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The article presents information on the development of Starship spacecraft by private rocket company SpaceX. Topics discussed include the plan of U.S. National Aeronautics and Space Administration's (NASA) Ames Research Center planetary scientist Jennifer Heldmann to use Starship to probe lunar ice, the failure of NASA's Lunar Crater Observation and Sensing Satellite mission, and the objectives of SpaceX chief executive officer Elon Musk in developing Starship.

Highlights • Experimental and theoretical methodology for data acquisition of unsteady thermal state. • Description of test equipment and the algorithm for solving the inverse heat conduction problem. • Study of carbon/carbon shell for its use as hot structure in application such as windward connections. Abstract The aim of this paper is to study a shell structure for space applications. The aim of the structure is to be reusable and thick. This first study will focus on thermal environment. Starting from a defined geometry, a prototype will be manufactured. Thermo-structural behavior of the structure will be analyzed by numerical analysis and tests. Thermal properties, such as thermal conductivity and heat capacity, will be studied by the use of the inverse method. A robust numerical approach, such as inverse method, is one of the best for this problem as many parameters concur for the determination of properties. Such approach permits to perform the parametric and structural identification of the model. These procedures are presented including both experimental investigation and methodical-numerical aspects. Special test equipment and the regularizing algorithm for solving the ill-posed inverse heat conduction problem are described. In the frame of thermal properties determination, a verification and prediction of thermocouple error will be performed. Developed in the frame of this work the experimental and theoretical methodology for complex data acquisition of unsteady thermal state of the prototype of shell structure is proposed. [ABSTRACT FROM AUTHOR]

This paper discusses spacecraft attitude control using variable-speed control moment gyroscopes. A new operational concept for variable-speed control moment gyroscopes is proposed, i.e., the flywheels and the gimbals of the variable-speed control moment gyroscopes do not always spin at high speed, they spin at high speed only when they need to. The dynamics of the variable-speed control moment gyroscopes are incorporated into the model for the control system design. The complex nonlinear system can be represented as an equivalent linear time-varying system. As a result, an effective control system design method, model predictive control using robust pole assignment, can be used to design the spacecraft attitude control system using variable-speed control moment gyroscopes. A combination of the new operational concept and a model that includes the variable-speed control moment gyroscopes’ dynamics leads to several nice features: 1) the control system does not have any singular point; 2) the design reduces energy consumption because the fly wheels and gimbals are most likely operated at low speed when the spacecraft is stabilized; and 3) the theoretical analysis shows that the designed system is uniformly exponentially stable. A design example is provided. The simulation result shows the effectiveness of the proposed method. [ABSTRACT FROM AUTHOR]

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AVANTI is a formation-flight experiment involving two noncooperative satellites. After a brief overview of the challenges that experiment design and scenario induce, this paper presents the safety concept retained to guarantee the safety of the formation. The peculiarity of the proposed approach is that it does not rely on the continuous availability of tracking data of the client spacecraft but rather exploits the concept of passive safety of special relative trajectories. To this end, the formation safety criterion based on the minimum distance normal to the flight direction has been extended in order to be applicable also to drifting relative orbits, resulting from non-vanishing relative semi-major axis encountered during a rendezvous or produced by the action of the differential aerodynamic drag. [ABSTRACT FROM AUTHOR]

The use of electric propulsion (EP) for geostationary Earth orbit (GEO) stationkeeping and geostationary transfer orbit offers satellite operators the opportunity to reduce mission costs and to increase revenue by enabling a higher dry mass to orbit compared with traditional chemical or EP systems. The penalty for such benefits comes, initially, in the form of an increased time-to-orbit whereby the low-thrust transfer orbit duration will range from a few months, at best, to possibly one year or more. During the low-thrust transfer, the spacecraft will experience prolonged exposure to the portions of the trapped radiation belts to which the GEO spacecraft otherwise would not be subjected—most notably the inner belt and slot region. Further, the spacecraft will also experience on the order of 10^5 h of high-density high-energy plume plasma—a more challenging operational environment compared with that typically applied to the GEO spacecraft. Applicable environments are introduced, an overview of specific operational effects is provided, and design guidelines are highlighted. [ABSTRACT FROM PUBLISHER]

