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Analysis has been carried out and a mathematical model has been developed that describes thermal state of a substrate made of a thin polymer film during deposition of nanostructured coatings on its surface using vacuum ion-plasma method. Some recommendations are given on choosing optimum parameters of the technological process to stabilize temperature during deposition of loose plies of the coating. Technology of obtaining nanostructured resistive layers of coatings on thin polymer films based on experimental data and mathematical modelling results has been developed. This technology is used to maintain and control thermal conditions in onboard spacecraft systems. [ABSTRACT FROM AUTHOR]

The research for global climate changes calls for high quality satellite data and imageries regarding the Polar Regions. In recent years, the emerging Earth-Observation micro/nano satellite technology provides new data sources for polar region observations. The STU-2A, also named TW-1A, is such a nano satellite designed for polar region observation activities. It is a 3U CubeSat of 2.9 kg with a size of 30 × 10 × 10 cm carrying an Earth observation camera, launched into a Sun Synchronous Orbit (SSO) at 481 km with an inclination of 97.3°, on September 25, 2015. During the Antarctic summer of 2015/16, it has acquired visible-light true color images with a resolution of 94 m, covering different sea and coastal regions including Amundsen Sea, Ross Sea and Vincennes Bay. These images were used to analyze the change of glacier and sea ice, compared and calibrated with reference to the publically available MODIS images with a resolution of 250 m. As the camera was specially designed for the Polar regions which have an environment of low solar elevation angle and high surface reflectance, it eliminates the oversaturation problem of the MODIS sensors and can provide high quality images. Based on data analysis and assessment, it is confirmed that this satellite data can meet the demand of glacier and sea ice observation. This paper presents the Cubesat system design and configuration, the payload camera design, and its application in Antarctic glacier and sea ice observation. [ABSTRACT FROM AUTHOR]

The extreme environmental challenges of deep space exploration force unique solutions to small satellite design in order to enable their use as scientifically viable spacecraft. The challenges of implementing small satellites within limited resources can be daunting when faced with radiation effects on delicate electronics that require shielding or unique adaptations for protection, or mass, power and volume limitations due to constraints placed by the carrier spacecraft, or even Planetary Protection compliant design techniques that drive assembly and testing. This paper will explore two concept studies where the environmental constraints and/or planetary protection mitigations drove the design of the Flight System. The paper will describe the key technical drivers on the Sylph mission concept to explore a plume at Europa as a secondary free-flyer as a part of the planned Europa Mission. Sylph is a radiation-hardened smallsat concept that would utilize terrain relative navigation to fly at low altitudes through a plume, if found, and relay the mass spectra data back through the flyby spacecraft during its 24-h mission. The second topic to be discussed will be the mission design constraints of the Near Earth Asteroid (NEA) Scout concept. NEAScout is a 6U cubesat that would utilize an 86 sq. m solar sail as propulsion to execute a flyby with a near-Earth asteroid and help retire Strategic Knowledge Gaps for future human exploration. NEAScout would cruise for 24 months to reach and characterize one Near-Earth asteroid that is representative of Human Exploration targets and telemeter that data directly back to Earth at the end of its roughly 2.5 year mission. [ABSTRACT FROM AUTHOR]

This paper presents a novel non-certainty-equivalence adaptive control method for the attitude tracking control problem of spacecraft with inertia uncertainties. The proposed immersion and invariance (I&I) based adaptation law provides a more direct and flexible approach to circumvent the limitations of the basic I&I method without employing any filter signal. By virtue of the adaptation high-gain equivalence property derived from the proposed adaptive method, the closed-loop adaptive system with a low adaptation gain could recover the high adaptation gain performance of the filter-based I&I method, and the resulting control torque demands during the initial transient has been significantly reduced. A special feature of this method is that the convergence of the parameter estimation error has been observably improved by utilizing an adaptation gain matrix instead of a single adaptation gain value. Numerical simulations are presented to highlight the various benefits of the proposed method compared with the certainty-equivalence-based control method and filter-based I&I control schemes. [ABSTRACT FROM AUTHOR]

