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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.

To help optimize the spacecraft design and reduce the risk of spacecraft mission failure, a new approach to assess the survivability of spacecraft in orbit is presented here, including the following three steps: 1) Sensitivity Analysis of spacecraft. A new sensitivity analysis method, a ray method based on virtual outer wall, is presented here. Using rays to simulate the debris cloud can effectively address the component shadowing issues. 2) Component Vulnerability analysis of spacecraft. A function "Component functional reduction degree - Component physical damage degree" is provided here to clearly describe the component functional reduction. 3) System-level Survivability Assessment of spacecraft. A new method based on expert knowledge reasoning, instead of traditional artificial failure tree method, is presented here to greatly improve the efficiency and accuracy of calculation. [ABSTRACT FROM AUTHOR]

It is unclear whether the common finding of improved organizational performance with increasing organizational experience is driven by learning from success, learning from failure, or some combination of the two. We disaggregate these types of experience and address their relative (and interactive) effects on organizational performance in the orbital launch vehicle industry. We find that organizations learn more effectively from failures than successes, that knowledge from failure depreciates more slowly than knowledge from success, and that prior stocks of experience and the magnitude of failure influence how effectively organizations can learn from various forms of experience. [ABSTRACT FROM AUTHOR]

Abstract This paper presents the methodology and the optimization strategy applied by Bertin Technologies for over 10 years to perform space launch vehicle design and implemented by using the property software platform HADES V15.0. The problem formulation consists of finding the best launch vehicle concept i.e. the one minimizing launch cost while satisfying technical, mission and architecture constraints (payload mass on final orbit, number of stages, propulsive technologies ...). The strategy is based on a Multidisciplinary Design Feasible (MDF) approach coupled with the use of Genetic Algorithms (GA) for global optimization, and Gradient-Based Algorithms for final tuning and results refining. The HADES V15.0 platform provides the associated software environment, integrating a number of technical and economic modules consistently interconnected within a system optimization loop. The main disciplines taken into account in the platform are related to the launcher's propulsion, structure, aerodynamics, trajectory optimization and costs. The use of an integrated platform for multi-objective and multi-disciplinary optimization enables an efficient process and quick optimization. This methodology is particularly well fitted to the design of a small space launch vehicle, allowing to take into account the multidisciplinary nature of such a complex system and to manage the inherent sensitivity for this kind of vehicle. This platform was used specifically to design Bertin Technologies' cost-effective expandable Space Launch Vehicle (SLV) for Microsatellites, ROXANE, and to design the space launch vehicle in the H2020 ALTAIR project. Highlights • Development of a software platform to preliminary design space launch vehicle. • Multidisciplinary design optimization applied to space launch vehicle. • Optimization of twelve space launch vehicles for a minimum launch cost. • Application to the designs of two space launch vehicle, ROXANE and ALTAIR. [ABSTRACT FROM AUTHOR]

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]

The rapid development of emerging concepts and technologies, e.g., modularity, standardization, and fractionation, has resulted in the miniaturization of satellites and simplification of mission design, as well as a move towards more responsive and economical systems. However, the current customized and labor-intensive design philosophy is not naturally appropriate for enabling the application of these concepts and technologies. Therefore, a value-centric architecture for spacecraft design and certification is proposed in this paper to address this need. Firstly, the characteristics of different types of spacecraft are compared and summarized. In parallel, a survey on the current design and certification frameworks of large and complex systems is carried out. Taking in the experience of the frameworks, a generic design and certification architecture is developed from a value-centric perspective. Apart from keeping the critical advantages from the traditional methodologies, the new approach is capable of solving the problems inherent themselves. The results of a preliminary case study clearly show that the approach proposed can effectively capture, analyze, and optimize the value of different system designs, appealing to the growth of the space market. [ABSTRACT FROM AUTHOR]

The article discusses the significance of the computerized checkout equipment, which is beneficial to the development of Saturn 5-class rockets and the Apollo-class spacecraft in the U.S. According to the author, this equipment represents a significant advance in automation, and its impact on ordinary production and the reliability of consumer products appears to be incalculable. Moreover, the Saturn/Apollo program of the National Aeronautics and Space Administration is the developmental work in the 1940s and 1950s that culminated in a marriage of operational and diagnostic checkout systems. Meanwhile, one of the inside stories of Apollo are presented further.

