Assessment of the Use of Nanofluids in Spacecraft Active Thermal Control Systems

Science.gov (United States)

Ungar, Eugene K.; Erickson, Lisa R.

2011-01-01

The addition of metallic nanoparticles to a base heat transfer fluid can dramatically increase its thermal conductivity. These nanofluids have been shown to have advantages in some heat transport systems. Their enhanced properties can allow lower system volumetric flow rates and can reduce the required pumping power. Nanofluids have been suggested for use as working fluids for spacecraft Active Thermal Control Systems (ATCSs). However, there are no studies showing the end-to-end effect of nanofluids on the design and performance of spacecraft ATCSs. In the present work, a parametric study is performed to assess the use of nanofluids in a spacecraft ATCSs. The design parameters of the current Orion capsule and the tabulated thermophysical properties of nanofluids are used to assess the possible benefits of nanofluids and how their incorporation affects the overall design of a spacecraft ATCS. The study shows that the unique system and component-level design parameters of spacecraft ATCSs render them best suited for pure working fluids. The addition of nanoparticles to typical spacecraft thermal control working fluids actually results in an increase in the system mass and required pumping power.

Dynamic Isotope Power System (DIPS) Applications Study. Volume II. Nuclear Integrated Multimission Spacecraft (NIMS) design definition. Final report

International Nuclear Information System (INIS)

1979-11-01

The design requirements for the Nuclear Integrated Multimission Spacecraft. (NIMS) are discussed in detail. The requirements are a function of mission specifications, payload, control system requirements, electric system specifications, and cost limitations

Hybrid spacecraft attitude control system

Directory of Open Access Journals (Sweden)

Renuganth Varatharajoo

2016-02-01

Full Text Available The hybrid subsystem design could be an attractive approach for futurespacecraft to cope with their demands. The idea of combining theconventional Attitude Control System and the Electrical Power System ispresented in this article. The Combined Energy and Attitude ControlSystem (CEACS consisting of a double counter rotating flywheel assemblyis investigated for small satellites in this article. Another hybrid systemincorporating the conventional Attitude Control System into the ThermalControl System forming the Combined Attitude and Thermal ControlSystem (CATCS consisting of a "fluid wheel" and permanent magnets isalso investigated for small satellites herein. The governing equationsdescribing both these novel hybrid subsystems are presented and theironboard architectures are numerically tested. Both the investigated novelhybrid spacecraft subsystems comply with the reference missionrequirements.The hybrid subsystem design could be an attractive approach for futurespacecraft to cope with their demands. The idea of combining theconventional Attitude Control System and the Electrical Power System ispresented in this article. The Combined Energy and Attitude ControlSystem (CEACS consisting of a double counter rotating flywheel assemblyis investigated for small satellites in this article. Another hybrid systemincorporating the conventional Attitude Control System into the ThermalControl System forming the Combined Attitude and Thermal ControlSystem (CATCS consisting of a "fluid wheel" and permanent magnets isalso investigated for small satellites herein. The governing equationsdescribing both these novel hybrid subsystems are presented and theironboard architectures are numerically tested. Both the investigated novelhybrid spacecraft subsystems comply with the reference missionrequirements.

Electrical design for origami solar panels and a small spacecraft test mission

Science.gov (United States)

Drewelow, James; Straub, Jeremy

2017-05-01

Efficient power generation is crucial to the design of spacecraft. Mass, volume, and other limitations prevent the use of traditional spacecraft support structures from being suitable for the size of solar array required for some missions. Folding solar panel / panel array systems, however, present a number of design challenges. This paper considers the electrical design of an origami system. Specifically, it considers how to provide low impedance, durable channels for the generated power and the electrical aspects of the deployment system and procedure. The ability to dynamically reconfigure the electrical configuration of the solar cells is also discussed. Finally, a small satellite test mission to demonstrate the technology is proposed, before concluding.

Multi-kilowatt modularized spacecraft power processing system development

International Nuclear Information System (INIS)

Andrews, R.E.; Hayden, J.H.; Hedges, R.T.; Rehmann, D.W.

1975-07-01

A review of existing information pertaining to spacecraft power processing systems and equipment was accomplished with a view towards applicability to the modularization of multi-kilowatt power processors. Power requirements for future spacecraft were determined from the NASA mission model-shuttle systems payload data study which provided the limits for modular power equipment capabilities. Three power processing systems were compared to evaluation criteria to select the system best suited for modularity. The shunt regulated direct energy transfer system was selected by this analysis for a conceptual design effort which produced equipment specifications, schematics, envelope drawings, and power module configurations

Propulsion Trade Studies for Spacecraft Swarm Mission Design

Science.gov (United States)

Dono, Andres; Plice, Laura; Mueting, Joel; Conn, Tracie; Ho, Michael

2018-01-01

Spacecraft swarms constitute a challenge from an orbital mechanics standpoint. Traditional mission design involves the application of methodical processes where predefined maneuvers for an individual spacecraft are planned in advance. This approach does not scale to spacecraft swarms consisting of many satellites orbiting in close proximity; non-deterministic maneuvers cannot be preplanned due to the large number of units and the uncertainties associated with their differential deployment and orbital motion. For autonomous small sat swarms in LEO, we investigate two approaches for controlling the relative motion of a swarm. The first method involves modified miniature phasing maneuvers, where maneuvers are prescribed that cancel the differential delta V of each CubeSat's deployment vector. The second method relies on artificial potential functions (APFs) to contain the spacecraft within a volumetric boundary and avoid collisions. Performance results and required delta V budgets are summarized, indicating that each method has advantages and drawbacks for particular applications. The mini phasing maneuvers are more predictable and sustainable. The APF approach provides a more responsive and distributed performance, but at considerable propellant cost. After considering current state of the art CubeSat propulsion systems, we conclude that the first approach is feasible, but the modified APF method of requires too much control authority to be enabled by current propulsion systems.

The Earth Observing System AM Spacecraft - Thermal Control Subsystem

Science.gov (United States)

Chalmers, D.; Fredley, J.; Scott, C.

1993-01-01

Mission requirements for the EOS-AM Spacecraft intended to monitor global changes of the entire earth system are considered. The spacecraft is based on an instrument set containing the Advanced Spaceborne Thermal Emission and Reflection radiometer (ASTER), Clouds and Earth's Radiant Energy System (CERES), Multiangle Imaging Spectro-Radiometer (MISR), Moderate-Resolution Imaging Spectrometer (MODIS), and Measurements of Pollution in the Troposphere (MOPITT). Emphasis is placed on the design, analysis, development, and verification plans for the unique EOS-AM Thermal Control Subsystem (TCS) aimed at providing the required environments for all the onboard equipment in a densely packed layout. The TCS design maximizes the use of proven thermal design techniques and materials, in conjunction with a capillary pumped two-phase heat transport system for instrument thermal control.

Definition of the topological structure of the automatic control system of spacecrafts

International Nuclear Information System (INIS)

KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" data-affiliation=" (Siberian State Aerospace University named after Academician M.F.Reshetnev 31 KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" >Zelenkov, P V; KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" data-affiliation=" (Siberian State Aerospace University named after Academician M.F.Reshetnev 31 KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" >Karaseva, M V; KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" data-affiliation=" (Siberian State Aerospace University named after Academician M.F.Reshetnev 31 KrasnoyarskiyRabochiy prospect, Krasnoyarsk, 660014 (Russian Federation))" >Tsareva, E A; Tsarev, R Y

2015-01-01

The paper considers the problem of selection the topological structure of the automated control system of spacecrafts. The integer linear model of mathematical programming designed to define the optimal topological structure for spacecraft control is proposed. To solve the determination problem of topological structure of the control system of spacecrafts developed the procedure of the directed search of some structure variants according to the scheme 'Branch and bound'. The example of the automated control system of spacecraft development included the combination of ground control stations, managing the spacecraft of three classes with a geosynchronous orbit with constant orbital periods is presented

Time maintenance system for the BMDO MSX spacecraft

Science.gov (United States)

Hermes, Martin J.

1994-01-01

The Johns Hopkins University Applied Physics Laboratory (APL) is responsible for designing and implementing a clock maintenance system for the Ballistic Missile Defense Organizations (BMDO) Midcourse Space Experiment (MSX) spacecraft. The MSX spacecraft has an on-board clock that will be used to control execution of time-dependent commands and to time tag all science and housekeeping data received from the spacecraft. MSX mission objectives have dictated that this spacecraft time, UTC(MSX), maintain a required accuracy with respect to UTC(USNO) of +/- 10 ms with a +/- 1 ms desired accuracy. APL's atomic time standards and the downlinked spacecraft time were used to develop a time maintenance system that will estimate the current MSX clock time offset during an APL pass and make estimates of the clock's drift and aging using the offset estimates from many passes. Using this information, the clock's accuracy will be maintained by uplinking periodic clock correction commands. The resulting time maintenance system is a combination of offset measurement, command/telemetry, and mission planning hardware and computing assets. All assets provide necessary inputs for deciding when corrections to the MSX spacecraft clock must be made to maintain its required accuracy without inhibiting other mission objectives. The MSX time maintenance system is described as a whole and the clock offset measurement subsystem, a unique combination of precision time maintenance and measurement hardware controlled by a Macintosh computer, is detailed. Simulations show that the system estimates the MSX clock offset to less than+/- 33 microseconds.

Spacecraft momentum control systems

CERN Document Server

Leve, Frederick A; Peck, Mason A

2015-01-01

The goal of this book is to serve both as a practical technical reference and a resource for gaining a fuller understanding of the state of the art of spacecraft momentum control systems, specifically looking at control moment gyroscopes (CMGs). As a result, the subject matter includes theory, technology, and systems engineering. The authors combine material on system-level architecture of spacecraft that feature momentum-control systems with material about the momentum-control hardware and software. This also encompasses material on the theoretical and algorithmic approaches to the control of space vehicles with CMGs. In essence, CMGs are the attitude-control actuators that make contemporary highly agile spacecraft possible. The rise of commercial Earth imaging, the advances in privately built spacecraft (including small satellites), and the growing popularity of the subject matter in academic circles over the past decade argues that now is the time for an in-depth treatment of the topic. CMGs are augmented ...

Design issues of the piezo motor for the spacecraft reflector control system

Directory of Open Access Journals (Sweden)

Azin Anton

2018-01-01

Full Text Available Creation of large-size reflectors for spacecrafts is a topical issue for the space industry. The accuracy of the reflecting surface form and the structure weight are the main criteria for the reflector design. The accuracy of the reflecting surface form during a long-term operation is provided by adjustment when using piezoelectric motors in the reflector design. These motors have small weight-size parameters and can reach great torque values. The piezo motor is a distributed mechanical-acoustic oscillation system. Mechanical-acoustic oscillations are generated in the piezo motor by a PZT-stack and transmitted to an oscillator element, and then from the oscillator element to a load action element. At high frequencies, when dimensions of the oscillator are proportionate to the wavelength, the energy is transmitted by means of acoustic waves. In this case, mechanical waves practically are not involved in the energy transmission process. This thesis shows a method for selecting the material of a mechanical-acoustic oscillation system according to the efficiency of the acoustic energy transmission via a piezoelectric layered structure.

A design for a reusable water-based spacecraft known as the spacecoach

CERN Document Server

McConnell, Brian

2016-01-01

 Based on components already in existence, this manual details a reference design for an interplanetary spacecraft that is simple, durable, fully reusable and comprised mostly of water. Using such an accessible material leads to a spacecraft architecture that is radically simpler, safer and cheaper than conventional capsule based designs. If developed, the potential affordability of the design will substantially open all of the inner solar system to human exploration. A spacecraft that is comprised mostly of water will be much more like a living cell or a terrarium than a conventional rocket and capsule design. It will use water for many purposes before it is superheated in electric engines for propulsion, purposes which include radiation shielding, heat management, basic life support, crew consumption and comfort. The authors coined the term "spacecoaches" to describe them, as an allusion to the Prairie Schooners of the Old West, which were simple, rugged, and could live off the land.

Preliminary thermal design of the COLD-SAT spacecraft

Science.gov (United States)

Arif, Hugh

1991-01-01

The COLD-SAT free-flying spacecraft was to perform experiments with LH2 in the cryogenic fluid management technologies of storage, supply and transfer in reduced gravity. The Phase A preliminary design of the Thermal Control Subsystem (TCS) for the spacecraft exterior and interior surfaces and components of the bus subsystems is described. The TCS was composed of passive elements which were augmented with heaters. Trade studies to minimize the parasitic heat leakage into the cryogen storage tanks are described. Selection procedure for the thermally optimum on-orbit spacecraft attitude was defined. TRASYS-2 and SINDA'85 verification analysis was performed on the design and the results are presented.

Electronic Systems for Spacecraft Vehicles: Required EDA Tools

Science.gov (United States)

Bachnak, Rafic

1999-01-01

The continuous increase in complexity of electronic systems is making the design and manufacturing of such systems more challenging than ever before. As a result, designers are finding it impossible to design efficient systems without the use of sophisticated Electronic Design Automation (EDA) tools. These tools offer integrated simulation of the electrical, mechanical, and manufacturing functions and lead to a correct by design methodology. This report identifies the EDA tools that would be needed to design, analyze, simulate, and evaluate electronic systems for spacecraft vehicles. In addition, the report presents recommendations to enhance the current JSC electronic design capabilities. This includes cost information and a discussion as to the impact, both positive and negative, of implementing the recommendations.

SHARP: A multi-mission AI system for spacecraft telemetry monitoring and diagnosis

Science.gov (United States)

Lawson, Denise L.; James, Mark L.

1989-01-01

The Spacecraft Health Automated Reasoning Prototype (SHARP) is a system designed to demonstrate automated health and status analysis for multi-mission spacecraft and ground data systems operations. Telecommunications link analysis of the Voyager II spacecraft is the initial focus for the SHARP system demonstration which will occur during Voyager's encounter with the planet Neptune in August, 1989, in parallel with real-time Voyager operations. The SHARP system combines conventional computer science methodologies with artificial intelligence techniques to produce an effective method for detecting and analyzing potential spacecraft and ground systems problems. The system performs real-time analysis of spacecraft and other related telemetry, and is also capable of examining data in historical context. A brief introduction is given to the spacecraft and ground systems monitoring process at the Jet Propulsion Laboratory. The current method of operation for monitoring the Voyager Telecommunications subsystem is described, and the difficulties associated with the existing technology are highlighted. The approach taken in the SHARP system to overcome the current limitations is also described, as well as both the conventional and artificial intelligence solutions developed in SHARP.

Solar Power System Options for the Radiation and Technology Demonstration Spacecraft

Science.gov (United States)

Kerslake, Thomas W.; Haraburda, Francis M.; Riehl, John P.

2000-01-01

The Radiation and Technology Demonstration (RTD) Mission has the primary objective of demonstrating high-power (10 kilowatts) electric thruster technologies in Earth orbit. This paper discusses the conceptual design of the RTD spacecraft photovoltaic (PV) power system and mission performance analyses. These power system studies assessed multiple options for PV arrays, battery technologies and bus voltage levels. To quantify performance attributes of these power system options, a dedicated Fortran code was developed to predict power system performance and estimate system mass. The low-thrust mission trajectory was analyzed and important Earth orbital environments were modeled. Baseline power system design options are recommended on the basis of performance, mass and risk/complexity. Important findings from parametric studies are discussed and the resulting impacts to the spacecraft design and cost.

42: An Open-Source Simulation Tool for Study and Design of Spacecraft Attitude Control Systems

Science.gov (United States)

Stoneking, Eric

2018-01-01

Simulation is an important tool in the analysis and design of spacecraft attitude control systems. The speaker will discuss the simulation tool, called simply 42, that he has developed over the years to support his own work as an engineer in the Attitude Control Systems Engineering Branch at NASA Goddard Space Flight Center. 42 was intended from the outset to be high-fidelity and powerful, but also fast and easy to use. 42 is publicly available as open source since 2014. The speaker will describe some of 42's models and features, and discuss its applicability to studies ranging from early concept studies through the design cycle, integration, and operations. He will outline 42's architecture and share some thoughts on simulation development as a long-term project.

Ulysses spacecraft control and monitoring system

Science.gov (United States)

Hamer, P. A.; Snowden, P. J.

1991-01-01

The baseline Ulysses spacecraft control and monitoring system (SCMS) concepts and the converted SCMS, residing on a DEC/VAX 8350 hardware, are considered. The main functions of the system include monitoring and displaying spacecraft telemetry, preparing spacecraft commands, producing hard copies of experimental data, and archiving spacecraft telemetry. The SCMS system comprises over 20 subsystems ranging from low-level utility routines to the major monitoring and control software. These in total consist of approximately 55,000 lines of FORTRAN source code and 100 VMS command files. The SCMS major software facilities are described, including database files, telemetry processing, telecommanding, archiving of data, and display of telemetry.

Design of RTPV generators integrated with new millennium spacecraft for outer solar system

International Nuclear Information System (INIS)

Schock, A.; Or, C.; Kumar, V.

1996-01-01

The National Aeronautics and Space Administration's recently inaugurated New Millennium program, with its emphasis on miniaturized spacecraft, has generated interest in a low-power (10- to 30-watt), low-mass, high-efficiency RTPV (Radioisotope Thermophotovoltaic) power system. This led to a Department of Energy (DOE)-sponsored design study of such a system. A 75-watt design employed two 250-watt General Purpose Heat Source (GPHS) modules that DOE had previously developed and safety-qualified for various space missions. These modules were too large for the small RTPVs described in this paper. To minimize the need for new development and safety verification studies, derivative designs for 125-watt and 62.5-watt heat source modules containing identical fuel pellets, clads, impact shell, and thermal insulation were generated along with a novel heat source support scheme to reduce the heat losses through the structural supports, and a new and much simpler radiator structure, employing no honeycombs or heat pipes. Previous RTPV study had been based on the use of GaSb PV cells and spectrally selective IR filters. Because of the very encouraging results of system design studies, in the fall of 1994 an experimental program was initiated to develop improved filters and cells, to demonstrate how much improvement can actually be achieved. First priority was given to filter improvements, because our system studies indicated that improved filters would have a much greater effect on system performance than cell improvements. By September 1995 about 94% of the filter performance improvement projected in 1993 had been achieved. (Abstract Truncated)

SHARP: A multi-mission artificial intelligence system for spacecraft telemetry monitoring and diagnosis

Science.gov (United States)

Lawson, Denise L.; James, Mark L.

1989-01-01

The Spacecraft Health Automated Reasoning Prototype (SHARP) is a system designed to demonstrate automated health and status analysis for multi-mission spacecraft and ground data systems operations. Telecommunications link analysis of the Voyager 2 spacecraft is the initial focus for the SHARP system demonstration which will occur during Voyager's encounter with the planet Neptune in August, 1989, in parallel with real time Voyager operations. The SHARP system combines conventional computer science methodologies with artificial intelligence techniques to produce an effective method for detecting and analyzing potential spacecraft and ground systems problems. The system performs real time analysis of spacecraft and other related telemetry, and is also capable of examining data in historical context. A brief introduction is given to the spacecraft and ground systems monitoring process at the Jet Propulsion Laboratory. The current method of operation for monitoring the Voyager Telecommunications subsystem is described, and the difficulties associated with the existing technology are highlighted. The approach taken in the SHARP system to overcome the current limitations is also described, as well as both the conventional and artificial intelligence solutions developed in SHARP.

Leo Spacecraft Charging Design Guidelines: A Proposed NASA Standard

Science.gov (United States)

Hillard, G. B.; Ferguson, D. C.

2004-01-01

Over the past decade, Low Earth Orbiting (LEO) spacecraft have gradually required ever-increasing power levels. As a rule, this has been accomplished through the use of high voltage systems. Recent failures and anomalies on such spacecraft have been traced to various design practices and materials choices related to the high voltage solar arrays. NASA Glenn has studied these anomalies including plasma chamber testing on arrays similar to those that experienced difficulties on orbit. Many others in the community have been involved in a comprehensive effort to understand the problems and to develop practices to avoid them. The NASA Space Environments and Effects program, recognizing the timeliness of this effort, commissioned and funded a design guidelines document intended to capture the current state of understanding. This document, which was completed in the spring of 2003, has been submitted as a proposed NASA standard. We present here an overview of this document and discuss the effort to develop it as a NASA standard.

Modeling Temporal Processes in Early Spacecraft Design: Application of Discrete-Event Simulations for Darpa's F6 Program

Science.gov (United States)

Dubos, Gregory F.; Cornford, Steven

2012-01-01

While the ability to model the state of a space system over time is essential during spacecraft operations, the use of time-based simulations remains rare in preliminary design. The absence of the time dimension in most traditional early design tools can however become a hurdle when designing complex systems whose development and operations can be disrupted by various events, such as delays or failures. As the value delivered by a space system is highly affected by such events, exploring the trade space for designs that yield the maximum value calls for the explicit modeling of time.This paper discusses the use of discrete-event models to simulate spacecraft development schedule as well as operational scenarios and on-orbit resources in the presence of uncertainty. It illustrates how such simulations can be utilized to support trade studies, through the example of a tool developed for DARPA's F6 program to assist the design of "fractionated spacecraft".

Estimation Model of Spacecraft Parameters and Cost Based on a Statistical Analysis of COMPASS Designs

Science.gov (United States)

Gerberich, Matthew W.; Oleson, Steven R.

2013-01-01

The Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) team at Glenn Research Center has performed integrated system analysis of conceptual spacecraft mission designs since 2006 using a multidisciplinary concurrent engineering process. The set of completed designs was archived in a database, to allow for the study of relationships between design parameters. Although COMPASS uses a parametric spacecraft costing model, this research investigated the possibility of using a top-down approach to rapidly estimate the overall vehicle costs. This paper presents the relationships between significant design variables, including breakdowns of dry mass, wet mass, and cost. It also develops a model for a broad estimate of these parameters through basic mission characteristics, including the target location distance, the payload mass, the duration, the delta-v requirement, and the type of mission, propulsion, and electrical power. Finally, this paper examines the accuracy of this model in regards to past COMPASS designs, with an assessment of outlying spacecraft, and compares the results to historical data of completed NASA missions.

Printed Spacecraft Separation System

Energy Technology Data Exchange (ETDEWEB)

Dehoff, Ryan R [ORNL; Holmans, Walter [Planetary Systems Corporation

2016-10-01

In this project Planetary Systems Corporation proposed utilizing additive manufacturing (3D printing) to manufacture a titanium spacecraft separation system for commercial and US government customers to realize a 90% reduction in the cost and energy. These savings were demonstrated via “printing-in” many of the parts and sub-assemblies into one part, thus greatly reducing the labor associated with design, procurement, assembly and calibration of mechanisms. Planetary Systems Corporation redesigned several of the components of the separation system based on additive manufacturing principles including geometric flexibility and the ability to fabricate complex designs, ability to combine multiple parts of an assembly into a single component, and the ability to optimize design for specific mechanical property targets. Shock absorption was specifically targeted and requirements were established to attenuate damage to the Lightband system from shock of initiation. Planetary Systems Corporation redesigned components based on these requirements and sent the designs to Oak Ridge National Laboratory to be printed. ORNL printed the parts using the Arcam electron beam melting technology based on the desire for the parts to be fabricated from Ti-6Al-4V based on the weight and mechanical performance of the material. A second set of components was fabricated from stainless steel material on the Renishaw laser powder bed technology due to the improved geometric accuracy, surface finish, and wear resistance of the material. Planetary Systems Corporation evaluated these components and determined that 3D printing is potentially a viable method for achieving significant cost and savings metrics.