In order to demonstrate the influence of gender differences on occupant dynamic response during spacecraft high-level landing impact, this study established a seat-dummy system model using Pro-E and HyperWorks software. The 50th Hybrid III male dummy and the 5th Hybrid III female dummy were used in the model. The seat-dummy system model was calibrated and validated according to the actual impact condition and impact test data of the simulated spacecraft landing. The gender differences on the impact response of key body segments were demonstrated and analysed under high-level landing impact (at a peak value of 26g). The simulation results show that the peak acceleration value of the female is larger than the male by 43.7% on the shoulder and by about 33% on the chest and pelvis, while the female is smaller than the male by 40.0% on the head in the chest-back (anterior–posterior, Gx) direction. In the head-foot (superior–inferior, Gz) direction, the peak acceleration value of the female is larger than the male by about 45.2% on the head, 120% on the shoulder, 9.0% on the chest and 37.3% on the pelvis. Therefore, it is recommended that spacecraft designers should pay more attention to gender differences on the head-neck and pelvis, and provide better protection for females during landing impact. [ABSTRACT FROM PUBLISHER]

In this paper, a novel approach is proposed for the mission planning of on-orbit servicing such as visual inspection, active debris removal and refueling through multiple servicing satellites (SSs). The scheduling has been done with the aim of minimization of fuel consumption and mission duration. So a multi-objective optimization problem is dealt with here which is solved by employing particle swarm optimization algorithm. Also, Taguchi technique is employed for robust design of effective parameters of optimization problem. The day that the SSs have to leave parking orbit, transfer duration from parking orbit to final orbit, transfer duration between one target to another, and time spent for the SS on each target are the decision parameters which are obtained from the optimization problem. The raised idea is that in addition to the aforementioned decision parameters, eccentricity and inclination related to the initial orbit and also phase difference between the SSs on the initial orbit are identified by means of optimization problem, so that the designer has not much role on determining them. Furthermore, it is considered that the SS and the target rendezvous at the servicing point and the SS does not perform any phasing maneuver to reach the target. It should be noted that Lambert theorem is used for determination of the transfer orbit. The results show that the proposed approach reduces the fuel consumption and the mission duration significantly in comparison with the conventional approaches. [ABSTRACT FROM AUTHOR]

We present a geometrical methodology for analyzing the formation flying of electric solar wind sail based spacecraft that operate in heliocentric, elliptic, displaced orbits. The spacecraft orbit is maintained by adjusting its propulsive acceleration modulus, whose value is estimated using a thrust model that takes into account a variation of the propulsive performance with the sail attitude. The properties of the relative motion of the spacecraft are studied in detail and a geometrical solution is obtained in terms of relative displaced orbital elements, assumed to be small quantities. In particular, for the small eccentricity case (i.e. for a near-circular displaced orbit), the bounds characterized by the extreme values of relative distances are analytically calculated, thus providing an useful mathematical tool for preliminary design of the spacecraft formation structure. [ABSTRACT FROM AUTHOR]

The design of National Aeronautics and Space Administration Orion's prelaunch navigation system is introduced, both for the first flight test, Exploration Flight Test 1, and for the first planned Orion mission, Exploration Mission 1. A detailed tradeoff of possible design decisions is discussed, and the choices made by Orion are presented. The actual performance of the navigation system during Exploration Flight Test 1 is presented together with the navigation flight-software data provided by Orion to the ground controllers in telemetry. [ABSTRACT FROM AUTHOR]

Spacecraft must withstand rigorous mechanical environment experiences such as acceleration, noise, vibration, and shock during the process of launching, satellite-vehicle separation, and so on. In this paper, a new spacecraft multifunctional structure concept designed by us is introduced. The multifunctional structure has the functions of not only load bearing, but also vibration reduction, energy source, thermal control, and so on, and we adopt a series of viscoelastic parts as connections between substructures. Especially in this paper, a vibration antiresonance design method is proposed to realize the vibration reduction. The complex zero-point equations of the vibration system are firstly established, and then the vibration antiresonance design for the system is achieved. For solving the difficulties due to viscoelastic characteristics of the connecting parts, we present the determining formulas to obtain the structural parameters, so that the complex zero-point equations can be satisfied. Numerical simulation and ground experiment demonstrate the correctness and effectiveness of the proposed method. This method can solve the structural vibration control problem under the function constraints of load bearing and energy supplying and will expand the performance of spacecraft functional modules. [ABSTRACT FROM AUTHOR]

