Elements Of Spacecraft Research Articles

Application of Micro-Arc Oxidation Methods in Making the Structure Elements of Spacecraft

Application of Micro-Arc Oxidation Methods in Making the Structure Elements of Spacecraft

Space debris and micrometeoroid impact on spacecraft elements: Experimental simulation

We study experimentally how optical, electrical, and physical properties of spacecraft surface elements (optical glasses, “metal-insulator-metal” structures, and solar arrays) change upon bombardment by high-speed particles of submicron and micron size. For particle acceleration, we use an electrodynamic accelerator.

The statistical treatment implemented to obtain the planetary protection bioburdens for the Mars Science Laboratory mission

NASA Planetary Protection Policy requires that Category IV missions such as those going to the surface of Mars include detailed assessment and documentation of the bioburden on the spacecraft at launch. In the prior missions to Mars, the approaches used to estimate the bioburden could easily be conservative without penalizing the project because spacecraft elements such as the descent and landing stages had relatively small surface areas and volumes. With the advent of a large spacecraft such as Mars Science Laboratory (MSL), it became necessary for a modified—still conservative but more pragmatic—statistical treatment be used to obtain the standard deviations and the bioburden densities at about the 99.9% confidence limits. This article describes both the Gaussian and Poisson statistics that were implemented to analyze the bioburden data from the MSL spacecraft prior to launch. The standard deviations were weighted by the areas sampled with each swab or wipe. Some typical cases are given and discussed.

Stability of an impulsive control scheme for spacecraft formations in eccentric orbits

An N-impulse control scheme for spacecraft formations in elliptical orbits is developed to regulate the differential elements of the deputy spacecraft in the presence of the J2 perturbation. The presented control scheme is an extension of an existing circular-orbit formation control scheme and is shown to perform well at large eccentricities where the circular control scheme fails. For the case of two impulses being applied at arbitrary firing times, a discrete-time approach for ascertaining the stability of the controlled spacecraft formation, under the assumption of small impulsive thrusts, is presented. It is found that stability is guaranteed for the majority of firing time pairs; however, the requisite ΔV can be prohibitive for some firing time pairs. The control scheme and stability predictions for formations in high eccentricity orbits are validated in numerical simulation.

Plasma particle simulations on stray photoelectron current flows around a spacecraft

In tenuous space plasmas, photoelectron flows produce complex current paths among multiple conducting elements of spacecraft, which may influence the current–voltage characteristics of double‐probe electric field sensors. We performed full‐particle simulations on this effect by assuming a sensor configuration that is typical of recent designs like those on Cluster, THEMIS, and BepiColombo/MMO; the spherical probe is separated from a conducting boom by biased electrodes known as the ‘stub’ and the ‘guard’. The assumed bias potential scheme corresponds to that planned for BepiColombo/MMO and is different from those used in the other satellites. The analysis focuses on stray photoelectron currents flowing from these electrodes and a spacecraft body. Photoelectrons approaching the probe are commonly repelled by the guard, the potential of which is strongly biased negatively, and are subsequently affected by the probe potential. Consequently, the photoelectron current magnitude increases with increasing probe potential regardless of their origins, when the probe operates between the plasma and floating spacecraft potentials. The result indicates that both photoelectron currents from the spacecraft body and biased electrodes can be minimized by selecting the probe working potential as close as possible to the plasma potential. We also examine the photoelectron current dependence on the presence or absence of the guard electrode operation and confirm a positive effect of reducing the photoelectron current from the spacecraft. However, negative side effects of the guard operation enhance the photoelectron currents from the stub and guard, when the probe operates nearly at the plasma potential.

Dynamic Stability Optimization of Laminated Composite Plates under Combined Boundary Loading

Dynamic stability and design optimization of laminated simply supported plates under planar conservative boundary loads are investigated in current study. Examples can be found in internal connecting elements of spacecraft and aerospace structures subjected to edge axial and shear loads. Designation of such elements is function of layup configuration, plate aspect ratio, loading combinations, and layup thickness. An optimum design aims maximum stability load satisfying a predefined stable vibration frequency. The interaction between compound loading and layup angle parameter affects the order of merging vibration modes and may stabilize the dynamic response. Laminated plates are assumed to be angle-plies symmetric to mid-plane surface. Dynamic equilibrium PDE has been solved using kernel integral transformation for modal frequency values and eigenvalue-based orthogonal functions for critical stability loads. The dictating dynamic stability mode is shown to be controlled by geometric stiffness distributions of composite plates. Solution of presented design optimization problem has been done using analytical approach combined with interior penalty multiplier algorithm. The results are verified by FEA approach and stability zones of original and optimized plates are stated as final data. Presented method can help designers to stabilize the dynamic response of composite plates by selecting an optimized layup orientation and thickness for prescribed design circumstances.