SEQ-POINTER: Next generation, planetary spacecraft remote sensing science observation design tool

Science.gov (United States)

Boyer, Jeffrey S.

1994-11-01

Since Mariner, NASA-JPL planetary missions have been supported by ground software to plan and design remote sensing science observations. The software used by the science and sequence designers to plan and design observations has evolved with mission and technological advances. The original program, PEGASIS (Mariners 4, 6, and 7), was re-engineered as POGASIS (Mariner 9, Viking, and Mariner 10), and again later as POINTER (Voyager and Galileo). Each of these programs were developed under technological, political, and fiscal constraints which limited their adaptability to other missions and spacecraft designs. Implementation of a multi-mission tool, SEQ POINTER, under the auspices of the JPL Multimission Operations Systems Office (MOSO) is in progress. This version has been designed to address the limitations experienced on previous versions as they were being adapted to a new mission and spacecraft. The tool has been modularly designed with subroutine interface structures to support interchangeable celestial body and spacecraft definition models. The computational and graphics modules have also been designed to interface with data collected from previous spacecraft, or on-going observations, which describe the surface of each target body. These enhancements make SEQ POINTER a candidate for low-cost mission usage, when a remote sensing science observation design capability is required. The current and planned capabilities of the tool will be discussed. The presentation will also include a 5-10 minute video presentation demonstrating the capabilities of a proto-Cassini Project version that was adapted to test the tool. The work described in this abstract was performed by the Jet Propulsion Laboratory, California Institute of Technology, under contract to the National Aeronautics and Space Administration.

System Critical Design Audit (CDA). Books 1, 2 and 3; [Small Satellite Technology Initiative (SSTI Lewis Spacecraft Program)

Science.gov (United States)

1995-01-01

Small Satellite Technology Initiative (SSTI) Lewis Spacecraft Program is evaluated. Spacecraft integration, test, launch, and spacecraft bus are discussed. Payloads and technology demonstrations are presented. Mission data management system and ground segment are also addressed.

Multi-objective optimisation for spacecraft design for demise and survivability

OpenAIRE

Trisolini, Mirko; Colombo, Camilla; Lewis, Hugh

2017-01-01

The paper presents the development of a multi-objective optimisation framework to study the effects that preliminary design choices have on the demisability and the survivability of a spacecraft. Building a spacecraft such that most of it will demise during the re-entry through design-for-demise strategies may lead to design that are more vulnerable to space debris impacts, thus compromising the reliability of the mission. The two models developed to analyse the demisability and the survivabi...

Application of Space Environmental Observations to Spacecraft Pre-Launch Engineering and Spacecraft Operations

Science.gov (United States)

Barth, Janet L.; Xapsos, Michael

2008-01-01

This presentation focuses on the effects of the space environment on spacecraft systems and applying this knowledge to spacecraft pre-launch engineering and operations. Particle radiation, neutral gas particles, ultraviolet and x-rays, as well as micrometeoroids and orbital debris in the space environment have various effects on spacecraft systems, including degradation of microelectronic and optical components, physical damage, orbital decay, biasing of instrument readings, and system shutdowns. Space climate and weather must be considered during the mission life cycle (mission concept, mission planning, systems design, and launch and operations) to minimize and manage risk to both the spacecraft and its systems. A space environment model for use in the mission life cycle is presented.

Vibration Antiresonance Design for a Spacecraft Multifunctional Structure

OpenAIRE

Li, Dong-Xu; Liu, Wang; Hao, Dong

2017-01-01

Spacecraft must withstand rigorous mechanical environment experiences such as acceleration, noise, vibration, and shock during the process of launching, satellite-vehicle separation, and so on. In this paper, a new spacecraft multifunctional structure concept designed by us is introduced. The multifunctional structure has the functions of not only load bearing, but also vibration reduction, energy source, thermal control, and so on, and we adopt a series of viscoelastic parts as connections b...

Materials and structures technology insertion into spacecraft systems: Successes and challenges

Science.gov (United States)

Rawal, Suraj

2018-05-01

Over the last 30 years, significant advancements have led to the use of multifunctional materials and structures technologies in spacecraft systems. This includes the integration of adaptive structures, advanced composites, nanotechnology, and additive manufacturing technologies. Development of multifunctional structures has been directly influenced by the implementation of processes and tools for adaptive structures pioneered by Prof. Paolo Santini. Multifunctional materials and structures incorporating non-structural engineering functions such as thermal, electrical, radiation shielding, power, and sensors have been investigated. The result has been an integrated structure that offers reduced mass, packaging volume, and ease of integration for spacecraft systems. Current technology development efforts are being conducted to develop innovative multifunctional materials and structures designs incorporating advanced composites, nanotechnology, and additive manufacturing. However, these efforts offer significant challenges in the qualification and acceptance into spacecraft systems. This paper presents a brief overview of the technology development and successful insertion of advanced material technologies into spacecraft structures. Finally, opportunities and challenges to develop and mature next generation advanced materials and structures are presented.

Spacecraft Conceptual Design Compared to the Apollo Lunar Lander

Science.gov (United States)

Young, C.; Bowie, J.; Rust, R.; Lenius, J.; Anderson, M.; Connolly, J.

2011-01-01

Future human exploration of the Moon will require an optimized spacecraft design with each sub-system achieving the required minimum capability and maintaining high reliability. The objective of this study was to trade capability with reliability and minimize mass for the lunar lander spacecraft. The NASA parametric concept for a 3-person vehicle to the lunar surface with a 30% mass margin totaled was considerably heavier than the Apollo 15 Lunar Module "as flown" mass of 16.4 metric tons. The additional mass was attributed to mission requirements and system design choices that were made to meet the realities of modern spaceflight. The parametric tool used to size the current concept, Envision, accounts for primary and secondary mass requirements. For example, adding an astronaut increases the mass requirements for suits, water, food, oxygen, as well as, the increase in volume. The environmental control sub-systems becomes heavier with the increased requirements and more structure was needed to support the additional mass. There was also an increase in propellant usage. For comparison, an "Apollo-like" vehicle was created by removing these additional requirements. Utilizing the Envision parametric mass calculation tool and a quantitative reliability estimation tool designed by Valador Inc., it was determined that with today?s current technology a Lunar Module (LM) with Apollo capability could be built with less mass and similar reliability. The reliability of this new lander was compared to Apollo Lunar Module utilizing the same methodology, adjusting for mission timeline changes as well as component differences. Interestingly, the parametric concept's overall estimated risk for loss of mission (LOM) and loss of crew (LOC) did not significantly improve when compared to Apollo.

Maintainability design criteria for packaging of spacecraft replaceable electronic equipment.

Science.gov (United States)

Kappler, J. R.; Folsom, A. B.

1972-01-01

Maintainability must be designed into long-duration spacecraft and equipment to provide the required high probability of mission success with the least cost and weight. The ability to perform repairs quickly and easily in a space environment can be achieved by imposing specific maintainability design criteria on spacecraft equipment design and installation. A study was funded to investigate and define design criteria for electronic equipment that would permit rapid removal and replacement in a space environment. The results of the study are discussed together with subsequent simulated zero-g demonstration tests of a mockup with new concepts for packaging.

Adaptation and Re-Use of Spacecraft Power System Models for the Constellation Program

Science.gov (United States)

Hojnicki, Jeffrey S.; Kerslake, Thomas W.; Ayres, Mark; Han, Augustina H.; Adamson, Adrian M.

2008-01-01

NASA's Constellation Program is embarking on a new era of space exploration, returning to the Moon and beyond. The Constellation architecture will consist of a number of new spacecraft elements, including the Orion crew exploration vehicle, the Altair lunar lander, and the Ares family of launch vehicles. Each of these new spacecraft elements will need an electric power system, and those power systems will need to be designed to fulfill unique mission objectives and to survive the unique environments encountered on a lunar exploration mission. As with any new spacecraft power system development, preliminary design work will rely heavily on analysis to select the proper power technologies, size the power system components, and predict the system performance throughout the required mission profile. Constellation projects have the advantage of leveraging power system modeling developments from other recent programs such as the International Space Station (ISS) and the Mars Exploration Program. These programs have developed mature power system modeling tools, which can be quickly modified to meet the unique needs of Constellation, and thus provide a rapid capability for detailed power system modeling that otherwise would not exist.

Contributions of microgravity test results to the design of spacecraft fire-safety systems

Science.gov (United States)

Friedman, Robert; Urban, David L.

1993-01-01

Experiments conducted in spacecraft and drop towers show that thin-sheet materials have reduced flammability ranges and flame-spread rates under quiescent low-gravity environments (microgravity) compared to normal gravity. Furthermore, low-gravity flames may be suppressed more easily by atmospheric dilution or decreasing atmospheric total pressure than their normal-gravity counterparts. The addition of a ventilating air flow to the low-gravity flame zone, however, can greatly enhance the flammability range and flame spread. These results, along with observations of flame and smoke characteristics useful for microgravity fire-detection 'signatures', promise to be of considerable value to spacecraft fire-safety designs. The paper summarizes the fire detection and suppression techniques proposed for the Space Station Freedom and discusses both the application of low-gravity combustion knowledge to improve fire protection and the critical needs for further research.

Contributions of Microgravity Test Results to the Design of Spacecraft Fire Safety Systems

Science.gov (United States)

Friedman, Robert; Urban, David L.

1993-01-01

Experiments conducted in spacecraft and drop towers show that thin-sheet materials have reduced flammability ranges and flame-spread rates under quiescent low-gravity environments (microgravity) as compared to normal gravity. Furthermore, low-gravity flames may be suppressed more easily by atmospheric dilution or decreasing atmospheric total pressure than their normal-gravity counterparts. The addition of a ventilating air flow to the low-gravity flame zone, however, can greatly enhance the flammability range and flame spread. These results, along with observations of flame and smoke characteristics useful for microgravity fire-detection 'signatures', promise to be of considerable value to spacecraft fire-safety designs. The paper summarizes the fire detection and suppression techniques proposed for the Space Station Freedom and discusses both the application of low-gravity combustion knowledge to improve fire protection and the critical needs for further research.

Contingency Trajectory Design for a Lunar Orbit Insertion Maneuver Failure by the LADEE Spacecraft

Science.gov (United States)

Genova, A. L.

2014-01-01

This paper presents results from a contingency trajectory analysis performed for the Lunar Atmosphere & Dust Environment Explorer (LADEE) mission in the event of a missed lunar-orbit insertion (LOI) maneuver by the LADEE spacecraft. The effects of varying solar perturbations in the vicinity of the weak stability boundary (WSB) in the Sun-Earth system on the trajectory design are analyzed and discussed. It is shown that geocentric recovery trajectory options existed for the LADEE spacecraft, depending on the spacecraft's recovery time to perform an Earth escape-prevention maneuver after the hypothetical LOI maneuver failure and subsequent path traveled through the Sun-Earth WSB. If Earth-escape occurred, a heliocentric recovery option existed, but with reduced science capacapability for the spacecraft in an eccentric, not circular near-equatorial retrograde lunar orbit.

Galileo spacecraft power management and distribution system

International Nuclear Information System (INIS)

Detwiler, R.C.; Smith, R.L.

1990-01-01

It has been twelve years since two Voyager spacecraft began the direct route to the outer planets. In October 1989 a single Galileo spacecraft started the return to Jupiter. Conceived as a simple Voyager look-alike, the Galileo power management and distribution (PMAD) system has undergone many iterations in configuration. Major changes to the PMAD resulted from dual spun slip ring limitations, variations in launch vehicle thrust capabilities, and launch delays. Lack of an adequate launch vehicle for an interplanetary mission of Galileo's size has resulted in an extremely long flight duration. A Venus-Earth-Earth Gravity Assist (VEEGA) tour, vital to attain the required energy, results in a 6 year trip to Jupiter and its moons. This paper provides a description of the Galileo PMAD and documents the design drivers that established the final as-built hardware

Interactive Spacecraft Trajectory Design Strategies Featuring Poincare Map Topology

Science.gov (United States)

Schlei, Wayne R.

spacecraft rerouting scenario incorporating a very limited Delta V budget. In the Earth-Moon system, a low-DeltaV transfer from low Earth orbit (LEO) to the distant retrograde orbit (DRO) vicinity is derived with interactive topology-based design tactics. Finally, Poincare map topology is exploited in the Saturn-Enceladus system to explore a possible ballistic capture scenario around Enceladus.

A Quantitative Human Spacecraft Design Evaluation Model for Assessing Crew Accommodation and Utilization

Science.gov (United States)

Fanchiang, Christine

Crew performance, including both accommodation and utilization factors, is an integral part of every human spaceflight mission from commercial space tourism, to the demanding journey to Mars and beyond. Spacecraft were historically built by engineers and technologists trying to adapt the vehicle into cutting edge rocketry with the assumption that the astronauts could be trained and will adapt to the design. By and large, that is still the current state of the art. It is recognized, however, that poor human-machine design integration can lead to catastrophic and deadly mishaps. The premise of this work relies on the idea that if an accurate predictive model exists to forecast crew performance issues as a result of spacecraft design and operations, it can help designers and managers make better decisions throughout the design process, and ensure that the crewmembers are well-integrated with the system from the very start. The result should be a high-quality, user-friendly spacecraft that optimizes the utilization of the crew while keeping them alive, healthy, and happy during the course of the mission. Therefore, the goal of this work was to develop an integrative framework to quantitatively evaluate a spacecraft design from the crew performance perspective. The approach presented here is done at a very fundamental level starting with identifying and defining basic terminology, and then builds up important axioms of human spaceflight that lay the foundation for how such a framework can be developed. With the framework established, a methodology for characterizing the outcome using a mathematical model was developed by pulling from existing metrics and data collected on human performance in space. Representative test scenarios were run to show what information could be garnered and how it could be applied as a useful, understandable metric for future spacecraft design. While the model is the primary tangible product from this research, the more interesting outcome of

SSTL based thermal and power efficient RAM design on 28nm FPGA for spacecraft

DEFF Research Database (Denmark)

Kalia, Kartik; Pandey, Bishwajeet; Hussain, D. M.A.

2016-01-01

In this paper, an approach is made to design a Thermal and Power efficient RAM for that reason we have used DDR4L memory and six different members of SSTL I/Os standards on 28nm technology. Every spacecraft requires most energy efficient electronic system and for that very purpose we have designe...

Attitude coordination for spacecraft formation with multiple communication delays

Directory of Open Access Journals (Sweden)

Guo Yaohua

2015-04-01

Full Text Available Communication delays are inherently present in information exchange between spacecraft and have an effect on the control performance of spacecraft formation. In this work, attitude coordination control of spacecraft formation is addressed, which is in the presence of multiple communication delays between spacecraft. Virtual system-based approach is utilized in case that a constant reference attitude is available to only a part of the spacecraft. The feedback from the virtual systems to the spacecraft formation is introduced to maintain the formation. Using backstepping control method, input torque of each spacecraft is designed such that the attitude of each spacecraft converges asymptotically to the states of its corresponding virtual system. Furthermore, the backstepping technique and the Lyapunov–Krasovskii method contribute to the control law design when the reference attitude is time-varying and can be obtained by each spacecraft. Finally, effectiveness of the proposed methodology is illustrated by the numerical simulations of a spacecraft formation.

Isolation, pointing, and suppression (IPS) system for high-performance spacecraft

Science.gov (United States)

Hindle, Tim; Davis, Torey; Fischer, Jim

2007-04-01

Passive mechanical isolation is often times the first step taken to remedy vibration issues on-board a spacecraft. In many cases, this is done with a hexapod of axial members or struts to obtain the desired passive isolation in all six degrees-of-freedom (DOF). In some instances, where the disturbance sources are excessive or the payload is particularly sensitive to vibration, additional steps are taken to improve the performance beyond that of passive isolation. Additional performance or functionality can be obtained with the addition of active control, using a hexapod of hybrid (passive/active) elements at the interface between the payload and the bus. This paper describes Honeywell's Isolation, Pointing, and Suppression (IPS) system. It is a hybrid isolation system designed to isolate a sensitive spacecraft payload with very low passive resonant break frequencies while affording agile independent payload pointing, on-board payload disturbance rejection, and active isolation augmentation. This system is an extension of the work done on Honeywell's previous Vibration Isolation, Steering, and Suppression (VISS) flight experiment. Besides being designed for a different size payload than VISS, the IPS strut includes a dual-stage voice coil design for improved dynamic range as well as improved low-noise drive electronics. In addition, the IPS struts include integral load cells, gap sensors, and payloadside accelerometers for control and telemetry purposes. The associated system-level control architecture to accomplish these tasks is also new for this program as compared to VISS. A summary of the IPS system, including analysis and hardware design, build, and single axis bipod testing will be reviewed.

An Integrated Vision-Based System for Spacecraft Attitude and Topology Determination for Formation Flight Missions

Science.gov (United States)

Rogers, Aaron; Anderson, Kalle; Mracek, Anna; Zenick, Ray

2004-01-01

With the space industry's increasing focus upon multi-spacecraft formation flight missions, the ability to precisely determine system topology and the orientation of member spacecraft relative to both inertial space and each other is becoming a critical design requirement. Topology determination in satellite systems has traditionally made use of GPS or ground uplink position data for low Earth orbits, or, alternatively, inter-satellite ranging between all formation pairs. While these techniques work, they are not ideal for extension to interplanetary missions or to large fleets of decentralized, mixed-function spacecraft. The Vision-Based Attitude and Formation Determination System (VBAFDS) represents a novel solution to both the navigation and topology determination problems with an integrated approach that combines a miniature star tracker with a suite of robust processing algorithms. By combining a single range measurement with vision data to resolve complete system topology, the VBAFDS design represents a simple, resource-efficient solution that is not constrained to certain Earth orbits or formation geometries. In this paper, analysis and design of the VBAFDS integrated guidance, navigation and control (GN&C) technology will be discussed, including hardware requirements, algorithm development, and simulation results in the context of potential mission applications.

SHARP - Automated monitoring of spacecraft health and status

Science.gov (United States)

Atkinson, David J.; James, Mark L.; Martin, R. G.

1990-01-01

Briefly discussed here are the spacecraft and ground systems monitoring process at the Jet Propulsion Laboratory (JPL). Some of the difficulties associated with the existing technology used in mission operations are highlighted. A new automated system based on artificial intelligence technology is described which seeks to overcome many of these limitations. The system, called the Spacecraft Health Automated Reasoning Prototype (SHARP), is designed to automate health and status analysis for multi-mission spacecraft and ground data systems operations. The system has proved to be effective for detecting and analyzing potential spacecraft and ground systems problems by performing real-time analysis of spacecraft and ground data systems engineering telemetry. Telecommunications link analysis of the Voyager 2 spacecraft was the initial focus for evaluation of the system in real-time operations during the Voyager spacecraft encounter with Neptune in August 1989.

SHARP: Automated monitoring of spacecraft health and status

Science.gov (United States)

Atkinson, David J.; James, Mark L.; Martin, R. Gaius

1991-01-01

Briefly discussed here are the spacecraft and ground systems monitoring process at the Jet Propulsion Laboratory (JPL). Some of the difficulties associated with the existing technology used in mission operations are highlighted. A new automated system based on artificial intelligence technology is described which seeks to overcome many of these limitations. The system, called the Spacecraft Health Automated Reasoning Prototype (SHARP), is designed to automate health and status analysis for multi-mission spacecraft and ground data systems operations. The system has proved to be effective for detecting and analyzing potential spacecraft and ground systems problems by performing real-time analysis of spacecraft and ground data systems engineering telemetry. Telecommunications link analysis of the Voyager 2 spacecraft was the initial focus for evaluation of the system in real-time operations during the Voyager spacecraft encounter with Neptune in August 1989.

A computer graphics system for visualizing spacecraft in orbit

Science.gov (United States)

Eyles, Don E.

1989-01-01

To carry out unanticipated operations with resources already in space is part of the rationale for a permanently manned space station in Earth orbit. The astronauts aboard a space station will require an on-board, spatial display tool to assist the planning and rehearsal of upcoming operations. Such a tool can also help astronauts to monitor and control such operations as they occur, especially in cases where first-hand visibility is not possible. A computer graphics visualization system designed for such an application and currently implemented as part of a ground-based simulation is described. The visualization system presents to the user the spatial information available in the spacecraft's computers by drawing a dynamic picture containing the planet Earth, the Sun, a star field, and up to two spacecraft. The point of view within the picture can be controlled by the user to obtain a number of specific visualization functions. The elements of the display, the methods used to control the display's point of view, and some of the ways in which the system can be used are described.

A Shaftless Magnetically Levitated Multifunctional Spacecraft Flywheel Storage System

Science.gov (United States)

Stevens, Ken; Thornton, Richard; Clark, Tracy; Beaman, Bob G.; Dennehy, Neil; Day, John H. (Technical Monitor)

2002-01-01

Presently many types of spacecraft use a Spacecraft Attitude Control System (ACS) with momentum wheels for steering and electrochemical batteries to provide electrical power for the eclipse period of the spacecraft orbit. Future spacecraft will use Flywheels for combined use in ACS and Energy Storage. This can be done by using multiple wheels and varying the differential speed for ACS and varying the average speed for energy storage and recovery. Technology in these areas has improved since the 1990s so it is now feasible for flywheel systems to emerge from the laboratory for spacecraft use. This paper describes a new flywheel system that can be used for both ACS and energy storage. Some of the possible advantages of a flywheel system are: lower total mass and volume, higher efficiency, less thermal impact, improved satellite integration schedule and complexity, simplified satellite orbital operations, longer life with lower risk, less pointing jitter, and greater capability for high-rate slews. In short, they have the potential to enable new types of missions and provide lower cost. Two basic types of flywheel configurations are the Flywheel Energy Storage System (FESS) and the Integrated Power and Attitude Control System (IPACS).

Low cost spacecraft computers: Oxymoron or future trend?

Science.gov (United States)

Manning, Robert M.

1993-01-01

Over the last few decades, application of current terrestrial computer technology in embedded spacecraft control systems has been expensive and wrought with many technical challenges. These challenges have centered on overcoming the extreme environmental constraints (protons, neutrons, gamma radiation, cosmic rays, temperature, vibration, etc.) that often preclude direct use of commercial off-the-shelf computer technology. Reliability, fault tolerance and power have also greatly constrained the selection of spacecraft control system computers. More recently, new constraints are being felt, cost and mass in particular, that have again narrowed the degrees of freedom spacecraft designers once enjoyed. This paper discusses these challenges, how they were previously overcome, how future trends in commercial computer technology will simplify (or hinder) selection of computer technology for spacecraft control applications, and what spacecraft electronic system designers can do now to circumvent them.