The traditional method used for handling qualities assessment of manned space vehicles is too time-consuming to meet the requirements of an increasingly fast design process. In this study, a rendezvous and docking inverse simulation system to assess the handling qualities of spacecraft is proposed using a previously developed model-predictive-control architecture. By considering the fixed discrete force of the thrusters of the system, the inverse model is constructed using the least squares estimation method with a hyper-ellipsoidal restriction, the continuous control outputs of which are subsequently dispersed by pulse width modulation with sensitivity factors introduced. The inputs in every step are deemed constant parameters, and the method could be considered as a general method for solving nominal, redundant, and insufficient inverse problems. The rendezvous and docking inverse simulation is applied to a nine-degrees-of-freedom platform, and a novel handling qualities evaluation scheme is established according to the operation precision and astronauts' workload. Finally, different nominal trajectories are scored by the inverse simulation and an established evaluation scheme. The scores can offer theoretical guidance for astronaut training and more complex operation missions. [ABSTRACT FROM AUTHOR]

The Small Mars System is a proposed mission to Mars. Funded by the European Space Agency, the project has successfully completed Phase 0. The contractor is ALI S.c.a.r.l., and the study team includes the University of Naples “Federico II”, the Astronomical Observatory of Capodimonte and the Space Studies Institute of Catalonia. The objectives of the mission are both technological and scientific, and will be achieved by delivering a small Mars lander carrying a dust particle analyser and an aerial drone. The former shall perform in situ measurements of the size distribution and abundance of dust particles suspended in the Martian atmosphere, whereas the latter shall demonstrate low-altitude flight in the rarefied planetary environment. The mission-enabling technology is an innovative umbrella-like heat shield, known as IRENE, developed and patented by ALI. The mission is also a technological demonstration of the shield in the upper atmosphere of Mars. The core characteristics of SMS are the low cost (120 M€) and the small size (320 kg of wet mass at launch, 110 kg at landing), features which stand out with respect to previous Mars landers. To comply with them is extremely challenging at all levels, and sets strict requirements on the choice of the materials, the sizing of payloads and subsystems, their arrangement inside the spacecraft and the launcher's selection. In this contribution, the mission and system concept and design are illustrated and discussed. Special emphasis is given to the innovative features and to the challenges faced in the development of the work. [ABSTRACT FROM AUTHOR]

The purpose of this work is to develop a method for solving inverse transient-state thermal and strength non-linear problems in complex shapes. Non-linearity is caused both by the material temperature-dependent properties and radiation. The proposed algorithm reconstructs the whole transient temperature and thermal stress distribution based on temperatures measured in the element selected points. Measured transient temperature values are generated during a numerical simulation of aerodynamic heating on the atmospheric reentry capsule. Both constant and temperature-dependent properties of the material are assumed. A comparison is presented between the transient temperature distributions obtained based on the material constant and temperature-dependent properties. Finally, the developed method is used to identify the transient temperature and stress distribution in the atmospheric reentry capsule assuming temperature-dependent properties of the material. The proposed approach is expected to be a good solution for improving spacecraft structures. [ABSTRACT FROM AUTHOR]

This paper investigates the collision-free distributed coordination control of six-DOF spacecraft formation flying. Dual quaternions are used to develop the kinematic and dynamic models of relative translational and rotational motion between spacecraft. By assuming that the reference relative orbit and attitude information is only available to a subset of members in formation, an observer is introduced to estimate the reference information in finite time. With the estimates provided by the observer, a sliding mode controller is then designed for the coordination of six-DOF formation flying. Meanwhile, the artificial potential function method is employed to design evasive maneuvers in case of any collisions during orbital maneuvering. The overall stability of the closed-loop system is guaranteed by a Lyapunov-based method. The feasibility of the proposed control scheme is then demonstrated by a typical formation reconfiguration mission in a low Earth orbit environment. [ABSTRACT FROM AUTHOR]

The formation of high-velocity compact elements of shaped charges with a liner of a combined hemisphere–cylinder shape has been analyzed by numerical simulations of a two-dimensional axisymmetric problem of continuum mechanics. This liner jet generator contains a part of a hemisphere, a truncated sphere or a slightly prolate ellipsoid and a cut-off part in the form of a cylinder. The abstract massive speed characteristics of the formed compact elements depend on changes in the geometric parameters of the combined cumulative liner. Variants combined the cumulative liner with an explosive device to ensure the formation of elements of a gradientless weight from 5 g to 15 g at velocities of 7.5–10 km/s. This simple explosive device can be used to simulate the conditions in the Earth's single and group impact of micrometeorites and space debris objects in rocket and space technology. [ABSTRACT FROM AUTHOR]

