Spacecraft Structures Research Articles

Change of mechanical structures of spacecraft with variable quantity of degrees of freedom in purposes of reaction/momentum wheels unloading

The using of the released (“defreezed”) additional degrees of freedom is considered to control by the relative angular momentum of reaction/momentum wheels and to unload this relative angular momentum. The change of the relative angular momentum of wheels is fulfilled with the help of its translation to additional internal rotors and main body of spacecraft.

Improvement of spacecraft structure Dynamic Characteristics by Using Honeycomb Sandwich panels

Improvement of spacecraft structure Dynamic Characteristics by Using Honeycomb Sandwich panels

Experimental investigation of the thermoelastic performance of an aerospace aluminium honeycomb composite panel

Aluminium composite sandwich panels are widely used to enhance the design of structures subjected to dynamic mechanical loading in thermally harsh environments. Spacecraft structures fall into this category because typical environmental conditions include combined and variable mechanical and thermal loading. Usually mechanical loadings arise as a consequence of localised structural dynamics and the thermal loadings are attributable principally to the effects of solar irradiation and eclipse during the vehicle’s orbit. Together these have the potential to influence satellite de-point in particular. Therefore, building a combined physics model which is representative of the thermal and mechanical loadings has emerged as an interesting and useful aim, which can be thought of as defining an important thermoelastic deformation problem in this application. The performance of such a structure loaded in this way could obviously be considered in the context of separate thermodynamic and mechanical interpretations. However, multiphysics modelling is currently in hand based on the premise that the pseudo-static thermal loadings and the mechanical loadings encountered in various operating environments are not necessarily decoupled processes, and this will be the subject of a separate publication. The analytical modelling fully represents both static and dynamic mechanical and thermal loading conditions.It has become clear that predictive accuracy may be compromised by separation of the phenomena, at least without the introduction of a judicious correction factor. Therefore, in this paper an attempt has been made to identify experimentally the presence, and then to understand the attendant effects, of the coupling between the thermal and mechanical effects in an aluminium composite sandwich panel under test. The authors have performed a series of experiments on an aluminium honeycomb composite panel under three-point mechanical bending and controlled environmental temperature. The panel was subjected to a controllable, centrally located, very slowly increasing mechanical load in conjunction with thermal loading in the form of precisely controlled lowered and elevated environmental temperature. The tests were performed on a computer controlled Instron 8801 100 kN test machine for which the rate of change of applied mechanical load was automatically linked through feedback control to the rate of change of displacement. This ensured that the exact load–deflection profile can be obtained, even for materials with highly nonlinear characteristics.Both forms of loading have been shown to influence the displacement of the panel in significant ways, thereby confirming the importance of a combined physics approach.

An integrated system for modeling hypervelocity interaction of solids

The main purpose of the article is to show feasibility and efficiency of combining different types of simulating a hypervelocity impact interaction of solids within a single system. It allows effective application of new mathematical apparatus for analysis and comparison of experimental results. The paper considers the features of an integrated system that combines the results of experiments reproducing the high-velocity impact interaction of solids. The experiments included different types of modeling. System structure and the tasks solved within individual functional modulus were discussed. Based on the theory of fuzzy sets, new algorithms were proposed for comparing edges of the images obtained in the physical and computational experiment. Mathematical models were verified by such procedure. Differences in the shape and size of a cloud of particles that appears after penetrating a single barrier were estimated in the framework of designed models, as well as the differences in the sizes of punched holes using fuzzy variables and related membership functions. When analyzing experimental results, fuzzy sets enable switching from data structures to knowledge structures that are necessary to create an intelligent system for analyzing and predicting the effects of the interaction between space debris and spacecraft structures.

Radar-Absorbing Materials for Spacecraft

This article presents the results of experimental studies of properties that are resistant to abnormal high and abnormally low temperatures of samples of the radio-absorbing material VTMV-1C. Based on the studies, it was concluded that the radar-absorbing material can be used to attenuate reflected waves of electromagnetic fields from various parts of the spacecraft structure and to reduce the levels of resonance phenomena of the electromagnetic field in the modules and cases of onboard instruments and devices of spacecraft.

