Transfer Trajectory Research Articles

State Theory, Grassroots Agency, and Global Policy Transfer: The Life and Death of Colombia’sEscuela Nuevain Brazil (1997–2012)

This article analyzes the transfer and 15-year policy trajectory of Colombia’s “global best practice” Escuela Nueva in Brazil. This program, initially transferred to Brazil in 1997 with the help of the World Bank, was largely unknown for the first decade of its life span. Then, between 2008 and 2011, after the World Bank stopped funding the program, Escuela Nueva / Escola Ativa suddenly became one of the most well funded and controversial programs in the Brazilian Ministry of Education. Continual protest and unrest concerning the program led to its termination in 2012. This article argues that it is only possible to understand these developments through an explicit theory of the “contested” state, wherein the state’s purpose is understood as both social reproduction and mediating class conflicts. Drawing on the global policy transfer literature, this framework emphasizes the role of elite actors, transnational agencies, and grassroots mobilization in determining educational policy trajectories.

Study of Arjuna-Type Asteroids for Low-Thrust Orbital Transfer

This paper investigates the accessible low-thrust transfer trajectories for a near-Earth asteroid transfer mission. The target asteroids considered are Arjuna-type asteroids, which are characterized by their Earth-like orbital paths including low eccentricity and low inclination. The asteroid range is characterized by a specific range of semimajor axes and transfer angles to provide an overall assessment of the potential Arjuna transfer domain. A single hovering ion beam spacecraft is employed for the task of asteroid redirection. The method uses a continuous thrust over the duration of the transfer maneuver to redirect the asteroid to an Earth-bound orbit. The transfer model employs a minimized form of Gauss’s variational equations to determine the available trajectories for asteroid redirection. The transfer model employs, in addition to the aforementioned orbital equations, spacecraft thruster and sizing metrics as well as mission cost analysis formulae. The system parameters and orbital transfer paths are assessed with regard to key mission parameters, namely, time frame for redirection, number of orbital revolutions, system mass, propellant mass, thrust, power, system cost, and financial return rate.

Application of the Sequential Gradient Restoration Algorithm to the Solution of the Low Thrust Transfer Problem

The Sequential Gradient Restoration Algorithm (SGRA) is an indirect optimisation method based on the iterative alternation of two phases: in the first one (gradient), the objective is the minimisation of the cost function; in the second phase (restoration) the objective is the fulfilment of the differential equations, equality and inequality path constraints and boundary conditions. This paper illustrates the application of the SGRA to the generation of very long transfer trajectories (up to some few thousand orbits during several months) in a typical GTO to GEO transfer scenario performed with electric propulsion. The reason for using electric propulsion in that case is the significant propellant mass savings if compared to the classical chemical propulsion strategies, at the expenses of an elongated transfer phase (from few days to some months). The electric propulsion system considered in this paper is characterised by thrust levels between 100 mN and 1 N (for some 3 to 5 tons launch mass) and high specific impulse, between 1000 s and 5000 s.

Trajectory design for a rendezvous mission to Earth’s Trojan asteroid 2010 [formula omitted

In this paper a rendezvous mission to the Earth’s Trojan asteroid 2010 TK7 is proposed, and preliminary transfer trajectories are designed. Due to the high inclination (∼20.9°) of the target asteroid relative to the ecliptic plane, direct transfers usually require large amounts of fuel consumption, which is beyond the capacity of current technology. As gravity assist technique could effectively change the inclination of spacecraft’s trajectory, it is adopted to reduce the launch energy and rendezvous velocity maneuver. In practical computation, impulsive and low-thrust, gravity-assisted trajectories are considered. Among all the trajectories computed, the low-thrust gravity-assisted trajectory with Venus–Earth–Venus (V–E–V) swingby sequence performs the best in terms of propellant mass. For a spacecraft with initial mass of 800kg, propellant mass of the best trajectory is 36.74kg. Numerical results indicate that both the impulsive and low-thrust, gravity-assisted trajectories corresponding to V–E–V sequence could satisfy mission constraints, and can be applied to practical rendezvous mission.