This paper investigates two robust finite-time controllers for the attitude control of spacecraft based on rotation matrix. The first controller can compensate external disturbances with known bounds, whereas the second one can deal with both external disturbances and input saturation by using the hyperbolic tangent function and auxiliary system. Both controllers can avoid the singularity and converge to zero in finite time by using novel fast nonsingular terminal sliding mode control. Because the controllers are designed based on a rotation matrix that represents the set of attitudes both globally and uniquely, the system can overcome the drawbacks of unwinding. Numerical simulations are presented to demonstrate the effectiveness of the proposed control schemes. [ABSTRACT FROM AUTHOR]

Force measurements and wake surveys have been conducted on two swept NACA 0021 wings. One wing had a smooth leading edge, while the other wing had a tubercled leading edge. The force measurements and the wake survey results were in good agreement. Between 0 and 8° angles of attack, tubercles reduced the lift coefficient by 4-6%. For the same range of angles of attack, tubercles reduced the drag coefficient by 7-9.5%. Tubercles increased the lift-to-drag ratio of this wing by 2-6%, and increased the maximum lift-to-drag ratio by 3%. At angles of attack higher than 8°, tubercles typically decreased the lift coefficient and the lift-to-drag ratio, while substantially increasing the drag coefficient. The wake surveys revealed that tubercles reduced the drag coefficient near the wingtip and that they also spatially modulated the drag coefficient into local maxima and minima in the spanwise direction. Typically, tubercles reduced the drag coefficient over the peaks where the tubercle vortices produced downwash. Conversely, tubercles increased the drag coefficient over the troughs, where upwash occurred. The majority of the drag coefficient reduction occurred over the tubercled wingspan. [ABSTRACT FROM AUTHOR]

In this paper, a spacecraft radiator formed in a honeycomb structure is designed to enhance the thermal performance while reducing its mass. Examples of design guidelines for radiator configurations, such as the distance between heat pipes, facesheet thickness, and honeycomb core density, are suggested. To derive the analytic solution of the governing equation, a linear approximation is used and the accuracies of the solutions are verified with a fourth-order finite-difference method. There exist optimal combinations of design parameters that minimize the radiator mass while maintaining its heat rejection capacity. The heat rejection rate that minimizes the mass per unit heat rejection and the pertinent radiator shape also is presented. The combinations of optimal design are different among the three surface treatments and their characteristics are investigated. [ABSTRACT FROM AUTHOR]

In this paper, a novel control strategy is developed based on artificial potential field for the on-orbit autonomous assembly of four flexible spacecraft without inter-member collision. Each flexible spacecraft is simplified as a hub-beam model with truncated beam modes in the floating frame of reference and the communication graph among the four spacecraft is assumed to be a ring topology. The four spacecraft are driven to a pre-assembly configuration first and then to the assembly configuration. In order to design the artificial potential field for the first step, each spacecraft is outlined by an ellipse and a virtual leader of circle is introduced. The potential field mainly depends on the attitude error between the flexible spacecraft and its neighbor, the radial Euclidian distance between the ellipse and the circle and the classical Euclidian distance between the centers of the ellipse and the circle. It can be demonstrated that there are no local minima for the potential function and the global minimum is zero. If the function is equal to zero, the solution is not a certain state, but a set. All the states in the set are corresponding to the desired configurations. The Lyapunov analysis guarantees that the four spacecraft asymptotically converge to the target configuration. Moreover, the other potential field is also included to avoid the inter-member collision. In the control design of the second step, only small modification is made for the controller in the first step. Finally, the successful application of the proposed control law to the assembly mission is verified by two case studies. [ABSTRACT FROM AUTHOR]

