Solar Electric Propulsion Spacecraft Research Articles

Trajectory options for the DART mission

This study presents interplanetary trajectory options for the Double Asteroid Redirection Test (DART) spacecraft to reach the near Earth object, Didymos binary system, during its 2022 Earth conjunction. DART represents a component of a joint NASA-ESA mission to study near Earth object kinetic impact deflection. The DART trajectory must satisfy mission objectives for arrival timing, geometry, and lighting while minimizing launch vehicle and spacecraft propellant requirements. Chemical propulsion trajectories are feasible from two candidate launch windows in late 2020 and 2021. The 2020 trajectories are highly perturbed by Earth's orbit, requiring post-launch deep space maneuvers to retarget the Didymos system. Within these windows, opportunities exist for flybys of additional near Earth objects: Orpheus in 2021 or 2007 YJ in 2022. A second impact attempt, in the event that the first impact is unsuccessful, can be added at the expense of a shorter launch window and increased (∼3x) spacecraft ΔV. However, the second impact arrival geometry has poor lighting, high Earth ranges, and would require additional degrees of freedom for solar panel and/or antenna gimbals. A low-thrust spacecraft configuration increases the trajectory flexibility. A solar electric propulsion spacecraft could be affordably launched as a secondary spacecraft in an Earth orbit and spiral out to target the requisite interplanetary departure condition. A sample solar electric trajectory was constructed from an Earth geostationary transfer using a representative 1.5kW thruster. The trajectory requires 9 months to depart Earth's sphere of influence, after which its interplanetary trajectory includes a flyby of Orpheus and a second Didymos impact attempt. The solar electric spacecraft implementation would impose additional bus design constraints, including large solar arrays that could pose challenges for terminal guidance. On the basis of this study, there are many feasible options for DART to meet its mission design objectives and enable this unique kinetic impact experiment.

Efficient Computation of Optimal Interplanetary Trajectories Using Solar Electric Propulsion

A new method has been developed for obtaining approximate near-optimal low-thrust interplanetary transfers using solar electric propulsion spacecraft. The trajectory-calculation method consists of analytic curve-fit functions that have been empirically derived from a database of minimum-propellant coplanar transfers between Earth’s orbit and circular target orbits. It is assumed that the transfer begins with a powered arc, followed by a single coast arc, and ends with a second powered arc. The curve-fit functions depend on the mean thrust acceleration and target radius. A scaling technique transforms the Earth-orbit-based solutions in order to provide outward and inward trajectories between circular orbits of arbitrary radius. The inputs to the trajectory tool are the initial and final circular radii, initial spacecraft mass, input power, specific impulse, and thruster efficiency; and the computed (output) values are trip time, transfer angle, coast arc angle, and . The curve-fitting method is extremely fast, with run times of a few milliseconds. Numerical trials show that the curve-fitting technique accurately predicts trip time, with the largest errors ranging from 2 to 4%. Velocity increment or errors (and hence delivered-mass errors) are negligible. Finally, because the curve-fitting method can predict transfer angle, the trajectory tool can estimate departure and arrival dates. This curve-fitting method provides very rapid low-thrust interplanetary trajectory analysis and is therefore well suited as a preliminary mission design tool.

Interplanetary Trajectories for Multiple Satellite-Aided Capture at Jupiter

Multiple-satellite-aided capture employs sequential gravity-assist flybys of more than one of Jupiter’s Galilean moons. These flybys substantially reduce the amount of propellant required for a spacecraft to capture into orbit around Jupiter. The “average” plane of the Galilean satellites constrains the arrival V-infinity vector, which in turn constrains the interplanetary trajectories from Earth to Jupiter. The solution space of interplanetary trajectories that permit multiple-satellite-aided capture is explored, and trajectories that start from Earth and end at Jupiter capture are integrated. I. Introduction Satellite-aided capture is a mission design technique that is used to decrease the V required to capture a spacecraft into orbit around a planetary body. The technique employs gravity-assist flybys of massive satellites of a planetary body, so only missions to planets with large moons can benefit from this technique. Hence, satellite-aided capture is only available to Earth-return missions and missions to Jupiter, Saturn, Uranus, or Neptune. Jupiter, in particular, has four massive Galilean moons that can be used for satellite-aided capture. These satellite-aided capture trajectories were first proposed for missions to Jupiter by Longman 1 and by Longman and Schneider. 2 Cline 3 determined the best use of a Ganymede gravity-assist to minimize the Jupiter insertion maneuver (JOI) V required for capture into Jupiter orbit. Nock and Uphoff 4 performed a tour-de-force of satellite-aided capture trajectories for the entire Solar System by varying several trajectory parameters, including perijove radius after flyby, flyby altitude, declination of the incoming satellite-centered hyperbola, and the distribution of V between powered satellite flybys and the JOI maneuver. They also briefly investigated double-satellite-aided capture by determining the phasing and transfer orbit parameters necessary to achieve capture. Malcolm and McInnes 5 employ a vectorial targeting approach to solve the satellite-aided capture problem. Yam 6 and Okutsu et al. 7 present a central-body-changing algorithm that can model a satellite-aided capture using a patched-conic method. The algorithm accepts a planet-centeredV1 vector and a moon’s radius vector as input and outputs the phase angle between the incoming Jupiter-centered asymptote and the flyby. Landau et al. 8 proposed using solar electric propulsion (SEP) to reduce the Jupiter arrivalV1 of interplanetary trajectories in order to ballistically capture an SEP spacecraft into orbit about Jupiter with gravity assists of one or two of Jupiter’s Galilean moons. The first implementation of a single-satellite-aided capture occurred during the Galileo mission to Jupiter. 9 Two proposed NASA missions to Jupiter (the Europa Orbiter Mission 10‐12 and the Jupiter Icy Moons Orbiter 13 ) had designed single-satellite-aided capture trajectories using Ganymede and Callisto, respectively. The planned Jupiter Europa Orbiter mission has a nominal trajectory that includes an Io-aided capture. 14 Lynam et al. 15 proposed the use of double-, triple-, or quadruple-satellite-aided capture as another method of capturing a spacecraft into orbit around Jupiter with still lower V cost. That study focuses on the Jupiter capture phase of these trajectories and only briefly discusses the interplanetary trajectory phase of these missions. In this paper, we connect several of the multiplesatellite-aided capture sequences designed by Lynam et al. with interplanetary trajectories to form complete trajectories from Earth launch to Jupiter capture. Figure 1 shows a complete trajectory that begins at Earth launch and then performs gravity-assist flybys of Ganymede and Io as it approaches Jupiter. We also determine how often some of these capture trajectories are available for use in a future mission and how much total mission V can be saved. As in Lynam et al., we assume that the (non-trivial) navigational challenges associated with these trajectories can be surmounted.

