Bernd Dachwald

Impact of Optical Degradation on Solar Sail Mission Performance

The optical properties of the thin metalized polymer films that are projected for solar sails are likely to be affected by the damaging effects of the space environment, but their real degradation behavior is to a great extent unknown. The standard solar sail force models that are currently used for solar sail mission analysis and design do not take these effects into account. In this paper we use a parametric model for describing the sail film’s optical degradation with its environmental history to estimate the impact of different degradation behaviors on solar sail mission performance for some example interplanetary missions: the Mercury rendezvous missions, fast missions to Neptune and to the heliopause, and artificial Lagrange-point missions.

Solar Sail Kinetic Energy Impactor Trajectory Optimization for an Asteroid-Deflection Mission

A fictional asteroid-mitigation problem created by AIAA assumes that a 200-m near-Earth asteroid, detected on 04 July 2004 and designated as 2004 WR, will impact the Earth on 14 January 2015. Adopting this example scenario, we show that solar-sail spacecraft that impact the asteroid with a very high relative velocity are a realistic near-term option for mitigating the impact threat from near-Earth asteroids. The proposed mission requires several kinetic energy impactor solar-sail spacecraft. Each kinetic energy impactor consists of a 160 x 160 m, 168-kg solar sail and a 150-kg impactor. Because of their large ΔV capability, solar sailcraft with a relatively modest characteristic acceleration of 0.5 mm/s 2 can achieve an orbit that is retrograde to the target orbit within less than about 4.5 years. Before impacting 2004 WR at its perihelion of about 0.75 AU, each impactor is to be separated from its solar sail. With a relative impact velocity of about 81 km/s, each impactor will cause a conservatively estimated Δv of about 0.35 cm/s in the trajectory of the target asteroid, largely due to the impulsive effect of material ejected from the newly formed crater. The deflection caused by a single impactor will increase the Earth-miss distance by about 0.7 Earth radii. Several sailcraft will therefore be required for consecutive impacts, to increase the total Earth-miss distance to a safe value. In this paper, we elaborate on a potential mission scenario and investigate trade-offs between different mission parameters; for example, characteristic acceleration, sail temperature limit, hyperbolic excess energy for interplanetary insertion, and optical solar-sail degradation.

Heliocentric Solar Sail Orbit Transfers with Locally Optimal Control Laws

Solar sailing is increasingly being considered by space agencies for future science missions. With the absence of reaction mass from the primary propulsion system arises the potential for new high-energy mission concepts in the mid to far term, such as a Solar Polar Orbiter or an Interstellar Heliopause Probe [1,2]. One of the most time consuming tasks of mission analysis is trajectory generation and optimization. Optimal trajectory generation is a complex field and many schemes exist; however, these are typically characterized as being computationally intensive systems requiring a good degree of engineering judgment [3-6].

Minimum Transfer Times for Nonperfectly Reflecting Solar Sailcraft

Introduction S OLAR sailcraft trajectory analyses are typically made for highperformance sailcraft under the assumption that the solar sail is either an ideal reflector or by consideration of the nonideal reflectivity through an overall efficiency factor that reduces the magnitude of the solar radiation pressure (SRP) force but leaves its direction unaltered.1−3 In both cases, the SRP force model is a model of perfect, that is, specular, reflection because the direction of the SRP force is always perpendicular to the sail surface. For a thorough mission analysis, however, one must consider the optical properties of the real nonperfectly reflecting sail film, where the SRP force also has a component parallel to the sail surface. Taking the current stateof-the-art in engineering of ultralightweight structures into account, solar sailcraft of the first generation will be of relatively moderate performance. The simplification of perfect reflectivity and the limitation on high-performance sailcraft seem both to be caused mainly by the difficulty of generation of an adequate initial guess for traditional local trajectory optimization methods. As a smart method for global trajectory optimization, artificial neural networks can be combined with evolutionary algorithms to form so-called evolutionary neurocontrollers (ENCs).4 Evolutionary neurocontrol (ENC) does not require an initial guess. By the use of ENC, near globally optimal trajectories can also be calculated for nonperfectly reflecting solar sailcraft of moderate performance. In the sequel, after the different SRP force models have been defined, minimal transfer times for rendezvous missions within the inner solar system will be presented for perfectly and nonperfectly reflecting solar sailcraft, including a near-Earth asteroid rendezvous (1996FG3) and a main belt asteroid rendezvous (Vesta).