Enabling Advanced Automation in Spacecraft Operations with the Spacecraft Emergency Response System

Science.gov (United States)

Breed, Julie; Fox, Jeffrey A.; Powers, Edward I. (Technical Monitor)

2001-01-01

True autonomy is the Holy Grail of spacecraft mission operations. The goal of launching a satellite and letting it manage itself throughout its useful life is a worthy one. With true autonomy, the cost of mission operations would be reduced to a negligible amount. Under full autonomy, any problems (no matter the severity or type) that may arise with the spacecraft would be handled without any human intervention via some combination of smart sensors, on-board intelligence, and/or smart automated ground system. Until the day that complete autonomy is practical and affordable to deploy, incremental steps of deploying ever-increasing levels of automation (computerization of once manual tasks) on the ground and on the spacecraft are gradually decreasing the cost of mission operations. For example, NASA's Goddard Space Flight Center (NASA-GSFC) has been flying spacecraft with low cost operations for several years. NASA-GSFC's SMEX (Small Explorer) and MIDEX (Middle Explorer) missions have effectively deployed significant amounts of automation to enable the missions to fly predominately in 'light-out' mode. Under light-out operations the ground system is run without human intervention. Various tools perform many of the tasks previously performed by the human operators. One of the major issues in reducing human staff in favor of automation is the perceived increased in risk of losing data, or even losing a spacecraft, because of anomalous conditions that may occur when there is no one in the control center. When things go wrong, missions deploying advanced automation need to be sure that anomalous conditions are detected and that key personal are notified in a timely manner so that on-call team members can react to those conditions. To ensure the health and safety of its lights-out missions, NASA-GSFC's Advanced Automation and Autonomy branch (Code 588) developed the Spacecraft Emergency Response System (SERS). The SERS is a Web-based collaborative environment that enables

Spacecraft Conceptual Design for the 8-Meter Advanced Technology Large Aperture Space Telescope (ATLAST)

Science.gov (United States)

Hopkins, Randall C.; Capizzo, Peter; Fincher, Sharon; Hornsby, Linda S.; Jones, David

2010-01-01

The Advanced Concepts Office at Marshall Space Flight Center completed a brief spacecraft design study for the 8-meter monolithic Advanced Technology Large Aperture Space Telescope (ATLAST-8m). This spacecraft concept provides all power, communication, telemetry, avionics, guidance and control, and thermal control for the observatory, and inserts the observatory into a halo orbit about the second Sun-Earth Lagrange point. The multidisciplinary design team created a simple spacecraft design that enables component and science instrument servicing, employs articulating solar panels for help with momentum management, and provides precise pointing control while at the same time fast slewing for the observatory.

Conceptual definition of Automated Power Systems Management. [for planetary spacecraft

Science.gov (United States)

Imamura, M. S.; Skelly, L.; Weiner, H.

1977-01-01

Automated Power Systems Management (APSM) is defined as the capability of a spacecraft power system to automatically perform monitoring, computational, command, and control functions without ground intervention. Power systems for future planetary spacecraft must have this capability because they must perform up to 10 years, and accommodate real-time changes in mission execution autonomously. Specific APSM functions include fault detection, isolation, and correction; system performance and load profile prediction; power system optimization; system checkout; and data storage and transmission control. This paper describes the basic method of implementing these specific functions. The APSM hardware includes a central power system computer and a processor dedicated to each major power system subassembly along with digital interface circuitry. The major payoffs anticipated are in enhancement of spacecraft reliability and life and reduction of overall spacecraft program cost.

Aerospace Vehicle Design, Spacecraft Section. Final Project Reports. Volume 2; Project Groups 6-8

Science.gov (United States)

1989-01-01

Three groups of student engineers in an aerospace vehicle design course present their designs for a vehicle that can be used to resupply the Space Station Freedam and provide emergency crew return to earth capability. The vehicle's requirements include a lifetime that exceeds six years, low cost, the capability for withstanding pressurization, launch, orbit, and reentry hazards, and reliability. The vehicle's subsystems are structures, communication and command data systems, attitude and articulation control, life support and crew systems, power and propulsion, reentry and recovery systems, and mission management, planning, and costing. Special attention is given to spacecraft communications.

A Hybrid Systems Strategy to Support Autonomous Spacecraft Trajectory Design and Optimization in Multiple Dynamical Regimes

Data.gov (United States)

National Aeronautics and Space Administration — With ever increasing numbers of near-Earth satellites and deep space missions, autonomous spacecraft guidance, navigation, and control (GNC) systems are increasingly...

Investigation of fast initialization of spacecraft bubble memory systems

Science.gov (United States)

Looney, K. T.; Nichols, C. D.; Hayes, P. J.

1984-01-01

Bubble domain technology offers significant improvement in reliability and functionality for spacecraft onboard memory applications. In considering potential memory systems organizations, minimization of power in high capacity bubble memory systems necessitates the activation of only the desired portions of the memory. In power strobing arbitrary memory segments, a capability of fast turn on is required. Bubble device architectures, which provide redundant loop coding in the bubble devices, limit the initialization speed. Alternate initialization techniques are investigated to overcome this design limitation. An initialization technique using a small amount of external storage is demonstrated.

Advanced Solar-propelled Cargo Spacecraft for Mars Missions

Science.gov (United States)

Auziasdeturenne, Jacqueline; Beall, Mark; Burianek, Joseph; Cinniger, Anna; Dunmire, Barbrina; Haberman, Eric; Iwamoto, James; Johnson, Stephen; Mccracken, Shawn; Miller, Melanie

1989-01-01

Three concepts for an unmanned, solar powered, cargo spacecraft for Mars support missions were investigated. These spacecraft are designed to carry a 50,000 kg payload from a low Earth orbit to a low Mars orbit. Each design uses a distinctly different propulsion system: A Solar Radiation Absorption (SRA) system, a Solar-Pumped Laser (SPL) system and a solar powered magnetoplasmadynamic (MPD) arc system. The SRA directly converts solar energy to thermal energy in the propellant through a novel process. In the SPL system, a pair of solar-pumped, multi-megawatt, CO2 lasers in sunsynchronous Earth orbit converts solar energy to laser energy. The MPD system used indium phosphide solar cells to convert sunlight to electricity, which powers the propulsion system. Various orbital transfer options are examined for these concepts. In the SRA system, the mother ship transfers the payload into a very high Earth orbit and a small auxiliary propulsion system boosts the payload into a Hohmann transfer to Mars. The SPL spacecraft and the SPL powered spacecraft return to Earth for subsequent missions. The MPD propelled spacecraft, however, remains at Mars as an orbiting space station. A patched conic approximation was used to determine a heliocentric interplanetary transfer orbit for the MPD propelled spacecraft. All three solar-powered spacecraft use an aerobrake procedure to place the payload into a low Mars parking orbit. The payload delivery times range from 160 days to 873 days (2.39 years).

Weight estimates and packaging techniques for the microwave radiometer spacecraft. [shuttle compatible design

Science.gov (United States)

Jensen, J. K.; Wright, R. L.

1981-01-01

Estimates of total spacecraft weight and packaging options were made for three conceptual designs of a microwave radiometer spacecraft. Erectable structures were found to be slightly lighter than deployable structures but could be packaged in one-tenth the volume. The tension rim concept, an unconventional design approach, was found to be the lightest and transportable to orbit in the least number of shuttle flights.

Attitude Estimation in Fractionated Spacecraft Cluster Systems

Science.gov (United States)

Hadaegh, Fred Y.; Blackmore, James C.

2011-01-01

An attitude estimation was examined in fractioned free-flying spacecraft. Instead of a single, monolithic spacecraft, a fractionated free-flying spacecraft uses multiple spacecraft modules. These modules are connected only through wireless communication links and, potentially, wireless power links. The key advantage of this concept is the ability to respond to uncertainty. For example, if a single spacecraft module in the cluster fails, a new one can be launched at a lower cost and risk than would be incurred with onorbit servicing or replacement of the monolithic spacecraft. In order to create such a system, however, it is essential to know what the navigation capabilities of the fractionated system are as a function of the capabilities of the individual modules, and to have an algorithm that can perform estimation of the attitudes and relative positions of the modules with fractionated sensing capabilities. Looking specifically at fractionated attitude estimation with startrackers and optical relative attitude sensors, a set of mathematical tools has been developed that specify the set of sensors necessary to ensure that the attitude of the entire cluster ( cluster attitude ) can be observed. Also developed was a navigation filter that can estimate the cluster attitude if these conditions are satisfied. Each module in the cluster may have either a startracker, a relative attitude sensor, or both. An extended Kalman filter can be used to estimate the attitude of all modules. A range of estimation performances can be achieved depending on the sensors used and the topology of the sensing network.

Centralized database for interconnection system design. [for spacecraft

Science.gov (United States)

Billitti, Joseph W.

1989-01-01

A database application called DFACS (Database, Forms and Applications for Cabling and Systems) is described. The objective of DFACS is to improve the speed and accuracy of interconnection system information flow during the design and fabrication stages of a project, while simultaneously supporting both the horizontal (end-to-end wiring) and the vertical (wiring by connector) design stratagems used by the Jet Propulsion Laboratory (JPL) project engineering community. The DFACS architecture is centered around a centralized database and program methodology which emulates the manual design process hitherto used at JPL. DFACS has been tested and successfully applied to existing JPL hardware tasks with a resulting reduction in schedule time and costs.

Thermal design of spacecraft solar arrays using a polyimide foam

International Nuclear Information System (INIS)

Bianco, N; Iasiello, M; Naso, V

2015-01-01

The design of the Thermal Control System (TCS) of spacecraft solar arrays plays a fundamental role. Indeed, the spacecraft components must operate within a certain range of temperature. If this doesn't occur, their performance is reduced and they may even break. Solar arrays, which are employed to recharge batteries, are directly exposed to the solar heat flux, and they need to be insulated from the earth's surface irradiation. Insulation is currently provided either with a white paint coating or with a Multi Layer Insulation (MLI) system [1]. A configuration based on an open-cell polyimide foam has also been recently proposed [2]. Using polyimide foams in TCSs looks very attractive in terms of costs, weight and assembling. An innovative thermal analysis of the above cited TCS configurations is carried out in this paper, by solving the porous media energy equation, under the assumption of Local Thermal Equilibrium (LTE) between the two phases. Radiation effects through the solar array are also considered by using the Rosseland approximation. Under a stationary daylight condition, temperature profiles are obtained by means of the finite-element based code COMSOL Multiphysics ® . Finally, since the weight plays an important role in aerospace applications, weights of the three TCS configurations are compared. (paper)

Thermal design of spacecraft solar arrays using a polyimide foam

Science.gov (United States)

Bianco, N.; Iasiello, M.; Naso, V.

2015-11-01

The design of the Thermal Control System (TCS) of spacecraft solar arrays plays a fundamental role. Indeed, the spacecraft components must operate within a certain range of temperature. If this doesn't occur, their performance is reduced and they may even break. Solar arrays, which are employed to recharge batteries, are directly exposed to the solar heat flux, and they need to be insulated from the earth's surface irradiation. Insulation is currently provided either with a white paint coating or with a Multi Layer Insulation (MLI) system [1]. A configuration based on an open-cell polyimide foam has also been recently proposed [2]. Using polyimide foams in TCSs looks very attractive in terms of costs, weight and assembling. An innovative thermal analysis of the above cited TCS configurations is carried out in this paper, by solving the porous media energy equation, under the assumption of Local Thermal Equilibrium (LTE) between the two phases. Radiation effects through the solar array are also considered by using the Rosseland approximation. Under a stationary daylight condition, temperature profiles are obtained by means of the finite-element based code COMSOL Multiphysics®. Finally, since the weight plays an important role in aerospace applications, weights of the three TCS configurations are compared.

Low Cost Rapid Response Spacecraft, (LCRRS): A Research Project in Low Cost Spacecraft Design and Fabrication in a Rapid Prototyping Environment

Science.gov (United States)

Spremo, Stevan; Bregman, Jesse; Dallara, Christopher D.; Ghassemieh, Shakib M.; Hanratty, James; Jackson, Evan; Kitts, Christopher; Klupar, Pete; Lindsay, Michael; Ignacio, Mas;

Adaptive Management of Computing and Network Resources for Spacecraft Systems

Science.gov (United States)

Pfarr, Barbara; Welch, Lonnie R.; Detter, Ryan; Tjaden, Brett; Huh, Eui-Nam; Szczur, Martha R. (Technical Monitor)

2000-01-01

It is likely that NASA's future spacecraft systems will consist of distributed processes which will handle dynamically varying workloads in response to perceived scientific events, the spacecraft environment, spacecraft anomalies and user commands. Since all situations and possible uses of sensors cannot be anticipated during pre-deployment phases, an approach for dynamically adapting the allocation of distributed computational and communication resources is needed. To address this, we are evolving the DeSiDeRaTa adaptive resource management approach to enable reconfigurable ground and space information systems. The DeSiDeRaTa approach embodies a set of middleware mechanisms for adapting resource allocations, and a framework for reasoning about the real-time performance of distributed application systems. The framework and middleware will be extended to accommodate (1) the dynamic aspects of intra-constellation network topologies, and (2) the complete real-time path from the instrument to the user. We are developing a ground-based testbed that will enable NASA to perform early evaluation of adaptive resource management techniques without the expense of first deploying them in space. The benefits of the proposed effort are numerous, including the ability to use sensors in new ways not anticipated at design time; the production of information technology that ties the sensor web together; the accommodation of greater numbers of missions with fewer resources; and the opportunity to leverage the DeSiDeRaTa project's expertise, infrastructure and models for adaptive resource management for distributed real-time systems.

Air Purification in Closed Environments: An Overview of Spacecraft Systems

Science.gov (United States)

Perry, Jay L.; LeVan, Douglas; Crumbley, Robert (Technical Monitor)

2002-01-01

The primary goal for a collective protection system and a spacecraft environmental control and life support system (ECLSS) are strikingly similar. Essentially both function to provide the occupants of a building or vehicle with a safe, habitable environment. The collective protection system shields military and civilian personnel from short-term exposure to external threats presented by toxic agents and industrial chemicals while an ECLSS sustains astronauts for extended periods within the hostile environment of space. Both have air quality control similarities with various aircraft and 'tight' buildings. This paper reviews basic similarities between air purification system requirements for collective protection and an ECLSS that define surprisingly common technological challenges and solutions. Systems developed for air revitalization on board spacecraft are discussed along with some history on their early development as well as a view of future needs. Emphasis is placed upon two systems implemented by the National Aeronautics and Space Administration (NASA) onboard the International Space Station (ISS): the trace contaminant control system (TCCS) and the molecular sieve-based carbon dioxide removal assembly (CDRA). Over its history, the NASA has developed and implemented many life support systems for astronauts. As the duration, complexity, and crew size of manned missions increased from minutes or hours for a single astronaut during Project Mercury to days and ultimately months for crews of 3 or more during the Apollo, Skylab, Shuttle, and ISS programs, these systems have become more sophisticated. Systems aboard spacecraft such as the ISS have been designed to provide long-term environmental control and life support. Challenges facing the NASA's efforts include minimizing mass, volume, and power for such systems, while maximizing their safety, reliability, and performance. This paper will highlight similarities and differences among air purification systems

Hybrid microtransmitter for free-space optical spacecraft communication: design, manufacturing, and characterization

Science.gov (United States)

Lotfi, Sara; Palmer, Kristoffer; Kratz, Henrik; Thornell, Greger

2009-02-01

Optical intra-communication links are investigated by several currently operational qualification missions. Compared with RF communication systems, the optical domain obtains a wider bandwidth, enables miniaturized spacecraft and reduced power consumption. In this project, a microtransmitter is designed and manufactured for formation flying spacecraft with transmission rates of 1 Gbit/s. Simulations in Matlab and Simulink show that a BER of 10-9 can be achieved with aperture sizes of 1 cm and a transmitter output peak power of 12 mW for a distance of 10 km. The results show that the performance of the communication link decreases due to mechanical vibrations in the spacecraft together with a narrow laser beam. A dual-axis microactuator designed as a deflectable mirror has been developed for the laser beam steering where the fabrication is based on a double-sided, bulk micromachining process. The mirror actuates by joints consisting of v-grooves filled with SU-8 polymer. The deflection is controlled by integrated resistive heaters in the joints causing the polymer to expand thermally. Results show that the mirror actuates 20-30° in the temperature interval 25-250°C. Flat Fresnel lenses made of Pyrex 7740 are used to collimate the laser beam. These lenses are simulated in the Comsol software and optimized for a 670 nm red VCSEL. The lenses are manufactured using lithography and reactive ion etching. All tests are made in a normal laboratory environment, but the effect of the space environment is discussed.

Intelligent Systems and Advanced User Interfaces for Design, Operation, and Maintenance of Command Management Systems

Science.gov (United States)

Mitchell, Christine M.

1998-01-01

Historically Command Management Systems (CMS) have been large, expensive, spacecraft-specific software systems that were costly to build, operate, and maintain. Current and emerging hardware, software, and user interface technologies may offer an opportunity to facilitate the initial formulation and design of a spacecraft-specific CMS as well as a to develop a more generic or a set of core components for CMS systems. Current MOC (mission operations center) hardware and software include Unix workstations, the C/C++ and Java programming languages, and X and Java window interfaces representations. This configuration provides the power and flexibility to support sophisticated systems and intelligent user interfaces that exploit state-of-the-art technologies in human-machine systems engineering, decision making, artificial intelligence, and software engineering. One of the goals of this research is to explore the extent to which technologies developed in the research laboratory can be productively applied in a complex system such as spacecraft command management. Initial examination of some of the issues in CMS design and operation suggests that application of technologies such as intelligent planning, case-based reasoning, design and analysis tools from a human-machine systems engineering point of view (e.g., operator and designer models) and human-computer interaction tools, (e.g., graphics, visualization, and animation), may provide significant savings in the design, operation, and maintenance of a spacecraft-specific CMS as well as continuity for CMS design and development across spacecraft with varying needs. The savings in this case is in software reuse at all stages of the software engineering process.

Analysis on coverage ability of BeiDou navigation satellite system for manned spacecraft

Science.gov (United States)

Zhao, Sihao; Yao, Zheng; Zhuang, Xuebin; Lu, Mingquan

2014-12-01

To investigate the service ability of the BeiDou Navigation Satellite System (BDS) for manned spacecraft, both the current regional and the future-planned global constellations of BDS are introduced and simulated. The orbital parameters of the International Space Station and China's Tiangong-1 spacelab are used to create the simulation scenario and evaluate the performance of the BDS constellations. The number of visible satellites and the position dilution (PDOP) of precision at the spacecraft-based receiver are evaluated. Simulation and analysis show quantitative results on the coverage ability and time percentages of both the current BDS regional and future global constellations for manned-space orbits which can be a guideline to the applications and mission design of BDS receivers on manned spacecraft.

Guidance, navigation, and control subsystem for the EOS-AM spacecraft

Science.gov (United States)

Linder, David M.; Tolek, Joseph T.; Lombardo, John

1992-01-01

This paper presents the preliminary design of the Guidance, Navigation, and Control (GN&C) subsystem for the EOS-AM spacecraft and specifically focuses on the GN&C Normal Mode design. First, a brief description of the EOS-AM science mission, instruments, and system-level spacecraft design is provided. Next, an overview of the GN&C subsystem functional and performance requirements, hardware, and operating modes is presented. Then, the GN&C Normal Mode attitude determination, attitude control, and navigation systems are detailed. Finally, descriptions of the spacecraft's overall jitter performance and Safe Mode are provided.

Spacecraft Data Simulator for the test of level zero processing systems

Science.gov (United States)

Shi, Jeff; Gordon, Julie; Mirchandani, Chandru; Nguyen, Diem

1994-01-01

The Microelectronic Systems Branch (MSB) at Goddard Space Flight Center (GSFC) has developed a Spacecraft Data Simulator (SDS) to support the development, test, and verification of prototype and production Level Zero Processing (LZP) systems. Based on a disk array system, the SDS is capable of generating large test data sets up to 5 Gigabytes and outputting serial test data at rates up to 80 Mbps. The SDS supports data formats including NASA Communication (Nascom) blocks, Consultative Committee for Space Data System (CCSDS) Version 1 & 2 frames and packets, and all the Advanced Orbiting Systems (AOS) services. The capability to simulate both sequential and non-sequential time-ordered downlink data streams with errors and gaps is crucial to test LZP systems. This paper describes the system architecture, hardware and software designs, and test data designs. Examples of test data designs are included to illustrate the application of the SDS.

Designing a Robust Nonlinear Dynamic Inversion Controller for Spacecraft Formation Flying

Directory of Open Access Journals (Sweden)

Inseok Yang

2014-01-01

Full Text Available The robust nonlinear dynamic inversion (RNDI control technique is proposed to keep the relative position of spacecrafts while formation flying. The proposed RNDI control method is based on nonlinear dynamic inversion (NDI. NDI is nonlinear control method that replaces the original dynamics into the user-selected desired dynamics. Because NDI removes nonlinearities in the model by inverting the original dynamics directly, it also eliminates the need of designing suitable controllers for each equilibrium point; that is, NDI works as self-scheduled controller. Removing the original model also provides advantages of ease to satisfy the specific requirements by simply handling desired dynamics. Therefore, NDI is simple and has many similarities to classical control. In real applications, however, it is difficult to achieve perfect cancellation of the original dynamics due to uncertainties that lead to performance degradation and even make the system unstable. This paper proposes robustness assurance method for NDI. The proposed RNDI is designed by combining NDI and sliding mode control (SMC. SMC is inherently robust using high-speed switching inputs. This paper verifies similarities of NDI and SMC, firstly. And then RNDI control method is proposed. The performance of the proposed method is evaluated by simulations applied to spacecraft formation flying problem.

Dips spacecraft integration issues

International Nuclear Information System (INIS)

Determan, W.R.; Harty, R.B.

1988-01-01

The Department of Energy, in cooperation with the Department of Defense, has recently initiated the dynamic isotope power system (DIPS) demonstration program. DIPS is designed to provide 1 to 10 kW of electrical power for future military spacecraft. One of the near-term missions considered as a potential application for DIPS was the boost surveillance and tracking system (BSTS). A brief review and summary of the reasons behind a selection of DIPS for BSTS-type missions is presented. Many of these are directly related to spacecraft integration issues; these issues will be reviewed in the areas of system safety, operations, survivability, reliability, and autonomy

Spacecraft Thermal Management

Science.gov (United States)

Hurlbert, Kathryn Miller

2009-01-01

In the 21st century, the National Aeronautics and Space Administration (NASA), the Russian Federal Space Agency, the National Space Agency of Ukraine, the China National Space Administration, and many other organizations representing spacefaring nations shall continue or newly implement robust space programs. Additionally, business corporations are pursuing commercialization of space for enabling space tourism and capital business ventures. Future space missions are likely to include orbiting satellites, orbiting platforms, space stations, interplanetary vehicles, planetary surface missions, and planetary research probes. Many of these missions will include humans to conduct research for scientific and terrestrial benefits and for space tourism, and this century will therefore establish a permanent human presence beyond Earth s confines. Other missions will not include humans, but will be autonomous (e.g., satellites, robotic exploration), and will also serve to support the goals of exploring space and providing benefits to Earth s populace. This section focuses on thermal management systems for human space exploration, although the guiding principles can be applied to unmanned space vehicles as well. All spacecraft require a thermal management system to maintain a tolerable thermal environment for the spacecraft crew and/or equipment. The requirements for human rating and the specified controlled temperature range (approximately 275 K - 310 K) for crewed spacecraft are unique, and key design criteria stem from overall vehicle and operational/programatic considerations. These criteria include high reliability, low mass, minimal power requirements, low development and operational costs, and high confidence for mission success and safety. This section describes the four major subsystems for crewed spacecraft thermal management systems, and design considerations for each. Additionally, some examples of specialized or advanced thermal system technologies are presented

Error Analysis System for Spacecraft Navigation Using the Global Positioning System (GPS)

Science.gov (United States)

Truong, S. H.; Hart, R. C.; Hartman, K. R.; Tomcsik, T. L.; Searl, J. E.; Bernstein, A.