There are different kinds of equipments distributed in different locations for a large complicated multi-shaker vibration test facility, so it is challenging to monitor the real state of test facility thoroughly during its operation. Long-term operation of this test facility will lead to the degradation of reliability and malfunction, and sometimes the emergency stop of the whole test system that threatens the safety of the spacecraft seriously. This paper presents in detail the design and application of a set of health monitoring device for a large multi-shaker vibration test facility which is capable of monitoring the operation state in real time and predicting the potential malfunction of the whole test facility to ensure the reliability of this large test system and safety of the spacecraft during its environmental vibration test. [ABSTRACT FROM AUTHOR]

With space developments for diversified missions, intelligent thermal design of spacecraft is essential. This report proposes an active thermal radiation control device, which consists of an array of thermal switch that is enhanced by the near-field radiation. Suspended diaphragms are snapped down by electrostatic force, when the driving voltage is applied to the top electrode. The present simulation for parallel plates of Au (lower electrode) and Cu (upper electrode) at 300 K shows that the effective emittance of 0.02 for a 1 μm separation is dramatically increased to 0.89 for a 10 nm gap. A MEMS (Micro Electro Mechanical Systems)-based active radiator was designed based on the requirements for the driving voltage and resonant frequency. The fill factor of the proposed radiator is as large as 61%. In this study, an analysis based on thermal resistance of various heat paths was conducted. It was found that the thermal resistance of the near-field effect was lower than that of contact heat conduction, indicating that the ON-OFF switching performance is largely improved by near-field radiation. The heat flux in the ON state can be enhanced by a factor of 28.5 ( T 1  = 400 K, T 3  = 3 K). [ABSTRACT FROM AUTHOR]

Accurate prediction of aerodynamic and thermal environment around a gap has a significant effect on the development of spacecraft. The implicit finite volume schemes are derived and programmed from Navier-Stokes equations. Taking the gap between thermal insulation tiles as an example, a numerical simulation is performed by the finite volume method to obtain the flow characteristic in a gap and then to analyze the heat transfer mechanism. The numerical results are consistent with the experimental ones, which prove the precision of the method used in this paper. Furthermore, the numerical results reveal that the heat convection plays a leading role in heat transfer around a gap. [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]

Among the mitigation measures introduced to cope with the space debris issue there is the de-orbiting of decommissioned satellites. Guidelines for re-entering objects call for a ground casualty risk no higher than 10 −4 . To comply with this requirement, satellites can be designed through a design-for-demise philosophy. Still, a spacecraft designed to demise through the atmosphere has to survive the debris-populated space environment for many years. The demisability and the survivability of a satellite can both be influenced by a set of common design choices such as the material selection, the geometry definition, and the position of the components inside the spacecraft. Within this context, two models have been developed to analyse the demise and the survivability of satellites. Given the competing nature of the demisability and the survivability requirements, a multi-objective optimisation framework was developed, with the aim to identify trade-off solutions for the preliminary design of satellites. As the problem is nonlinear and involves the combination of continuous and discrete variables, classical derivative based approaches are unsuited and a genetic algorithm was selected instead. The genetic algorithm uses the developed demisability and survivability criteria as the fitness functions of the multi-objective algorithm. The paper presents a test case, which considers the preliminary optimisation of tanks in terms of material, geometry, location, and number of tanks for a representative Earth observation mission. The configuration of the external structure of the spacecraft is fixed. Tanks were selected because they are sensitive to both design requirements: they represent critical components in the demise process and impact damage can cause the loss of the mission because of leaking and ruptures. The results present the possible trade off solutions, constituting the Pareto front obtained from the multi-objective optimisation. [ABSTRACT FROM AUTHOR]