Hollow cathodes are electron sources used for the gas ionization and the beam neutralization in both ion and Hall effect thrusters (HETs). A reduction of power and propellant consumption from the cathode is particularly needed in small satellite applications, where power and mass budgets are inherently limited. Concurrently, the interest in high-power HETs is increasingly fostered for a number of space applications, including final positioning and station-keeping of Geostationary Earth Orbit (GEO) satellites, spacecraft transfers from Low Earth Orbit (LEO) to GEO, and deep-space exploration missions. As such, several hollow cathodes have been developed and tested at Sitael, each conceived for a specific power class of thrusters. A numerical model was used during the cathode design to define the geometry, in accordance with the thruster unit specifications in terms of discharge current, mass flow rate, and lifetime. Lanthanum hexaboride (LaB6) hollow cathodes were successfully developed for HETs with discharge power ranging from 100 W to 20 kW. Experimental campaigns were carried out in both stand-alone and coupled configurations, to verify the operation of the cathodes and validate the numerical model. The comparison between experimental and theoretical results are presented, offering a sound framework to drive the design of future hollow cathodes. [ABSTRACT FROM AUTHOR]

In the design process, when placing the on-board equipment of a spacecraft in the compartments of a minimum volume, along with the design and technological constraints the possible impact of the space environment must be taken into account. These are classified in the literature as a means of impact energy directed at the spacecraft transmission during its intended operation. As such, damaging factors most often considered are a means of mechanical action of meteoric and man-made particles. The energy released in the surface layer structure from the effects of other power transmission means or penetrating its material (different types of beam funds), as well as the energy of the electromagnetic pulse can disrupt the normal functioning of the onboard spacecraft equipment Protection of aircraft equipment by means of energy transfer napavlennoy determines the reliability of the spacecraft functioning and efficiency of the performance of the flight plan. In this paper the mathematical formulation and method of solving the problem of finding dense layouts spacecraft through proper allocation of the onboard equipment inside compartments for the effect on the spacecraft factors directed energy transfer. [ABSTRACT FROM AUTHOR]

The problem of near Earth space debris is studied in this article. A special spacecraft debris collector, equipped with electro rocket engine of low-thrust, for large space debris disposal is introduced. The mass model of spacecraft debris collector, one-off or reusable, is obtained. Ballistic scheme, spacecraft debris collector orbital transfer from parking orbit to space debris orbit, its disposal in Earth atmosphere and reusable spacecraft return to the parking orbit are studied. For the introduced criteria of transport transfer efficiency and assumption about consistency of acceleration from thrust, analytical relations for spacecraft design parameters calculation are obtained. The design parameters calculation results are shown in a generalized form. [ABSTRACT FROM AUTHOR]

This article deals with the problem of developing the statistical model of spacecraft star trackers mass. Cluster analysis is involved in order to obtain adequate results. Some up-to-date star trackers types are examined and the regression model for a star tracker mass dependent on their main parameters is put forward. The model gives an opportunity to estimate star tracker mass at the initial stages of spacecraft designing. [ABSTRACT FROM AUTHOR]

A Multilevel approach of modeling the stiffness and strength of ultra-light aerospace structures from short reinforced composite materials is presented. The object of the research is the strength elements of aerospace structures and specifically, the lugs for the transfer of concentrated forces in the gateway of unit places. The material is anisotropic. The properties of this material depend on the organization of the casting process of the plates from which the lugs are cut. The first level of the model is the descriptive model of the casting process of the plate from PEEK material. The casting model of plates is based on the geometry of the gating system and the tool geometry which defines the characteristics of the high-viscosity heat exchange, reinforced during binding with the environment. The preprocessor for the first level of model production is Moldex Designer, with which the geometrical characteristics of the gating system of heat exchange are set. The finite-element mesh for calculation of the hydrodynamic task is constructed. Calculation of casting process of the gate is carried out in Moldex 3D system. The initial data are the finite-element model, the parameters defining the casting mode: temperature, pressure, material volume, and also the parameters of modes of hold pressure and cooling of detail. The orientation of the file of fibers, which is used for the description of the anisotropy of the products considered is the result of casting modeling of the plate. The second level of the model is the model of anisotropic material with the characteristics defined taking the orientation of reinforcing fibers into account, obtained from the results of the casting process in the DIGIMAT system. The possibility of detailing the material characteristics received on the basis of the processing of strength tests of material samples is considered. The third level of model is the finite-element model of the product considering anisotropy of material. The model is constructed in the ANSYS Workbench system. The strength characteristics of the anisotropic material are defined in the model by the DIGIMAT module connected to ANSYS through the material parameter setting DIGIMAT Material. The multilevel model allows calculation of the strain-stress state of products of irregular shape cast from composite materials reinforced by short high-strength fibers. The results of the multilevel model production are verified with field research of the considered products. [ABSTRACT FROM AUTHOR]