Analytical low-jerk reorientation maneuvers for multi-body spacecraft structures

A low-jerk attitude guidance method is developed, based on an analytical smoothing of a bang-off-bang maneuver. A set of closed-form equations are derived and used to plan constrained low-jerk maneuvers, with prescribed boundary conditions, inertia, time and maximum control torque. The guidance law is first developed for one-dimensional and then three-dimensional rotations using two different approaches: (i) by designing a rotation about the Euler axis and (ii) by using the inverse kinematics equations. A generic model of the torque induced by a multi-body appendage is derived using the lumped-parameter method. This method can also be used to approximate the dynamic behavior of flexible appendages. The simulations results show that the smoothing techniques reduce the excitation of multi-body and flexible structures during a slew maneuver.

Impact force identification in structures using time-domain spectral finite elements

During the launch and subsequent life of a spacecraft, there are various transient loads due to stage separation and micro-meteorite impact which a spacecraft structure must be designed to withstand. However, the force histories for these transients at these locations are not available to design the structure due to practical considerations like mounting of instrumentation. To address this problem, a new algorithm for performing impact force identification based on time-domain responses measured at various locations is proposed. The time-domain spectral finite element model is adopted as it requires a fewer number of measured responses which clearly has an advantage with respect to conventional finite element method. Hence, we use time-domain spectral finite element method, to generate the mass, stiffness, and damping matrices, which are required to perform force identification. Experiments are performed to obtain the time-domain responses on the beam and portal frame structures on which force identification is performed. The efficiency of the new algorithm is demonstrated using a variety of responses and different one-dimensional structures. However, the proposed algorithm is general in nature and can be used for one-, two-, and three-dimensional structures and with conventional generalized finite element model. The results from the proposed algorithm show an excellent match between the reconstructed force histories and the experimentally measured force. © 2020, Springer-Verlag GmbH Austria, part of Springer Nature.

Experiment and numerical investigation of the aero-thermal demise of pristine and pre-damaged GLARE™ fiber-metal laminates

Since many decades researchers have been interrogating the re-entry physics of lunar objects in the Earth’s atmosphere. A request like ‘demise + re-entry’ in a web search engine gives hundreds if not thousands of studies concerned with the breakup and the disappearance of space flotsam in the Earth’s atmosphere. However, it is hard to comprehend a rational reason why the engineering physics involved in the thermal degradation of spacecraft structures still can be explained with inherent uncertainties. Many derelict satellites and spacecraft are orbiting the Earth. They will gradually enter the Earth’s atmosphere attributed to the gravitational force and may will not survive the aero-thermal load induced by the dense Earth’s atmosphere. This study, therefore, investigated the propensity to aero-thermal demise of a pristine and a pre-damaged spacecraft shielding structure during atmospheric re-entry. The shield was made of a GLAss fiber REinforced aluminum panel, which accommodated five quasi-isotropic S2-glass fiber reinforced FM94-epoxy laminates alternately stacked between six aluminum 2024-T3 alloy sheets. The pre-damage of the shielding system was akin to a damage can be generated by an impact of a micrometeoroid. To impart the pre-damage, a GLAss fiber REinforced aluminum plate was first impacted with a 2 mm diameter stainless steel spherical projectile at 5.49 km/s. The projectile perforated through the plate. Next, the perforated GLAss fiber REinforced aluminum plate was exposed to an air plasma-jet (oxygen dissociated) of a specific enthalpy of 5.4 × 103 kJ/kg at standard air pressure conditions. The high enthalpy plasma thermally degraded the pre-damaged plate, while debris was discharged uprange and downrange out of the plate. The dynamic pressure of the plasma flow exacerbated the debris discharge. The endothermic destruction of a pristine GLAss fiber REinforced aluminum specimen followed several steps: first, layer-after-layer oxidation and melting of the aluminum sheets; next, decomposition and sublimation of epoxy accompanied by thermal and viscous diffusion of molten aluminum and molten glass fibers; finally, the onset of a hole in the pristine specimen. By contrast, the plasma-jet that streamed through the punched hole of the pre-damaged GLAss fiber REinforced aluminum specimen, had thermally decomposed the frontal material layers of the specimen to a large extent in the initial few seconds. The punched hole, however, did not succumb to enlarge after the surface temperature of the GLAss fiber REinforced aluminum specimen reached an asymptote. To reproduce the experimental findings, a new Moving Mesh Front procedure had been developed and implemented in a Lagrangian finite element framework. The numerical analysis considered the temperature-dependent thermal properties of orthotropic S2-glass/FM94-epoxy composites and isotropic aluminum alloys. For a comparable heating condition, the agreement between the predicted and the measured temperatures of the specimen front surface was within 4.4%, while the corresponding agreement was within 6.7% for the rear surface temperature of the test specimen.