1997-01-01

The Flight Dynamics Division (FDD) at the National Aeronautics and Space Administration (NASA) Goddard Space Flight Center (GSFC) is currently developing improved space-navigation filtering algorithms to use the Global Positioning System (GPS) for autonomous real-time onboard orbit determination. In connection with a GPS technology demonstration on the Small Satellite Technology Initiative (SSTI)/Lewis spacecraft, FDD analysts and programmers have teamed with the GSFC Guidance, Navigation, and Control Branch to develop the GPS Enhanced Orbit Determination Experiment (GEODE) system. The GEODE system consists of a Kalman filter operating as a navigation tool for estimating the position, velocity, and additional states required to accurately navigate the orbiting Lewis spacecraft by using astrodynamic modeling and GPS measurements from the receiver. A parallel effort at the FDD is the development of a GPS Error Analysis System (GEAS) that will be used to analyze and improve navigation filtering algorithms during development phases and during in-flight calibration. For GEAS, the Kalman filter theory is extended to estimate the errors in position, velocity, and other error states of interest. The estimation of errors in physical variables at regular intervals will allow the time, cause, and effect of navigation system weaknesses to be identified. In addition, by modeling a sufficient set of navigation system errors, a system failure that causes an observed error anomaly can be traced and accounted for. The GEAS software is formulated using Object Oriented Design (OOD) techniques implemented in the C++ programming language on a Sun SPARC workstation. The Phase 1 of this effort is the development of a basic system to be used to evaluate navigation algorithms implemented in the GEODE system. This paper presents the GEAS mathematical methodology, systems and operations concepts, and software design and implementation. Results from the use of the basic system to evaluate

Equations of Motion of Free-Floating Spacecraft-Manipulator Systems: An Engineer's Tutorial

Directory of Open Access Journals (Sweden)

Markus Wilde

2018-04-01

Full Text Available The paper provides a step-by-step tutorial on the Generalized Jacobian Matrix (GJM approach for modeling and simulation of spacecraft-manipulator systems. The General Jacobian Matrix approach describes the motion of the end-effector of an underactuated manipulator system solely by the manipulator joint rotations, with the attitude and position of the base-spacecraft resulting from the manipulator motion. The coupling of the manipulator motion with the base-spacecraft are thus expressed in a generalized inertia matrix and a GJM. The focus of the paper lies on the complete analytic derivation of the generalized equations of motion of a free-floating spacecraft-manipulator system. This includes symbolic analytic expressions for all inertia property matrices of the system, including their time derivatives and joint-angle derivatives, as well as an expression for the generalized Jacobian of a generic point on any link of the spacecraft-manipulator system. The kinematics structure of the spacecraft-manipulator system is described both in terms of direction-cosine matrices and unit quaternions. An additional important contribution of this paper is to propose a new and more detailed definition for the modes of maneuvering of a spacecraft-manipulator. In particular, the two commonly used categories free-flying and free-floating are expanded by the introduction of five categories, namely floating, rotation-floating, rotation-flying, translation-flying, and flying. A fully-symbolic and a partially-symbolic option for the implementation of a numerical simulation model based on the proposed analytic approach are introduced and exemplary simulation results for a planar four-link spacecraft-manipulator system and a spatial six-link spacecraft manipulator system are presented.

In-orbit evaluation of the control system/structural mode interactions of the OSO-8 spacecraft

Science.gov (United States)

Slafer, L. I.

1979-01-01

The Orbiting Solar Observatory-8 experienced severe structural mode/control loop interaction problems during the spacecraft development. Extensive analytical studies, using the hybrid coordinate modeling approach, and comprehensive ground testing were carried out in order to achieve the system's precision pointing performance requirements. A recent series of flight tests were conducted with the spacecraft in which a wide bandwidth, high resolution telemetry system was utilized to evaluate the on-orbit flexible dynamics characteristics of the vehicle along with the control system performance. The paper describes the results of these tests, reviewing the basic design problem, analytical approach taken, ground test philosophy, and on-orbit testing. Data from the tests was used to determine the primary mode frequency, damping, and servo coupling dynamics for the on-orbit condition. Additionally, the test results have verified analytically predicted differences between the on-orbit and ground test environments, and have led to a validation of both the analytical modeling and servo design techniques used during the development of the control system.

Modeling the fundamental characteristics and processes of the spacecraft functioning

Science.gov (United States)

Bazhenov, V. I.; Osin, M. I.; Zakharov, Y. V.

1986-01-01

The fundamental aspects of modeling of spacecraft characteristics by using computing means are considered. Particular attention is devoted to the design studies, the description of physical appearance of the spacecraft, and simulated modeling of spacecraft systems. The fundamental questions of organizing the on-the-ground spacecraft testing and the methods of mathematical modeling were presented.

Iterative Repair Planning for Spacecraft Operations Using the Aspen System

Science.gov (United States)

Rabideau, G.; Knight, R.; Chien, S.; Fukunaga, A.; Govindjee, A.

2000-01-01

This paper describes the Automated Scheduling and Planning Environment (ASPEN). ASPEN encodes complex spacecraft knowledge of operability constraints, flight rules, spacecraft hardware, science experiments and operations procedures to allow for automated generation of low level spacecraft sequences. Using a technique called iterative repair, ASPEN classifies constraint violations (i.e., conflicts) and attempts to repair each by performing a planning or scheduling operation. It must reason about which conflict to resolve first and what repair method to try for the given conflict. ASPEN is currently being utilized in the development of automated planner/scheduler systems for several spacecraft, including the UFO-1 naval communications satellite and the Citizen Explorer (CX1) satellite, as well as for planetary rover operations and antenna ground systems automation. This paper focuses on the algorithm and search strategies employed by ASPEN to resolve spacecraft operations constraints, as well as the data structures for representing these constraints.

Water electrolysis system - H2 and O2 generation. [for spacecraft atmosphere revitalization

Science.gov (United States)

Schubert, F. H.; Lee, M. K.; Davenport, R. J.; Quattrone, P. D.

1978-01-01

An oxygen generation system design based on the static feed water electrolysis concept is described. The system is designed to generate 4.20 kg/d of oxygen to satisfy the metabolic needs of a three-person crew, to compensate for spacecraft leakage, and to provide the oxygen required by the electrochemical depolarized CO2 concentrator. The system has a fixed hardware weight of 75 kg, occupies a volume of 0.11 cu m, and requires only 1.1 kw of electrical power. The static feed electrolysis concept is discussed, and experimental data on the high-performance electrode are presented.

Preliminary design of the thermal protection system for solar probe

Science.gov (United States)

Dirling, R. B., Jr.; Loomis, W. C.; Heightland, C. N.

1982-01-01

A preliminary design of the thermal protection system for the NASA Solar Probe spacecraft is presented. As presently conceived, the spacecraft will be launched by the Space Shuttle on a Jovian swing-by trajectory and at perihelion approach to three solar radii of the surface of the Earth's sun. The system design satisfies maximum envelope, structural integrity, equipotential, and mass loss/contamination requirements by employing lightweight carbon-carbon emissive shields. The primary shield is a thin shell, 15.5-deg half-angle cone which absorbs direct solar flux at up to 10-deg off-nadir spacecraft pointing angles. Secondary shields of sandwich construction and low thickness-direction thermal conductivity are used to reduce the primary shield infrared radiation to the spacecraft payload.

Case Studies in Crewed Spacecraft Environmental Control and Life Support System Process Compatibility and Cabin Environmental Impact

Science.gov (United States)

Perry, J. L.

2017-01-01

Contamination of a crewed spacecraft's cabin environment leading to environmental control and life support system (ECLSS) functional capability and operational margin degradation or loss can have an adverse effect on NASA's space exploration mission figures of merit-safety, mission success, effectiveness, and affordability. The role of evaluating the ECLSS's compatibility and cabin environmental impact as a key component of pass trace contaminant control is presented and the technical approach is described in the context of implementing NASA's safety and mission success objectives. Assessment examples are presented for a variety of chemicals used in vehicle systems and experiment hardware for the International Space Station program. The ECLSS compatibility and cabin environmental impact assessment approach, which can be applied to any crewed spacecraft development and operational effort, can provide guidance to crewed spacecraft system and payload developers relative to design criteria assigned ECLSS compatibility and cabin environmental impact ratings can be used by payload and system developers as criteria for ensuring adequate physical and operational containment. In additional to serving as an aid for guiding containment design, the assessments can guide flight rule and procedure development toward protecting the ECLSS as well as approaches for contamination event remediation.

Proposed Modifications to Engineering Design Guidelines Related to Resistivity Measurements and Spacecraft Charging

Science.gov (United States)

Dennison, J. R.; Swaminathan, Prasanna; Jost, Randy; Brunson, Jerilyn; Green, Nelson; Frederickson, A. Robb

2005-01-01

A key parameter in modeling differential spacecraft charging is the resistivity of insulating materials. This determines how charge will accumulate and redistribute across the spacecraft, as well as the time scale for charge transport and dissipation. Existing spacecraft charging guidelines recommend use of tests and imported resistivity data from handbooks that are based principally upon ASTM methods that are more applicable to classical ground conditions and designed for problems associated with power loss through the dielectric, than for how long charge can be stored on an insulator. These data have been found to underestimate charging effects by one to four orders of magnitude for spacecraft charging applications. A review is presented of methods to measure the resistive of highly insulating materials, including the electrometer-resistance method, the electrometer-constant voltage method, the voltage rate-of-change method and the charge storage method. This is based on joint experimental studies conducted at NASA Jet Propulsion Laboratory and Utah State University to investigate the charge storage method and its relation to spacecraft charging. The different methods are found to be appropriate for different resistivity ranges and for different charging circumstances. A simple physics-based model of these methods allows separation of the polarization current and dark current components from long duration measurements of resistivity over day- to month-long time scales. Model parameters are directly related to the magnitude of charge transfer and storage and the rate of charge transport. The model largely explains the observed differences in resistivity found using the different methods and provides a framework for recommendations for the appropriate test method for spacecraft materials with different resistivities and applications. The proposed changes to the existing engineering guidelines are intended to provide design engineers more appropriate methods for

A historical overview of the electrical power systems in the US manned and some US unmanned spacecraft

Science.gov (United States)

Maisel, J. E.

1984-01-01

A historical overview of electrical power systems used in the U.S. manned spacecraft and some of the U.S. unmanned spacecraft is presented in this investigation. A time frame of approximately 25 years, the period for 1959 to 1984, is covered in this report. Results indicate that the nominal bus voltage was 28 volts dc in most spacecraft and all other voltage levels were derived from this voltage through such techniques as voltage inversion or rectification, or a combination. Most spacecraft used solar arrays for the main source of power except for those spacecraft that had a relatively short flight duration, or deep spaceprobes that were designed for very long flight duration. Fuel cells were used on Gemini, Apollo, and Space Shuttle (short duration flights) while radioisotope thermoelectric generators were employed on the Pioneer, Jupiter/Saturn, Viking Lander, and Voyager spacecraft (long duration flights). The main dc bus voltage was unregulated on the manned spacecraft with voltage regulation provided at the user loads. A combination of regulated, semiregulated, and unregulated buses were used on the unmanned spacecraft depending on the type of load. For example, scientific instruments were usually connected to regulated buses while fans, relays, etc. were energized from an unregulated bus. Different forms of voltage regulation, such as shunt, buck/boot, and pulse-width modulated regulators, were used. This report includes a comprehensive bibliography on spacecraft electrical power systems for the space programs investigated.

A simple method to design non-collision relative orbits for close spacecraft formation flying

Science.gov (United States)

Jiang, Wei; Li, JunFeng; Jiang, FangHua; Bernelli-Zazzera, Franco

2018-05-01

A set of linearized relative motion equations of spacecraft flying on unperturbed elliptical orbits are specialized for particular cases, where the leader orbit is circular or equatorial. Based on these extended equations, we are able to analyze the relative motion regulation between a pair of spacecraft flying on arbitrary unperturbed orbits with the same semi-major axis in close formation. Given the initial orbital elements of the leader, this paper presents a simple way to design initial relative orbital elements of close spacecraft with the same semi-major axis, thus preventing collision under non-perturbed conditions. Considering the mean influence of J 2 perturbation, namely secular J 2 perturbation, we derive the mean derivatives of orbital element differences, and then expand them to first order. Thus the first order expansion of orbital element differences can be added to the relative motion equations for further analysis. For a pair of spacecraft that will never collide under non-perturbed situations, we present a simple method to determine whether a collision will occur when J 2 perturbation is considered. Examples are given to prove the validity of the extended relative motion equations and to illustrate how the methods presented can be used. The simple method for designing initial relative orbital elements proposed here could be helpful to the preliminary design of the relative orbital elements between spacecraft in a close formation, when collision avoidance is necessary.

Toward a new spacecraft optimal design lifetime? Impact of marginal cost of durability and reduced launch price

Science.gov (United States)

Snelgrove, Kailah B.; Saleh, Joseph Homer

2016-10-01

The average design lifetime of satellites continues to increase, in part due to the expectation that the satellite cost per operational day decreases monotonically with increased design lifetime. In this work, we challenge this expectation by revisiting the durability choice problem for spacecraft in the face of reduced launch price and under various cost of durability models. We first provide a brief overview of the economic thought on durability and highlight its limitations as they pertain to our problem (e.g., the assumption of zero marginal cost of durability). We then investigate the merging influence of spacecraft cost of durability and launch price, and we identify conditions that give rise cost-optimal design lifetimes that are shorter than the longest lifetime technically achievable. For example, we find that high costs of durability favor short design lifetimes, and that under these conditions the optimal choice is relatively robust to reduction in launch prices. By contrast, lower costs of durability favor longer design lifetimes, and the optimal choice is highly sensitive to reduction in launch price. In both cases, reduction in launch prices translates into reduction of the optimal design lifetime. Our results identify a number of situations for which satellite operators would be better served by spacecraft with shorter design lifetimes. Beyond cost issues and repeat purchases, other implications of long design lifetime include the increased risk of technological slowdown given the lower frequency of purchases and technology refresh, and the increased risk for satellite operators that the spacecraft will be technologically obsolete before the end of its life (with the corollary of loss of value and competitive advantage). We conclude with the recommendation that, should pressure to extend spacecraft design lifetime continue, satellite manufacturers should explore opportunities to lease their spacecraft to operators, or to take a stake in the ownership

Artificial Intelligence and Spacecraft Power Systems

Science.gov (United States)

Dugel-Whitehead, Norma R.

1997-01-01

This talk will present the work which has been done at NASA Marshall Space Flight Center involving the use of Artificial Intelligence to control the power system in a spacecraft. The presentation will include a brief history of power system automation, and some basic definitions of the types of artificial intelligence which have been investigated at MSFC for power system automation. A video tape of one of our autonomous power systems using co-operating expert systems, and advanced hardware will be presented.

Application of Modern Fortran to Spacecraft Trajectory Design and Optimization

Science.gov (United States)

Williams, Jacob; Falck, Robert D.; Beekman, Izaak B.

2018-01-01

In this paper, applications of the modern Fortran programming language to the field of spacecraft trajectory optimization and design are examined. Modern object-oriented Fortran has many advantages for scientific programming, although many legacy Fortran aerospace codes have not been upgraded to use the newer standards (or have been rewritten in other languages perceived to be more modern). NASA's Copernicus spacecraft trajectory optimization program, originally a combination of Fortran 77 and Fortran 95, has attempted to keep up with modern standards and makes significant use of the new language features. Various algorithms and methods are presented from trajectory tools such as Copernicus, as well as modern Fortran open source libraries and other projects.

Design and Implementation of Hitl Simulator Coupleing Communications Payload and Software Spacecraft Bus

Directory of Open Access Journals (Sweden)

In-Jun Kim

2003-12-01

Full Text Available Engineering qualification model payload for a communications and broadcasting satellite(CBS was developed by ETRI from May, 2000 to April, 2003. For the purpose of functional test and verification of the payload, a real-time hardware-in-the-loop(HITL CBS simulator(CBSSIM was also developed. We assumed that the spacecraft platform for the CBSSIM is a geostationary communication satellite using momentum bias three-axis stabilization control technique based on Koreasat. The payload hardware is combined with CBSSIM via Power, Command and Telemetry System(PCTS of Electrical Ground Support Equipment(EGSE. CBSSIM is connected with PCTS by TCP/IP and the payload is combined with PCTS by MIL-STD-1553B protocol and DC harness. This simulator runs under the PC-based simulation environment with Windows 2000 operating system. The satellite commands from the operators are transferred to the payload or bus subsystem models through the real-time process block in the simulator. Design requirements of the CBSSIM are to operate in real-time and generate telemetry. CBSSIM provides various graphic monitoring interfaces and control functions and supports both pre-launch and after-launch of a communication satellite system. In this paper, the HITL simulator system including CBSSIM, communications payload and PCTS as the medium of interface between CBSSIM and communications payload will be described in aspects of the system architecture, spacecraft models, and simulator operation environment.

Design of a mission network system using SpaceWire for scientific payloads onboard the Arase spacecraft

Science.gov (United States)

Takashima, Takeshi; Ogawa, Emiko; Asamura, Kazushi; Hikishima, Mitsuru

2018-05-01

Arase is a small scientific satellite program conducted by the Institute of Space and Astronautical Science/Japan Aerospace Exploration Agency, which is dedicated to the detailed study of the radiation belts around Earth through in situ observations. In particular, the goal is to directly observe the interaction between plasma waves and particles, which cause the generation of high-energy electrons. To observe the waves and particles in detail, we must record large volumes of burst data with high transmission rates through onboard mission network systems. For this purpose, we developed a high-speed and highly reliable mission network based on SpaceWire, as well as a new and large memory data recorder equipped with a data search function based on observation time (the time index, TI, is the satellite time starting from when the spacecraft is powered on.) with respect to the orbital data generated in large quantities. By adopting a new transaction concept of a ring topology network with SpaceWire, we could secure a redundant mission network system without using large routers and having to suppress the increase in cable weight. We confirmed that their orbit performs as designed.[Figure not available: see fulltext.

Powersail High Power Propulsion System Design Study

Science.gov (United States)

Gulczinski, Frank S., III

2000-11-01

A desire by the United States Air Force to exploit the space environment has led to a need for increased on-orbit electrical power availability. To enable this, the Air Force Research Laboratory Space Vehicles Directorate (AFRL/ VS) is developing Powersail: a two-phased program to demonstrate high power (100 kW to 1 MW) capability in space using a deployable, flexible solar array connected to the host spacecraft using a slack umbilical. The first phase will be a proof-of-concept demonstration at 50 kW, followed by the second phase, an operational system at full power. In support of this program, the AFRL propulsion Directorate's Spacecraft Propulsion Branch (AFRL/PRS ) at Edwards AFB has commissioned a design study of the Powersail High Power Propulsion System. The purpose of this study, the results of which are summarized in this paper, is to perform mission and design trades to identify potential full-power applications (both near-Earth and interplanetary) and the corresponding propulsion system requirements and design. The design study shall farther identify a suitable low power demonstration flight that maximizes risk reduction for the fully operational system. This propulsion system is expected to be threefold: (1) primary propulsion for moving the entire vehicle, (2) a propulsion unit that maintains the solar array position relative to the host spacecraft, and (3) control propulsion for maintaining proper orientation for the flexible solar array.

Reliability considerations in long-life outer planet spacecraft system design

Science.gov (United States)

Casani, E. K.

1975-01-01

A Mariner Jupiter/Saturn mission has been planned for 1977. System reliability questions are discussed, taking into account the actual and design lifetime, causes of mission termination, in-flight failures and their consequences for the mission, and the use of redundancy to avoid failures. The design process employed optimizes the use of proven subsystem and system designs and then makes the necessary improvements to increase the lifetime as required.

The electrical power subsystem design for the high energy solar physics spacecraft concepts

Science.gov (United States)

Kulkarni, Milind

1993-01-01

This paper discusses the Electrical Power Subsystem (EPS) requirements, architecture, design description, performance analysis, and heritage of the components for two spacecraft concepts for the High Energy Solar Physics (HESP) Mission. It summarizes the mission requirements and the spacecraft subsystems and instrument power requirements, and it describes the EPS architecture for both options. A trade study performed on the selection of the solar cells - body mounted versus deployed panels - and the optimum number of panels is also presented. Solar cell manufacturing losses, array manufacturing losses, and the radiation and temperature effects on the GaAs/Ge and Si solar cells were considered part of the trade study and are included in this paper. Solar cell characteristics, cell circuit description, and the solar array area design are presented, as is battery sizing analysis performed based on the power requirements during launch and initial spacecraft operations. This paper discusses Earth occultation periods and the battery power requirements during this period as well as shunt control, battery conditioning, and bus regulation schemes. Design margins, redundancy philosophy, and predicted on-orbit battery and solar cell performance are summarized. Finally, the heritage of the components and technology risk assessment are provided.

A fuzzy logic intelligent diagnostic system for spacecraft integrated vehicle health management

Science.gov (United States)

Wu, G. Gordon

1995-01-01

Due to the complexity of future space missions and the large amount of data involved, greater autonomy in data processing is demanded for mission operations, training, and vehicle health management. In this paper, we develop a fuzzy logic intelligent diagnostic system to perform data reduction, data analysis, and fault diagnosis for spacecraft vehicle health management applications. The diagnostic system contains a data filter and an inference engine. The data filter is designed to intelligently select only the necessary data for analysis, while the inference engine is designed for failure detection, warning, and decision on corrective actions using fuzzy logic synthesis. Due to its adaptive nature and on-line learning ability, the diagnostic system is capable of dealing with environmental noise, uncertainties, conflict information, and sensor faults.

Standardization and Economics of Nuclear Spacecraft, Final Report, Phase I, Sense Study

Energy Technology Data Exchange (ETDEWEB)

1973-03-01

Feasibility and cost benefits of nuclear-powered standardized spacecraft are investigated. The study indicates that two shuttle-launched nuclear-powered spacecraft should be able to serve the majority of unmanned NASA missions anticipated for the 1980's. The standard spacecraft include structure, thermal control, power, attitude control, some propulsion capability and tracking, telemetry, and command subsystems. One spacecraft design, powered by the radioisotope thermoelectric generator, can serve missions requiring up to 450 watts. The other spacecraft design, powered by similar nuclear heat sources in a Brayton-cycle generator, can serve missions requiring up to 21000 watts. Design concepts and trade-offs are discussed. The conceptual designs selected are presented and successfully tested against a variety of missions. The thermal design is such that both spacecraft are capable of operating in any earth orbit and any orientation without modification. Three-axis stabilization is included. Several spacecraft can be stacked in the shuttle payload compartment for multi-mission launches. A reactor-powered thermoelectric generator system, operating at an electric power level of 5000 watts, is briefly studied for applicability to two test missions of divers requirements. A cost analysis indicates that use of the two standardized spacecraft offers sizable savings in comparison with specially designed solar-powered spacecraft. There is a duplicate copy.