Solar Sail Trajectory Optimization for Intercepting, Impacting, and Deflecting Near-Earth Asteroids

A fictional asteroid mitigation problem posed by AIAA assumes that a 200m nearEarth asteroid (NEA), detected on 04 July 2004 and designated as 2004WR, will impact the Earth on 14 January 2015. Adopting this exemplary scenario, we show that solar sail spacecraft that impact the asteroid with very high velocity are a realistic near-term option for mitigating the impact threat from NEAs. The proposed mission requires several Kinetic Energy Interceptor (KEI) solar sail spacecraft. Each sailcraft consists of a 160m × 160m, 168 kg solar sail and a 150 kg impactor. Because of their large ∆V-capability, solar sailcraft with a characteristic acceleration of 0.5mm/s can achieve an orbit that is retrograde to the target orbit within less than about 4.5 years. Prior to impacting 2004WR at its perihelion of about 0.75AU, each impactor is to be separated from its solar sail. With a relative impact velocity of about 81 km/s, each impactor will cause a conservatively estimated ∆v of about 0.35 cm/s in the trajectory of the target asteroid, largely due to the impulsive effect of material ejected from the newly formed crater. The deflection caused by a single impactor will increase the Earth-miss distance by about 0.7 Earth radii. Several sailcraft will therefore be required for consecutive impacts to increase the total Earthmiss distance to a safe value. In this paper, we elaborate a potential mission scenario and investigate trade-offs between different mission parameters, e.g. characteristic acceleration, sail temperature limit, hyperbolic excess energy for interplanetary insertion, and optical solar sail degradation.

Refined Solar Sail Force Model with Mission Application

The aim of this paper is to propose a refined mathematical model for describing the acceleration experienced by a solar sail. Unlike the conventional model characterized by constant coefficients, the force coefficients of the sail are now assumed to depend on the light incidence angle, the sail surface roughness, and the sun–sail distance. The new model is elaborated with the support of experimental data that show how the main variable affecting the force coefficients is the light incidence angle. To emphasize the differences between the refined force model with respect to the conventional one, a comparison is established through the analysis of a circle-to-circle interplanetary rendezvous problem between coplanar orbits. The problem is solved using an indirect approach and the optimal steering law is approximated in polynomial form. A number of optimal trajectories toward Mars and Venus are simulated and the results obtained are discussed as a function of the dimensionless sail loading parameter and the sail surface roughness.

Potential Solar Sail Degradation Effects on Trajectory and Attitude Control

Paper on the potential for solar sail degradation effects on trajectory and altitude controls.

Evolutionary Neurocontrol: A Novel Method for Low-Thrust Gravity-Assist Trajectory Optimization

The combination of low-thrust propulsion and gravity assists to enhance deep-space missions has proven to be a remarkable task. In this paper, we present a novel method that is based on evolutionary neurocontrollers. The main advantage in the use of a neurocontroller is the generation of a control law with a limited number of decision variables. On the other hand, the evolutionary algorithm allows one to look for globally optimal solutions more efficiently than with a systematic search. In addition, a steepest-ascent algorithm is introduced that acts as a navigator during the planetary encounter, providing the neurocontroller with the optimal insertion parameters. Results are presented for a Mercury rendezvous with a Venus gravity assist and for a Pluto flyby with a Jupiter gravity assist.

Flight Opportunities from Mars to Earth for Piloted Missions Using Continuous Thrust Propulsion

For a piloted Mars mission, the inbound flight (Mars to Earth) is actually the most restricting space-part of the mission – especially if short flight times and short stay times at Mars are required. In this paper, different electric propulsion systems (different thrust levels, specific impulses, and thrust to weight ratios) and different flight strategies for a crewed return from Mars are analyzed and compared to high thrust propulsion systems (chemical and nuclear thermal). This is done with respect to feasibility, flight times, propellant consumption, and influence on the roundtrip problem for launch opportunities within 2016-2031. It is demonstrated that with moderate continuous thrust levels and specific impulses (100N, 3000 s) – even for short stays at Mars – Earth return trips are feasible with moderate propellant needs and within reasonable inbound flight times of 200 to 350 days.

Optimization of low-thrust Earth-Moon transfers using evolutionary neurocontrol

Although low-thrust propulsion is an interesting option for scientific and reconnaissance missions to targets in planetary space, like the Moon, associated transfer strategies pose challenging requirements in terms of optimal control. The method of Evolutionary Neurocontrol (ENC), which has been applied very successfully to interplanetary low-thrust transfer problems, is now used for solving this type of steering problem. For exemplary validation, two low-thrust transfers from an Earth-bound geostationary transfer orbit into a Moon-bound orbit are optimized with respect to minimum flight time.