Representative Spacecraft Research Articles

EHFE refinement and extension for variable physical Earth radiation in spacecraft uncertainty thermal analysis

Spacecraft uncertainty thermal analysis is an important part of ensuring robust and reliable spacecraft thermal control system design. The calculation of the Earth external heat flux is an important part of it, however, the current model for calculating the Earth external heat flux is the constant physical Earth radiation model. In this paper, a model of the Earth’s variable physical radiation is developed based on the long-term averaging method. We refine and extend the application of the external heat flux expansion (EHFE) formula to accommodate our proposed model. This study further includes an uncertainty analysis of Earth radiative external heat flux, utilizing the enhanced EHFE formula in the context of a representative spacecraft. We provide statistical data, such as mean values, standard deviation, and probability density functions, derived from our analysis. The augmented EHFE formula, when properly generalized, can be applied to compute the celestial body’s variable physical radiative external heat flux affecting the spacecraft. This methodological advancement offers theoretical underpinning for the thermal design of spacecraft intended for deep space exploration.

Estimating Fields in Spacecraft Cavities: Experimental Validation and Optimization of Finite-Difference Time-Domain and Power Balance Computational Tools

Electromagnetic fields in representative spacecraft cavities were successfully predicted using finite-difference time-domain and power balance computational tools. Results were validated with measurements of two test articles, showing excellent correlation in shielding effectiveness from 300 MHz to 18 GHz. The validated tools were then extended to predict fields inside representative, to-scale payload fairings including common systems and components like satellite payloads, antennas, acoustic blankets, and a cable harness. Various computational techniques were used to compare their speed and accuracy. Ultimately, we conclude that a multi-fidelity approach – combining full-wave, statistical, and hybrid solutions – is beneficial and necessary for complex and large cavity problems. The tools and techniques presented here can serve as part of a toolkit to rapidly estimate shielding effectiveness, the impact of payloads, and overall fields in spacecraft cavities.

Re-entry prediction and demisability analysis for the atmospheric disposal of geosynchronous satellites

The paper presents a re-entry analysis of Geosynchronous Orbit (GSO) satellites on disposal trajectories that enhance the effects of the Earth oblateness and lunisolar perturbations. These types of trajectories can lead to a natural re-entry of the spacecraft within 20 years. An analysis was performed to characterise the entry conditions for these satellites and the risk they can pose for people on the ground if disposal via re-entry is used. The paper first proposes a methodology to interface the long-term propagation used to study the evolution of the disposal trajectories and the destructive re-entry simulations used to assess the spacecraft casualty risk. This is achieved by revisiting the concept of overshoot boundary. The paper also presents the demisability and casualty risk analysis for a representative spacecraft configuration, showing that the casualty risk is greater than the 10$^{-4}$ threshold and that further actions should be taken to improve the compliance of these satellites in case of disposal via re-entry is used.

Co-Design of Strain-Actuated Solar Arrays for Spacecraft Precision Pointing and Jitter Reduction

This work presents a novel spacecraft attitude control architecture using strain-actuated solar arrays that does not require the use of conventional attitude control hardware. A strain-actuated solar array enables attitude slewing maneuvers and precision pointing (image acquisition) stares, while simultaneously suppressing structural vibrations. Distributed piezoelectric actuators help achieve higher precision, higher bandwidth, and quieter operation than reaction wheels. To understand the design tradeoffs for this architecture, a framework for the integrated design of distributed structural geometry and distributed control is presented. The physical properties of the array are modeled and designed with respect to a piecewise linear distributed thickness profile. The distributed control is a voltage profile across the array modeled as a spatially continuous function. The dynamics of the system are modeled using a coupled ordinary differential equation–partial differential equation system using extended generalizations for hybrid coordinate systems. The combined physical and control system design, or co-design problem is investigated to understand the optimal performance of the system. Single-axis slew maneuvers of 7.2 milliradians or 1485 arcsec are achieved for a representative spacecraft model without increasing array mass or reducing array planform area. From additional tradeoff studies, a design criteria is revealed for the array structure and control strategy based on the optimal design tradeoff between large array inertia and fast structural dynamics. Moreover, the fundamental limits on strain-actuated solar arrays slew angle magnitude are demonstrated using an intuitive pseudorigid body dynamic model.

