While many satellite operators are aware of the possibility of a collision between their satellite and another object in earth orbit, most seem unaware of the frequency of near misses occurring each day. Until recently, no service existed to advise satellite operators of an impending conjunction of a satellite payload with another satellite, putting the responsibility for determining these occurrences squarely on the satellite operator's shoulders. This problem has been further confounded by the lack of a timely, comprehensive data set of satellite orbital element sets and computationally efficient tools to provide predictions using industry-standard software. As a result, hundreds of conjunctions within 1 km occur each week, with little or no intervention, putting billions of dollars of space hardware at risk, along with their associated missions. As a service to the satellite operator community, the Center for Space Standards & Innovation (CSSI) offers SOCRATES-Satellite Orbital Conjunction Reports Assessing Threatening Encounters in Space. Twice each day, CSSI runs a list of all satellite payloads on orbit against a list of all objects on orbit using the catalog of all unclassified NORAD two-line element sets to look for conjunctions over the next seven days. The runs are made using STK/CAT-Satellite Tool Kit's Conjunction Analysis Tools-together with the NORAD SGP4 propagator in STK. This paper will discuss how SOCRATES works and how it can help satellite operators avoid undesired close approaches through advanced mission planning.

This paper describes a general method of identifying the key parameters of multiple-point contact-dynamics models and a robotics-based testbed for experimentally verifying the new method. Some of the current and future flight systems are required to make physical contact on orbit for on-orbit servicing such as docking, refueling, repairing, etc. Because of the high risks associated with contact operations, the design and operation of such a flight system must be thoroughly analyzed and verified in advance by hardware testing and/or high-fidelity computer simulation. Computer simulations are increasingly playing a major role in system verification because it is extreme difficult to test 6-DOF microgravity contact dynamics on the ground. However, the accuracy of computer simulation depends not only on the mathematical model (i.e., formulation, algorithms, and computer code) but also on the values of model parameters. It is, therefore, desirable to have a systematic method which can identify multiple model parameters directly from routine physical tests of the contact components. The robotics-based experiment testbed introduced in this paper is specially designed to test and verify such a method of identifying contact parameters. The method is capable of identifying the key stiffness, damping, and friction parameters of a contact dynamics model all together from hardware test of contacting components having complicated geometries and multiple contacts. It can also be used to extract contact-dynamics model parameters of a dynamic system from its routine test of complex contact hardware. The paper discusses the major design requirements of this experimental testbed and how they are met by the specific design of the system.

This paper addresses contact dynamics simulations for a Hubble Robotic Servicing mission. First, the modeling of the robotic system is introduced. The simulation models of the robotic system include: flexible body dynamics, control system models and geometric models of the contacting bodies. These models are incorporated into MDA's simulation facility, the multibody dynamics simulator "Space Station Portable Operations Training Simulator (SPOTS)". Three contact dynamics simulation examples of the robotic servicing operations are presented: (1) capture of the Hubble Space Telescope, (2) berthing the Hubble Space Telescope to the Hubble Robotic Vehicle and (3) inserting the Wide Field Camera into the Hubble Space Telescope.

This paper presents the concept of a cable-manipulator based 6-DOF hardware-in-the-loop (HIL) dynamics simulation system for testing and verification of microgravity contact-dynamics behavior of a space system. It then focuses on the inverse dynamics problem of the 6-DOF cable-driven manipulator which is designed for the simulation system. Accurate modeling and solution of the inverse dynamics is a key requirement for the control and high-fidelity performance of the complex simulation system. The inverse dynamics problem is solved completely under the basic operational conditions of a cable manipulator - all the cables must be always in tension for any possible end-effector motion of the manipulator. It is the first time that a systematic method of determining whether or not the inverse dynamics problem has a solution is proposed with full mathematical proof. Based upon this proven method, two numerical examples are presented to demonstrate the inverse dynamics solution of a 6-DOF cable manipulator. The study results support the feasibility of using such a manipulator for hardware-in-the-loop simulation of microgravity contact-dynamics.

