A ballistic limit equation for hypervelocity impacts on composite honeycomb sandwich panel satellite structures

During a recent experimental test campaign performed in the framework of ESA Contract 16721, the ballistic performance of multiple satellite-representative Carbon Fibre Reinforced Plastic (CFRP)/Aluminium honeycomb sandwich panel structural configurations (GOCE, Radarsat-2, Herschel/Planck, BeppoSax) was investigated using the two-stage light-gas guns at EMI. The experimental results were used to develop and validate a new empirical Ballistic Limit Equation (BLE), which was derived from an existing Whipple-shield BLE. This new BLE provided a good level of accuracy in predicting the ballistic performance of stand-alone sandwich panel structures. Additionally, the equation is capable of predicting the ballistic limit of a thin Al plate located at a standoff behind the sandwich panel structure. This thin plate is the representative of internal satellite systems, e.g. an Al electronic box cover, a wall of a metallic vessel, etc. Good agreement was achieved with both the experimental test campaign results and additional test data from the literature for the vast majority of set-ups investigated. For some experiments, the ballistic limit was conservatively predicted, a result attributed to shortcomings in correctly accounting for the presence of high surface density multi-layer insulation on the outer facesheet. Four existing BLEs commonly applied for application with stand-alone sandwich panels were reviewed using the new impact test data. It was found that a number of these common approaches provided non-conservative predictions for sandwich panels with CFRP facesheets.     

The YORP effect on the GOES 8 and GOES 10 satellites: A case study

The Yarkovsky-O’Keefe-Radzievskii-Paddack (YORP) effect is a proposed explanation for the observed rotation behavior of inactive satellites in Earth orbit. This paper further explores the YORP effect for highly asymmetric inactive satellites. Satellite models are developed to represent the GOES 8 and GOES 10 satellites, both of which are currently inactive in geosynchronous Earth orbit (GEO). A simple satellite model for the GOES 8 satellite is used to analyze the short period variations of the angular velocity and obliquity as a result of the YORP effect. A more complex model for the rotational dynamics of the GOES 8 and GOES 10 satellites are developed to probe their sensitivity and to match observed spin periods and states of these satellites. The simulated rotation periods are compared to observations for both satellites. The comparison between YORP theory and observed rotation rates for both satellites show that the YORP effect could be the cause for the observed rotational behavior. The YORP model also predicts a novel state for the GOES 8 satellite, namely that it could periodically fall into a tumbling rotation state. Recent observations of this satellite are consistent with this prediction.     

Phobos regolith simulants PGI-1 and PCA-1

The Martian moon Phobos is the target of the upcoming JAXA Martian Moons eXploration (MMX) mission. There are currently no known samples of Phobos, so spacecraft hardware testing and scientific studies require Phobos regolith simulants. Here, we present two new Phobos regolith simulants: Phobos Captured Asteroid-1 (PCA-1) and Phobos Giant Impact (PGI-1). These two simulants reflect the two hypotheses for Phobos’s formation, and thus the two broad possibilities for composition. This work follows previous efforts to document the development of mineralogically accurate simulants, in an effort to overcome past pitfalls with inappropriate uses of simulants. We report physical and geotechnical properties of PCA-1 and PGI-1, including their reflectance spectra, grain size distributions, abrasivity, cohesion, strength, and hardness. Our intent is for PCA-1 and PGI-1 to be open standards for Phobos simulants, and we present the simulant recipes and production methodology for use and modification by the community.     

Solar Wind Conditions and Composition During the Genesis Mission as Measured by in situ Spacecraft

We describe the Genesis mission solar-wind sample collection period and the solar wind conditions at the L1 point during this 2.3-year period. In order to relate the solar wind samples to solar composition, the conditions under which the samples were collected must be understood in the context of the long-term solar wind. We find that the state of the solar wind was typical of conditions over the past four solar cycles. However, Genesis spent a relatively large fraction of the time in coronal-hole flow as compared to what might have been expected for the declining phase of the solar cycle. Data from the Solar Wind Ion Composition Spectrometer (SWICS) on the Advanced Composition Explorer (ACE) are used to determine the effectiveness of the Genesis solar-wind regime selection algorithm. The data collected by SWICS confirm that the Genesis algorithm successfully separated and collected solar wind regimes having distinct solar origins, particularly in the case of the coronal hole sample. The SWICS data also demonstrate that the different regimes are elementally fractionated. When compared with Ulysses composition data from the previous solar cycle, we find a similar degree of fractionation between regimes as well as fractionation relative to the average photospheric composition. The Genesis solar wind samples are under long-term curation at NASA Johnson Space Center so that as sample analysis techniques evolve, pristine solar wind samples will be available to the scientific community in the decades to come. This article and a companion paper (Wiens et al. 2013, this issue) provide post-flight information necessary for the analysis of the Genesis array and foil solar wind samples and the Genesis solar wind ion concentrator samples, and thus serve to complement the Space Science Review volume, The Genesis Mission (v. 105, 2003).     