Improved Craig–Bampton stochastic method for spacecraft vibroacoustic analysis

This paper addresses the susceptibility of finite element models to uncertainty in frequency ranges with relatively high modal density, particularly in the context of vibroacoustic analysis. The principal idea is based on a stochastic finite element method (FEM) technique called Craig–Bampton stochastic method (CBSM). It is a parametric Monte Carlo simulation (MCS) approach that can be performed at a fraction of the otherwise potentially impractical computational cost, due to the use of reduced rather than full system matrices. An enhanced formulation of the CBSM, significantly improving its efficiency by exploiting the block structure of the condensed model’s stiffness and mass matrices is derived. The improved method is adapted for use with distributed loads, such as diffuse sound field excitation. Its practical implementation is illustrated through a simple theoretical example followed by a high-complexity spacecraft structure case. In both cases solutions are compared to those for a classic MCS of the non-condensed models. Through an extensive parametric survey, recommendations are given on the ideal perturbation levels and underlying statistical distributions for the improved CBSM’s random variables. The proposed technique shows a very strong agreement with the benchmark MCS results. Computational time reductions of over 1 and 3 orders of magnitude against the original CBSM and the MCS, respectively, are demonstrated.

Analysis of Eddy Current Generation on the Juno Spacecraft in Jupiter's Magnetosphere

AbstractThe Juno mission to Jupiter, in polar orbit around the gas giant since 4 July 2016, samples the planet's environment with dedicated particle and fields instruments. Juno's magnetometer investigation employs a pair of boom‐mounted vector fluxgate magnetometers colocated with a set of star cameras to map Jupiter's magnetic field with high accuracy. Juno is a spinning spacecraft, rotating at approximately two rotations per minute. In strong magnetic field environments experienced near periapsis, Eddy currents are generated within electrically conductive material near the magnetic sensors. These currents adversely affect measurements of the environment, most evident in the appearance of a spin modulation in the field magnitude. We demonstrate, by finite element modeling and laboratory measurements, that the spin modulation is caused by a physical signal due to Eddy currents generated by the rotation of the conductive spacecraft structure in the presence of a strong magnetic field. We present a finite element model of the induced field and develop a matrix method for removing the Eddy current contribution to the measured field. Juno magnetic field measurements in strong fields are corrected for Eddy current contributions using this model of the interaction.

Thermomechanical Fatigue of Unidirectional Carbon Fiber/Epoxy Composite in Space

Thermomechanical fatigue is one of the challenges for spacecrafts during space missions. As a result of the extreme temperature variation in space, thermal cycles are created, and due to the imposed mechanical loads to spacecrafts such as engine loads while they turn on and off, mechanical cycles are created. The worst structural fatigue for spacecrafts occurs when both thermal and mechanical cycles happen, simultaneously. The reason is that both thermal and mechanical cycles could cause stress concentration in the spacecraft structure. Therefore, the probability of crack initiation or propagation in the spacecraft structure increases. In this study, for the first time, novel thermomechanical fatigue relations are introduced to evaluate the safety of unidirectional carbon fiber/epoxy composite in the space environment. This composite material has been used in the spacecraft structure due to its lightweight and high strength. Furthermore, with applying thermal fatigue relations, thermally-safe planets, moons, and asteroids in the solar system for unidirectional carbon fiber/epoxy composite are identified.

Potential and challenges of additive manufacturing for topology optimized spacecraft structures

This study focused on the potential of topology optimization (TO) for metallic tertiary structures of spacecrafts produced by the additive manufacturing (AM) technique laser powder bed fusion. First, a screening of existing conventionally manufactured products was carried out to evaluate the benefits of a redesign concerning product performance and the associated economic impact. As a result of the study, the most suitable demonstrator was selected. This reference structure was redesigned by TO taking into consideration the AM process constraints. Another major aim of this work was to evaluate the possibilities and challenges of AM (accuracies, surface quality, process parameters, postmachining, and mechanical properties) in addition to the redesign process. A comprehensive approach was implemented including detailed analysis of the powder, mechanical properties, in-process parameters, and nondestructive inspection (NDI). All measured values were used for a back loop to the design process, thereby providing a final robust redesign. Finally, the fine-tuned demonstrator was built up in an iterative process. The parts were tested under representative conditions for the application to verify the performance. The demonstrator qualification test campaign contained thermal cycling, vibration testing, static load testing, and NDI. Thus, an improvement in technology readiness level up to “near flight qualified” was reached.