Modelling ion propulsion plume interactions with spacecraft in formation flight

AbstractThis paper presents a simulation study of ion thruster plume effects on formation flying spacecraft. Two formation flight applications using micro-ion propulsion are considered: an L2-Halo orbit interferometer formation and a LEO micro-satellite formation. Worst case scenarios in both missions have been investigated. We focus our study on thruster configurations resulting in possible indirect plume impingement on satellites outside of the direct impingement zone. Indirect impingement which cannot be predicted except through plasma simulation or in-flight measurements might expose critical spacecraft elements such as optical sensors to a harsh contamination environment. A high-fidelity electrostatic plasma simulation code for parallel computing platforms was used in the study.In our study, we found that using miniature scale ion propulsion in both formation missions creates a very low charge-exchange plasma environment which results in tolerable contamination environment for other spacecraft in the formation.

Experimental and theoretical investigation of encounter of a group of particles with protecting elements of spacecraft

We present the results of comprehensive experimental and theoretical investigations of encounter with a barrier of a group of bodies thrown with a high-velocity. Throwing of a group of particles (from two to twelve bodies) was realized on a ballistic route using powder and light-gas units of different calibers in the range of velocities 500–3500 m/s. The process of particle throwing was controlled by acting aerodynamic forces. In experiments on collisions with barriers of a finite thickness (which imitates the protective shield of spacecraft) the number of particles in a homogeneous stream was varied from 2 to 7 at changing the flux density (distances between particles). Experimental data are obtained on variations of the area and mass of back-surface splinters. Numerical calculations simulated a knock of 2 to 4 particles against a barrier in the cases of normal impact and at an angle. The calculations were performed in three-dimensional formulation and applying criteria of complete destruction of material. The appearance of additional destruction centers in the barrier due to mutual influence of particles is revealed. Simple criteria are obtained for estimating the degree of interference of particles and the character of barrier destruction.

Dynamic interaction of rotating momentum wheels with spacecraft elements

In modern spacecraft with the requirement of increased accuracy of payloads, the on-orbit structural dynamic behavior of spacecraft is increasingly influencing the design and performance of spacecraft. During the integrated spacecraft testing of one of the satellites, a strong coupling between rotating momentum wheels and an earth sensor was detected. This resulted in corruption of the earth sensor data at certain wheel speeds. This paper deals with the dynamic coupling problem of a rotating momentum wheel with its support brackets affecting other subsystems of spacecraft. As part of this investigation, extensive modal tests and vibration tests were carried out on the momentum wheel bracket assembly with wheels in stationary and rotating conditions. It was found that the effects of gyroscopic forces arising out of rotating wheels are significant and this aspect needs to be taken into account while designing the mounting brackets. Results of analysis and tests were used to redesign the bracket leading to a significant reduction in the interaction and associated problems. A procedure for design of a support structure using a low-order mathematical model is also shown.

Comparison of interball-2 spacecraft potential from IESP and KM-7 experiments in high-latitude regions of the magnetosphere at altitudes of 2–3 R E