The problem of creating ultra-light composite structures is considered. An overview of the main modern methods of manufacturing composite materials and methods of increasing the weight efficiency of constructions in view of the structure destination, its force work and operating conditions are given. The main design and technological solutions to reduce the weight of a regular construction zone, connection zone and transition zone are offered. The importance of taking into account the anisotropy of mechanical and thermo-elastic properties of composite materials for structures with special requirements for dimensional stability is noted. Guidelines for choice of materials components and manufacturing technology depending on the design application, operating conditions and the nature of load distribution are given. Examples create ultralight structures and their properties are given. [ABSTRACT FROM AUTHOR]

This paper discusses the application of computer-assisted design systems to development of low thrust spacecraft intended to take an exploratory flight to potentially hazardous asteroids. The design model of the low thrust spacecraft that is created with computer-aided design systems is described. The mathematical motion model within the heliocentric system of coordinates for such type of spacecraft is considered and used for a motion simulation session. The simulation of flight to the potentially hazardous asteroid is performed with the aid of special complex software that is developed for the purpose of the work. The results of the paper consist of a detailed three-dimensional model of the low thrust spacecraft, heliocentric trajectory of the spacecraft, the values of flight duration and propellant consumption during exploratory flight. [ABSTRACT FROM AUTHOR]

The problem of electro rocket engine type and its main design parameters optimization for spacecraft located at low circular Earth orbit is studied. The efficiency of chemical (impulse) and electro rocket engine of low-thrust is analyzed. Introduced criteria of efficiency – payload mass maximization, introduced assumption that engine exhaust velocity is constant. In accordance with criteria and assumption, the analytical solution for spacecraft design parameters of spacecraft is obtained. Also, analytical solution for area of efficiency of electro rocket low-thrust engine is obtained. Calculation of spacecraft design parameters equipped with electro rocket low-thrust engine and area of its efficiency applying is carried out. [ABSTRACT FROM AUTHOR]

The solar wind is a highly turbulent medium, with a high level of field fluctuations throughout a broad range of scales. These include an inertial range where a turbulent cascade is assumed to be active. The solar wind cascade shows intermittency, which however may depend on the wind conditions. Recent observations have shown that ion-scale magnetic turbulence is almost self-similar, rather than intermittent. A similar result was observed for the high resolution measurements of proton density provided by the spacecraft Spektr-R. Intermittency may be interpreted as the result of the multifractal properties of the turbulent cascade. In this perspective, this paper is devoted to the description of the multifractal properties of the high resolution density measurements. In particular, we have used the standard coarse-graining technique to evaluate the generalized dimensions D q , and from these the multifractal spectrum f ( α ) , in two ranges of scale. A fit with the p-model for intermittency provided a quantitative measure of multifractality. Such indicator was then compared with alternative measures: the width of the multifractal spectrum, the peak of the kurtosis, and its scaling exponent. The results indicate that the small-scale fluctuations are multifractal, and suggest that different measures of intermittency are required to fully understand the small scale cascade. [ABSTRACT FROM AUTHOR]

Coronal radio-sounding experiments were carried out using two-way coherent dual-frequency carrier signals of the ESA spacecraft ROSETTA in 2010 and MARS EXPRESS in 2010/2011. Differential frequency measurements recorded at both NASA and ESA tracking stations (sample rate: 1 Hz) are analyzed in this paper. Spectral analysis of the S-band, X-band, and differential frequency records has shown that the r.m.s. frequency fluctuation of each signal can be described by a radial power-law function of the form σ i = A i ( R / R ⊙ ) − βi , where i = s , x , sx . The ratio of the coefficients A s and A x differs from the expected theoretical value A s / A x = f s / f x . This occurs because the X-band fluctuations underlie two-way propagation conditions while the S-band fluctuations are essentially the product of a one-way propagation experiment. The intensity of the frequency fluctuations decreases sharply at high heliolatitudes. The asymmetry of the frequency fluctuation intensity between ingress and egress is exploited to determine the solar wind speed at small heliocentric distances. [ABSTRACT FROM AUTHOR]