Self-sealing Polymeric Materials: Mechanism and Applications

Self-sealing polymers possess the virtue to seal small punctures/cracks/damages automatically whenever occurred in the materials/components during their service. This behaviour of self-sealing in polymeric materials is inspired by the biological system in which self-sealing is a usual phenomenon. There are many self-sealing approaches demonstrated by researchers which are being applied in many polymeric components and systems to prevent them from catastrophic failure. The sealant materials in these systems are the main parts which play a crucial role in the sealing phenomenon. This review article describes self-sealing phenomenon, sealant materials and their applications in the various strategic materials and systems namely self-sealing spacecraft structures in the meteoroid environments, self-sealing fuel tank for fighter aircrafts, self-sealing hemostatic syringe to prevent bleeding, self-sealing tires to prevent punctures and concrete structures to prevent water leakages etc.

Моделирование динамики движения спускаемого аппарата на заключительном этапе посадки

When designing promising spacecrafts and missions to explore the planets of the Solar system, preliminary testing of all the important systems affecting the success of the space mission is necessary. Among the other systems responsible for the functioning of the spacecraft, it is necessary to distinguish the landing system. The landing stage is important and complex, since this stage is accompanied by large overloads and forces of the stick-slip nature acting on the spacecraft, their effect can be critical for the spacecraft structure. When landing the descent vehicle, it is necessary to study separately the moment of approach to the surface and the movement in the ground. The purpose of this article is to describe the methodology for modeling the motion dynamics at the final stage of movement and perform numerical studies of the descent vehicle motion dynamics at the final stage of movement in the case of a hard landing. Studies were conducted for different initial conditions of approach to the surface. As a result, the values of the arising overloads acting on the descent vehicle were obtained. The data provided by the simulation allow making conclusions about the most dangerous landing options that should be considered for the successful completion of future missions.

Research on High Temperature Protection Technology of Infrared Temperature Measuring Equipment in Space Environment

With the increasing complexity of spacecraft structure and thermal design, it is more difficult to implement surface temperature measurement technology, and the temperature measurement area tends to diverse. Therefore, there is an increasing demand for the application of non-contact temperature measurement technology in spacecraft thermal test. This paper takes the application of infrared temperature measuring equipment in vacuum and high temperature environment as the research object, designs the thermal protection scheme and device of the equipment, and simulates and analyses the thermal protection of infrared temperature measuring equipment based on node network method. Through physical test, the device can effectively realize the thermal protection of equipment in vacuum and high temperature environment, and ensure that the equipment is in the normal working temperature range and its temperature measurement algorithm model is not affected, the accuracy of temperature measurement is better than ±2°C, which meets the use requirements in spacecraft thermal test.

Digital twin for the structural health management of reusable spacecraft: A case study

Reusable spacecraft can significantly reduce the cost of space travel, while evaluation of the structural health of the craft between flights becomes one of the key issues. A digital twin framework is proposed in this paper to track the life of spacecraft structures. A digital twin is a digital representation of an engineering system. It can simulate, monitor, diagnose, predict states and optimize operations of the real engineering system in real time. The proposed framework can be divided into offline and online stages. It has the functions of diagnosis, model updating, performance evaluation and data storage. To demonstrate the prognosis and decision support capabilities of the framework, a numerical example considering fatigue crack growth in a load-bearing frame is carried out. The method of manufactured solutions is employed for validation. Information entropy and relative entropy are used for measuring the uncertainties in crack length prediction. The results show that through the framework, crack growth model can be updated to have a lower uncertainty. Future crack growth and reusable life can be predicted more accurately using the improved model. With the structural life of the spacecraft quantified by the framework, mission success rates for repeated flights can be maximized at a lower cost.