The MARSIS antenna booms are constructed using lenticular hinges between straight boom segments in a novel design which allows the booms to be extremely lightweight while retaining a high stiffness and well defined structural properties once they are deployed. Lenticular hinges are elegant in form but are complicated to model as they deploy dynamically and require highly specialized nonlinear techniques founded on carefully measured mechanical properties. Results from component level testing were incorporated into a highly specialized ADAMS model which employed an automated damping algorithm to account for the discontinuous boom lengths formed during the deployment. Additional models with more limited capabilities were also developed in both DADS and ABAQUS to verify the ADAMS model computations and to help better define the numerical behavior of the models at the component and system levels. A careful comparison is made between the ADAMS and DADS models in a series of progressive steps in order to verify their numerical results. Different trade studies considered in the model development are outlined to demonstrate a suitable level of model fidelity. Some model sensitivities to various parameters are explored using subscale and full system models. Finally, some full system DADS models are exercised to illustrate the limitations of traditional modeling techniques for variable geometry systems which were overcome in the ADAMS model.

This paper introduces a 3-step approach for the identification of a linear structure that is controlled by nonlinear damping devices. First, the structure with the integrated nonlinear damper is subjected to random vibration test and the frequency response function (FRF) of the structure is calculated from the input-output data of the physical system. Based on the frequency domain data, a state space model is then estimated using a recently developed FRF curve-fitting technique that is designed especially for lightly damped structures with control inputs. Finally an iterative process is used to optimize the model performance in the time domain and an integrated model of the nonlinearly controlled structure is derived by interconnecting the structure model with that of the nonlinear damper. The complete approach is illustrated by the modeling of a base-isolated structure controlled by a magnetorheological (MR) fluid damper.

There is a critical need, not just in the Department of Defense (DOD) but the entire space industry, to reduce the development time and overall cost of satellite missions. To that end, the DOD is actively pursuing the capability to reduce the deployment time of a new system from years to weeks or even days. The goal is to provide the advantages space affords not just to the strategic planner but also to the battlefield commanders. One of the most challenging aspects of this problem is the satellite's thermal control system (TCS). Traditionally the TCS must be vigorously designed, analyzed, tested, and optimized from the ground up for every satellite mission. This "reinvention of the wheel" is costly and time intensive. The next generation satellite TCS must be modular and scalable in order to cover a wide range of applications, orbits, and mission requirements. To meet these requirements a robust thermal control system utilizing forced convection thermal switches was investigated. The problem was investigated in two separate stages. The first focused on the overall design of the bus. The second stage focused on the overarching bus architecture and the design impacts of employing a thermal switch based TCS design. For the hot case, the fan provided additional cooling to increase the heat transfer rate of the subsystem. During the cold case, the result was a significant reduction in survival heater power.

There has been a recent increase in emphasis on small satellites because of their low cost, short development times, relative simplicity, and cost efficiency. However, these satellites do have drawbacks. Their small size results in small surface areas which often translate into thermal and power constraints. A small satellite may not have enough surface area for radiators and/or solar panels. The radiators are used to release internal heat during hot environments, and solar panels create necessary power for the heaters during cold environments. Because of the surface area and power limitations, a passive thermal design was then selected for the Formation Autonomy Spacecraft with Thrust, Relative Navigation, Attitude, and Crosslink Program (FASTRAC) twin satellites, built by students at the University of Texas at Austin. Thermal cycling and thermal analysis were performed. The thermal cycling was done in Chamber-N at Johnson Space Center, Texas, using worst case hot and cold scenarios. The thermal analysis was conducted using Finite Elements (FE), and the results were compared to the test data and validated. FASTRAC is planned to be in a LEO orbit which ranges between 300km and 500km in altitude. The orbits were then simulated to determine the characteristics of the LEO orbits in which FASTRAC can survive.

Thermal design for space systems is an iterative process that balances the temperature requirements for all mission phases with the available resources. Secondary payloads often have to be designed for a wide range of conditions available on various launch platforms, without the benefit of additional resources such as power or thermal shielding. This paper will discuss the thermal design, analysis, and thermal vacuum testing of a small satellite payload that was initially intended for launch from the US Space Shuttle and eventually launched on the EELV Heavy demonstration in December 2004.

Automated rendezvous and docking (AR&D) is an activity that is vital to NASA's future endeavors. A team of NASA engineers has created a test bed that allows sensor systems to be tested but also incorporates operator interfaces in order to evaluate the information required by an operator to oversee automated systems. Several sensors have been run through the test bed and more are scheduled in the future. The test bed will be described along with the operator interfaces and results of sensor testing.