This paper examines the velocity-free attitude stabilization of an underactuated nadir-pointing rigid spacecraft in a circular orbit in the presence of time-varying disturbances. A nonlinear observer is used to estimate the angular velocity of the spacecraft to complete an output feedback control law. A sliding-mode control strategy is proposed for this stabilization problem in conjunction with a nonlinear observer. The local ultimate boundedness of the closed-loop system is rigorously proven. Numerical examples demonstrate the performance of the controller in the presence of time-varying disturbances, nonzero orbital eccentricity, inertia matrix uncertainties, and measurement noise in a sampled data implementation. In particular, even though the theoretical ultimate boundedness guarantees are local, the numerical examples demonstrate ultimate boundedness with large initial attitude errors. [ABSTRACT FROM AUTHOR]

Copyright of Computer Measurement & Control is the property of Magazine Agency of Computer Measurement & Control and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Acquiring good adhesion of ITO thin films on polymer substrate is a major concern, especially for space related applications. Delamination of ITO coating on these polymers can seriously damage the spacecraft. This paper presents the development of highly transparent and conducting ITO thin films on as-received and surface treated fluorinated ethylene propylene (FEP) and Kapton ® substrates by reactive direct current magnetron sputtering. Stability of the ITO coating on FEP and Kapton ® substrates was studied in simulated space environments. Environmental tests such as: relative humidity, thermal cycling and thermo vacuum were performed. Thermo-optical properties and sheet resistance of ITO coated FEP and Kapton ® substrates were studied before and after environmental tests. Optimized ITO coating with thickness of ~ 15 nm on FEP and Kapton ® substrates showed sheet resistance in the range of 2–4 kΩ/sq. with high average transmittance and high IR emittance. Adhesion of ITO coating on FEP substrate was improved by Ar plasma etching. X-ray photoelectron spectroscopy and field emission scanning electron microscopic studies of etched FEP substrate showed defluorination and high roughness of the etched surface which helped for better adhesion of ITO coating. We demonstrated that ITO coated plasma etched FEP substrate showed no change in the sheet resistance and thermo-optical properties. Moreover, ITO coated etched FEP substrate showed good environmental stability than ITO coated untreated FEP substrates. [ABSTRACT FROM AUTHOR]

Lunar IceCube is a 6U CubeSat that is designed to detect and observe lunar volatiles from a highly inclined orbit. This spacecraft, equipped with a low-thrust engine, is expected to be deployed from the upcoming Exploration Mission-1 vehicle. However, significant uncertainty in the deployment conditions for secondary payloads impacts both the availability and geometry of transfers that deliver the spacecraft to the lunar vicinity. A framework that leverages dynamical systems techniques is applied to a recently updated set of deployment conditions and spacecraft parameter values for the Lunar IceCube mission, demonstrating the capability for rapid trajectory design. [ABSTRACT FROM AUTHOR]

This paper presents a distributed adaptive control framework for multiple spacecraft formation flying around Lagrange point orbits, which account for unmeasurable velocities and (spacecraft) mass uncertainties. The nominal trajectory for the formation system is a halo orbit parameterized by Fourier series expansions. Such an explicit, albeit approximate, description of the nominal trajectory facilitates each spacecraft in formation to include the relative state information into a cooperative feedback control system design, so that the relative motion can be driven towards a desired trajectory while maintaining a group synchronization during the maneuver. The developed distributed control strategies rely on the protocols formulated on an undirected topology with mutual information interactions, utilizing every available neighbor-to-neighbor communication data couplings, in order to improve the reliability of the formation. Numerical simulations show that the proposed adaptive control laws guarantee global asymptotic convergence and system robustness. [ABSTRACT FROM AUTHOR]