Locally-optimal electric sail transfer

This paper analyzes the locally-optimal heliocentric transfer of a spacecraft propelled by an electric solar wind sail, an innovative propellantless propulsion system that generates a propulsive acceleration exploiting the momentum of solar wind particles. The potentialities of such an advanced thruster are investigated in terms of flight times required to achieve a given heliocentric orbit. The problem is addressed using a locally-optimal formulation, by minimizing a scalar performance index that depends on the time derivatives of the osculating orbital elements. The proposed algorithm gives an estimate of the globally-optimal flight time with reduced computational efforts compared to a traditional optimization approach. Also, when the performance index involves a single orbital parameter and the transfer trajectory is two-dimensional, the proposed approach provides an analytical solution to the locally-optimal control problem. The procedure discussed in the paper is used to quantify the near-optimal performance of an electric solar wind sail in some advanced mission scenarios, such as the design of a heliocentric non-Keplerian orbit for solar activity monitoring, the exploration of the Solar System boundaries, and the rendezvous with comets 1P Halley and 67P/Churyumov-Gerasimenko.

Analysis of direct transfer trajectories from LL2 halo orbits to LLOs

A convenient procedure for designing the direct transfer trajectory from lunar L2 point (LL2) halo orbit to a low lunar orbit (LLO) is presented in this paper. The trajectory characteristics are analyzed to support the manned lunar missions design aimed at lunar surface global access. The concise procedure is established based on the circular restricted three-body problem (CR3BP) model. An analytical algorithm is employed to estimate an initial maneuver vector for approaching the Moon in its close vicinity. An iteration process is adopted to generate favorable trajectory that satisfies the constraints at perilune. By introducing a number of intermediate coordinate frames, an algorithm to compute the arriving LLO inclination and right ascension of ascending node (RAAN) is proposed. The orbital inclination and RAAN in this paper are defined and established in the J2000 frame rather than in the synodical frame. Numerical results show that, regardless of value of out-of-plane amplitude ( $Az$ ) of the halo orbit, the overall maneuver cost of the trajectory largely depends on departure position, and it has two minima around 0.65 km/s. Further study shows that the values of the arriving LLO inclination and RAAN largely depend on the choices of the departure time and the value of $Az$ . The periodicity, due to the natural motion of the Moon, is discovered to play a role in this time dependency. It is concluded that the fuel optimal trajectory permits access to almost any final lunar orbit, including a polar orbit, by means of varying the departure time and $Az$ value.

Study of Time-Free Transfers into Libration Point Orbits with Multiple Constraints

Preliminary designs of time-free transfers connecting a low Earth orbit with a halo orbit around the sun–Earth point are developed with impulse maneuver and multiple orbital element constraints. As a practical application, the constraints, including orbital altitude, inclination, right ascension of ascending node, and flight-path angle, are considered to describe the circular Earth parking orbit. The problem is then solved directly by a differential correction, using the constraint partial derivatives relative to state vector and Julian day. To reduce the sensitivity of the initial condition for differential correction, initial value functions are constructed based on the relationship of free variables and constraints. Numerical results show that for halo orbit insertion ranges from a few meters per second to hundreds of meters per second, depending on the specific inclination and right ascension of ascending node of Earth parking orbit, and demonstrate that the method for designing transfer trajectories with multiple constraints is feasible.

Алгоритм нахождения траектории перелёта между двумя эллиптическими орбитами

Алгоритм нахождения траектории перелёта между двумя эллиптическими орбитами

Low-energy transfers to cislunar periodic orbits visiting triangular libration points

This paper investigates the cislunar periodic orbits that pass through triangular libration points of the Earth–Moon system and studies the techniques on design low-energy transfer trajectories. In order to compute periodic orbits, families of impulsive transfers between triangular libration points are taken to generate the initial guesses of periodic orbits, and multiple shooting techniques are applied to solving the problem. Then, varieties of periodic orbits in cislunar space are obtained, and stability analysis shows that the majority of them are unstable. Among these periodic orbits, an unstable periodic orbit in near 3:2 resonance with the Moon is taken as the nominal orbit of an assumed mission. As the stable manifolds of the target orbit could approach the Moon, low-energy transfer trajectories can be designed by combining lunar gravity assist with the invariant manifold structure of the target orbit. In practice, both the natural and perturbed invariant manifolds are considered to obtain the low-energy transfers, which are further refined to the Sun-perturbed Earth–Moon system. Results indicate that (a) compared to the case of natural invariant manifolds, the optimal transfers using perturbed invariant manifolds could reduce flight time at least 50days, (b) compared to the cheapest direct transfer, the optimal low-energy transfer obtained by combining lunar gravity assist and invariant manifolds could save on-board fuel consumption more than 200m/s, and (c) by taking advantage of the gravitational perturbation of the Sun, the low-energy transfers could save more fuel consumption than the corresponding ones obtained in the Earth–Moon system.