Generalized Frequency Domain State-Space Models for Analyzing Flexible Rotating Spacecraft

The mathematical model for a flexible spacecraft that is rotating about a single axis rotation is described by coupled rigid and flexible body degrees-of-freedom, where the equations of motion are modeled by integro-partial differential equations. Beam-like structures are often useful for analyzing boom-like flexible appendages. The equations of motion are analyzed by introducing generalized Fourier series that transform the governing equations into a system of ordinary differential equations. Though technically straightforward, two problems arise with this approach: (1) the model is frequency-truncated because a finite number of series terms are retained in the model, and (2) computationally intense matrix-valued transfer function calculations are required for understanding the frequency domain behavior of the system. Both of these problems are resolved by: (1) computing the Laplace transform of the governing integro-partial differential equation of motion; and (2) introducing a generalized state space (consisting of the deformational coordinate and three spatial partial derivatives, as well as single and double spatial integrals of the deformational coordinate). The resulting math model is cast in the form of a linear state-space differential equation that is solved in terms of a matrix exponential and convolution integral. The structural boundary conditions defined by Hamilton’s principle are enforced on the closed-form solution for the generalized state space. The generalized state space model is then manipulated to provide analytic scalar transfer function models for original integro-partial differential system dynamics. Symbolic methods are used to obtain closed-form eigen decomposition–based solutions for the matrix exponential/convolution integral algorithm. Numerical results are presented that compare the classical series based approach with the generalized state space approach for computing representative spacecraft transfer function models.

A Comprehensive Diagnosis Methodology for Complex Hybrid Systems: A Case Study on Spacecraft Power Distribution Systems

The application of model-based diagnosis schemes to real systems introduces many significant challenges, such as building accurate system models for heterogeneous systems with complex behaviors, dealing with noisy measurements and disturbances, and producing valuable results in a timely manner with limited information and computational resources. The Advanced Diagnostics and Prognostics Testbed (ADAPT), which was deployed at the NASA Ames Research Center, is a representative spacecraft electrical power distribution system that embodies a number of these challenges. ADAPT contains a large number of interconnected components, and a set of circuit breakers and relays that enable a number of distinct power distribution configurations. The system includes electrical dc and ac loads, mechanical subsystems (such as motors), and fluid systems (such as pumps). The system components are susceptible to different types of faults, i.e., unexpected changes in parameter values, discrete faults in switching elements, and sensor faults. This paper presents Hybrid Transcend, which is a comprehensive model-based diagnosis scheme to address these challenges. The scheme uses the hybrid bond graph modeling language to systematically develop computational models and algorithms for hybrid state estimation, robust fault detection, and efficient fault isolation. The computational methods are implemented as a suite of software tools that enable diagnostic analysis and testing through simulation, diagnosability studies, and deployment on the experimental testbed. Simulation and experimental results demonstrate the effectiveness of the methodology.

Vulnerability of spacecraft harnesses to hypervelocity impacts

Abstract During a study performed in framework of a European Space Agency contract, the vulnerability of spacecraft harnesses has been assessed. The harnesses consisted of three different space-grade cable types: power cables, screened twisted pair data cables and radio frequency (RF) cables. They were alternately shielded by two different types of representative spacecraft structure walls. Ten hypervelocity impact (HVI) tests at 0° incidence have been performed with impact velocities ranging from 6.4 km/s to 7.7 km/s. Projectiles have been aluminium spheres with diameters ranging from 1.5 mm to 4.0 mm. During the tests, all cables were operated at their representative conditions and the disturbances, induced by the impacts, were measured. The malfunction observed could be related to two physical failure mechanisms: (1) short circuits caused by a conducting cloud of molten and evaporated aluminium and, for the data cables, (2) strands being bent by impacting fragments creating a short circuit between screen and signal cable. The influence of a more complex structure wall to the failure mechanism in (1) is shown. Malfunction was dependent on mechanical damage, but no clear correlation between severity of malfunction and mechanical damage could be established. The data cables were the most vulnerable cable, while the RF cables were the most robust. The disturbances recorded could pose a significant threat to connected electronic equipment. Examples of electrical performance are given.

Generating requirements for complex embedded systems using State Analysis

It has become clear that spacecraft system complexity is reaching a threshold where customary methods of control are no longer affordable or sufficiently reliable. At the heart of this problem are the conventional approaches to systems and software engineering based on subsystem-level functional decomposition, which fail to scale in the tangled web of interactions typically encountered in complex spacecraft designs. Furthermore, there is a fundamental gap between the requirements on software specified by systems engineers and the implementation of these requirements by software engineers. Software engineers must perform the translation of requirements into software code, hoping to accurately capture the systems engineer's understanding of the system behavior, which is not always explicitly specified. This gap opens up the possibility for misinterpretation of the systems engineer's intent, potentially leading to software errors. This problem is addressed by a systems engineering methodology called State Analysis, which provides a process for capturing system and software requirements in the form of explicit models. This paper describes how requirements for complex aerospace systems can be developed using State Analysis, using representative spacecraft examples.

Engineering Complex Embedded Systems with State Analysis and the Mission Data System

It has become clear that spacecraft system complexity is reaching a threshold where customary methods of control are no longer affordable or sufficiently reliable. At the heart of this problem are the conventional approaches to systems and software engineering based on subsystem-level functional decomposition, which fail to scale in the tangled web of interactions typically encountered in complex spacecraft designs. Furthermore, there is a fundamental gap between the requirements on software specified by systems engineers and the implementation of these requirements by software engineers. Software engineers must perform the translation of requirements into software code, hoping to accurately capture the systems engineer's understanding of the system behavior, which is not always explicitly specified. This gap opens up the possibility for misinterpretation of the systems engineer s intent, potentially leading to software errors. This problem is addressed by a systems engineering methodology called State Analysis, which provides a process for capturing system and software requirements in the form of explicit models. This paper describes how requirements for complex aerospace systems can be developed using State Analysis and how these requirements inform the design of the system software, using representative spacecraft examples.