We have reviewed the basic results in the radiation-induced conductivity and the bulk charging of polymers obtained by our group during the past 30 years, which are mostly unknown in the West. Special attention is given to a new experimental technique extensively used in our investigations based on an electron-gun technology allowing combined induced conductivity and carrier mobility measurements. Quasi-band and hopping theories of the carrier transport are critically discussed. [ABSTRACT FROM PUBLISHER]

There are two competing tendencies in spacecraft (SC) design and manufacturing. On the one hand, it is a never-ceasing upgrading of the SC electronics to lower energy consumption and increase its functionality, which in turn makes it more susceptible to electrostatic discharges (ESDs). On the other hand, designers constantly seek to improve electronics protection against ESDs. Spectr-R SC features on-board equipment, which is highly sensitive to possible ESDs. To exclude ESD-related failures, the following protection measures have been taken. Based on the Satellite-MIEM computer code, we developed the lumped-element model of the Spectr-R body currents and cable-affected electronic response. This model made it possible to compute SC body currents initiated by ESDs originating at various structural sites. For various cable harnesses, we measured transformation coefficients converting body currents into the amplitude of the noise voltages appearing in a cable (a patented methodology was used for this purpose). Afterward, these transformation coefficients together with the ESD-induced currents have been fed into the Satellite-MIEM software to compute noise voltages appearing in electronic input circuits, connected via these cables to the satellite exterior surface. Comparing protection efficiency of cables and electronic equipment allowed us to select suitable cable harnesses and the appropriate equipment protection to insure reliable in-flight performance of the space vehicle. From the launch in July 2011 and up to the present time, there were no failures of the Spectr-R electronics, which could be traced to ESDs. [ABSTRACT FROM PUBLISHER]

On July 4th, 2005, in celebration of our nation's birthday, NASA's Deep Impact Impactor spacecraft collided with comet Tempel1 at 10 km/sec – marking the first hypervelocity impact of a celestial body by a human-made spacecraft. With closing speeds of 23,000 mph, the Impactor's active guidance system steered it to impact on a sunlit portion of the comet's surface. As it closed in on Tempel 1, the Impactor's camera relayed close-up images of the comet's surface to the Flyby spacecraft for downlink to Earth. Meanwhile, the Flyby spacecraft used its two instruments to image the impact and then continued to photograph the comet as it followed its orbital path around the Sun. The primary science data was returned to Earth in near real-time, and all data was returned to Earth within 24 hours of the encounter. For the NASA Discovery-class Deep Impact mission, a two-part Deep Impact spacecraft was constructed: the Impactor spacecraft and the impact characterization (flyby) spacecraft, and an associated suite of surveillance instruments. These instruments included one high resolution visible imager, two identical medium-resolution visible imagers (one on the flyby and one on the Impactor) and one infrared spectrometer. The two-part spacecraft launched together in January 12, 2005 and separated on July 3 rd , 24 hours before reaching its Tempel 1 target. The Impactor separated from the flyby spacecraft and autonomously positioned itself directly in front of the encroaching Tempel 1 comet for a spectacular hypervelocity impact. This paper highlights portions of the article Deep Impact: Excavating Comet Tempel 1 written by the Deep Impact science team and published in Science as well as on-line by Science Express, 09/08/05. [ABSTRACT FROM AUTHOR]

This paper describes a novel framework to determine the relative pose and range of a solid-of-revolution-shaped spacecraft from a single image without any artificial beacons. The translation and the symmetry axis of the spacecraft can be estimated from the imaged cross sections of the spacecraft body. Then the pose and range of the spacecraft are fully determined by means of the images of its solar panels and asymmetric feature. Our method has been validated by both synthetic and real images. [ABSTRACT FROM AUTHOR]

