pointed out on basis of the RIT15S thruster. The function principle of gridded ion thrusters can be described as a two step process. In the first step the propellant is ionized and after that, in the second step it is accelerated in the electrostatic fields of an ion extraction and acceleration grid system. Both tasks a nearly independent from each other. The acceleration has no effect on the ionisation process and therefore the beam velocity can simply be controlled by the selection of the voltages applied to the grid system. Amoung different types of electric thrusters this is a unique feature of all gridded ion engines. In contrast, the use of other electric propulsion concepts requires often deep changes in the thruster design for an increased beam velocity. For example, it was necessary to invent a two stage concept for SPTs to reach a beam velocity comparable with that of gridded thrusters. Such efforts are not required in case of the later ones. Presently, the beam velocity of gridded ion thrusthers is merely limited by their electric input power consumption and not in a conceptional way. It is important to be aware that the specific input power consumption (the total consumption of electric power Ptot over the thrust F, usually given in>luWatt per Millinewton') of any kind of thruster is a simple function of the beam velocity Vbeam and the electric efficiency of the thrusters rjel:

The advantages of using electric propulsion for propulsion are well‐known in the aerospace community. The high specific impulse and, therefore, lower propellant requirements make it a very attractive propulsion option for the Space Exploration Initiative (SEI). Recent studies have shown that nuclear electric propulsion (NEP) is not only attractive for the transport of cargo but that fast piloted missions to Mars are possible as well, with alphas on the order of 7.5 kg/kW. An advanced NEP system with a pecific power (alpha) of 2.5 kg/kW or less would significantly enhance the manned mission option of NEP by reducing the trip time even further. This paper describes an advanced system that combines the PEGASUS Drive with systems of the Rotating Multimegawatt Boiling Liquid Metal (RMBLR) power system that was developed as part of the DOE multimegawatt program and just recently declassified. In its original configuration, the PEGASUS Drive was a 10‐MWe propulsion system. The RMBLR was a 20‐MW electric power system. By combining the two, a second‐generation PEGASUS Drive can be developed with an alpha less than 2.5 kg/kW. This paper will address the technology advancements incorporated into the PEGASUS Drive, the analysis of a fast piloted mission and an unmanned cargo transport Mars mission, and the integration of laser power beaming to provide surface power.