Concurrent rendezvous control of underactuated spacecraft

The concurrent control of spacecraft equipped with one-axis unilateral thruster and three-axis attitude actuator is considered in this paper. The proposed control law utilizes attitude control channels along with the single thrust force concurrently, for three-dimensional trajectory tracking and rendezvous with a target object. The concurrent controller also achieves orbital transfer to low Earth orbits with long range separation. To demonstrate the orbit transfer capabilities of the concurrent controller, a smooth elliptical orbit transfer trajectory for co-planar circular orbits is designed. The velocity change and energy consumption of the designed orbit transfer trajectory is observed to be equivalent to that of Hohmann transfer.

Transfer of Training: The Known and the Unknown

Transfer of training is one of the oldest topics of interest to industrial and organizational (I/O) psychologists. Drawing on several meta-analytic studies and recent empirical work, we first synthesize what is now reliably known with respect to the generalization and retention of learned knowledge and skills to work contexts. The second part of our review focuses on what is unknown—the significant gaps in our knowledge where we believe new directions in our research strategies are warranted. We offer three prescriptions: (a) going one step beyond most existing studies to offer greater precision in our specification and measurement of variables and interventions, (b) connecting the dots by focusing on transfer criteria and transfer trajectories, and (c) shifting the operative paradigm of research to examine contemporary learning from a problem-centered perspective. There is ample opportunity to increase the yield on enormous organizational investments in training if transfer scholars and practitioners are fully informed of what is known and prepared to systematically confront the unknown in new and innovative ways.

Direct and indirect capture of near-Earth asteroids in the Earth–Moon system

Near-Earth asteroids have attracted attention for both scientific and commercial mission applications. Due to the fact that the Earth–Moon $$\hbox {L}_{1}$$ and $$\hbox {L}_{2}$$ points are candidates for gateway stations for lunar exploration, and an ideal location for space science, capturing asteroids and inserting them into periodic orbits around these points is of significant interest for the future. In this paper, we define a new type of lunar asteroid capture, termed direct capture. In this capture strategy, the candidate asteroid leaves its heliocentric orbit after an initial impulse, with its dynamics modeled using the Sun–Earth–Moon restricted four-body problem until its insertion, with a second impulse, onto the $$\hbox {L}_{2}$$ stable manifold in the Earth–Moon circular restricted three-body problem. A Lambert arc in the Sun-asteroid two-body problem is used as an initial guess and a differential corrector used to generate the transfer trajectory from the asteroid’s initial obit to the stable manifold associated with Earth–Moon $$\hbox {L}_{2}$$ point. Results show that the direct asteroid capture strategy needs a shorter flight time compared to an indirect asteroid capture, which couples capture in the Sun–Earth circular restricted three-body problem and subsequent transfer to the Earth–Moon circular restricted three-body problem. Finally, the direct and indirect asteroid capture strategies are also applied to consider capture of asteroids at the triangular libration points in the Earth–Moon system.

Invariant manifolds around artificial equilibrium points for low-thrust propulsion spacecraft

Low-thrust propulsion is incorporated into circular restricted three-body problem to balance the gravitational and centrifugal forces, and then artificial equilibrium points can be generated. The linear dynamics indicates that there are stable and unstable artificial equilibrium points. Around the unstable artificial equilibrium points, there are center and hyperbolic invariant manifolds. In this work, invariant manifolds around artificial equilibrium points are expressed as formal series of amplitudes corresponding to hyperbolic and center dynamics, and high-order series solutions are constructed up to an arbitrary order. By taking advantage of the series expansions constructed, the motions around unstable artificial equilibrium points can be parameterized. In order to check the validity, the practical convergence of series solutions truncated at different orders is considered. Finally, series expansions of invariant manifolds are applied to designing transfer trajectories from the primary to periodic orbits around artificial equilibrium points which are located inside $L_{1}$ and beyond $L_{2}$ points.