Development of solid-gas equilibrium propulsion system for small spacecraft

Science.gov (United States)

Chujo, Toshihiro; Mori, Osamu; Kubo, Yuki

2017-11-01

A phase equilibrium propulsion system is a kind of cold-gas jet in which the phase equilibrium state of the fuel is maintained in a tank and its vapor is ejected when a valve is opened. One such example is a gas-liquid equilibrium propulsion system that uses liquefied gas as fuel. This system was mounted on the IKAROS solar sail and has been demonstrated in orbit. The system has a higher storage efficiency and a lighter configuration than a high-pressure cold-gas jet because the vapor pressure is lower, and is suitable for small spacecraft. However, the system requires a gas-liquid separation device in order to avoid leakage of the liquid, which makes the system complex. As another example of a phase equilibrium propulsion system, we introduce a solid-gas equilibrium propulsion system, which uses a sublimable substance as fuel and ejects its vapor. This system has an even lower vapor pressure and does not require such a separation device, instead requiring only a filter to keep the solid inside the tank. Moreover, the system is much simpler and lighter, making it more suitable for small spacecraft, especially CubeSat-class spacecraft, and the low thrust of the system allows spacecraft motion to be controlled precisely. In addition, the thrust level can be controlled by controlling the temperature of the fuel, which changes the vapor pressure. The present paper introduces the concept of the proposed system, and describes ejection experiments and its evaluation. The basic function of the proposed system is demonstrated in order to verify its usefulness.

Multi-Mission System Architecture Platform: Design and Verification of the Remote Engineering Unit

Science.gov (United States)

Sartori, John

2005-01-01

The Multi-Mission System Architecture Platform (MSAP) represents an effort to bolster efficiency in the spacecraft design process. By incorporating essential spacecraft functionality into a modular, expandable system, the MSAP provides a foundation on which future spacecraft missions can be developed. Once completed, the MSAP will provide support for missions with varying objectives, while maintaining a level of standardization that will minimize redesign of general system components. One subsystem of the MSAP, the Remote Engineering Unit (REU), functions by gathering engineering telemetry from strategic points on the spacecraft and providing these measurements to the spacecraft's Command and Data Handling (C&DH) subsystem. Before the MSAP Project reaches completion, all hardware, including the REU, must be verified. However, the speed and complexity of the REU circuitry rules out the possibility of physical prototyping. Instead, the MSAP hardware is designed and verified using the Verilog Hardware Definition Language (HDL). An increasingly popular means of digital design, HDL programming provides a level of abstraction, which allows the designer to focus on functionality while logic synthesis tools take care of gate-level design and optimization. As verification of the REU proceeds, errors are quickly remedied, preventing costly changes during hardware validation. After undergoing the careful, iterative processes of verification and validation, the REU and MSAP will prove their readiness for use in a multitude of spacecraft missions.

The mechanical design and dynamic testing of the IBEX-H1 electrostatic analyzer spacecraft instrument

Energy Technology Data Exchange (ETDEWEB)

Bernardin, John D [Los Alamos National Laboratory; Baca, Allen G [SNL

2009-01-01

This paper presents the mechanical design, fabrication and dynamic testing of an electrostatic analyzer spacecraft instrument. The functional and environmental requirements combined with limited spacecraft accommodations, resulted in complex component geometries, unique material selections, and difficult fabrication processes. The challenging aspects of the mechanical design and several of the more difficult production processes are discussed. In addition, the successes, failures, and lessons learned from acoustic and random vibration testing of a full-scale prototype instrument are presented.

The Space Technology-7 Disturbance Reduction System Precision Control Flight Validation Experiment Control System Design

Science.gov (United States)

O'Donnell, James R.; Hsu, Oscar C.; Maghami, Peirman G.; Markley, F. Landis

2006-01-01

As originally proposed, the Space Technology-7 Disturbance Reduction System (DRS) project, managed out of the Jet Propulsion Laboratory, was designed to validate technologies required for future missions such as the Laser Interferometer Space Antenna (LISA). The two technologies to be demonstrated by DRS were Gravitational Reference Sensors (GRSs) and Colloidal MicroNewton Thrusters (CMNTs). Control algorithms being designed by the Dynamic Control System (DCS) team at the Goddard Space Flight Center would control the spacecraft so that it flew about a freely-floating GRS test mass, keeping it centered within its housing. For programmatic reasons, the GRSs were descoped from DRS. The primary goals of the new mission are to validate the performance of the CMNTs and to demonstrate precise spacecraft position control. DRS will fly as a part of the European Space Agency (ESA) LISA Pathfinder (LPF) spacecraft along with a similar ESA experiment, the LISA Technology Package (LTP). With no GRS, the DCS attitude and drag-free control systems make use of the sensor being developed by ESA as a part of the LTP. The control system is designed to maintain the spacecraft s position with respect to the test mass, to within 10 nm/the square root of Hz over the DRS science frequency band of 1 to 30 mHz.

How Spacecraft Fly Spaceflight Without Formulae

CERN Document Server

Swinerd, Graham

2009-01-01

About half a century ago a small satellite, Sputnik 1, was launched. The satellite did very little other than to transmit a radio signal to announce its presence in orbit. However, this humble beginning heralded the dawn of the Space Age. Today literally thousands of robotic spacecraft have been launched, many of which have flown to far-flung regions of the Solar System carrying with them the human spirit of scientific discovery and exploration. Numerous other satellites have been launched in orbit around the Earth providing services that support our technological society on the ground. How Spacecraft Fly: Spaceflight Without Formulae by Graham Swinerd focuses on how these spacecraft work. The book opens with a historical perspective of how we have come to understand our Solar System and the Universe. It then progresses through orbital flight, rocket science, the hostile environment within which spacecraft operate, and how they are designed. The concluding chapters give a glimpse of what the 21st century may ...

Automated constraint checking of spacecraft command sequences

Science.gov (United States)

Horvath, Joan C.; Alkalaj, Leon J.; Schneider, Karl M.; Spitale, Joseph M.; Le, Dang

1995-01-01

Robotic spacecraft are controlled by onboard sets of commands called "sequences." Determining that sequences will have the desired effect on the spacecraft can be expensive in terms of both labor and computer coding time, with different particular costs for different types of spacecraft. Specification languages and appropriate user interface to the languages can be used to make the most effective use of engineering validation time. This paper describes one specification and verification environment ("SAVE") designed for validating that command sequences have not violated any flight rules. This SAVE system was subsequently adapted for flight use on the TOPEX/Poseidon spacecraft. The relationship of this work to rule-based artificial intelligence and to other specification techniques is discussed, as well as the issues that arise in the transfer of technology from a research prototype to a full flight system.

Design feasibility via ascent optimality for next-generation spacecraft

Science.gov (United States)

Miele, A.; Mancuso, S.

This paper deals with the optimization of the ascent trajectories for single-stage-sub-orbit (SSSO), single-stage-to-orbit (SSTO), and two-stage-to-orbit (TSTO) rocket-powered spacecraft. The maximum payload weight problem is studied for different values of the engine specific impulse and spacecraft structural factor. The main conclusions are that: feasibility of SSSO spacecraft is guaranteed for all the parameter combinations considered; feasibility of SSTO spacecraft depends strongly on the parameter combination chosen; not only feasibility of TSTO spacecraft is guaranteed for all the parameter combinations considered, but the TSTO payload is several times the SSTO payload. Improvements in engine specific impulse and spacecraft structural factor are desirable and crucial for SSTO feasibility; indeed, aerodynamic improvements do not yield significant improvements in payload. For SSSO, SSTO, and TSTO spacecraft, simple engineering approximations are developed connecting the maximum payload weight to the engine specific impulse and spacecraft structural factor. With reference to the specific impulse/structural factor domain, these engineering approximations lead to the construction of zero-payload lines separating the feasibility region (positive payload) from the unfeasibility region (negative payload).

Gravity Probe B spacecraft description

International Nuclear Information System (INIS)

Bennett, Norman R; Burns, Kevin; Katz, Russell; Kirschenbaum, Jon; Mason, Gary; Shehata, Shawky

2015-01-01

The Gravity Probe B spacecraft, developed, integrated, and tested by Lockheed Missiles and Space Company and later Lockheed Martin Corporation, consisted of structures, mechanisms, command and data handling, attitude and translation control, electrical power, thermal control, flight software, and communications. When integrated with the payload elements, the integrated system became the space vehicle. Key requirements shaping the design of the spacecraft were: (1) the tight mission timeline (17 months, 9 days of on-orbit operation), (2) precise attitude and translational control, (3) thermal protection of science hardware, (4) minimizing aerodynamic, magnetic, and eddy current effects, and (5) the need to provide a robust, low risk spacecraft. The spacecraft met all mission requirements, as demonstrated by dewar lifetime meeting specification, positive power and thermal margins, precision attitude control and drag-free performance, reliable communications, and the collection of more than 97% of the available science data. (paper)

Conceptual Design of an Electric Sail Technology Demonstration Mission Spacecraft

Science.gov (United States)

Wiegmann, Bruce M.

2017-01-01

There is great interest in examining the outer planets of our solar system and Heliopause region (edge of Solar System) and beyond regions of interstellar space by both the Planetary and Heliophysics communities. These needs are well docu-mented in the recent National Academy of Sciences Decadal Surveys. There is significant interest in developing revolutionary propulsion techniques that will enable such Heliopause scientific missions to be completed within 10 to15 years of the launch date. One such enabling propulsion technique commonly known as Electric Sail (E-Sail) propulsion employs positively charged bare wire tethers that extend radially outward from a rotating spacecraft spinning at a rate of one revolution per hour. Around the positively charged bare-wire tethers, a Debye Sheath is created once positive voltage is applied. This sheath stands off of the bare wire tether at a sheath diameter that is proportional to the voltage in the wire coupled with the flux density of solar wind ions within the solar system (or the location of spacecraft in the solar system. The protons that are expended from the sun (solar wind) at 400 to 800 km/sec are electrostatically repelled away from these positively charged Debye sheaths and propulsive thrust is produced via the resulting momentum transfer. The amount of thrust produced is directly proportional to the total wire length. The Marshall Space Flight Center (MSFC) Electric Sail team is currently funded via a two year Phase II NASA Innovative Advanced Concepts (NIAC) awarded in July 2015. The team's current activities are: 1) Developing a Particle in Cell (PIC) numeric engineering model from the experimental data collected at MSFC's Solar Wind Facility on the interaction between simulated solar wind interaction with a charged bare wire that can be applied to a variety of missions, 2) The development of the necessary tether deployers and tethers to enable successful de-ployment of multiple, multi km length bare tethers

MIDN: A spacecraft Micro-dosimeter mission

International Nuclear Information System (INIS)

Pisacane, V. L.; Ziegler, J. F.; Nelson, M. E.; Caylor, M.; Flake, D.; Heyen, L.; Youngborg, E.; Rosenfeld, A. B.; Cucinotta, F.; Zaider, M.; Dicello, J. F.

2006-01-01

MIDN (Micro-dosimetry instrument) is a payload on the MidSTAR-I spacecraft (Midshipman Space Technology Applications Research) under development at the United States Naval Academy. MIDN is a solid-state system being designed and constructed to measure Micro-dosimetric spectra to determine radiation quality factors for space environments. Radiation is a critical threat to the health of astronauts and to the success of missions in low-Earth orbit and space exploration. The system will consist of three separate sensors, one external to the spacecraft, one internal and one embedded in polyethylene. Design goals are mass <3 kg and power <2 W. The MidSTAR-I mission in 2006 will provide an opportunity to evaluate a preliminary version of this system. Its low power and mass makes it useful for the International Space Station and manned and unmanned interplanetary missions as a real-time system to assess and alert astronauts to enhanced radiation environments. (authors)

Advanced dependent pressure vessel (DPV) nickel-hydrogen spacecraft battery design

Energy Technology Data Exchange (ETDEWEB)

Coates, D.K.; Grindstaff, B.; Swaim, O.; Fox, C. [Eagle-Picher Industries, Inc., Joplin, MO (United States). Advanced Systems Operation

1995-12-31

The dependent pressure vessel (DPV) nickel-hydrogen (NiH{sub 2}) battery is being developed as a potential spacecraft battery design for both military and commercial satellites. The limitations of standard NiH{sub 2} individual pressure vessel (IPV) flight battery technology are primarily related to the internal cell design and the battery packaging issues associated with grouping multiple cylindrical cells. The DPV cell design offers higher energy density and reduced cost, while retaining the established IPV technology flight heritage and database. The advanced cell design offers a more efficient mechanical, electrical and thermal cell configuration and a reduced parts count. The geometry of the DPV cell promotes compact, minimum volume packaging and weight efficiency. The DPV battery design offers significant cost and weight savings advantages while providing minimal design risks.

Singular formalism and admissible control of spacecraft with rotating flexible solar array

Directory of Open Access Journals (Sweden)

Lu Dongning

2014-02-01

Full Text Available This paper is concerned with the attitude control of a three-axis-stabilized spacecraft which consists of a central rigid body and a flexible sun-tracking solar array driven by a solar array drive assembly. Based on the linearization of the dynamics of the spacecraft and the modal identities about the flexible and rigid coupling matrices, the spacecraft attitude dynamics is reduced to a formally singular system with periodically varying parameters, which is quite different from a spacecraft with fixed appendages. In the framework of the singular control theory, the regularity and impulse-freeness of the singular system is analyzed and then admissible attitude controllers are designed by Lyapunov’s method. To improve the robustness against system uncertainties, an H∞ optimal control is designed by optimizing the H∞ norm of the system transfer function matrix. Comparative numerical experiments are performed to verify the theoretical results.

The use of twin-screen-based WIMPS in spacecraft control

Science.gov (United States)

Klim, R. D.

1990-10-01

The ergonomic problems of designing a sophisticated Windows Icons Mouse Pop-up (WIMP) based twin screen workstation are outlined. These same problems will be encountered by future spacecraft controllers. The design of a modern, advanced workstation for use on a distributed multicontrol center in a multisatellite control system is outlined. The system uses access control mechanisms to ensure that only authorized personnel can undertake certain operations on the workstation. Rules governing the use of windowing features, screen attributes, icons, keyboard and mouse in spacecraft control are discussed.

Development status of solid polymer electrolyte water electrolysis for manned spacecraft life support systems

Science.gov (United States)

Nuttall, L. J.; Titterington, W. A.

1974-01-01

Details of the design and system verification test results are presented for a six-man-rated oxygen generation system. The system configuration incorporates components and instrumentation for computer-controlled operation with automatic start-up/shutdown sequencing, fault detection and isolation, and with self-contained sensors and controls for automatic safe emergency shutdown. All fluid and electrical components, sensors, and electronic controls are designed to be easily maintainable under zero-gravity conditions. On-board component spares are utilized in the system concept to sustain long-term operation (six months minimum) in a manned spacecraft application. The system is centered on a 27-cell solid polymer electrolyte water electrolysis module which, combined with the associated system components and controls, forms a total system envelope 40 in. high, 40 in. wide, and 30 in. deep.

Control system design and analysis using the INteractive Controls Analysis (INCA) program

Science.gov (United States)

Bauer, Frank H.; Downing, John P.

1987-01-01

The INteractive Controls Analysis (INCA) program was developed at the Goddard Space Flight Center to provide a user friendly efficient environment for the design and analysis of linear control systems. Since its inception, INCA has found extensive use in the design, development, and analysis of control systems for spacecraft, instruments, robotics, and pointing systems. Moreover, the results of the analytic tools imbedded in INCA have been flight proven with at least three currently orbiting spacecraft. This paper describes the INCA program and illustrates, using a flight proven example, how the package can perform complex design analyses with relative ease.

JEM-EUSO Design for Accommodation on the SpaceX Dragon Spacecraft

Science.gov (United States)

Christl, Mark

2013-01-01

The JEM-EUSO mission has been planned for launch on JAXA's H2 Launch Vehicle. Recently, the SpaceX Dragon spacecraft has emerged as an alternative payload carrier for JEM-EUSO. This paper will discuss a concept for the re-design of JEM-EUSO so that it can be launched on Dragon.

Design of a gigawatt space solar power satellite using optical concentrator system

Science.gov (United States)

Dessanti, B.; Komerath, N.; Shah, S.

A 1-gigawatt space solar power satellite using a large array of individually pointable optical elements is identified as the key mass element of a large scale space solar power architecture using the Space Power Grid concept. The proposed satellite design enables a significant increase in specific power. Placed in sun-synchronous dynamic orbits near 2000km altitude, these satellites can maintain the constant solar view requirement of GEO-based architectures, while greatly reducing the beaming distance required, decreasing the required antenna size and in turn the overall system mass. The satellite uses an array of individually pointable optical elements (which we call a Mirasol Concentrator Array) to concentrate solar energy to an intensified feed target that feeds into the main heater of the spacecraft, similar conceptually to heliostat arrays. The spacecraft then utilizes Brayton cycle conversion to take advantage of non-linear power level scaling in order to generate high specific power values. Using phase array antennas, the power is then beamed at a millimeter wave frequency of 220GHz down to Earth. The design of the Mirasol concentrator system will be described and a detailed mass estimation of the system is developed. The technical challenges of pointing the elements and maintaining constant solar view is investigated. An end-to-end efficiency analysis is performed. Subsystem designs for the spacecraft are outlined. A detailed mass budget is refined to reflect reductions in uncertainty of the spacecraft mass, particularly in the Mirasol system. One of the key mass drivers of the spacecraft is the active thermal control system. The design of a lightweight thermal control system utilizing graphene sheets is also detailed.

Systems Engineering Using Heritage Spacecraft Technology: Lessons Learned from Discovery and New Frontiers Deep Space Missions

Science.gov (United States)

Barley, Bryan; Newhouse, Marilyn; Clardy, Dennon

2011-01-01

In the design and development of complex spacecraft missions, project teams frequently assume the use of advanced technology or heritage systems to enable a mission or reduce the overall mission risk and cost. As projects proceed through the development life cycle, increasingly detailed knowledge of the advanced or heritage systems and the system environment identifies unanticipated issues that result in cost overruns or schedule impacts. The Discovery & New Frontiers (D&NF) Program Office recently studied cost overruns and schedule delays resulting from advanced technology or heritage assumptions for 6 D&NF missions. The goal was to identify the underlying causes for the overruns and delays, and to develop practical mitigations to assist the D&NF projects in identifying potential risks and controlling the associated impacts to proposed mission costs and schedules. The study found that the cost and schedule growth did not result from technical hurdles requiring significant technology development. Instead, systems engineering processes did not identify critical issues early enough in the design cycle to ensure project schedules and estimated costs address the inherent risks. In general, the overruns were traceable to: inadequate understanding of the heritage system s behavior within the proposed spacecraft design and mission environment; an insufficient level of experience with the heritage system; or an inadequate scoping of the system-wide impacts necessary to implement the heritage or advanced technology. This presentation summarizes the study s findings and offers suggestions for improving the project s ability to identify and manage the risks inherent in the technology and heritage design solution.

Wheel speed management control system for spacecraft

Science.gov (United States)

Goodzeit, Neil E. (Inventor); Linder, David M. (Inventor)

1991-01-01

A spacecraft attitude control system uses at least four reaction wheels. In order to minimize reaction wheel speed and therefore power, a wheel speed management system is provided. The management system monitors the wheel speeds and generates a wheel speed error vector. The error vector is integrated, and the error vector and its integral are combined to form a correction vector. The correction vector is summed with the attitude control torque command signals for driving the reaction wheels.

Value-informed space systems design and acquisition

Science.gov (United States)

Brathwaite, Joy

Investments in space systems are substantial, indivisible, and irreversible, characteristics that make them high-risk, especially when coupled with an uncertain demand environment. Traditional approaches to system design and acquisition, derived from a performance- or cost-centric mindset, incorporate little information about the spacecraft in relation to its environment and its value to its stakeholders. These traditional approaches, while appropriate in stable environments, are ill-suited for the current, distinctly uncertain, and rapidly changing technical and economic conditions; as such, they have to be revisited and adapted to the present context. This thesis proposes that in uncertain environments, decision-making with respect to space system design and acquisition should be value-based, or at a minimum value-informed. This research advances the value-centric paradigm by providing the theoretical basis, foundational frameworks, and supporting analytical tools for value assessment of priced and unpriced space systems. For priced systems, stochastic models of the market environment and financial models of stakeholder preferences are developed and integrated with a spacecraft-sizing tool to assess the system's net present value. The analytical framework is applied to a case study of a communications satellite, with market, financial, and technical data obtained from the satellite operator, Intelsat. The case study investigates the implications of the value-centric versus the cost-centric design and acquisition choices. Results identify the ways in which value-optimal spacecraft design choices are contingent on both technical and market conditions, and that larger spacecraft for example, which reap economies of scale benefits, as reflected by their decreasing cost-per-transponder, are not always the best (most valuable) choices. Market conditions and technical constraints for which convergence occurs between design choices under a cost-centric and a value

(abstract) ARGOS: a System to Monitor Ulysses Nutation and Thruster Firings from Variations of the Spacecraft Radio Signal

Science.gov (United States)

McElrath, T. P.; Cangahuala, L. A.; Miller, K. J.; Stravert, L. R.; Garcia-Perez, Raul

1995-01-01

Ulysses is a spin-stabilized spacecraft that experienced significant nutation after its launch in October 1990. This was due to the Sun-spacecraft-Earth geometry, and a study of the phenomenon predicted that the nutation would again be a problem during 1994-95. The difficulty of obtaining nutation estimates in real time from the spacecraft telemetry forced the ESA/NASA Ulysses Team to explore alternative information sources. The work performed by the ESA Operations Team provided a model for a system that uses the radio signal strength measurements to monitor the spacecraft dynamics. These measurements (referred to as AGC) are provided once per second by the tracking stations of the DSN. The system was named ARGOS (Attitude Reckoning from Ground Observable Signals) after the ever-vigilant, hundred-eyed giant of Greek Mythology. The ARGOS design also included Doppler processing, because Doppler shifts indicate thruster firings commanded by the active nutation control carried out onboard the spacecraft. While there is some visibility into thruster activity from telemetry, careful processing of the high-sample-rate Doppler data provides an accurate means of detecting the presence and time of thruster firings. DSN Doppler measurements are available at a ten-per-second rate in the same tracking data block as the AGC data.

Lifetime of a spacecraft around a synchronous system of asteroids using a dipole model

Science.gov (United States)

dos Santos, Leonardo Barbosa Torres; de Almeida Prado, Antonio F. Bertachini; Sanchez, Diogo Merguizo

2017-11-01

Space missions allow us to expand our knowledge about the origin of the solar system. It is believed that asteroids and comets preserve the physical characteristics from the time that the solar system was created. For this reason, there was an increase of missions to asteroids in the past few years. To send spacecraft to asteroids or comets is challenging, since these objects have their own characteristics in several aspects, such as size, shape, physical properties, etc., which are often only discovered after the approach and even after the landing of the spacecraft. These missions must be developed with sufficient flexibility to adjust to these parameters, which are better determined only when the spacecraft reaches the system. Therefore, conducting a dynamic investigation of a spacecraft around a multiple asteroid system offers an extremely rich environment. Extracting accurate information through analytical approaches is quite challenging and requires a significant number of restrictive assumptions. For this reason, a numerical approach to the dynamics of a spacecraft in the vicinity of a binary asteroid system is offered in this paper. In the present work, the equations of the Restricted Synchronous Four-Body Problem (RSFBP) are used to model a binary asteroid system. The main objective of this work is to construct grids of initial conditions, which relates semi-major axis and eccentricity, in order to quantify the lifetime of a spacecraft when released close to the less massive body of the binary system (modeled as a rotating mass dipole). We performed an analysis of the lifetime of the spacecraft considering several mass ratios of a binary system of asteroids and investigating the behavior of a spacecraft in the vicinity of this system. We analyze direct and retrograde orbits. This study investigated orbits that survive for at least 500 orbital periods of the system (which is approximately one year), then not colliding or escaping from the system during this

Galileo spacecraft inertial sensors in-flight calibration design

Science.gov (United States)

Jahanshahi, M. H.; Lai, J. Y.