Highly-autonomous event-driven spacecraft control

Future JPL missions will continue to be scientifically and technically more ambitious, and will demand more autonomy to accomplish complex tasks in uncertain environments and in close proximity to extraterrestrial surfaces. A prime example is small body rendezvous and sample return. In addition to mission demands, affordability is now a primary driver. The call is for smaller missions with greatly reduced cost of operation and less expensive spacecraft designs. Spacecraft with highly-autonomous, goal-directed control systems are proposed to meet these challenges. This paper discusses the plan to design and develop a proof-of-concept attitude and control subsystem (ACS) that has the ability to capture science events and enable a small body rendezvous and sample return mission while requiring only one person level-of-effort for ACS ground operation support. The technology to be developed and demonstrated includes: on-board sequence generation and execution, precision closed-loop maneuver and attitude control, target acquisition and tracking, and sensing and representation of spacecraft system state. A reference mission development scenario, complete with a representative mission, spacecraft design concept, and development process, will be used in order to incorporate all the nuances of a real mission, where experience has shown that the real problems lurk. Previously-conducted JPL studies will be used to develop the scenarios. The representative spacecraft will be small to micro and consistent with a Discovery class mission. Possible approaches for the new paradigms in system architecture, ground commanding and test and verification that will be necessary for highly-autonomous event-driven controls will be addressed.

An Integrated Systems Approach for Design of Aerospace Structures

This paper is primarily concerned with the analysis and design of structural systems relating to aeronautical and aerospace engineering. Some numerical formulations are presented that deal with the synthesis of these systems integrating such disciplines as structures, unsteady aero-dynamics and control engineering. A general-purpose finite element computer program developed for associated large-scale computations is also described in some detail. Finally, results of relevant numerical analyses pertaining to some representative aircraft and spacecraft, obtained by utilizing the program, are presented that testify to the efficacy of the currently developed solution procedures.

Attitude stability of dual-spin spacecraft with unsymmetrical bodies

This paper investigates the attitude stability of dual-spin spacecraft with unsymmetrical bodies. Of particular interest is the unstable attitude behavior that can result during the despin maneuver of the spacecraft's platform or rotor. The attitude stability of several representative spacecraft configurations is investigated. The results indicate that the likelihood of unstable attitude behavior decreases as the spin-up body becomes more symmetric and the spin-down body becomes less so. The change in the relative rotation rate between the spacecraft's platform and rotor is assumed to be slow compared to the precession rate. The attitude stability information presented evolves from the application of Hill's infinite determinant method, with the results confirmed by Floquet theory. Parameter selections for the Floquet study were guided by the results predicted by Hill's method. Further, the analysis results are compared to those of Tsuchiya, which are based on the method of averaging. The present work shows that the stability regions are rich in details which have not been identified by previous investigations.

Control of Large Flexible Space Structures Using Pole Placement Design Techniques

A design approach via the pole placement techniques for the class of large flexible space structures is developed. The numerical problems in pole placement algorithm, arising from large dimension systems and the extremely low frequency eigenvalues which occur in large space structure models are examined. It shows these numerical difficulties may be overcome by properly selecting the sensor/actuator locations and introducing a frequency scaling scheme. The concepts of this paper are illustrated by some numerical studies on the linear feedback control design of a representative large spacecraft consisting of a small rigid core with ten radial booms (five booms 1000 ft long and five shorter booms 700 ft long) lying in a plane.

93-GHz radar cross section measurements of satellite elemental scatterers

The radar cross section (RCS) of representative spacecraft materials and components was measured at 93 GHz. Measurements of solar cells, thermal blanket material, structural components, and sensors are included. These measurements indicate the highly specular nature of scattering at millimeter wavelengths. The measurements generally agree with results predicted on the basis of common asymptotic formulations.

Interaction of the solar wind with the outer planets

The hypersonic analog for the interaction of the solar wind with Jupiter, Saturn, Uranus, Neptune and Pluto is used to provide estimates of shock shapes and locations as well as average magnetosheath and/or ionosheath properties for these planets. Several representative spacecraft flyby trajectories (designed for outer-planet ‘Grand Tour’ simulations) are superimposed upon a series of figures in order to provide estimates of potential plasma and field parameters which may be encountered. Consideration is given first to the possibility that several of these planets have intrinsic magnetic fields and, secondly, to the interaction of the solar wind directly on the ionosphere should there actually be no intrinsic field. Saturn and Pluto are chosen as examples of this latter case.