The MetNet vehicle: a lander to deploy environmental stations for local and global investigations of Mars

Abstract. Investigations of global and related local phenomena on Mars such as atmospheric circulation patterns, boundary layer phenomena, water, dust and climatological cycles and investigations of the planetary interior would benefit from simultaneous, distributed in situ measurements. Practically, such an observation network would require low-mass landers, with a high packing density, so a large number of landers could be delivered to Mars with the minimum number of launchers.The Mars Network Lander (MetNet Lander; MNL), a small semi-hard lander/penetrator design with a payload mass fraction of approximately 17 %, has been developed, tested and prototyped. The MNL features an innovative Entry, Descent and Landing System (EDLS) that is based on inflatable structures. The EDLS is capable of decelerating the lander from interplanetary transfer trajectories down to a surface impact speed of 50–70 m s−1 with a deceleration of < 500 g for < 20 ms. The total mass of the prototype design is ≈ 24 kg, with ≈ 4 kg of mass available for the payload.The EDLS is designed to orient the penetrator for a vertical impact. As the payload bay will be embedded in the surface materials, the bay's temperature excursions will be much less than if it were fully exposed on the Martian surface, allowing a reduction in the amount of thermal insulation and savings on mass.The MNL is well suited for delivering meteorological and atmospheric instruments to the Martian surface. The payload concept also enables the use of other environmental instruments. The small size and low mass of a MNL makes it ideally suited for piggy-backing on larger spacecraft. MNLs are designed primarily for use as surface networks but could also be used as pathfinders for high-value landed missions.

Design and evaluation of a variable stiffness manual operating platform for laparoendoscopic single site surgery (LESS).

Most of the existing robotic platforms for LESS have workspace and load capacity weaknesses, because of the limitation of one single incision. We have developed a LESS manual operating platform of which the stiffness of the insertion tube is controllable. The system included two dexterous tool manipulators, a stereo-vision module and a variable stiffness insertion tube (VSIT), which was designed using phase-change material (mixed indium, gallium and stannum). Experiments to evaluate the effectiveness of the VSIT were set up. Peg transfer tasks and trajectory tracking tasks were conducted to assess the initial performance of the overall system. The experimental results for stiffness characteristic suggested that the rigidity of the VSIT with a straight-forward pose was considerably increased by about four times in the rigid mode. Peg transfer tasks and trajectory tracking tasks were performed successfully with an average time of 97s and 52s, respectively. The experimental results for stiffness characteristic showed that the manual operating platform had great promise for solving large workspace, high manipulation force and stability problems in LESS. The tool manipulators had the ability to achieve basic operations.

Ecological Land Fragmentation Evaluation and Dynamic Change of a Typical Black Soil Farming Area in Northeast China

Ecological land is a land use category provided with considerable ecological value and a vital indicator reflecting regional eco-environmental quality. However, it has experienced severe fragmentation during the rapid urbanization in China which strongly threatened the regional ecological security, land use pattern and human living environment. Therefore, analysis of spatiotemporal change of ecological land use and ecological landscape pattern is particularly essential. In this paper, a case study was made in Nong’an County, which is a typical black soil farming area located in northeast China facing severe conflicts among cultivated land protection, urban expansion and ecological security. A landscape fragmentation evaluation model was proposed to measure the degree of regional ecological land fragmentation. We also determined the land use change features through the methods of dynamic change information exploration and by performing transfer trajectory analysis during the period from 1996 to 2014. The results showed that the ecological land in Nong’an County has experienced increasing fragmentation during the past 18 years. The statistical results showed that the land transition between ecological land and other land categories was quite frequent, and it especially appeared as a dramatic decline of grassland and severe increase of saline-alkali land. In addition, human interferences especially construction activities and cultivated land occupation were still the dominant factors to the fragmentation of ecological land and the frequent transition among the land use categories. The fragmentation degree showed a downward tendency at the end of the study, which indicated noticeable benefits of land use regulation and land protection policies directed towards land ecological value. This study aims to provide a scientific evaluation model for measuring ecological land fragmentation degree, and figure out the regional land use transition relationships to offer suggestions for decision-making and provide a practical case in a typical region.