1983-01-01

The successful navigation of Galileo depends on accurate trajectory correction maneuvers (TCM's) performed during the mission. A set of Inertial Sensor (INS) units, comprised of gyros and accelerometers, mounted on the spacecraft, are utilized to control and monitor the performance of the TCM's. To provide the optimum performance, in-flight calibrations of INS are planned. These calibrations will take place on a regular basis. In this paper, a mathematical description is given of the data reduction technique used in analyzing a typical set of calibration data. The design of the calibration and the inertial sensor error models, necessary for the above analysis, are delineated in detail.

Operator Station Design System - A computer aided design approach to work station layout

Science.gov (United States)

Lewis, J. L.

1979-01-01

The Operator Station Design System is resident in NASA's Johnson Space Center Spacecraft Design Division Performance Laboratory. It includes stand-alone minicomputer hardware and Panel Layout Automated Interactive Design and Crew Station Assessment of Reach software. The data base consists of the Shuttle Transportation System Orbiter Crew Compartment (in part), the Orbiter payload bay and remote manipulator (in part), and various anthropometric populations. The system is utilized to provide panel layouts, assess reach and vision, determine interference and fit problems early in the design phase, study design applications as a function of anthropometric and mission requirements, and to accomplish conceptual design to support advanced study efforts.

Automating Trend Analysis for Spacecraft Constellations

Science.gov (United States)

Davis, George; Cooter, Miranda; Updike, Clark; Carey, Everett; Mackey, Jennifer; Rykowski, Timothy; Powers, Edward I. (Technical Monitor)

2001-01-01

missions such as DRACO with the intent that mission operations costs be significantly reduced. The goal of the Constellation Spacecraft Trend Analysis Toolkit (CSTAT) project is to serve as the pathfinder for a fully automated trending system to support spacecraft constellations. The development approach to be taken is evolutionary. In the first year of the project, the intent is to significantly advance the state of the art in current trending systems through improved functionality and increased automation. In the second year, the intent is to add an expert system shell, likely through the adaptation of an existing commercial-off-the-shelf (COTS) or government-off-the-shelf (GOTS) tool to implement some level of the trending intelligence that humans currently provide in manual operations. In the third year, the intent is to infuse the resulting technology into a near-term constellation or formation-flying mission to test it and gain experience in automated trending. The lessons learned from the real missions operations experience will then be used to improve the system, and to ultimately incorporate it into a fully autonomous, closed-loop mission operations system that is truly capable of supporting large constellations. In this paper, the process of automating trend analysis for spacecraft constellations will be addressed. First, the results of a survey on automation in spacecraft mission operations in general, and in trending systems in particular will be presented to provide an overview of the current state of the art. Next, a rule-based model for implementing intelligent spacecraft subsystem trending will be then presented, followed by a survey of existing COTS/GOTS tools that could be adapted for implementing such a model. The baseline design and architecture of the CSTAT system will be presented. Finally, some results obtained from initial software tests and demonstrations will be presented.

Robust Parametric Control of Spacecraft Rendezvous

Directory of Open Access Journals (Sweden)

Dake Gu

2014-01-01

Full Text Available This paper proposes a method to design the robust parametric control for autonomous rendezvous of spacecrafts with the inertial information with uncertainty. We consider model uncertainty of traditional C-W equation to formulate the dynamic model of the relative motion. Based on eigenstructure assignment and model reference theory, a concise control law for spacecraft rendezvous is proposed which could be fixed through solving an optimization problem. The cost function considers the stabilization of the system and other performances. Simulation results illustrate the robustness and effectiveness of the proposed control.

Development of the solar array deployment and drive system for the XTE spacecraft

Science.gov (United States)

Farley, Rodger; Ngo, Son

1995-01-01

The X-ray Timing Explorer (XTE) spacecraft is a NASA science low-earth orbit explorer-class satellite to be launched in 1995, and is an in-house Goddard Space Flight Center (GSFC) project. It has two deployable aluminum honeycomb solar array wings with each wing being articulated by a single axis solar array drive assembly. This paper will address the design, the qualification testing, and the development problems as they surfaced of the Solar Array Deployment and Drive System.

Automating a spacecraft electrical power system using expert systems

Science.gov (United States)

Lollar, L. F.

1991-01-01

Since Skylab, Marshall Space Flight Center (MSFC) has recognized the need for large electrical power systems (EPS's) in upcoming Spacecraft. The operation of the spacecraft depends on the EPS. Therefore, it must be efficient, safe, and reliable. In 1978, as a consequence of having to supply a large number of EPS personnel to monitor and control Skylab, the Electrical power Branch of MSFC began the autonomously managed power system (AMPS) project. This project resulted in the assembly of a 25-kW high-voltage dc test facility and provided the means of getting man out of the loop as much as possible. AMPS includes several embedded controllers which allow a significant level of autonomous operation. More recently, the Electrical Division at MSFC has developed the space station module power management and distribution (SSM/PMAD) breadboard to investigate managing and distributing power in the Space Station Freedom habitation and laboratory modules. Again, the requirement for a high level of autonomy for the efficient operation over the lifetime of the station and for the benefits of enhanced safety has been demonstrated. This paper describes the two breadboards and the hierarchical approach to automation which was developed through these projects.

A small spacecraft for multipoint measurement of ionospheric plasma

Science.gov (United States)

Roberts, T. M.; Lynch, K. A.; Clayton, R. E.; Weiss, J.; Hampton, D. L.

2017-07-01

Measurement of ionospheric plasma is often performed by a single in situ device or remotely using cameras and radar. This article describes a small, low-resource, deployed spacecraft used as part of a local, multipoint measurement network. A B-field aligned sounding rocket ejects four of these spin-stabilized spacecraft in a cross pattern. In this application, each spacecraft carries two retarding potential analyzers which are used to determine plasma density, flow, and ion temperature. An inertial measurement unit and a light-emitting diode array are used to determine the position and orientation of the devices after deployment. The design of this spacecraft is first described, and then results from a recent test flight are discussed. This flight demonstrated the successful operation of the deployment mechanism and telemetry systems, provided some preliminary plasma measurements in a simple mid-latitude environment, and revealed several design issues.

Mechanical design and vibro-acoustic testing of ultrathin carbon foils for a spacecraft instrument

Energy Technology Data Exchange (ETDEWEB)

Bernardin, John D [Los Alamos National Laboratory; Baca, Allen G [SNL

2009-01-01

IBEX-Hi is an electrostatic analyzer spacecraft instrument designed to measure the energy and flux distribution of energetic neutral atoms (ENAs) emanating from the interaction zone between the Earth's solar system and the Milky Way galaxy. A key element to this electro-optic instrument is an array of fourteen carbon foils that are used to ionize the ENAs. The foils are comprised of an ultrathin (50-100 {angstrom} thick) layer of carbon suspended across the surface of an electroformed Nickel wire screen, which in turn is held taught by a metal frame holder. The electro formed orthogonal screen has square wire elements, 12.7 {micro}m thick, with a pitch of 131.1 wires/cm. Each foil holder has an open aperture approximately 5 cm by 2.5 cm. Designing and implementing foil holders with such a large surface area has not been attempted for spaceflight in the past and has proven to be extremely challenging. The delicate carbon foils are subject to fatigue failure from the large acoustic and vibration loads that they will be exposed to during launch of the spacecraft. This paper describes the evolution of the foil holder design from previous space instrument applications to a flight-like IBEX-Hi prototype. Vibro-acoustic qualification tests of the IBEX-Hi prototype instrument and the resulting failure of several foils are summarized. This is followed by a discussion of iterative foil holder design modifications and laser vibrometer modal testing to support future fatigue failure analyses, along with additional acoustic testing of the IBEX-Hi prototype instrument. The results of these design and testing activities are merged and the resulting flight-like foil holder assembly is proposed.

Attitude tracking control of flexible spacecraft with large amplitude slosh

Science.gov (United States)

Deng, Mingle; Yue, Baozeng

2017-12-01

This paper is focused on attitude tracking control of a spacecraft that is equipped with flexible appendage and partially filled liquid propellant tank. The large amplitude liquid slosh is included by using a moving pulsating ball model that is further improved to estimate the settling location of liquid in microgravity or a zero-g environment. The flexible appendage is modelled as a three-dimensional Bernoulli-Euler beam, and the assumed modal method is employed. A hybrid controller that combines sliding mode control with an adaptive algorithm is designed for spacecraft to perform attitude tracking. The proposed controller has proved to be asymptotically stable. A nonlinear model for the overall coupled system including spacecraft attitude dynamics, liquid slosh, structural vibration and control action is established. Numerical simulation results are presented to show the dynamic behaviors of the coupled system and to verify the effectiveness of the control approach when the spacecraft undergoes the disturbance produced by large amplitude slosh and appendage vibration. Lastly, the designed adaptive algorithm is found to be effective to improve the precision of attitude tracking.

Disinfection of Spacecraft Potable Water Systems by Passivation with Ionic Silver

Science.gov (United States)

Birmele, Michele N.; McCoy, LaShelle e.; Roberts, Michael S.

2011-01-01

Microbial growth is common on wetted surfaces in spacecraft environmental control and life support systems despite the use of chemical and physical disinfection methods. Advanced control technologies are needed to limit microorganisms and increase the reliability of life support systems required for long-duration human missions. Silver ions and compounds are widely used as antimicrobial agents for medical applications and continue to be used as a residual biocide in some spacecraft water systems. The National Aeronautics and Space Administration (NASA) has identified silver fluoride for use in the potable water system on the next generation spacecraft. Due to ionic interactions between silver fluoride in solution and wetted metallic surfaces, ionic silver is rapidly depleted from solution and loses its antimicrobial efficacy over time. This report describes research to prolong the antimicrobial efficacy of ionic silver by maintaining its solubility. Three types of metal coupons (lnconel 718, Stainless Steel 316, and Titanium 6AI-4V) used in spacecraft potable water systems were exposed to either a continuous flow of water amended with 0.4 mg/L ionic silver fluoride or to a static, pre-treatment passivation in 50 mg/L ionic silver fluoride with or without a surface oxidation pre-treatment. Coupons were then challenged in a high-shear, CDC bioreactor (BioSurface Technologies) by exposure to six bacteria previously isolated from spacecraft potable water systems. Continuous exposure to 0.4 mg/L ionic silver over the course of 24 hours during the flow phase resulted in a >7-log reduction. The residual effect of a 24-hour passivation treatment in 50 mg/L of ionic silver resulted in a >3-log reduction, whereas a two-week treatment resulted in a >4-log reduction. Results indicate that 0.4 mg/L ionic silver is an effective biocide against many bacteria and that a prepassivation of metal surfaces with silver can provide additional microbial control.

Spacecraft Attitude Control in Hamiltonian Framework

DEFF Research Database (Denmark)

Wisniewski, Rafal

2000-01-01

The objective of this paper is to give a design scheme for attitude control algorithms of a generic spacecraft. Along with the system model formulated in the Hamilton's canonical form the algorithm uses information about a required potential energy and a dissipative term. The control action...

Space power systems--''Spacecraft 2000''

International Nuclear Information System (INIS)

Faymon, K.A.

1985-01-01

The National Space programs of the 21st century will require abundant and relatively low cost power and energy produced by high reliability-low mass systems. Advancement of current power system related technologies will enable the U.S. to realize increased scientific payload for government missions or increased revenue producing payload for commercial space endeavors. Autonomous, unattended operation will be a highly desirable characteristic of these advanced power systems. Those space power-energy related technologies, which will comprise the space craft of the late 1990's and the early 2000's, will evolve from today's state-of-the-art systems and those long term technology development programs presently in place. However, to foster accelerated development of the more critical technologies which have the potential for high-payoffs, additional programs will be proposed and put in place between now and the end of the century. Such a program is ''Spacecraft 2000'', which is described in this paper

Mesh Network Architecture for Enabling Inter-Spacecraft Communication

Science.gov (United States)

Becker, Christopher; Merrill, Garrick

2017-01-01

To enable communication between spacecraft operating in a formation or small constellation, a mesh network architecture was developed and tested using a time division multiple access (TDMA) communication scheme. The network is designed to allow for the exchange of telemetry and other data between spacecraft to enable collaboration between small spacecraft. The system uses a peer-to-peer topology with no central router, so that it does not have a single point of failure. The mesh network is dynamically configurable to allow for addition and subtraction of new spacecraft into the communication network. Flight testing was performed using an unmanned aerial system (UAS) formation acting as a spacecraft analogue and providing a stressing environment to prove mesh network performance. The mesh network was primarily devised to provide low latency, high frequency communication but is flexible and can also be configured to provide higher bandwidth for applications desiring high data throughput. The network includes a relay functionality that extends the maximum range between spacecraft in the network by relaying data from node to node. The mesh network control is implemented completely in software making it hardware agnostic, thereby allowing it to function with a wide variety of existing radios and computing platforms..

Optimal Electrical Energy Slewing for Reaction Wheel Spacecraft

Science.gov (United States)

Marsh, Harleigh Christian

The results contained in this dissertation contribute to a deeper level of understanding to the energy required to slew a spacecraft using reaction wheels. This work addresses the fundamental manner in which spacecrafts are slewed (eigenaxis maneuvering), and demonstrates that this conventional maneuver can be dramatically improved upon in regards to reduction of energy, dissipative losses, as well as peak power. Energy is a fundamental resource that effects every asset, system, and subsystem upon a spacecraft, from the attitude control system which orients the spacecraft, to the communication subsystem to link with ground stations, to the payloads which collect scientific data. For a reaction wheel spacecraft, the attitude control system is a particularly heavy load on the power and energy resources on a spacecraft. The central focus of this dissertation is reducing the burden which the attitude control system places upon the spacecraft in regards to electrical energy, which is shown in this dissertation to be a challenging problem to computationally solve and analyze. Reducing power and energy demands can have a multitude of benefits, spanning from the initial design phase, to in-flight operations, to potentially extending the mission life of the spacecraft. This goal is approached from a practical standpoint apropos to an industry-flight setting. Metrics to measure electrical energy and power are developed which are in-line with the cost associated to operating reaction wheel based attitude control systems. These metrics are incorporated into multiple families of practical high-dimensional constrained nonlinear optimal control problems to reduce the electrical energy, as well as the instantaneous power burdens imposed by the attitude control system upon the spacecraft. Minimizing electrical energy is shown to be a problem in L1 optimal control which is nonsmooth in regards to state variables as well as the control. To overcome the challenge of nonsmoothness, a

Space Environments and Spacecraft Effects Organization Concept

Science.gov (United States)

Edwards, David L.; Burns, Howard D.; Miller, Sharon K.; Porter, Ron; Schneider, Todd A.; Spann, James F.; Xapsos, Michael

2012-01-01

The National Aeronautics and Space Administration (NASA) is embarking on a course to expand human presence beyond Low Earth Orbit (LEO) while also expanding its mission to explore the solar system. Destinations such as Near Earth Asteroids (NEA), Mars and its moons, and the outer planets are but a few of the mission targets. Each new destination presents an opportunity to increase our knowledge of the solar system and the unique environments for each mission target. NASA has multiple technical and science discipline areas specializing in specific space environments disciplines that will help serve to enable these missions. To complement these existing discipline areas, a concept is presented focusing on the development of a space environments and spacecraft effects (SENSE) organization. This SENSE organization includes disciplines such as space climate, space weather, natural and induced space environments, effects on spacecraft materials and systems and the transition of research information into application. This space environment and spacecraft effects organization will be composed of Technical Working Groups (TWG). These technical working groups will survey customers and users, generate products, and provide knowledge supporting four functional areas: design environments, engineering effects, operational support, and programmatic support. The four functional areas align with phases in the program mission lifecycle and are briefly described below. Design environments are used primarily in the mission concept and design phases of a program. Engineering effects focuses on the material, component, sub-system and system-level selection and the testing to verify design and operational performance. Operational support provides products based on real time or near real time space weather to mission operators to aid in real time and near-term decision-making. The programmatic support function maintains an interface with the numerous programs within NASA, other federal

Benefits of Spacecraft Level Vibration Testing

Science.gov (United States)

Gordon, Scott; Kern, Dennis L.

2015-01-01

NASA-HDBK-7008 Spacecraft Level Dynamic Environments Testing discusses the approaches, benefits, dangers, and recommended practices for spacecraft level dynamic environments testing, including vibration testing. This paper discusses in additional detail the benefits and actual experiences of vibration testing spacecraft for NASA Goddard Space Flight Center (GSFC) and Jet Propulsion Laboratory (JPL) flight projects. JPL and GSFC have both similarities and differences in their spacecraft level vibration test approach: JPL uses a random vibration input and a frequency range usually starting at 5 Hz and extending to as high as 250 Hz. GSFC uses a sine sweep vibration input and a frequency range usually starting at 5 Hz and extending only to the limits of the coupled loads analysis (typically 50 to 60 Hz). However, both JPL and GSFC use force limiting to realistically notch spacecraft resonances and response (acceleration) limiting as necessary to protect spacecraft structure and hardware from exceeding design strength capabilities. Despite GSFC and JPL differences in spacecraft level vibration test approaches, both have uncovered a significant number of spacecraft design and workmanship anomalies in vibration tests. This paper will give an overview of JPL and GSFC spacecraft vibration testing approaches and provide a detailed description of spacecraft anomalies revealed.

Distributed Control Architectures for Precision Spacecraft Formations, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — LaunchPoint Technologies, Inc. (LaunchPoint) proposes to develop synthesis methods and design architectures for distributed control systems in precision spacecraft...

Autonomous spacecraft rendezvous and docking

Science.gov (United States)

Tietz, J. C.; Almand, B. J.

A storyboard display is presented which summarizes work done recently in design and simulation of autonomous video rendezvous and docking systems for spacecraft. This display includes: photographs of the simulation hardware, plots of chase vehicle trajectories from simulations, pictures of the docking aid including image processing interpretations, and drawings of the control system strategy. Viewgraph-style sheets on the display bulletin board summarize the simulation objectives, benefits, special considerations, approach, and results.

A Spacecraft Housekeeping System-on-Chip in a Radiation Hardened Structured ASIC

Science.gov (United States)

Suarez, George; DuMonthier, Jeffrey J.; Sheikh, Salman S.; Powell, Wesley A.; King, Robyn L.

2012-01-01

Housekeeping systems are essential to health monitoring of spacecraft and instruments. Typically, sensors are distributed across various sub-systems and data is collected using components such as analog-to-digital converters, analog multiplexers and amplifiers. In most cases programmable devices are used to implement the data acquisition control and storage, and the interface to higher level systems. Such discrete implementations require additional size, weight, power and interconnect complexity versus an integrated circuit solution, as well as the qualification of multiple parts. Although commercial devices are readily available, they are not suitable for space applications due the radiation tolerance and qualification requirements. The Housekeeping System-o n-A-Chip (HKSOC) is a low power, radiation hardened integrated solution suitable for spacecraft and instrument control and data collection. A prototype has been designed and includes a wide variety of functions including a 16-channel analog front-end for driving and reading sensors, analog-to-digital and digital-to-analog converters, on-chip temperature sensor, power supply current sense circuits, general purpose comparators and amplifiers, a 32-bit processor, digital I/O, pulse-width modulation (PWM) generators, timers and I2C master and slave serial interfaces. In addition, the device can operate in a bypass mode where the processor is disabled and external logic is used to control the analog and mixed signal functions. The device is suitable for stand-alone or distributed systems where multiple chips can be deployed across different sub-systems as intelligent nodes with computing and processing capabilities.

Operational Philosophy Concerning Manned Spacecraft Cabin Leaks

Science.gov (United States)

DeSimpelaere, Edward

2011-01-01

The last thirty years have seen the Space Shuttle as the prime United States spacecraft for manned spaceflight missions. Many lessons have been learned about spacecraft design and operation throughout these years. Over the next few decades, a large increase of manned spaceflight in the commercial sector is expected. This will result in the exposure of commercial crews and passengers to many of the same risks crews of the Space Shuttle have encountered. One of the more dire situations that can be encountered is the loss of pressure in the habitable volume of the spacecraft during on orbit operations. This is referred to as a cabin leak. This paper seeks to establish a general cabin leak response philosophy with the intent of educating future spacecraft designers and operators. After establishing a relative definition for a cabin leak, the paper covers general descriptions of detection equipment, detection methods, and general operational methods for management of a cabin leak. Subsequently, all these items are addressed from the perspective of the Space Shuttle Program, as this will be of the most value to future spacecraft due to similar operating profiles. Emphasis here is placed upon why and how these methods and philosophies have evolved to meet the Space Shuttle s needs. This includes the core ideas of: considerations of maintaining higher cabin pressures vs. lower cabin pressures, the pros and cons of a system designed to feed the leak with gas from pressurized tanks vs. using pressure suits to protect against lower cabin pressures, timeline and consumables constraints, re-entry considerations with leaks of unknown origin, and the impact the International Space Station (ISS) has had to the standard Space Shuttle cabin leak response philosophy. This last item in itself includes: procedural management differences, hardware considerations, additional capabilities due to the presence of the ISS and its resource, and ISS docking/undocking considerations with a

Low-Cost, Class D Testing of Spacecraft Photovoltaic Systems Can Reduce Risk

Science.gov (United States)

Forgione, Joshua B.; Kojima, Gilbert K.; Hanel, Robert; Mallinson, Mark V.

2014-01-01

The end-to-end verification of a spacecraft photovoltaic power generation system requires light! Specifically, the standard practice for doing so is the Large Area Pulsed Solar Simulation (LAPSS). A LAPSS test can characterize a photovoltaic system's efficiency via its response to rapidly applied impulses of simulated sunlight. However, a Class D program on a constrained budget and schedule may not have the resources to ship an entire satellite for a LAPSS test alone. Such was the case with the Lunar Atmospheric and Dust Environment Explorer (LADEE) program, which was also averse to the risk of hardware damage during shipment. When the Electrical Power System (EPS) team was denied a spacecraft-level LAPSS test, the lack of an end-to-end power generation test elevated to a project-level technical risk. The team pulled together very limited resources to not only eliminate the risk, but build a process to monitor the health of the system through mission operations. We discuss a process for performing a low-cost, end-to-end test of the LADEE photovoltaic system. The approach combines system-level functional test, panel-level performance results, and periodic inspection (and repair) up until launch. Following launch, mission operations tools are utilized to assess system performance based on a scant amount of data. The process starts in manufacturing at the subcontractor. The panel manufacturer provides functional test and LAPSS data on each individual panel. We apply an initial assumption that the per-panel performance is sufficient to meet the power generation requirements. The manufacturer's data is also carried as the performance allocation for each panel during EPS system modeling and initial mission operations. During integration and test, a high-power, professional theater lamp system provides simulated sunlight to each panel on the spacecraft, thereby permitting a true end-to-end system test. A passing test results in a step response to nearly full-rated current

A Return to Innovative Engineering Design, Critical Thinking and Systems Engineering

Science.gov (United States)

Camarda, Charles J.