This paper proposes a multiobjective optimization method for the control-structure integrated design of flexible spacecraft to reduce the total mass and optimize the control performance. The equations of motion for flexible spacecraft are derived from the Lagrange's principle and the assumed modes method. The design variables are the structural dimensions of the flexible structure and controller parameters. The objectives and constraints are derived from structure and control performance indexes. The objectives include total mass, control cost, and vibrational energy, and the constraints include the stability of the closed-loop system, settling time, overshoot, maximum control, and maximum vibrational displacement of the tip. A modified version of the multiobjective evolutionary algorithm based on decomposition (MOEA/D) with our proposed hybrid constraint handling method is proposed for optimization. As a case study, it has been applied to a spacecraft with symmetrically installed flexible appendages to find optimal tradeoffs in control-structure design. The simulation results show that the multiobjective optimization method for the control-structure integrated design of flexible spacecraft is feasible and effective, and could give an improvement of structural and control designs. [ABSTRACT FROM AUTHOR]

In order to reduce steps in developments of spacecraft and onboard instruments, we developed two tools SIB2PlanEditor and SIB2Viewer. SIB2PlanEditor is dedicated editor for test procedure described in the ISAS command plan language. SIB2Viewer is a program to enhance capability of user interface for SIB2 (SIB2UI). Both tools reduce simple work in developments and contribute designer of spacecraft and onboard instruments to concentrate on preparation of content of test procedure and design itself. This paper describes function of these tools and evaluated merits through trial usage. By usage of SIB2PlanEditor, time needed for input of test procedure is reduced by factor 1/3~1/2. These tools are also expected to contribute to reduce time of execution of electrical integration and test of the spacecraft through promotion of SIB2 input in earlier stage of the development. [ABSTRACT FROM AUTHOR]

The article offers information related to the Dragon that is a partially reusable spacecraft developed by American private space transportation company Space Exploration Technologies Corp. It reports that Google Inc. has launched the Lunar X-Prize that is a largest international incentive based prize. . It also offers information on a prototype robotic lander called Mighty Eagle developed by the U.S. National Aeronautics & Space Administration.

In this paper, the problem of attitude control of a flexible spacecraft is investigated. Three controllers are presented. The first controller is based on dynamic inversion, while the second is based on the µ-synthesis method and the third approach is composed of dynamic inversion and µ-synthesis schemes. It is assumed that only one torque on the hub is used. Actuator saturation is also considered in the design of controllers. To evaluate the performance of the proposed controllers, an extensive number of simulations on the model of the spacecraft are performed. The performances of the proposed controllers are compared in terms of nominal performance, robustness to uncertainties, suppression of panel vibration, sensitivity to measurement noise, environment disturbance, and nonlinearity in large maneuvers. Simulation results confirm the ability of the proposed controller in tracking the attitude trajectory while damping the panel vibration. It is also verified that the perturbations, environment disturbances, and measurement errors have only slight effects on the tracking and damping performances. It is notable that the composite method (dynamic inversion and µ-synthesis) is not new; however, this application is new. [ABSTRACT FROM PUBLISHER]

We propose controllers for leader-follower attitude synchronization of spacecraft formations in the presence of disturbances, that is, the leader spacecraft is controlled to follow a given reference, while a follower spacecraft is controlled to synchronize its motion with the leader's. In the ideal disturbance-free scenario, we show that synchronization takes place asymptotically. Moreover, we show the property of uniform practical asymptotic stability which implies that the synchronization is robust to bounded disturbances. [ABSTRACT FROM AUTHOR]

Most spacecraft are usually assembled from some simple substructures by various joints. These joints are generally nonlinear and it can strongly affect the dynamics of spacecraft. Finite Element (FE) models of spacecraft usually have many degrees of freedom (DOFs), and the nonlinear joints are generally spatially localized. It is difficult to calculate the forced harmonic response of spacecraft by using current methods. In the present paper, a new approach is proposed to overcome the difficulties. The approach can greatly reduce the computational efforts by only keeping DOFs associated with nonlinear elements, DOFs on which external forces act and DOFs that are of interest. Hence it can be applied to large-scale FE models with local nonlinearities. A FE model of a satellite with some nonlinear joints is used to demonstrate the concept and show the effectiveness of the proposed procedure. The effects of the joints on the acceleration response of payloads in the frequency domain are also discussed. [ABSTRACT FROM AUTHOR]

