Systems Design of a Hybrid Sail Pole - Sitter

This paper presents the preliminary systems design of a pole-sitter. This is a spacecraft that hovers over an Earth pole, creating a platform for full hemispheric observation of the polar regions, as well as direct-link telecommunications. To provide the necessary thrust, a hybrid propulsion system combines a solar sail with a more mature solar electric propulsion (SEP) thruster. Previous work by the authors showed that the combination of the two allows lower propellant mass fractions, at the cost of increased system complexity. This paper compares the pure SEP spacecraft with the hybrid spacecraft in terms of the launch mass necessary to deliver a certain payload for a given mission duration. A mass budget is proposed, and the conditions investigated under which the hybrid sail saves on the initial spacecraft initial mass. It is found that the hybrid spacecraft with nearto mid-term sail technology has a lower initial mass than the SEP case if the mission duration is 7 years or more, with greater benefits for longer duration missions. The hybrid spacecraft with far-term sail technology outperforms the pure SEP case even for short missions. INTRODUCTION Solar electric propulsion (SEP) is a mature technology that provides a spacecraft with a relatively low thrust (of the order of a fraction of a Newton (Wallace, 2004)). Despite its high specific impulse, the thrust duration is always limited by the amount of propellant on-board. In contrast, solar sailing (McInnes, 1999b) is a propellant-less spacecraft propulsion system: it exploits the solar radiation pressure due to solar photons impinging on a large, highly reflecting surface (the sail) to generate thrust. Despite the original idea of solar sailing is dating back to 1924 (Tsander, 1969), only recently has a spacecraft successfully deployed a solar sail for the first time (Yamaguchi et al. , 2010). Studies are ongoing (Baoyin and McInnes, 2006, Mengali and Quarta, 2009), including mission design (Kawaguchi et al. , 2009), due to the interesting potential of enabling missions that are not constrained by propellant mass availability. Solar sails appear to be suitable for potentially long duration missions that need a small, but continuous, amount of thrust. One example of such applications are interplanetary transfers (Mengali and Quarta, 2005), but solar sails have also been investigated to generate artificial equilibrium points, for example in the proximity of the Lagrange points of the Sun-Earth (Baoyin and McInnes, 2005) or Sun-Moon (Simo and McInnes, 2009b) system: the effect of the sail is that of creating regions in which the spacecraft can be stationary with respect to the two main bodies (McInnes, 1999a). These regions can be consistently far from the classical Lagrange points, thus enabling a whole range of new applications, from telecommunications and data relay to Earth and Moon observation. Considering advantages and limitations of both SEP and solar sailing, the idea of a hybrid propulsion spacecraft, combining a solar sail and SEP arises. At the cost of increased spacecraft complexity, the two propulsion systems complement each other, cancelling their reciprocal disadvantages and limitations. In principle, SEP can provide thrust in any direction (through attitude maneuvers or a gimbal-mounted thruster), thus it can provide the missing acceleration component towards the Sun, that the sail cannot generate. Similarly, the hybrid spacecraft can be seen as an SEP spacecraft, in which an auxiliary solar sail provides part of the acceleration, enabling saving of propellant, and lower demand on the electric thruster, possibly with some intervals in which it could be turned off. In this sense, the hybrid spacecraft can be seen as a way to gradually introduce solar sails for space applications, and hence to reduce the advancement degree of difficulty (AD) (Macdonald and McInnes, 2010) in the technology readiness level scale. Hybridizing the two propulsion systems is a recent idea (Leipold and Götz, 2002), nevertheless research is flourishing in this almost completely unexplored field, investigating its potential for novel, interesting applications. Baig and McInnes (2008); proposed the use of a hybrid sail for generating artificial equilibria above L1 in the Sun-Earth system for Earth observation; Mengali and Quarta (Mengali and Quarta, 2007a, b) investigated optimal interplanetary transfers to Venus and Mars using an indirect optimization method; Quarta et al. (2010) also considered