The Contact Dynamics Simulation Laboratory (CDSL) of the Marshall Space Flight Center provides for refined hardware-in-the-loop real-time simulation of docking and berthing mechanisms and associated control systems. This facility is employed to verify the performance of docking/berthing mechanisms during Earth-orbit operations, determine the capture envelope of docking/berthing devices, measure contact loads at vehicle interfaces, and evaluate visual cues for man-in-the-loop operations. The CDSL has developed test verified analytical models of such systems as the International Space Station (ISS) Common Berthing Mechanism (CBM) and the Hubble Space Telescope (HST) Three Point Docking Mechanism. This paper will describe the modeling and test techniques employed at the CDSL and present results from recent programs.

To perform realistic demonstrations of autonomous docking maneuvers using micro-satellites, the MIT Space Systems Laboratory (SSL) developed a miniature universal docking port along with an optical sensing system for relative state estimation. The docking port has an androgynous design and is universal since any two identical ports can be connected together. After a rigid connection is made, it is capable of passing electrical loads between the connected micro-satellites. The optical sensor uses a set of infrared LED's, a miniature CCD-based video camera, and an Extended Kalman Filter to determine the six relative degrees of freedom of the docking satellite. The SPHERES testbed, also developed by the MIT SSL, was used to demonstrate the integrated docking port and sensor system. This study focuses on the development of the optical docking sensor, and presents test results collected to date during fully autonomous docking experiments performed at the MIT SSL 2-D laboratory. Tests were performed to verify the validity of the docking sensor by taking measurements at known distances. These results give an estimate of the sensor accuracy, and are compared with a theoretical model to understand the sources of error in the state measurements.

The rendezvous problem was analyzed by the guidance method of qualitative differential game for space vehicle while the target is maneuvering, based upon the perturbation equation of orbital elements. By taking into account of fuel assumption for rendezvous control process under radial thrust, we introduced the performance index for quantitative differential game into Hamiltonian function to resolve the guidance problems for maneuvering space vehicle for orbit approaching, and gave a set of theoretical solution, barriers and an example for simulation. From mathematically simulation it is verified that there are the barriers in the method for the target maneuvering.

Optical communications systems are vital to allow high speed satellite-to-satellite and satellite-to-ground-based communication links with low power consumption and low weight. To predict the performance of such systems it is essential to have an accurate simulation model which allows to predict the experimental results. We have implemented a coherent optical communications system which can be used for ultra long free-space distances. It incorporates a challenging optical phase lock loop (PLL). We also developed a simulation model for this advanced optical telecommunication system. It is shown that the experimental and numerical results obtained are in excellent agreement. By changing the parameters of the simulation model we can predict which of those parameters are most important to achieve a reliable high speed intersatellite optical link over a long free-space distance. One of the key parameters is the performance of our optical PLL. This is most important for systems which use the highly sensitive phase-shift keying (PSK) modulation format. Our developed optical PLL with a linewidth of as low as 130Hz shows excellent results both in simulation and experiments.

This paper presents a novel analytical method for comparing simulation data with measured data, with the goal of proving Equivalence and Consistency between the model and the real data. Our method overcomes the problems of disparity in the inputs to the simulation and varying number of parameters between the simulation and the measured flight data. Our method derives analytical Data Models that are analyzed in frequency space, yielding quantitative assessment values for the model performance relative to the measured data. The model output for a sensor and its "real world" measured data are collected for comparison.

The extended use of microelectromechanical systems (MEMS) in the development of new microinstrumentation for aerospatial applications, which combine extreme sensitivity, accuracy and compactness, introduced the need to simplify their design process in order to reduce the design time and cost. The recent apparition of analogue and mixed signal extensions of hardware descriptions languages (VHDL-AMS, Verilog-AMS and SystemC-AMS) permits to co-simulate the HDL (VHDL and Verilog) design models for the digital signal processing and communication circuitry with behavioral models for the non digital parts (analog and mixed signal processing, RF circuitry and MEMS components). Since the beginning of the microinstrumentation design process the modeling and simulation could help to define better the specifications and in the architecture selection and in the SoC design process in a more realistic environment. We will present our experience in the application of these languages in the design of microinstruments by using behavioral modeling of MEMS.