Abstract: The space cable consists of evacuated tubes supported by fast-moving objects, called bolts, travelling inside them. The bolts are connected to the tubes via permanent magnets stabilized electronically for minimal power consumption. Version 1 was designed to replace the role of a first-stage rocket by lifting space vehicles from the ground to 50km at 5800km/h. Version 2 rises to 140km and has wider applications. Comparable proposals include the space fountain and the launch loop. In the long term, such a fixed infrastructure is more economical than rockets. A key issue with these dynamically supported structures is lateral stability, particularly in the presence of varying cross winds in the stratosphere. The relevant partial differential equations are derived below with a set of solutions. A proposal is presented for providing stability by means of tethers, pipes, and a support structure at each end. Each support is a miniature space cable rising to 15km, and it bears the load of the tethers and pipes that would otherwise have to be carried by the main structure. [Copyright &y& Elsevier]

Abstract: More than ten years experience in hands-on space education has been achieved at Scuola di Ingegneria Aerospaziale of Università di Roma “la Sapienza”, where the UNISAT program was established in the early nineties. The students participating in this program are involved in a microsatellite design, manufacturing, test, launch and operation in orbit activity, from initial mission concept to operation in orbit. The microsatellite program develops in a two years timeline, fitting with the graduate student program curricular activity at Scuola di Ingegneria Aerospaziale. Four microsatellites have been launched every other year since 2000 from the Baikonour Cosmodrome by the DNEPR LV. In this way there was the opportunity to exploit the UNISAT platform to perform small scientific and technological experiments in orbit. Besides education, a main goal of the UNISAT program is testing in orbit commercial off-the-shelf components, which allow to keep the program cost low and compatible with the University research budget. The main spacecraft subsystems, including the in orbit technological and scientific experiments, and the ground station operations are briefly described in the paper, focussing on the education and research aspects. [Copyright &y& Elsevier]

Abstract: In the given article scientific and technical complex for modeling, researching and testing of rocket-space vehicles’ power installations which was created in Power Source Laboratory of National Aerospace University “KhAI” is described. This scientific and technical complex gives the opportunity to replace the full-sized tests on model tests and to reduce financial and temporary inputs at modeling, researching and testing of rocket-space vehicles’ power installations. Using the given complex it is possible to solve the problems of designing and researching of rocket-space vehicles’ power installations efficiently, and also to provide experimental researches of physical processes and tests of solar and chemical batteries of rocket-space complexes and space vehicles. Scientific and technical complex also allows providing accelerated tests, diagnostics, life-time control and restoring of chemical accumulators for rocket-space vehicles’ power supply systems. [Copyright &y& Elsevier]

A multi-functional structure saves mass from a spacecraft by incorporating other functional subsystems into the structure. By using the structural properties of a non-structural element, inert structure may be eliminated, and the requirement to allot internal volume to the subsystem in question is removed. The current paper describes a multi-functional structure based on the secondary power system. By using commercially available plastic lithium-ion cells to form the core of a sandwich panel, inert mass is eliminated from both the structure and from the battery enclosure. The feasibility of the proposed multi-functional structure is demonstrated though vibration testing on a single cell, and the successful manufacture of a test panel. The work goes on to quantify the potential mass savings that may be achieved by using a multi-functional structure of this type. By varying a set of spacecraft attributes, the study identifies that small spacecraft with high power requirements have the potential to gain the most benefit from using a multi-functional structure of this type. [ABSTRACT FROM AUTHOR]

We present a new European Mars mission proposal to build on the UK-led Beagle2 Mars mission and continue its astrobiology-focussed investigation of Mars. The small surface element to be delivered to the Martian surface—Vanguard—is designed to be carried by a Mars Express-type spacecraft bus to Mars and adopts a similar entry, descent and landing system as Beagle2. The surface element comprises a triad of robotic devices—a lander, a micro-rover of the Sojourner class for surface mobility, and three ground-penetrating moles mounted onto the rover for sub-surface penetration to 5 m depth. The major onboard instruments on the rover include a Raman spectrometer/imager, a laser plasma spectrometer, an infrared spectrometer—these laser instruments provide the basis for in situ “remote” sensing of the sub-surface Martian environment within a powerful scientific package. The moles carry the instruments’ sensor head array to the sub-surface. The moles are thus required to undergo a one-way trip down the boreholes without the need for recovery of moles or samples, eliminating much of the robotic complexity invoked by such operations. [Copyright &y& Elsevier]