hybridizing high thrust and an electric sail; Simo and McInnes exploited hybrid propulsion to find displaced periodic orbits in the Earth-Moon system (Simo and McInnes, 2009a); finally, JAXA has developed a hybrid solar sail demonstrator, IKAROS (Mori et al. , 2009). In this work, hybrid propulsion is exploited to enable a mission in which the spacecraft is constantly above one of the Earth’s poles, i.e. lying on the Earth’s polar axis (Driver, 1980). This type of mission is known as pole-sitter. The pole-sitter provides a platform for continuous, real-time, medium-resolution observation of the Earth poles, with a full hemispheric view, as well as direct-link telecommunication and visibility of one of the Earth poles. It is well known that line-of-sight telecommunications to conventional spacecraft in geostationary orbits is not possible at high latitudes and polar regions, and telecommunication with polar regions will be a key issue in the future as changes to the arctic ice pack opens navigation channels for shipping. Also, the pole-sitter would offer real-time observation of the poles for climate science at modest resolution, in contrast to the periodic images that can be obtained by classical high-inclination orbits. This mission concept has already been proposed in the literature. A stationary spacecraft in the rotating Sun-Earth system for Earth observation and data relay was investigated using a pure solar sail by McInnes and Mulligan (2003). This work was successively extended by Baig and McInnes (2008): in that paper, it was proposed to place a spacecraft in the same artificial equilibrium points of the Sun-Earth rotating system, but using hybrid SEP/sail propulsion. Finally, the orbital dynamics of a hybrid propulsion continuous pole-sitter was presented in a work by Ceriotti and McInnes (2010a), enabling a practical realization of the solar sail pole-sitter orbits proposed earlier by Forward (Forward, 1991). Ceriotti and McInnes (2010a) designed optimal orbits that follow the polar axis of the Earth in the restricted three-body problem. It was also shown that the hybrid propulsion spacecraft enabled consistent savings in propellant mass fraction. However, the hybrid propulsion spacecraft is inevitably more complex with respect to pure SEP: for example, the sail needs to be pointed to the right attitude to provide the correct force, as does the SEP thruster: this requires a gimbal system, that is unnecessary in the pure SEP spacecraft, as the whole spacecraft can be tilted using standard attitude control maneuvers. In this paper, the objective is to study the preliminary system design of a pole-sitter spacecraft, both in the SEP and hybrid propulsion cases, necessary to carry a given payload. The work will assess in which cases the hybrid spacecraft enables a lower initial wet mass with respect to the SEP, for the same payload mass. The starting point for the mission design are the optimal orbits that were designed by Ceriotti and McInnes (2010a), to which we refer the reader for a complete description of the method and the results. In this paper, we will just present a brief overview of the equations of motion (section 1) and the optimization procedure (section 2). In section 3, the mass breakdown for the spacecraft will be explained in detail, followed by a comparison between the SEP case and the hybrid case, for a range of different technological parameters and mission durations. 1. EQUATIONS OF MOTION The circular restricted three-body problem (CR3BP) framework is considered (Sun-Earth-spacecraft). As is common, a synodic reference frame is used (Fig. 1). The mass of the Sun and the Earth are denoted and respectively, and 1 m 2 m ˆ   ω z the angular velocity of the system. The equations that describe the motion of the spacecraft of mass m in this system are:   1 2 1 2 s T r r                   r ω r ω ω r a a   (1) where is the position vector, r s a is the acceleration due to the solar radiation pressure on the spacecraft sail, T m  a is the acceleration from the solar electric propulsion (SEP) system, provided by thrust vector . T T Equation (1) will be used in its canonical non-dimensional form, which assumes 1   ,   2 1 2 m m m    , and the unit of distance is the separation of the two primaries. With these assumptions, the position along the x -axis of is ˆ 1 m   , and the position of is 1 2 m   . For the Earth-Sun system, 6 3.0404 10     . The two vectors r and r represent the position of the spacecraft with respect to the Sun and the Earth, respectively (see 1 2