Present study reports for the first time laser cutting of multilayered coatings on both side of ultra thin (i.e., 75 µm) glass substrate based rigid optical solar reflector (OSR) for spacecraft thermal control application. The optimization of cutting parameters was carried out as a function of laser power, cutting speed and number of cutting passes and their effect on cutting edge quality. Systematic and in-detail microstructural characterizations were carried out by optical and scanning electron microscopy techniques to study the laser affected zone and cutting edge quality. Sheet resistance and water contact angle experiments were also conducted locally both prior and after laser cut to investigate the changes of electrical and surface properties, if any. [ABSTRACT FROM AUTHOR]

The paper proposes a concept of a microrocket system meant for the injection of nano-, pico-, and femtoscale satellites into near-Earth orbit. Requirements for the appearance and basic design parameters of the micro-rocket system are substantiated. Characteristics of possible prototypes and analogues of this system are analyzed. [ABSTRACT FROM AUTHOR]

This article presents some types of planetary descent vehicles, the shape of which varies in different flight phases. The advantages of such vehicles over those with unchangeable form (from launch to landing) are discussed. It is shown that the use of transformable descent vehicles widens the scope of possible tasks to solve. [ABSTRACT FROM AUTHOR]

The article analyzes problems of design and development of advanced superheavy launch vehicles. Mass and energy characteristics and design layout of launch vehicles are substantiated. Delivery methods of bulky superheavy launch vehicle components to the spacecraft launch site are discussed. Methods of reduction of financial and technical risks of development and operation of superheavy launch vehicles are analyzed. The problem of environmental impacts of superheavy launch vehicle launches is posed. [ABSTRACT FROM AUTHOR]

Aspects of the design of small spacecraft with electric propulsion power plants for investigating minor bodies in the Solar System are examined. The results of design and ballistic analysis of transfer into an orbit of terrestrial asteroids using electric propulsion thrusters are given. The possible concept design of the spacecraft is determined and the structure of a small spacecraft with an electric propulsion power plant is presented. Parameters of the electric propulsion power plant of a small spacecraft for a flight to the minor bodies of the Solar System are estimated. [ABSTRACT FROM AUTHOR]

The article focuses on the lessons learned by the American people and its leaders from the space program dubbed as Project Mercury. The twenty-two orbit flight of Major Leroy Gordon Cooper, Jr. has been one of the more expensive learning experiences of America's peacetime history. In the final review of the program, the Americans learned that a trained pilot can function in an orbiting spacecraft in null gravity about as well as he can in an airplane for at least thirty-four hours without lasting physical or mental effects. And also, American technology, geared to the mass production of consumer articles that break down quickly, was not good enough for rockets and spacecraft, in which breakdowns, even minor ones, are catastrophic.

The hybrid structures show excellent performance on vibration reduction for ship, aircraft and spacecraft designs. Meanwhile, the topology optimization is widely used for structure vibration reduction and weight control. The design of hybrid structures considering simultaneous materials selection and topology optimization are big challenges in theoretical study and engineering applications. In this paper, according to the proposed laminate component method (LCM) and solid isotropic microstructure with penalty (SIMP) method, the mathematical formulations are presented for concurrent materials selection and topology optimizations of hybrid structures. Thickness distributions of the plies in laminate components are defined as materials selection design variables by LCM method. Relative densities of elements in the components are defined as topology design variables by SIMP method. Design examples of hybrid 3-bar truss structures and hybrid floating raft with vibration reduction requirements verified the effectiveness of the presented optimization models. [ABSTRACT FROM AUTHOR]

The average design lifetime of satellites continues to increase, in part due to the expectation that the satellite cost per operational day decreases monotonically with increased design lifetime. In this work, we challenge this expectation by revisiting the durability choice problem for spacecraft in the face of reduced launch price and under various cost of durability models. We first provide a brief overview of the economic thought on durability and highlight its limitations as they pertain to our problem (e.g., the assumption of zero marginal cost of durability). We then investigate the merging influence of spacecraft cost of durability and launch price, and we identify conditions that give rise cost-optimal design lifetimes that are shorter than the longest lifetime technically achievable. For example, we find that high costs of durability favor short design lifetimes, and that under these conditions the optimal choice is relatively robust to reduction in launch prices. By contrast, lower costs of durability favor longer design lifetimes, and the optimal choice is highly sensitive to reduction in launch price. In both cases, reduction in launch prices translates into reduction of the optimal design lifetime. Our results identify a number of situations for which satellite operators would be better served by spacecraft with shorter design lifetimes. Beyond cost issues and repeat purchases, other implications of long design lifetime include the increased risk of technological slowdown given the lower frequency of purchases and technology refresh, and the increased risk for satellite operators that the spacecraft will be technologically obsolete before the end of its life (with the corollary of loss of value and competitive advantage). We conclude with the recommendation that, should pressure to extend spacecraft design lifetime continue, satellite manufacturers should explore opportunities to lease their spacecraft to operators, or to take a stake in the ownership of the asset on orbit. [ABSTRACT FROM AUTHOR]