A star-photon sailcraft mission in the Alpha Centauri system

The recent findings of exoplanets have renewed interest in exploration missions of extra-solar star systems. In this context the binary star system of Alpha Centauri A/B plays a scientifically important role. The current paper aims at investigating the capture and the transfer trajectories from A to B performed by a stellar-photon sailcraft, which gives the best performance in long duration missions. The mission has been divided into four phases and minimum transfer time trajectories have been obtained by using a direct optimisation approach. Finally, a preliminary analysis in terms of Zero Velocity Curves has been carried out.

Dephasing Times in the Phycobiliprotein Antenna Complexes PE545 and PE555

Since long-lived quantum coherence have not only been observed in the Fenna-Matthews-Olson (FMO) complex of green sulfur bacteria but also in the phycoerythrin 545 (PE545) photosynthetic antenna system, the interest in these systems has been increased significantly. Moreover, similar results have been found for the PE555 complex in which the two alpha-beta monomers are rotated compared to the PE545 structure leading to a water filled channel. By now, all three complexes have been investigated by a sequential combination of molecular dynamics (MD) simulations, calculations of vertical excitation energies along the trajectory and quantum dynamics for the exciton transfer [1-3]. A comparison of the obtained spectral densities and dynamics will be provided. An interesting property for the systems is the dephasing time which can give an estimate for how long quantum coherences might survive in the respective system. In a first step, the relationship between dephasing time and energy gap fluctuations of the individual pigments has been determined [4]. For the above mentioned complex but also additional exciton and charge transfer systems it can be confirmed that that an inverse proportionality exists between dephasing time and average gap energy fluctuation. Interestingly, an interestingly very similar behavior has been found for all these systems. In a subsequent step, the relationship between dephasing time and excitonic energy gap fluctuations for entire complexes including the respective inter-molecular couplings has been studied.[1] C. Olbrich, J. Strumpfer, K. Schulten, U. Kleinekathofer, J. Phys. Chem. Lett. 2, 1771 (2011).[2] M. Aghtar, J. Strumpfer, C. Olbrich, K. Schulten, U. Kleinekathofer, J. Phys. Chem. Lett. 5, 3131 (2014).[3] S. Chandrasekaran, K. R. Pothula, U. Kleinekathofer, J. Phys. Chem. B, DOI: 10.1021/acs.jpcb.6b05803 (2016).[4] M. I. Mallus, M. Aghtar, S. Chandrasekaran, G. Ludemann, M. Elstner, U. Kleinekathofer, J. Phys. Chem. Lett. 7, 1102 (2016).

Patched Conic Section Maneuver Trajectory Planning For Two-Craft Coulomb Formation

Actively controlled electrostatic forces are utilized to control the relative motion of the two-craft formation. The maneuver problem envisioned requires that the maneuver time, the final position, and final velocity vectors match the prescribed values. Two inequality constraints, the minimum distance requirement and the maximum charge level constraint, are taken into consideration. Three-phase patched conic section trajectory approach is proposed to accomplish the maneuver mission. The circular transfer trajectory and the general patched conic section trajectory programming algorithms are developed. The existence of solutions of these two types of trajectories are investigated. Two special cases, which are equally charged trajectory and the joint periapsis trajectory, are taken as examples to show the effectiveness of the trajectory design method.

Direct transfer trajectory design options for interplanetary orbiter missions using an iterative patched conic method

In a direct interplanetary transfer, the spacecraft moves from a parking orbit of the departure planet to a parking orbit of the arrival planet. The transfer trajectory must be designed such that the specified arrival parking orbit conditions are achieved. For a fixed departure epoch and flight duration, there are four distinct transfer trajectory design options in a direct transfer. The conventional patched conic method, the most widely used analytical trajectory design method, does not identify these design options. An iterative patched conic method that identifies these distinct design options is developed and presented in this paper. This method involves two iterative processes: (i) iteration on the hyperbolic orbit characteristics using an analytical tuning strategy to achieve the hyperbolic excess velocity vector at the patch point, (ii) iteration on the patch points at the sphere of influence. The performance of the proposed method is compared with the conventional and V-infinity tuned patched conic methods. A design analysis tool, based on the proposed method, is developed and tested in various orbiter mission scenarios.