Dynamics and stabilization of flexible spacecraft structures with magnetic coil-pair actuators

The dynamics of flexible spacecraft structures with embedded magnetic actuators are studied. Discretized equations of motion for a constrained plate model and an unconstrained two-plate/rigid bus model are derived for actuators that create in-plane forces via current-carrying coils. Taking inputs to be the magnetic forces generated between pairs of coils, the goal of vibration suppression is shown to result in a bilinear control problem. The selection of a generalized rate output variable is used to obtain a passive map from coil-pair force inputs, yielding $$\mathscr {L}_2$$ stability for strictly passive feedback laws. The implementation of one such feedback corresponds to using collocated sensing at the actuator locations, and given mild assumptions about the actuator location, is a robust stabilizer with respect to unstructured uncertainties. Simulations of the closed-loop system are completed to support findings throughout.

Research on Leakage Location of Spacecraft in Orbit Based on Frequency Weighting Matrix Beamforming Algorithm by Lamb Waves

Clashes between space debris and spacecraft in orbit may cause air leakages, which pose a substantial danger to the crew and the spacecraft. Lamb wave dispersion in spacecraft structures and the randomness of leak holes are the difficulties in leak location. To solve these problems, a frequency weighting matrix beamforming algorithm is proposed in this paper. The elastic Lamb waves that are caused by leakages are acquired by an ‘L’ shaped sensor array consisting of eight acoustic emission sensors. The angle of a leak can be obtained through the superposition of different time delays, and the intersection of two angles can be used to find the location of the leak. Traditional beamforming is improved by matching the wave speeds in different frequency bands and weightings according to the energy distribution. Narrowband filtering is used to delay overlay different signal speeds with different frequency bands via a dispersion curve. The weighting method is used to compensate the frequency band response of different leak holes. The detailed location algorithm process is introduced and verified by experiments. For 1.5 and 2 mm leak holes, location direction accuracies of 1.33° and 1.93° for one sensor array were obtained, respectively.

Numerical investigation of debris impact on spacecraft structure at hyper-high velocity

In this paper, our analysis concentrated on the mechanical behavior of sandwich panel subjected to impact loading. Thus, to improve the resistance of satellite structures from debris collision consequences; we proceed a performance investigation numerically of different materials based on the evaluation of contact force, total energy and kinetic energy generated from the impact of spherical steel debris on sandwich panel. Furthermore, a comparison of deformations and stresses levels in sandwich panel collide by a projectile at hyper-high velocity using finite element with ANSYS code was made. The difference between our models is the type of skins materials which are aluminum and functionally graded material FGM (alumina/nickel). The impact results a total perforation of sandwich panel with aluminum facesheets creating a fragments; contrary to FGM which the projectile penetrates only the front facesheet and honeycomb core. However, for an effective protection of sandwich panel against debris perforation, the introduction of FGM skins technique is considered as a candidate solution.

Penetration Mechanics of GLARE™ Fiber-Metal Laminates upon Collision with Micrometeoroids

The spacecraft designed for long-duration service are susceptible to hypervelocity impacts of micrometeoroids and orbiting space debris. Such impacts on spacecraft structures can cause spacecraft failure and loss of life. In order to adequately protect the spacecraft bulkhead and flight critical systems, many high-strength composite materials have been developed for debris bumper. GLAss fiber REinforced aluminum is one of the high-performance composites. The review of articles, however, yielded no single study, which has been dedicated to interrogate the damage mechanics of GLAss fiber REinforced aluminum upon collision with micrometeoroids. This study, therefore, aimed at the numerical investigation of penetration phenomena of thick GLAss fiber REinforced aluminum 5-6/5-0.4 laminates upon hypervelocity impact of a projectile. The numerical study employed a predictive model that merged the smoothed particle hydrodynamics and the finite element methods. The model could predict the colossal damage modes of GLAss fiber REinforced aluminum. As seen, the normal impact of a 2 mm diameter spherical 2024-T3 aluminum projectile on GLAss fiber REinforced aluminum, at a relative velocity of 7.11 km/s, resulted into membrane stretching and fiber failure of the glass fiber reinforced epoxy composite laminates. By contrast, the aluminum layers experienced an enormous strain-rate and consequently, suffered thinning, fracture and large mass erosion. The perpetual release waves fragmented the projectile and dispersed the projectile mass prior to the further penetration of GLAss fiber REinforced aluminum. To verify the accuracy of the numerical model, experiments had been conducted by using a two-stage light-gas gun. The experiments generated damage modes of GLAss fiber REinforced aluminum in good correspondence to that of the predictive model. Yet, disparity between the estimations of experiments and simulations had been apparent, which is anticipated due to the phase change of material had not been accounted for in the analysis.