Measurements of Interball-2 spacecraft potential by two instruments, IESP and KM-7, are reviewed and simultaneous measurements are compared. Unacceptable discrepancy between results of spacecraft potential measurements, including opposite signs, was found. Actually, both experiments are methodically identical: they used the same type of sensor - spherical Langmuir probes operating in the “floating” mode, and they have measured the same parameter - the voltage difference between the probe and the satellite structure. The IESP instrument measured one value of this parameter at fixed bias current to the probe. The KM-7 measured the whole current-voltage characteristic (the probe current as a function of the probe potential), which consists of 11 IESP-type measurements at different values of bias current. The difference lies only in the way of technical implementation, as the probes were operating in different ambient conditions. The IESP probes were mounted at the ends of long booms and thus were affected by the solar UV emission, while the KM-7 probe was rather well protected against UV emission by proper mounting and screening the head of the sensor from both direct UV emission and those reflected from spacecraft elements. The comparison of two data sets and variations along the orbit has shown, that KM-7 correctly measures the spacecraft potential. In high-latitude inner magnetosphere (in auroral region and polar cap) at altitudes 2–3 R E the spacecraft potential was predominantly negative down to −10 V, increasing up to about +5 V in some locations at rather extended parts of the satellite orbit. Reasonably good agreement between two experiments was obtained only at spacecraft potential ≤ −2 V and after inversion the sign of available calibration curve of IESP experiment; in this case both experiments very accurately reflect variation of potential data even in the small details. The operation mode of IESP experiment was not optimal and requires further analysis. The wave results obtained with IESP may be considered as reliable only when both experiments show, after inversion of available IESP calibration curve, coincident data, but the phase relations, if any, will be inversed in any case. The method of spacecraft potential measurements adopted in IESP should be used only for positive potential; if realspacecraft potential is negative, the method gives wrong and misleading results.

On the distribution of control functions in new-generation on-board embedded real-time systems

The opening of the space business to the commercial market causes the cost and criticality of the on-board system to shift towards the payload component that delivers the value-added of the mission. As the cost of the payload increases, every effort is made to reduce the cost of the other elements of the spacecraft. The integration of the traditional on-board real-time control functions into a single platform component, denominated 'spacecraft bus', is one of the main effects of this cost-reduction effort. The new scenario is likely to require the allocation of additional real-time control functions to payload components. The immediate consequence of this phenomenon is the distribution of on-board services across the spacecraft. In earlier work we have presented an engineering approach to the construction of the real-time software embedded on-board integrated platform components. In this paper, we extend that approach to address the future scenario in which real-time control functions may be distributed across multiple processing nodes in the system.

Optimization of Solutions for Ensuring a Long Service Life for Spacecraft

An approach is proposed to ensuring reliability of spacecraft with a long active service life. Models are developed to determine the spacecraft elements which are critical as regards failure under uncertain operating conditions. The problem of optimizing the means of ensuring reliability is formulated.

Deployment of Reconfigurable Spacecraft Under Earth Conditions in the Out-Of-Plane Rotation Mode

The cable-suspension method is used to simulate weightlessness on Earth in the general case. By an example of mathematical simulation of a system of rigid bodies connected via hinged joints and suspended from a special facility, it is shown that the method developed provides realistic simulation of orbital processes under earth conditions. The validation criterion for the simulation is the closeness of the values of the forces and moments in joints and the attitude and angular velocities to their theoretical values in orbital motion. Some methods for the simulation of weightlessness conditions on Earth are known: free falling, an airplane flying along a parabolic trajectory, and a pool. Each of these methods has advantages and shortcomings. A method for simulating the weightlessness of a transformable spacecraft is also known. In this case, the spacecraft elements are suspended by cables from points movable in a horizontal plane. This method allows simulating only plane motions of the main module arising in changing the geometry of the spacecraft. If the reconfiguration of an orbiting spacecraft causes the spatial rotation of the ma in module, then the existing cable-suspension methods are inapplicable. Here, we consider the cable-suspension method as applied to precisely this general case. By mathematically simulating the behavior of a system of solids connected via hinged joints and suspended from a special facility, it is shown that the method developed provides realistic simulation of orbital processes under earth conditions. As a validation criterion for the simulation, we check how close the values of the forces and moments in joints and the orientation angles and angular velocities are to the analogous theoretical values in orbital motion. Within the framework of analytical mechanics, a rigorous mathematical model of the facility, including the object under investigation, was created. It allows us to accumulate a great body of data on the spacecraft dynamics. An analysis of the behavior of a system of bodies spatially rotating under the action of some perturbing factors in the laboratory environment allowed us to establish the operational reliability of the simulation technique proposed and its low sensitivity to various perturbations. The problem of simulating weightlessness conditions on Earth for spacecraft of variable (in order to change over from the transportable to operational state) configuration arose simultaneously with spacecraft themselves. A technique is known for simulation of these conditions at a ground laboratory for objects changing their geometry through the elastic vibrations of their structural elements. Certain structural elements are suspended by cables and springs from fixed supports [3]. A shortcoming of this and other known methods and facilities is that they allow modeling only the plane motion of structural elements. In the overwhelming majority of cases, the motion of actual space systems is spatial. This is because the spacecraft and all its separate elements execute complex spatial motion about the center of mass under the internal forces that ensure the reconfiguration. The angular motion of a spacecraft and the weightlessness state simulated under plane-motion conditions are far from real. Such a technique allows us to simulate the motion of an object only in the internal degrees of freedom, transforming relative motion to absolute motion. The new weightlessness simulation technique considered here implies simulation of the motion of a spacecraft not only in internal but also in three external angular degrees of freedom.