Generation of Optimal Trajectories for Earth Hybrid Pole Sitters

A pole-sitter orbit is a closed path that is constantly above one of the Earth's poles, by means of continuous low thrust. This work proposes to hybridize solar sail propulsion and solar electric propulsion (SEP) on the same spacecraft, to enable such a pole-sitter orbit. Locally-optimal control laws are found with a semi-analytical inverse method, starting from a trajectory that satisfies the pole-sitter condition in the Sun-Earth circular restricted three-body problem. These solutions are subsequently used as first guess to find optimal orbits, using a direct method based on pseudospectral transcription. The orbital dynamics of both the pure SEP case and the hybrid case are investigated and compared. It is found that the hybrid spacecraft allows savings on propellant mass fraction. Finally, it is shown that for sufficiently long missions, a hybrid pole-sitter, based on mid-term technology, enables a consistent reduction in the launch mass for a given payload, with respect to a pure SEP spacecraft.

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 near- to 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.

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.

Effect of Voltage Level on Power System Design for Solar Electric Propulsion Missions

This paper presents study results quantifying the benefits of higher voltage, electric power system designs for a typical solar electric propulsion spacecraft Earth orbiting mission. A conceptual power system architecture was defined and design points were generated for several system voltages using state-of-the-art or advanced technologies. A 300-V “direct-drive” architecture was also analyzed to assess the benefits of directly powering the electric thruster from the photovoltaic array without up-conversion. Computational models were exercised to predict the performance and size power system components to meet spacecraft mission requirements. Pertinent space environments were calculated for the mission trajectory and an electron current collection model was developed to estimate photovoltaic array losses due to natural and induced plasma environments. The secondary benefits of power system mass savings for spacecraft propulsion and attitude control systems were also quantified. Results indicate that considerable spacecraft wet mass savings were achieved by the 300-V and 300-V direct-drive architectures.

Planetary mission capability of small low power Solar Electric Propulsion systems

Planetary mission performance is presented for small, low power Solar Electric Propulsion spacecraft launched on a Delta II (7925) launch vehicle. The planetary missions presented in this paper are those that appear most attractive for a small, lowcost solar electric propulsion mission. These missions include comet and main belt asteroid rendezvous and high energy outer solar system flyby missions, all requiring a large ΔV commitment from the propulsion system. The outer solar system missions include solar probe missions using a Jupiter gravity assist and flyby missions to small outer solar system bodies. Included in this paper are examples of heliocentric transfer trajectories and estimates of performance for selected targets for each class of mission.

Robust low order dynamic controller for flexible spacecraft

The design of an optimal low-order dynamic controller with α-shift (alpha-shift) for attitude control is attempted for a flexible satellite which is a solar electric propulsion spacecraft (SEPS). The satellite consists of a central rigid body to which two flexible solar panels are attached. The design of a third-order dynamic controller is accomplished using modified linear quadratic Gaussian theory. The improvement in the satellite attitude response with a small uniform α-shift is demonstrated by numerical simulation. A novel concept of block-shift is introduced to move different sets of closed-loop eigenvalues by different amounts. The advantages of this type of block-shift in the design of the controller are shown by simulation. The robustness of the controller with respect to parameter variations is shown to be ample.

An optimal dynamic output feedback controller for a flexible spacecraft

AbstractThe design of a low‐order dynamic output feedback controller for a flexible spacecraft is carried out using modified linear quadratic Gaussian (LQG) theory. The necessary conditions for optimization are given. The linear equations governing a flexible spacecraft with a central rigid body and two sun‐tracking solar panels are presented. The parameters of the Solar Electric Propulsion Spacecraft (SEPS) are selected for analysis and simulation. The optimal gains for the dynamic controller are estimated using an iterative algorithm. The sequential procedure which assures convergence is selected. The initial gains which stabilize the system are chosen on the basis of the principle of a PD controller. A third‐order controller for pitch, roll and yaw axes of the 18th‐order system, which includes sensor and actuator dynamics, is designed. Numerical simulations carried out to ascertain the performance of the controller show the performance to be satisfactory.

A Mechanism for Three-Axis Control of an Ion Thruster Array

Three axis attitude ion thrust vector control mechanism for solar electric propulsion spacecraft, discussing gimbal and translation actuators, array and cabling