2007-01-01

I believe we are facing a critical time where innovative engineering design is of paramount importance to the success of our aerospace industry. However, the very qualities and attributes necessary for enhancing, educating, and mentoring a creative spirit are in decline in important areas. The importance of creativity and innovation in this country was emphasized by a special edition of the Harvard Business Review OnPoint entitled: "The Creative Company" which compiled a series of past and present articles on the subject of creativity and innovation and stressed its importance to our national economy. There is also a recognition of a lack of engineering, critical thinking and problem-solving skills in our education systems and a trend toward trying to enhance those skills by developing K-12 educational programs such as Project Lead the Way, "Science for All Americans", Benchmarks 2061 , etc. In addition, with respect to spacecraft development, we have a growing need for young to mid-level engineers with appropriate experience and skills in spacecraft design, development, analysis, testing, and systems engineering. As the Director of Engineering at NASA's Johnson Space Center, I realized that sustaining engineering support of an operational human spacecraft such as the Space Shuttle is decidedly different than engineering design and development skills necessary for designing a new spacecraft such as the Crew Exploration Vehicle of the Constellation Program. We learned a very important lesson post Columbia in that the Space Shuttle is truly an experimental and not an operational vehicle and the strict adherence to developed rules and processes and chains of command of an inherently bureaucratic organizational structure will not protect us from a host of known unknowns let alone unknown unknowns. There are no strict rules, processes, or procedures for understanding anomalous results of an experiment, anomalies with an experimental spacecraft like Shuttle, or in the

Adaptive Jacobian Fuzzy Attitude Control for Flexible Spacecraft Combined Attitude and Sun Tracking System

Science.gov (United States)

Chak, Yew-Chung; Varatharajoo, Renuganth

2016-07-01

Many spacecraft attitude control systems today use reaction wheels to deliver precise torques to achieve three-axis attitude stabilization. However, irrecoverable mechanical failure of reaction wheels could potentially lead to mission interruption or total loss. The electrically-powered Solar Array Drive Assemblies (SADA) are usually installed in the pitch axis which rotate the solar arrays to track the Sun, can produce torques to compensate for the pitch-axis wheel failure. In addition, the attitude control of a flexible spacecraft poses a difficult problem. These difficulties include the strong nonlinear coupled dynamics between the rigid hub and flexible solar arrays, and the imprecisely known system parameters, such as inertia matrix, damping ratios, and flexible mode frequencies. In order to overcome these drawbacks, the adaptive Jacobian tracking fuzzy control is proposed for the combined attitude and sun-tracking control problem of a flexible spacecraft during attitude maneuvers in this work. For the adaptation of kinematic and dynamic uncertainties, the proposed scheme uses an adaptive sliding vector based on estimated attitude velocity via approximate Jacobian matrix. The unknown nonlinearities are approximated by deriving the fuzzy models with a set of linguistic If-Then rules using the idea of sector nonlinearity and local approximation in fuzzy partition spaces. The uncertain parameters of the estimated nonlinearities and the Jacobian matrix are being adjusted online by an adaptive law to realize feedback control. The attitude of the spacecraft can be directly controlled with the Jacobian feedback control when the attitude pointing trajectory is designed with respect to the spacecraft coordinate frame itself. A significant feature of this work is that the proposed adaptive Jacobian tracking scheme will result in not only the convergence of angular position and angular velocity tracking errors, but also the convergence of estimated angular velocity to

Earth Observatory Satellite system definition study. Report 3: Design cost trade-off studies and recommendations

Science.gov (United States)

1974-01-01

An analysis of the design and cost tradeoff aspects of the Earth Observatory Satellite (EOS) development is presented. The design/cost factors that affect a series of mission/system level concepts are discussed. The subjects considered are as follows: (1) spacecraft subsystem cost tradeoffs, (2) ground system cost tradeoffs, and (3) program cost summary. Tables of data are provided to summarize the results of the analyses. Illustrations of the various spacecraft configurations are included.

Dual shear plate power processor packaging design. [for Solar Electric Propulsion spacecraft

Science.gov (United States)

Franzon, A. O.; Fredrickson, C. D.; Ross, R. G.

1975-01-01

The use of solar electric propulsion (SEP) for spacecraft primary propulsion imposes an extreme range of operational and environmental design requirements associated with the diversity of missions for which solar electric primary propulsion is advantageous. One SEP element which is particularly sensitive to these environmental extremes is the power processor unit (PPU) which powers and controls the electric ion thruster. An improved power processor thermal-mechanical packaging approach, referred to as dual shear plate packaging, has been designed to accommodate these different requirements with minimum change to the power processor design. Details of this packaging design are presented together with test results obtained from thermal-vacuum and structural-vibration tests conducted with prototype hardware.

Airborne particulate matter in spacecraft

Science.gov (United States)

1988-01-01

Acceptability limits and sampling and monitoring strategies for airborne particles in spacecraft were considered. Based on instances of eye and respiratory tract irritation reported by Shuttle flight crews, the following acceptability limits for airborne particles were recommended: for flights of 1 week or less duration (1 mg/cu m for particles less than 10 microns in aerodynamic diameter (AD) plus 1 mg/cu m for particles 10 to 100 microns in AD); and for flights greater than 1 week and up to 6 months in duration (0.2 mg/cu m for particles less than 10 microns in AD plus 0.2 mg/cu m for particles 10 to 100 microns in AD. These numerical limits were recommended to aid in spacecraft atmosphere design which should aim at particulate levels that are a low as reasonably achievable. Sampling of spacecraft atmospheres for particles should include size-fractionated samples of 0 to 10, 10 to 100, and greater than 100 micron particles for mass concentration measurement and elementary chemical analysis by nondestructive analysis techniques. Morphological and chemical analyses of single particles should also be made to aid in identifying airborne particulate sources. Air cleaning systems based on inertial collection principles and fine particle collection devices based on electrostatic precipitation and filtration should be considered for incorporation into spacecraft air circulation systems. It was also recommended that research be carried out in space in the areas of health effects and particle characterization.

Use of two-phase flow heat transfer method in spacecraft thermal system

Science.gov (United States)

Hye, A.

1985-01-01

In space applications, weight, volume and power are critical parameters. Presently liquid freon is used in the radiator planels of the Space Shuttle to dissipate heat. This requires a large amount of freon, large power for pumps, large volume and weight. Use of two-phase flow method to transfer heat can reduce them significantly. A modified commercial vapor compression refrigerator/freezer was sucessfully flown in STS-4 to study the effect of zero-gravity on the system. The duty cycle was about 5 percent higher in flight as compared to that on earth due to low flow velocity in condenser. The vapor Reynolds number at exit was about 4000 as compared to about 12,000. Efforts are underway to design a refrigerator/freezer using an oil-free compressor for Spacelab Mission 4 scheduled to fly in January 1986. A thermal system can be designed for spacecraft using the two-phase flow to transfer heat economically.

Global Precipitation Measurement (GPM) Spacecraft Lithium Ion Battery Micro-Cycling Investigation

Science.gov (United States)

Dakermanji, George; Lee, Leonine; Spitzer, Thomas

2016-01-01

The Global Precipitation Measurement (GPM) spacecraft was jointly developed by NASA and JAXA. It is a Low Earth Orbit (LEO) spacecraft launched on February 27, 2014. The power system is a Direct Energy Transfer (DET) system designed to support 1950 watts orbit average power. The batteries use SONY 18650HC cells and consist of three 8s by 84p batteries operated in parallel as a single battery. During instrument integration with the spacecraft, large current transients were observed in the battery. Investigation into the matter traced the cause to the Dual-Frequency Precipitation Radar (DPR) phased array radar which generates cyclical high rate current transients on the spacecraft power bus. The power system electronics interaction with these transients resulted in the current transients in the battery. An accelerated test program was developed to bound the effect, and to assess the impact to the mission.

Comprehension of Spacecraft Telemetry Using Hierarchical Specifications of Behavior

Science.gov (United States)

Havelund, Klaus; Joshi, Rajeev

2014-01-01

A key challenge in operating remote spacecraft is that ground operators must rely on the limited visibility available through spacecraft telemetry in order to assess spacecraft health and operational status. We describe a tool for processing spacecraft telemetry that allows ground operators to impose structure on received telemetry in order to achieve a better comprehension of system state. A key element of our approach is the design of a domain-specific language that allows operators to express models of expected system behavior using partial specifications. The language allows behavior specifications with data fields, similar to other recent runtime verification systems. What is notable about our approach is the ability to develop hierarchical specifications of behavior. The language is implemented as an internal DSL in the Scala programming language that synthesizes rules from patterns of specification behavior. The rules are automatically applied to received telemetry and the inferred behaviors are available to ground operators using a visualization interface that makes it easier to understand and track spacecraft state. We describe initial results from applying our tool to telemetry received from the Curiosity rover currently roving the surface of Mars, where the visualizations are being used to trend subsystem behaviors, in order to identify potential problems before they happen. However, the technology is completely general and can be applied to any system that generates telemetry such as event logs.

“天宫一号”目标飞行器信息管理策略%Design of information management strategy for Tiangong-1 target spacecraft

Institute of Scientific and Technical Information of China (English)

柏林厚; 李智勇; 南洪涛; 程伟

2013-01-01

“天宫一号”目标飞行器信息系统作为目标飞行器的重要功能系统,主要完成目标飞行器内部信息管理以及与地面和载人飞船的通信管理,确保目标飞行器的健康和稳定运行,支持目标飞行器各项任务的完成.文章对“天宫一号”目标飞行器信息系统的信息管理策略进行了概述,包括信息管理需求、设计原则、系统组成、主要技术特点以及可靠性安全性设计.“天宫一号”目标飞行器信息系统的信息管理策略经过首次无人和有人交会对接任务验证,满足任务要求,为后续空间站信息系统详细设计奠定了基础.%As the key functional system of the Tiangong-1 target spacecraft, the information management system is mainly responsible for the management of the internal information and the communication with the earth station and the Shenzhou spaceship, to ensure the target spacecraft's stable working and to support the accomplishment of the mission. This paper reviews the design of this target spacecraft information management strategy, including the information requirements, the design principle, the system configuration, the technical characteristics, and the reliability and safety design. With the accomplishment of China's first unmanned and manned rendezvous and docking mission, the information management strategy is demonstrated to fulfill the requirements. The information management system design of the Tiangong-1 target spacecraft lays the foundation for the specific design of China's future space station.

Future spacecraft propulsion systems. Enabling technologies for space exploration. 2. ed.

Energy Technology Data Exchange (ETDEWEB)

Czysz, Paul A. [St. Louis Univ., MO (United States). Oliver L. Parks Endowed Chair in Aerospace Engineering; Bruno, Claudio [Univ. degli Studi di Roma (Italy). Dipt. di Meccanica e Aeronautica

2009-07-01

In this second edition of Future Spacecraft Propulsion Systems, the authors demonstrate the need to break free from the old established concepts of expendable rockets, using chemical propulsion, and to develop new breeds of launch vehicle capable of both launching payloads into orbit at a dramatically reduced cost and for sustained operations in low-Earth orbit. The next steps to establishing a permanent 'presence' in the Solar System beyond Earth are the commercialisation of sustained operations on the Moon and the development of advanced nuclear or high-energy space propulsion systems for Solar System exploration out to the boundary of interstellar space. In the future, high-energy particle research facilities may one day yield a very high-energy propulsion system that will take us to the nearby stars, or even beyond. Space is not quiet: it is a continuous series of nuclear explosions that provide the material for new star systems to form and provide the challenge to explore. This book provides an assessment of the industrial capability required to construct and operate the necessary spacecraft. Time and distance communication and control limitations impose robotic constraints. Space environments restrict human sustained presence and put high demands on electronic, control and materials systems. This comprehensive and authoritative book puts spacecraft propulsion systems in perspective, from earth orbit launchers to astronomical/space exploration vehicles. It includes new material on fusion propulsion, new figures and updates and expands the information given in the first edition. (orig.)

Standardized spacecraft: a methodology for decision making. AMS report No. 1199

International Nuclear Information System (INIS)

Greenberg, J.S.; Nichols, R.A.

1974-01-01

As the space program matures, more and more attention is being focused on ways to reduce the costs of performing space missions. Standardization has been suggested as a way of providing cost reductions. The question of standardization at the system level, in particular, the question of the desirability of spacecraft standardization for geocentric space missions is addressed. The spacecraft is considered to be a bus upon which mission oriented equipment, the payload, is mounted. Three basic questions are considered: (1) is spacecraft standardization economically desirable; (2) if spacecraft standardization is economically desirable, what standardized spacecraft configuration or mix of configurations and technologies should be developed; and (3) if standardized spacecraft are to be developed, what power levels should they be designed for. A methodology which has been developed and which is necessary to follow if the above questions are to be answered and informed decisions made relative to spacecraft standardization is described. To illustrate the decision making problems and the need for the developed methodology and the data requirements, typical standardized spacecraft have been considered. Both standardized solar and nuclear-powered spacecraft and mission specialized spacecraft have been conceptualized and performance and cost estimates have been made. These estimates are not considered to be of sufficient accuracy to allow decisions regarding spacecraft mix and power levels to be made at this time. The estimates are deemed of sufficient accuracy so as to demonstrate the desirability of spacecraft standardization and the methodology (as well as the need for the methodology) which is necessary to decide upon the best mix of standardized spacecraft and their design power levels. (U.S.)

Design of systemic autonomous safety for Tiangong-I target spacecraft%“天宫一号”,目标飞行器系统级自主安全设计

Institute of Scientific and Technical Information of China (English)

李智勇

2011-01-01

为确保载人飞行器在长期飞行中的设备安全以及短期飞行中航天员的安全,需要从系统层面进行自主安全设计,使航天器在出现地面无法快速反应的故障时能够启动安全模式进行自我保护.文章以能源安全设计为主对“天宫一号”目标飞行器系统级自主安全设计进行了论述,总结了设计经验,对后续型号的设计提出了建议.%In space flights, the safety of hardware during a long-time flight and that of astronaut during a short-time flight are important issues in the implementation of the mission goal. The capability of independent safety control should be designed for spacecraft in the system level against major faults beyond instant ground responses. This paper discusses the system design of autonomous safety for manned spacecraft, with primary concern on the power system. Some engineering practices and suggestions are summarized for further studies.

Low power arcjet system spacecraft impacts

Science.gov (United States)

Pencil, Eric J.; Sarmiento, Charles J.; Lichtin, D. A.; Palchefsky, J. W.; Bogorad, A. L.

1993-01-01

Potential plume contamination of spacecraft surfaces was investigated by positioning spacecraft material samples relative to an arcjet thruster. Samples in the simulated solar array region were exposed to the cold gas arcjet plume for 40 hrs to address concerns about contamination by backstreaming diffusion pump oil. Except for one sample, no significant changes were measured in absorptance and emittance within experimental error. Concerns about surface property degradation due to electrostatic discharges led to the investigation of the discharge phenomenon of charged samples during arcjet ignition. Short duration exposure of charged samples demonstrated that potential differences are consistently and completely eliminated within the first second of exposure to a weakly ionized plume. The spark discharge mechanism was not the discharge phenomenon. The results suggest that the arcjet could act as a charge control device on spacecraft.

Simulation of Tomographic Reconstruction of Magnetosphere Plasma Distribution By Multi-spacecraft Systems.

Science.gov (United States)

Kunitsyn, V.; Nesterov, I.; Andreeva, E.; Zelenyi, L.; Veselov, M.; Galperin, Y.; Buchner, J.

A satellite radiotomography method for electron density distributions was recently proposed for closely-space multi-spacecraft group of high-altitude satellites to study the physics of reconnection process. The original idea of the ROY project is to use a constellation of spacecrafts (one main and several sub-satellites) in order to carry out closely-spaced multipoint measurements and 2D tomographic reconstruction of elec- tron density in the space between the main satellite and the subsatellites. The distances between the satellites were chosen to vary from dozens to few hundreds of kilometers. The easiest data interpretation is achieved when the subsatellites are placed along the plasma streamline. Then, whenever a plasma density irregularity moves between the main satellite and the subsatellites it will be scanned in different directions and we can get 2D distribution of plasma using these projections. However in general sub- satellites are not placed exactly along the plasma streamline. The method of plasma velocity determination relative to multi-spacecraft systems is considered. Possibilities of 3D tomographic imaging using multi-spacecraft systems are analyzed. The model- ing has shown that efficient scheme for 3D tomographic imaging would be to place spacecrafts in different planes so that the angle between the planes would make not more then ten degrees. Work is supported by INTAS PROJECT 2000-465.

Diagnosing Faults in Electrical Power Systems of Spacecraft and Aircraft

Data.gov (United States)

National Aeronautics and Space Administration — Electrical power systems play a critical role in spacecraft and aircraft, and they exhibit a rich variety of failure modes. This paper discusses electrical power...

High precision relative position sensing system for formation flying spacecraft

Data.gov (United States)

National Aeronautics and Space Administration — We propose to develop and test an optical sensing system that provides high precision relative position sensing for formation flying spacecraft.  A high precision...

Application of software technology to a future spacecraft computer design

Science.gov (United States)

Labaugh, R. J.

1980-01-01

A study was conducted to determine how major improvements in spacecraft computer systems can be obtained from recent advances in hardware and software technology. Investigations into integrated circuit technology indicated that the CMOS/SOS chip set being developed for the Air Force Avionics Laboratory at Wright Patterson had the best potential for improving the performance of spaceborne computer systems. An integral part of the chip set is the bit slice arithmetic and logic unit. The flexibility allowed by microprogramming, combined with the software investigations, led to the specification of a baseline architecture and instruction set.

Spacecraft Fire Safety Research at NASA Glenn Research Center

Science.gov (United States)

Meyer, Marit

2016-01-01

Appropriate design of fire detection systems requires knowledge of both the expected fire signature and the background aerosol levels. Terrestrial fire detection systems have been developed based on extensive study of terrestrial fires. Unfortunately there is no corresponding data set for spacecraft fires and consequently the fire detectors in current spacecraft were developed based upon terrestrial designs. In low gravity, buoyant flow is negligible which causes particles to concentrate at the smoke source, increasing their residence time, and increasing the transport time to smoke detectors. Microgravity fires have significantly different structure than those in 1-g which can change the formation history of the smoke particles. Finally the materials used in spacecraft are different from typical terrestrial environments where smoke properties have been evaluated. It is critically important to detect a fire in its early phase before a flame is established, given the fixed volume of air on any spacecraft. Consequently, the primary target for spacecraft fire detection is pyrolysis products rather than soot. Experimental investigations have been performed at three different NASA facilities which characterize smoke aerosols from overheating common spacecraft materials. The earliest effort consists of aerosol measurements in low gravity, called the Smoke Aerosol Measurement Experiment (SAME), and subsequent ground-based testing of SAME smoke in 55-gallon drums with an aerosol reference instrument. Another set of experiments were performed at NASAs Johnson Space Center White Sands Test Facility (WSTF), with additional fuels and an alternate smoke production method. Measurements of these smoke products include mass and number concentration, and a thermal precipitator was designed for this investigation to capture particles for microscopic analysis. The final experiments presented are from NASAs Gases and Aerosols from Smoldering Polymers (GASP) Laboratory, with selected

Implementing model-based system engineering for the whole lifecycle of a spacecraft

Science.gov (United States)

Fischer, P. M.; Lüdtke, D.; Lange, C.; Roshani, F.-C.; Dannemann, F.; Gerndt, A.

2017-09-01

Design information of a spacecraft is collected over all phases in the lifecycle of a project. A lot of this information is exchanged between different engineering tasks and business processes. In some lifecycle phases, model-based system engineering (MBSE) has introduced system models and databases that help to organize such information and to keep it consistent for everyone. Nevertheless, none of the existing databases approached the whole lifecycle yet. Virtual Satellite is the MBSE database developed at DLR. It has been used for quite some time in Phase A studies and is currently extended for implementing it in the whole lifecycle of spacecraft projects. Since it is unforeseeable which future use cases such a database needs to support in all these different projects, the underlying data model has to provide tailoring and extension mechanisms to its conceptual data model (CDM). This paper explains the mechanisms as they are implemented in Virtual Satellite, which enables extending the CDM along the project without corrupting already stored information. As an upcoming major use case, Virtual Satellite will be implemented as MBSE tool in the S2TEP project. This project provides a new satellite bus for internal research and several different payload missions in the future. This paper explains how Virtual Satellite will be used to manage configuration control problems associated with such a multi-mission platform. It discusses how the S2TEP project starts using the software for collecting the first design information from concurrent engineering studies, then making use of the extension mechanisms of the CDM to introduce further information artefacts such as functional electrical architecture, thus linking more and more processes into an integrated MBSE approach.

The Interaction of Spacecraft Cabin Atmospheric Quality and Water Processing System Performance

Science.gov (United States)

Perry, Jay L.; Croomes, Scott D. (Technical Monitor)

2002-01-01

Although designed to remove organic contaminants from a variety of waste water streams, the planned U.S.- and present Russian-provided water processing systems onboard the International Space Station (ISS) have capacity limits for some of the more common volatile cleaning solvents used for housekeeping purposes. Using large quantities of volatile cleaning solvents during the ground processing and in-flight operational phases of a crewed spacecraft such as the ISS can lead to significant challenges to the water processing systems. To understand the challenges facing the management of water processing capacity, the relationship between cabin atmospheric quality and humidity condensate loading is presented. This relationship is developed as a tool to determine the cabin atmospheric loading that may compromise water processing system performance. A comparison of cabin atmospheric loading with volatile cleaning solvents from ISS, Mir, and Shuttle are presented to predict acceptable limits to maintain optimal water processing system performance.

Spacecraft Attitude Determination

DEFF Research Database (Denmark)

Bak, Thomas

This thesis describes the development of an attitude determination system for spacecraft based only on magnetic field measurements. The need for such system is motivated by the increased demands for inexpensive, lightweight solutions for small spacecraft. These spacecraft demands full attitude...... determination based on simple, reliable sensors. Meeting these objectives with a single vector magnetometer is difficult and requires temporal fusion of data in order to avoid local observability problems. In order to guaranteed globally nonsingular solutions, quaternions are generally the preferred attitude...... is a detailed study of the influence of approximations in the modeling of the system. The quantitative effects of errors in the process and noise statistics are discussed in detail. The third contribution is the introduction of these methods to the attitude determination on-board the Ørsted satellite...

A spacecraft computer repairable via command.

Science.gov (United States)

Fimmel, R. O.; Baker, T. E.

1971-01-01

The MULTIPAC is a central data system developed for deep-space probes with the distinctive feature that it may be repaired during flight via command and telemetry links by reprogramming around the failed unit. The computer organization uses pools of identical modules which the program organizes into one or more computers called processors. The interaction of these modules is dynamically controlled by the program rather than hardware. In the event of a failure, new programs are entered which reorganize the central data system with a somewhat reduced total processing capability aboard the spacecraft. Emphasis is placed on the evolution of the system architecture and the final overall system design rather than the specific logic design.