Copyright of Computer Measurement & Control is the property of Magazine Agency of Computer Measurement & Control and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Tethered space-tug system is regarded as one of the most promising active debris removal technologies to effectively decrease the steep increasing population of space debris. In order to suppress the spin of space debris, single-tethered space-tug system is employed by regulating the tether. Unfortunately, this system is underactuated as tether length is the only input, and there are two control objectives: the spinning debris and the vibration of tether. Thus, it may suffer great oscillations and result in failure in space debris removal. This paper presents the study of attitude stabilization of the single-tethered space-tug system using not only tether length but also the offset of tether attachment point to suppress the spin of debris, so as to accomplish the space debris removal mission. Firstly, a precise 3D mathematical model in which the debris and tug are both treated as rigid bodies is developed to study the dynamical evolution of the tethered space-tug system. The relative motion equation of the system is described using Lagrange method. Secondly, the dynamic characteristic of the system is analyzed and an offset control law is designed to stabilize the spin of debris by exploiting the variation of tether offset and the regulation of tether length. Besides, an estimation formula is proposed to evaluate the capability of tether for suppressing spinning debris. Finally, the effectiveness of attitude stabilization by the utilization of the proposed scheme is demonstrated via numerical case studies. [ABSTRACT FROM AUTHOR]

Purpose This paper aims to address the problem of formation control for spacecraft formation in elliptic orbits by using local relative measurements.Design/methodology/approach A decentralized formation control law is proposed to solve the aforementioned problem. The control law for each spacecraft uses only its relative state with respect to the neighboring spacecraft it can sense. These relative states can be acquired by local relative measurements. The formation control problem is converted to n stabilization problems of a single spacecraft by using algebraic graph theories. The resulting relative motion model is described by a linear time-varying system with uncertain parameters. An optimal guaranteed cost control scheme is subsequently used to obtain the desired control performance.Findings Numerical simulations show the effectiveness of the proposed formation control law.Practical implications The proposed control law can be considered as an alternative to global positioning system-based relative navigation and control system for formation flying missions.Originality/value The proposed decentralized formation control architecture needs only local relative measurements. Fuel consumption is considered by using an optimal guaranteed cost control scheme. [ABSTRACT FROM AUTHOR]

The landers of the Soviet Venera series--from Venera-9 to Venera-14--designed at the Lavochkin Association are a man-made monument to spectacular achievements of Soviet space research. For more than 40 years, they have remained the uneclipsed Soviet results in space studies of the Solar System. Within the last almost half a century, the experiments carried out by the Venera-9 to Venera-14 probes for studying the surface of the planet have not been repeated by any space agency in the world, mainly due to quite substantial technical problems. Since that time, no Russian missions with landers have been sent to Venus either. On Venus, there is an anoxic carbon dioxide atmosphere, where the pressure is 9.2 MPa and the temperature is 735 K near the surface. A long-lived lander should experience these conditions for an appreciable length of time. What technical solutions could provide a longer operation time for a new probe investigating the surface of Venus, if its thermal scheme is constructed similar to that of the Venera series? Onboard new landers, there should be a sealed module, where the physical conditions required for operating scientific instruments are maintained for a long period. At the same time, new high-temperature electronic equipment that remains functional under the above-mentioned conditions have appeared. In this paper, we consider and discuss different variants of the system for a long-lived sealed lander, in particular, the absorption of the penetrating heat due to water evaporation and the thermal protection construction for the instruments with intermediate characteristics. [ABSTRACT FROM AUTHOR]

In this paper we developed a multi-disciplinary, multi-objective optimization procedure for the shape generation of re-entry spacecrafts performing conventional landing from a Low Earth Orbit return mission. A special free-form parametric model, able to define complex vehicle shapes with no explicit support surfaces, was defined and used for this purpose. Model capabilities have been preliminary validated by emulating the HOPE-X vehicle prototype and computing the aerodynamic coefficients at Mach number 2 and 10. Multi-objective optimization has been performed by considering a multidisciplinary approach comprising aerodynamic analysis, trajectory estimation, and heating analysis starting by fixed waypoints along the descent path. A Pareto front based on mass and cross range objective functions was generated, highlighting the existence of several design scenarios: minimum mass, maximum winglet, maximum cross range. The existing trade-offs between the objective functions were related mainly to bank angle values and vehicle length, featuring main design trends. [ABSTRACT FROM AUTHOR]