Aging of polymer composite materials exposed to the conditions in outer space

A comprehensive investigation is made of glass, carbon, organic fiber-reinforced plastics, and epoxy-based hybrid composite materials employed in Salyut-type spacecraft which remained in space for up to 1501 1501 days. In particular, the properties, aging mechanism, and strain-strength variations in these materials due to exposure to the conditions in outer space were studied. After a series of tests were performed in space the standard strain and strength parameters as well as the mass, density, and thickness changes in the composite materials were estimated. Electron-microscopic and dynamic-mechanical analyses were performed, and the thermal expansion was estimated for a wide range of temperatures. The principal, dominant process occurring due to the continuous presence in outer space was found to be post-curing of the resin materials, which in turn affected the mechanical characteristics of the composite materials. After 456–1501 days in space the room-temperature strength of the composite materials (except for organic plastics) did not decrease, while at high temperatures it even increased. The post-curing and restructuring of some composite materials lowered their dynamic shear moduli in the glassy state of the resin. Due to consolidation of the surface layer of hybrid composite materials irradiated and subjected to thermal cycles, failure during bending varied from transverse fracture to delamination. The negative effect of the post-curing process was expressed as higher internal tension in the hybrid composite materials with different linear thermal expansion coefficients. The magnitude of this effect depended on the amplitude of the thermal cycles. The unprotected surface of the composites bombarded by atomic oxygen, microparticles, and space garbage were subjected to pickling and microerosion, the maximum effect occurring at the initial stage of exposure. Desorption of moisture and low-molecular products during the first 100–200 days of thermal cycling in the vacuum of near-earth orbit must be considered when estimating the total mass loss of composite materials. Data from microscopic, dynamic-mechanical, and other types of analyses revealed that the outer-space factors improved the supermolecular order of the resin volume, while the subsurface layer structure of the composite materials had loosened. Microcracks formed in the plastic's surface during 1501 days in outer space did not, in general, affect the mechanical parameters of the composite materials. Most of the observed effects of exposure to conditions in outer space were less pronounced for plastics protected by aluminum foil or other plastic coatings. The data obtained can be used for designing external elements of spacecraft by selecting materials with specified and predictable properties for long-term service.

Mathematical modeling of nonstationary thermophysical processes in spacecraft onboard-apparatus bays

Propagation of heat due to heat conduction and operation of the low-temperature heat pipes in the structural elements of spacecraft are modeled numerically. The multidimensional nonstationary temperature fields in honeycombed panels with spatially distributed and locally concentrated heat sources were calculated.

The 1981 Jupiter orbiter probe mission

Extensive planning for an orbiter and probe mission to Jupiter has been underway at the Jet Propulsion Laboratory and the Ames Research Center for several years. Recent efforts have focused on the development of a mission design and spacecraft system for the late 1981-early 1982 launch opportunity to Jupiter. Designated the Jupiter Orbiter Probe (JOP) mission, this mission will allow scientists to conduct extensive surveys of the planet, its environment, and satellites by using elliptic orbits whose inclinations and orientations will vary during the mission. Close satellite flybys will perturb the orbit of the spacecraft greatly. During the many orbits of varying eccentricity over the lifetime of the mission, extensive investigations of the Jovian magnetosphere and imaging, and other remote sensing of Jupiter and its satellites, will be possible. This paper presents a synopsis of the IOP science, mission and spacecraft elements developed during the current Pre-Project Phase. It is planned that the JOP Project begin in fiscal year 1978.

Establishment of a redundancy priority for spacecraft elements : E. R. Carrubba, Proc. 1970 Ann. Symp. Reliab., Los Angeles. IEEE Cat. No. 70 C 2-R, 3–5 February (1970), p. 187

Establishment of a redundancy priority for spacecraft elements : E. R. Carrubba, Proc. 1970 Ann. Symp. Reliab., Los Angeles. IEEE Cat. No. 70 C 2-R, 3–5 February (1970), p. 187