The Scientific Measurement System of the Gravity Recovery and Interior Laboratory (GRAIL) Mission

The Gravity Recovery and Interior Laboratory (GRAIL) mission to the Moon utilized an integrated scientific measurement system comprised of flight, ground, mission, and data system elements in order to meet the end-to-end performance required to achieve its scientific objectives. Modeling and simulation efforts were carried out early in the mission that influenced and optimized the design, implementation, and testing of these elements. Because the two prime scientific observables, range between the two spacecraft and range rates between each spacecraft and ground stations, can be affected by the performance of any element of the mission, we treated every element as part of an extended science instrument, a science system. All simulations and modeling took into account the design and configuration of each element to compute the expected performance and error budgets. In the process, scientific requirements were converted to engineering specifications that became the primary drivers for development and testing. Extensive simulations demonstrated that the scientific objectives could in most cases be met with significant margin. Errors are grouped into dynamic or kinematic sources and the largest source of non-gravitational error comes from spacecraft thermal radiation. With all error models included, the baseline solution shows that estimation of the lunar gravity field is robust against both dynamic and kinematic errors and a nominal field of degree 300 or better could be achieved according to the scaled Kaula rule for the Moon. The core signature is more sensitive to modeling errors and can be recovered with a small margin.     

Membrane structure active flatness control using genetic algorithm with online objective reweighting

This paper presents experimental studies on active flatness control of a membrane structure using genetic algorithm (GA). Different from the previous algorithms with a single objective function, a genetic algorithm with online objective reweighting capability is proposed here. This genetic algorithm implements an objective function that re-weights its objective online and the objective function is the flatness plus tension. The goal of the reweighting genetic algorithm is to allow for a computationally efficient method by which membrane wrinkling can be reduced, while maintaining a consideration for control authority. Experiments are conducted on our membrane test facility with two implementations: a static mutation rate and an adaptive mutation rate. The experimental results show that the proposed genetic algorithm is able to converge on flatness while reducing the total amount of tension required.     

Slot Length and Line-of-Sight Protection

The reasoning which led to the particular slot structure defined for the collision avoidance system specified by the Air Transportation Association Collision Avoidance System Technical Working Group is discussed. The objectives were to choose a slot length that would 1) minimize the probability of interference, 2) maximize the capacity of the information channel (as measured by the number of messages received in a unit of time). Interference can be manifested by the garbling of either the collision avoidance message or the synchronization signals.     

Optimization of low-energy resonant hopping transfers between planetary moons

In response to the scientific interest in Jupiter's Galilean moons, NASA and ESA have plans to send orbiting missions to Europa and Ganymede, respectively. The inter-moon transfers of the Jovian system offer obvious advantages in terms of scientific return, but are also challenging to design and optimize due in part to the large, often chaotic, sensitivities associated with repeated close encounters of the planetary moons. The approach outlined in this paper confronts this shortcoming by exploiting the multi-body dynamics with a patched three-body model to enable multiple “resonant-hopping” gravity assists. Initial conditions of unstable resonant orbits are pre-computed and provide starting points for the elusive initial guess associated with the highly nonlinear optimization problem. The core of the optimization algorithm relies on a fast and robust multiple-shooting technique to provide better controllability and reduce the sensitivities associated with the close approach trajectories. The complexity of the optimization problem is also reduced with the help of the Tisserand–Poincaré (T–P) graph that provides a simple way to target trajectories in the patched three-body problem. Preliminary numerical results of inter-moon transfers in the Jovian system are presented. For example, using only 59 m/s and 158 days, a spacecraft can transfer between a close resonant orbit of Ganymede and a close resonant orbit of Europa.     

Flight Operations in the New Millennium

New approaches are being studied for real-time interaction, and related supporting processes, with spacecraft and instruments in deep space. Spacecraft are evolving, improving in many ways, and generally becoming more robust. Operations is changing also, and will be more automated in the future. However, there is a challenge. Deep space missions are not all alike. The operations phases of discovery and exploration are an extension of the research that creates the mission; they are the time of obtaining results. This examines the historical role of flight operations and its evolving processes to develop an understanding of the operational methods that will be effective in the future. It takes people, equipment, software, space, and connectivity for operations success. A balance has to be struck between improving technology, gaining knowledge, automation, and realistic expectations. Finally, the recommended methods to gain efficiency in operations are system-wide services and shared resources. These common processes will meet the challenge of varied missions.     

Online identification and control of linearized aircraft dynamics

An approach for online identification and control that requires weaker excitation than the existing approaches based on least-squares schemes and closed-loop systems is examined. It uses multiple-objective optimization theory to resolve the conflict between identification and controller performance as they compete for the only available resource, the inputs to the aircraft. The approach is applied to a longitudinal model of a representative linearized high-performance aircraft model. Simulation results compare the final controller with a conventional gain-scheduled pitch command augmentation system. It is demonstrated that by allowing some control input to be given to the identification process, the controller's overall performance is improved