The design of space systems is a complex and multidisciplinary process. In this study, two deterministic and nondeterministic approaches are applied to the system design optimization of a spacecraft which is actually a small satellite in low Earth orbit with remote sensing mission. These approaches were then evaluated and compared. Different disciplines such as mission analysis, payload, electrical power supply, mass model, and launch manifest were properly combined for further use. Furthermore, genetic algorithm and sequential quadratic programming were employed as the system-level and local-level optimizers. The main optimization objective of this study is to minimize the resolution of the satellite imaging payload while there are several constraints. A probabilistic analysis was performed to compare the results of the deterministic and nondeterministic approaches. Analysis of the results showed that the deterministic approaches may lead to an unreliable design (because of leaving little or no room for uncertainties), while using the reliability-based multidisciplinary design optimization architecture, all probabilistic constraints were satisfied. [ABSTRACT FROM AUTHOR]

According to the requirement of small satellite, this paper designed a digital sun sensor which diaphragm is a V-shaped cross-section structure. Using Position Sensitive Detector (PSD) as the light detector, we designed the V-shaped cross-section structure based on the pinhole imaging principle. The sun sensor realized the accurate calculation for two axis sun angle of the sun sensor. The mechanical test, thermal test and testing of the sun sensor are designed and carried out. The mechanical test and thermal test results verify the stability of the sun sensor. Testing result shows that the detection angle can reach (120°) (120°), and the attitude determination accuracy is better than 6" in the entire viewing field. The mass, volume and power consumption of the sun sensor is 0.177 kg, 78 mm×77 mm×21 mm and 0.25 W. The sun sensor has low power consumption, large viewing angle and high precision characteristics, which realized the sun sensor the miniaturization and meet the requirements of the micro satellite. Its performance has been verified in orbit. [ABSTRACT FROM AUTHOR]

The article presents a study related to spacecraft attitude stabilization with the use of magnetorquers with separation between measurement and actuation. Topics discussed include actuation mechanisms related to spacecraft attitude stabilization; ways to spacecraft attitude control; and use of electromagnetic actuators for generation of attitude control torques on satellites flying in low earth orbits.

The article presents a study related to mass ratio of electrodynamics tether to spacecraft on deorbit stability and efficiency. Topics discussed include ways in which space activities in low Earth orbit have produced a large population of dead bodies, or debris; guidelines to remove or mitigate the space debris; and parametrical analysis of deorbit performance.

The article presents a study related to spacecraft attitude stabilization with the use of magnetorquers with separation between measurement and actuation. Topics discussed include actuation mechanisms related to spacecraft attitude stabilization; ways to spacecraft attitude control; and use of electromagnetic actuators for generation of attitude control torques on satellites flying in low earth orbits.

The article presents a study related to adaptive attitude stabilization control design for spacecraft under physical limitations. Topics discussed include role of Attitude stabilization control is an important orbital mission for spacecraft; the saturation problem for attitude control; and use of adaptive attitude tracking controller.

The satellite battery pack is one of the important and vital parts that are required conducting special design with multiple capabilities. Some parameters like the minimum weight, lack of outgassing of parts, avoiding from short-circuit in batteries inside battery pack, the possibility for making series connection in batteries, and prevention from destroying battery pack structure caused by vibrations due to displacement of missile upon launching of a satellite are included in some requirements for design and manufacture of satellite battery pack where in the present essay we will study on design and manufacture of battery pack structure by considering the above- mentioned requirements in mind. In order to determine the authenticity of design and manufacture of battery pack in real conditions for launching and vacuum conditions, several environmental test of thoroughness such as Thermal Vacuum Testing and Vibration Tests have been carried out on the manufactured structure so the results of batteries performance will be purposed within these tests. The results obtained from environmental test conducted on satellite battery pack structure suggest the fully successful and innovative achievement in design and manufacture of battery pack structure. [ABSTRACT FROM AUTHOR]