Spacecraft relative motion planning is concerned with the design and execution of maneuvers relative to a nominal target. These types of maneuvers are frequently used in missions such as rendezvous and docking, satellite inspection, and formation flight, where exclusion zones representing spacecraft or other obstacles must be avoided. The presence of these exclusion zones leads to nonlinear and nonconvex constraints that must be satisfied. In this paper, a novel approach to spacecraft relative motion planning with obstacle avoidance and thrust constraints is developed. This approach is based on a graph search applied to a virtual net of closed (periodic) natural motion trajectories, where the natural motion trajectories represent virtual net nodes (vertices), and adjacency and connection information is determined by conditions defined in terms of safe, positively invariant lubes built around each trajectory. These conditions guarantee that transitions from one natural motion trajectory to another natural motion trajectory can be completed without constraint violations. The proposed approach improves the flexibility of a previous approach based on the use of forced equilibria and has other advantages in terms of reduced fuel consumption and passive safety. The resulting maneuvers, if planned onboard, can be executed directly or, if planned off-board, can be used to warm start trajectory optimizers to generate further improvements. [ABSTRACT FROM AUTHOR]

One of the primary design considerations of earth-orbiting spacecraft is the mitigation of the damage that might occur from an on-orbit MMOD impact. Traditional damage-resistant design consists of a ‘bumper’ that is placed a small distance away from a spacecraft component or from the wall of the element in which it is housed. The performance of such a multi-wall structural element is typically characterized by its ballistic limit equation (BLE), which defines the threshold particle size that results in a failure of the spacecraft element. BLEs are also key components of any micro-meteoroid/orbital debris (MMOD) risk assessment calculations. However, these assessments often call for BLEs to predict impact response for projectiles made of materials not used in the development of those BLEs. The question naturally arises regarding how close are the predictions of such BLEs when used in impact scenarios involving projectiles made of materials not necessarily considered in their development. In an effort to address this issue, a study was performed with the objective of assessing the validity of the NNO BLE for non-aluminum particles. Particle materials considered included steel, copper, and Al 2 O 3 (i.e. particles that are made of materials that are more dense than aluminum). Comparisons are made between actual test results involving these non-aluminum projectiles and the predictions of the NNO BLE. In nearly all cases, the NNO BLE was found not to work very well in the predicting failure / no failure response of these non-aluminum projectiles. A new NNO-type BLE is then developed that can be used to more reliably predict the response of dual-wall systems under the hypervelocity impact of such “heavier” non-aluminum projectiles. [ABSTRACT FROM AUTHOR]

Excitation design is one of the important contents in the identification of inertia parameters and the form of excitation has a great influence on the identification result. This paper presents a new method to design and calculate the optimal excitation. Firstly for a spacecraft equipped with momentum wheels, the identification problem is established based on conversation of angular momentum and an inverse operating. A performance index which is similar to but not the condition number is first defined as the benchmark for designing the optimal excitation. Because the performance index only depends on performances of the actuator, such as the angular momentum of the wheel, a simple direct-search method is applied to calculate the optimal excitation for the case without terminal angular velocity constraints and a two-step direct-search method for the case with terminal angular velocity constraints. While the initial angular momentum of spacecraft system is considered, the optimal excitation is obtained based on the difference of two successive measurements. Finally, the optimal excitation for a spacecraft using thrusters is designed according to the same design process. Simulation results show that the calculated optimal excitation has the good performance index and can produce accurate identification results even when some perturbations are considered. [ABSTRACT FROM AUTHOR]