An Orbit Propagation Software for Mars Orbiting Spacecraft

Directory of Open Access Journals (Sweden)

Young-Joo Song

2004-12-01

Full Text Available An orbit propagation software for the Mars orbiting spacecraft has been developed and verified in preparations for the future Korean Mars missions. Dynamic model for Mars orbiting spacecraft has been studied, and Mars centered coordinate systems are utilized to express spacecraft state vectors. Coordinate corrections to the Mars centered coordinate system have been made to adjust the effects caused by Mars precession and nutation. After spacecraft enters Sphere of Influence (SOI of the Mars, the spacecraft experiences various perturbation effects as it approaches to Mars. Every possible perturbation effect is considered during integrations of spacecraft state vectors. The Mars50c gravity field model and the Mars-GRAM 2001 model are used to compute perturbation effects due to Mars gravity field and Mars atmospheric drag, respectively. To compute exact locations of other planets, JPL's DE405 ephemerides are used. Phobos and Deimos's ephemeris are computed using analytical method because their informations are not released with DE405. Mars Global Surveyor's mapping orbital data are used to verify the developed propagator performances. After one Martian day propagation (12 orbital periods, the results show about maximum ±5 meter errors, in every position state components(radial, cross-track and along-track, when compared to these from the Astrogator propagation in the Satellite Tool Kit. This result shows high reliability of the developed software which can be used to design near Mars missions for Korea, in future.

Nuclear-Powered GPS Spacecraft Design Study

Energy Technology Data Exchange (ETDEWEB)

Raab, Bernard

1977-05-01

This is the final report of a study to investigate the potential benefits of a nuclear (radioisotope) - powered satellite for advanced phases of the Global Positioning System (GPS) program. The critical parameters were: power to user; mean mission duration; orbital predictability; thermal control of on-board frequency standards; and vulnerability. The reference design approach is described, and input data are given for two power systems that are under development: an organic Rankine system and a Brayton cycle system. Reference design details are provided and structural design and analysis are discussed, as well as thermal design and analysis. A higher altitude version is also considered.

REQUIREMENTS FOR IMAGE QUALITY OF EMERGENCY SPACECRAFTS

Directory of Open Access Journals (Sweden)

A. I. Altukhov

2015-05-01

Full Text Available The paper deals with the method for formation of quality requirements to the images of emergency spacecrafts. The images are obtained by means of remote sensing of near-earth space orbital deployment in the visible range. of electromagnetic radiation. The method is based on a joint taking into account conditions of space survey, characteristics of surveillance equipment, main design features of the observed spacecrafts and orbital inspection tasks. Method. Quality score is the predicted linear resolution image that gives the possibility to create a complete view of pictorial properties of the space image obtained by electro-optical system from the observing satellite. Formulation of requirements to the numerical value of this indicator is proposed to perform based on the properties of remote sensing system, forming images in the conditions of outer space, and the properties of the observed emergency spacecraft: dimensions, platform construction of the satellite, on-board equipment placement. For method implementation the authors have developed a predictive model of requirements to a linear resolution for images of emergency spacecrafts, making it possible to select the intervals of space shooting and get the satellite images required for quality interpretation. Main results. To verify the proposed model functionality we have carried out calculations of the numerical values for the linear resolution of the image, ensuring the successful task of determining the gross structural damage of the spacecrafts and identifying changes in their spatial orientation. As input data were used with dimensions and geometric primitives corresponding to the shape of deemed inspected spacecrafts: Resurs-P", "Canopus-B", "Electro-L". Numerical values of the linear resolution images have been obtained, ensuring the successful task solution for determining the gross structural damage of spacecrafts.

An expert system for diagnosing environmentally induced spacecraft anomalies

Science.gov (United States)

Rolincik, Mark; Lauriente, Michael; Koons, Harry C.; Gorney, David

1992-01-01

A new rule-based, machine independent analytical tool was designed for diagnosing spacecraft anomalies using an expert system. Expert systems provide an effective method for saving knowledge, allow computers to sift through large amounts of data pinpointing significant parts, and most importantly, use heuristics in addition to algorithms, which allow approximate reasoning and inference and the ability to attack problems not rigidly defined. The knowledge base consists of over two-hundred (200) rules and provides links to historical and environmental databases. The environmental causes considered are bulk charging, single event upsets (SEU), surface charging, and total radiation dose. The system's driver translates forward chaining rules into a backward chaining sequence, prompting the user for information pertinent to the causes considered. The use of heuristics frees the user from searching through large amounts of irrelevant information and allows the user to input partial information (varying degrees of confidence in an answer) or 'unknown' to any question. The modularity of the expert system allows for easy updates and modifications. It not only provides scientists with needed risk analysis and confidence not found in algorithmic programs, but is also an effective learning tool, and the window implementation makes it very easy to use. The system currently runs on a Micro VAX II at Goddard Space Flight Center (GSFC). The inference engine used is NASA's C Language Integrated Production System (CLIPS).

Passive Devices for Advanced Fluid Management aboard Spacecraft, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Acute challenges are faced by the designers of fluid systems for spacecraft because of the persistently unfamiliar and unforgiving low-g environment. For example,...

Thrusting maneuver control of a small spacecraft via only gimbaled-thruster scheme

Science.gov (United States)

Kabganian, Mansour; Kouhi, Hamed; Shahravi, Morteza; Fani Saberi, Farhad

2018-05-01

The thrust vector control (TVC) scheme is a powerful method in spacecraft attitude control. Since the control of a small spacecraft is being studied here, a solid rocket motor (SRM) should be used instead of a liquid propellant motor. Among the TVC methods, gimbaled-TVC as an efficient method is employed in this paper. The spacecraft structure is composed of a body and a gimbaled-SRM where common attitude control systems such as reaction control system (RCS) and spin-stabilization are not presented. A nonlinear two-body model is considered for the characterization of the gimbaled-thruster spacecraft where, the only control input is provided by a gimbal actuator. The attitude of the spacecraft is affected by a large exogenous disturbance torque which is generated by a thrust vector misalignment from the center of mass (C.M). A linear control law is designed to stabilize the spacecraft attitude while rejecting the mentioned disturbance torque. A semi-analytical formulation of the region of attraction (RoA) is developed to ensure the local stability and fast convergence of the nonlinear closed-loop system. Simulation results of the 3D maneuvers are included to show the applicability of this method for use in a small spacecraft.

On TTEthernet for Integrated Fault-Tolerant Spacecraft Networks

Science.gov (United States)

Loveless, Andrew

2015-01-01

There has recently been a push for adopting integrated modular avionics (IMA) principles in designing spacecraft architectures. This consolidation of multiple vehicle functions to shared computing platforms can significantly reduce spacecraft cost, weight, and de- sign complexity. Ethernet technology is attractive for inclusion in more integrated avionic systems due to its high speed, flexibility, and the availability of inexpensive commercial off-the-shelf (COTS) components. Furthermore, Ethernet can be augmented with a variety of quality of service (QoS) enhancements that enable its use for transmitting critical data. TTEthernet introduces a decentralized clock synchronization paradigm enabling the use of time-triggered Ethernet messaging appropriate for hard real-time applications. TTEthernet can also provide two forms of event-driven communication, therefore accommodating the full spectrum of traffic criticality levels required in IMA architectures. This paper explores the application of TTEthernet technology to future IMA spacecraft architectures as part of the Avionics and Software (A&S) project chartered by NASA's Advanced Exploration Systems (AES) program.

Performance Testing of a Photocatalytic Oxidation Module for Spacecraft Cabin Atmosphere Revitalization

Science.gov (United States)

Perry, Jay L.; Abney, Morgan B.; Frederick, Kenneth R.; Scott, Joseph P.; Kaiser, Mark; Seminara, Gary; Bershitsky, Alex

2011-01-01

Photocatalytic oxidation (PCO) is a candidate process technology for use in high volumetric flow rate trace contaminant control applications in sealed environments. The targeted application for PCO as applied to crewed spacecraft life support system architectures is summarized. Technical challenges characteristic of PCO are considered. Performance testing of a breadboard PCO reactor design for mineralizing polar organic compounds in a spacecraft cabin atmosphere is described. Test results are analyzed and compared to results reported in the literature for comparable PCO reactor designs.

Practical Applications of Cosmic Ray Science: Spacecraft, Aircraft, Ground Based Computation and Control Systems and Human Health and Safety

Science.gov (United States)

Atwell, William; Koontz, Steve; Normand, Eugene

2012-01-01

In this paper we review the discovery of cosmic ray effects on the performance and reliability of microelectronic systems as well as on human health and safety, as well as the development of the engineering and health science tools used to evaluate and mitigate cosmic ray effects in earth surface, atmospheric flight, and space flight environments. Three twentieth century technological developments, 1) high altitude commercial and military aircraft; 2) manned and unmanned spacecraft; and 3) increasingly complex and sensitive solid state micro-electronics systems, have driven an ongoing evolution of basic cosmic ray science into a set of practical engineering tools (e.g. ground based test methods as well as high energy particle transport and reaction codes) needed to design, test, and verify the safety and reliability of modern complex electronic systems as well as effects on human health and safety. The effects of primary cosmic ray particles, and secondary particle showers produced by nuclear reactions with spacecraft materials, can determine the design and verification processes (as well as the total dollar cost) for manned and unmanned spacecraft avionics systems. Similar considerations apply to commercial and military aircraft operating at high latitudes and altitudes near the atmospheric Pfotzer maximum. Even ground based computational and controls systems can be negatively affected by secondary particle showers at the Earth's surface, especially if the net target area of the sensitive electronic system components is large. Accumulation of both primary cosmic ray and secondary cosmic ray induced particle shower radiation dose is an important health and safety consideration for commercial or military air crews operating at high altitude/latitude and is also one of the most important factors presently limiting manned space flight operations beyond low-Earth orbit (LEO).

Cassini Spacecraft In-Flight Swap to Backup Attitude Control Thrusters

Science.gov (United States)

Bates, David M.

2010-01-01

NASA's Cassini Spacecraft, launched on October 15th, 1997 and arrived at Saturn on June 30th, 2004, is the largest and most ambitious interplanetary spacecraft in history. In order to meet the challenging attitude control and navigation requirements of the orbit profile at Saturn, Cassini is equipped with a monopropellant thruster based Reaction Control System (RCS), a bipropellant Main Engine Assembly (MEA) and a Reaction Wheel Assembly (RWA). In 2008, after 11 years of reliable service, several RCS thrusters began to show signs of end of life degradation, which led the operations team to successfully perform the swap to the backup RCS system, the details and challenges of which are described in this paper. With some modifications, it is hoped that similar techniques and design strategies could be used to benefit other spacecraft.

Trajectory Design for a Cislunar Cubesat Leveraging Dynamical Systems Techniques: The Lunar Icecube Mission

Science.gov (United States)

Bosanac, Natasha; Cox, Andrew; Howell, Kathleen C.; Folta, David

2017-01-01

Lunar IceCube is a 6U CubeSat that is designed to detect and observe lunar volatiles from a highly inclined orbit. This spacecraft, equipped with a low-thrust engine, will be deployed from the upcoming Exploration Mission-1 vehicle in late 2018. However, significant uncertainty in the deployment conditions for secondary payloads impacts both the availability and geometry of transfers that deliver the spacecraft to the lunar vicinity. A framework that leverages dynamical systems techniques is applied to a recently updated set of deployment conditions and spacecraft parameter values for the Lunar IceCube mission, demonstrating the capability for rapid trajectory design.

The Near Earth Object (NEO) Scout Spacecraft: A Low-cost Approach to In-situ Characterization of the NEO Population

Science.gov (United States)

Woeppel, Eric A.; Balsamo, James M.; Fischer, Karl J.; East, Matthew J.; Styborski, Jeremy A.; Roche, Christopher A.; Ott, Mackenzie D.; Scorza, Matthew J.; Doherty, Christopher D.; Trovato, Andrew J.;

Probabilistic Risk Assessment for Decision Making During Spacecraft Operations

Science.gov (United States)

Meshkat, Leila

2009-01-01

Decisions made during the operational phase of a space mission often have significant and immediate consequences. Without the explicit consideration of the risks involved and their representation in a solid model, it is very likely that these risks are not considered systematically in trade studies. Wrong decisions during the operational phase of a space mission can lead to immediate system failure whereas correct decisions can help recover the system even from faulty conditions. A problem of special interest is the determination of the system fault protection strategies upon the occurrence of faults within the system. Decisions regarding the fault protection strategy also heavily rely on a correct understanding of the state of the system and an integrated risk model that represents the various possible scenarios and their respective likelihoods. Probabilistic Risk Assessment (PRA) modeling is applicable to the full lifecycle of a space mission project, from concept development to preliminary design, detailed design, development and operations. The benefits and utilities of the model, however, depend on the phase of the mission for which it is used. This is because of the difference in the key strategic decisions that support each mission phase. The focus of this paper is on describing the particular methods used for PRA modeling during the operational phase of a spacecraft by gleaning insight from recently conducted case studies on two operational Mars orbiters. During operations, the key decisions relate to the commands sent to the spacecraft for any kind of diagnostics, anomaly resolution, trajectory changes, or planning. Often, faults and failures occur in the parts of the spacecraft but are contained or mitigated before they can cause serious damage. The failure behavior of the system during operations provides valuable data for updating and adjusting the related PRA models that are built primarily based on historical failure data. The PRA models, in turn

Toward autonomous spacecraft

Science.gov (United States)

Fogel, L. J.; Calabrese, P. G.; Walsh, M. J.; Owens, A. J.

1982-01-01

Ways in which autonomous behavior of spacecraft can be extended to treat situations wherein a closed loop control by a human may not be appropriate or even possible are explored. Predictive models that minimize mean least squared error and arbitrary cost functions are discussed. A methodology for extracting cyclic components for an arbitrary environment with respect to usual and arbitrary criteria is developed. An approach to prediction and control based on evolutionary programming is outlined. A computer program capable of predicting time series is presented. A design of a control system for a robotic dense with partially unknown physical properties is presented.

Dissertation Defense Computational Fluid Dynamics Uncertainty Analysis for Payload Fairing Spacecraft Environmental Control Systems

Science.gov (United States)

Groves, Curtis Edward

2014-01-01

Spacecraft thermal protection systems are at risk of being damaged due to airflow produced from Environmental Control Systems. There are inherent uncertainties and errors associated with using Computational Fluid Dynamics to predict the airflow field around a spacecraft from the Environmental Control System. This paper describes an approach to quantify the uncertainty in using Computational Fluid Dynamics to predict airflow speeds around an encapsulated spacecraft without the use of test data. Quantifying the uncertainty in analytical predictions is imperative to the success of any simulation-based product. The method could provide an alternative to traditional "validation by test only" mentality. This method could be extended to other disciplines and has potential to provide uncertainty for any numerical simulation, thus lowering the cost of performing these verifications while increasing the confidence in those predictions. Spacecraft requirements can include a maximum airflow speed to protect delicate instruments during ground processing. Computational Fluid Dynamics can be used to verify these requirements; however, the model must be validated by test data. This research includes the following three objectives and methods. Objective one is develop, model, and perform a Computational Fluid Dynamics analysis of three (3) generic, non-proprietary, environmental control systems and spacecraft configurations. Several commercially available and open source solvers have the capability to model the turbulent, highly three-dimensional, incompressible flow regime. The proposed method uses FLUENT, STARCCM+, and OPENFOAM. Objective two is to perform an uncertainty analysis of the Computational Fluid Dynamics model using the methodology found in "Comprehensive Approach to Verification and Validation of Computational Fluid Dynamics Simulations". This method requires three separate grids and solutions, which quantify the error bars around Computational Fluid Dynamics

Dissertation Defense: Computational Fluid Dynamics Uncertainty Analysis for Payload Fairing Spacecraft Environmental Control Systems

Science.gov (United States)

Groves, Curtis Edward

2014-01-01

Spacecraft thermal protection systems are at risk of being damaged due to airflow produced from Environmental Control Systems. There are inherent uncertainties and errors associated with using Computational Fluid Dynamics to predict the airflow field around a spacecraft from the Environmental Control System. This paper describes an approach to quantify the uncertainty in using Computational Fluid Dynamics to predict airflow speeds around an encapsulated spacecraft without the use of test data. Quantifying the uncertainty in analytical predictions is imperative to the success of any simulation-based product. The method could provide an alternative to traditional validation by test only mentality. This method could be extended to other disciplines and has potential to provide uncertainty for any numerical simulation, thus lowering the cost of performing these verifications while increasing the confidence in those predictions.Spacecraft requirements can include a maximum airflow speed to protect delicate instruments during ground processing. Computational Fluid Dynamics can be used to verify these requirements; however, the model must be validated by test data. This research includes the following three objectives and methods. Objective one is develop, model, and perform a Computational Fluid Dynamics analysis of three (3) generic, non-proprietary, environmental control systems and spacecraft configurations. Several commercially available and open source solvers have the capability to model the turbulent, highly three-dimensional, incompressible flow regime. The proposed method uses FLUENT, STARCCM+, and OPENFOAM. Objective two is to perform an uncertainty analysis of the Computational Fluid Dynamics model using the methodology found in Comprehensive Approach to Verification and Validation of Computational Fluid Dynamics Simulations. This method requires three separate grids and solutions, which quantify the error bars around Computational Fluid Dynamics predictions

On the concept of survivability, with application to spacecraft and space-based networks

International Nuclear Information System (INIS)

Castet, Jean-Francois; Saleh, Joseph H.

2012-01-01

Survivability is an important attribute and requirement for military systems. Recently, survivability has become increasingly important for public infrastructure systems as well. In this work, we bring considerations of survivability to bear on space systems. We develop a conceptual framework and quantitative analyses based on stochastic Petri nets (SPN) to characterize and compare the survivability of different space architectures. The architectures here considered are a monolith spacecraft and a space-based network. To build the stochastic Petri net models for the degradations and failures of these two architectures, we conducted statistical analyses of historical multi-state failure data of spacecraft subsystems, and we assembled these subsystems, and their SPN models, in ways to create our monolith and networked systems. Preliminary results indicate, and quantify the extent to which, a space-based network is more survivable than the monolith spacecraft with respect to on-orbit anomalies and failures. For space systems, during the design and acquisition process, different architectures are benchmarked against several metrics; we argue that if survivability is not accounted for, then the evaluation process is likely to be biased in favor of the traditional dominant design, namely the monolith spacecraft. If however in a given context, survivability is a critical requirement for a customer, the survivability framework here proposed, and the stochastic modeling capability developed, can demonstrate the extent to which a networked space architecture may better satisfy this requirement than a monolith spacecraft. These results should be of interest to operators whose space assets require high levels of survivability, especially in the light of emerging threats.

Spacecraft Angular Rates Estimation with Gyrowheel Based on Extended High Gain Observer

Directory of Open Access Journals (Sweden)

Xiaokun Liu

2016-04-01

Full Text Available A gyrowheel (GW is a kind of electronic electric-mechanical servo system, which can be applied to a spacecraft attitude control system (ACS as both an actuator and a sensor simultaneously. In order to solve the problem of two-dimensional spacecraft angular rate sensing as a GW outputting three-dimensional control torque, this paper proposed a method of an extended high gain observer (EHGO with the derived GW mathematical model to implement the spacecraft angular rate estimation when the GW rotor is working at large angles. For this purpose, the GW dynamic equation is firstly derived with the second kind Lagrange method, and the relationship between the measurable and unmeasurable variables is built. Then, the EHGO is designed to estimate and calculate spacecraft angular rates with the GW, and the stability of the designed EHGO is proven by the Lyapunov function. Moreover, considering the engineering application, the effect of measurement noise in the tilt angle sensors on the estimation accuracy of the EHGO is analyzed. Finally, the numerical simulation is performed to illustrate the validity of the method proposed in this paper.

Quick Spacecraft Thermal Analysis Tool, Phase II

Data.gov (United States)

National Aeronautics and Space Administration — For spacecraft design and development teams concerned with cost and schedule, the Quick Spacecraft Thermal Analysis Tool (QuickSTAT) is an innovative software suite...

Design and validation of a GNC system for missions to asteroids: the AIM scenario

Science.gov (United States)

Pellacani, A.; Kicman, P.; Suatoni, M.; Casasco, M.; Gil, J.; Carnelli, I.

2017-12-01

Deep space missions, and in particular missions to asteroids, impose a certain level of autonomy that depends on the mission objectives. If the mission requires the spacecraft to perform close approaches to the target body (the extreme case being a landing scenario), the autonomy level must be increased to guarantee the fast and reactive response which is required in both nominal and contingency operations. The GNC system must be designed in accordance with the required level of autonomy. The GNC system designed and tested in the frame of ESA's Asteroid Impact Mission (AIM) system studies (Phase A/B1 and Consolidation Phase) is an example of an autonomous GNC system that meets the challenging objectives of AIM. The paper reports the design of such GNC system and its validation through a DDVV plan that includes Model-in-the-Loop and Hardware-in-the-Loop testing. Main focus is the translational navigation, which is able to provide online the relative state estimation with respect to the target body using exclusively cameras as relative navigation sensors. The relative navigation outputs are meant to be used for nominal spacecraft trajectory corrections as well as to estimate the collision risk with the asteroid and, if needed, to command the execution of a collision avoidance manoeuvre to guarantee spacecraft safety

Link Analysis of High Throughput Spacecraft Communication Systems for Future Science Missions

Science.gov (United States)

Simons, Rainee N.

2015-01-01

NASA's plan to launch several spacecrafts into low Earth Orbit (LEO) to support science missions in the next ten years and beyond requires down link throughput on the order of several terabits per day. The ability to handle such a large volume of data far exceeds the capabilities of current systems. This paper proposes two solutions, first, a high data rate link between the LEO spacecraft and ground via relay satellites in geostationary orbit (GEO). Second, a high data rate direct to ground link from LEO. Next, the paper presents results from computer simulations carried out for both types of links taking into consideration spacecraft transmitter frequency, EIRP, and waveform; elevation angle dependent path loss through Earths atmosphere, and ground station receiver GT.

Relating MBSE to Spacecraft Development: A NASA Pathfinder

Science.gov (United States)

Othon, Bill

2016-01-01

The NASA Engineering and Safety Center (NESC) has sponsored a Pathfinder Study to investigate how Model Based Systems Engineering (MBSE) and Model Based Engineering (MBE) techniques can be applied by NASA spacecraft development projects. The objectives of this Pathfinder Study included analyzing both the products of the modeling activity, as well as the process and tool chain through which the spacecraft design activities are executed. Several aspects of MBSE methodology and process were explored. Adoption and consistent use of the MBSE methodology within an existing development environment can be difficult. The Pathfinder Team evaluated the possibility that an "MBSE Template" could be developed as both a teaching tool as well as a baseline from which future NASA projects could leverage. Elements of this template include spacecraft system component libraries, data dictionaries and ontology specifications, as well as software services that do work on the models themselves. The Pathfinder Study also evaluated the tool chain aspects of development. Two chains were considered: 1. The Development tool chain, through which SysML model development was performed and controlled, and 2. The Analysis tool chain, through which both static and dynamic system analysis is performed. Of particular interest was the ability to exchange data between SysML and other engineering tools such as CAD and Dynamic Simulation tools. For this study, the team selected a Mars Lander vehicle as the element to be designed. The paper will discuss what system models were developed, how data was captured and exchanged, and what analyses were conducted.

Development of preliminary design concept for a multifunction display and control system for the Orbiter crew station. Task 4: Design concept recommendation

Science.gov (United States)

Spiger, R. J.; Farrell, R. J.; Holcomb, G. A.

1982-01-01

Application of multifunction display and control systems to the NASA Orbiter spacecraft offers the potential for reducing crew workload and improving the presentation of system status and operational data to the crew. A design concept is presented for the application of a multifunction display and control system (MFDCS) to the Orbital Maneuvering System and Electrical Power Distribution and Control System on the Orbiter spacecraft. The MFDCS would provide the capability for automation of procedures, fault prioritization and software reconfiguration of the MFDCS data base. The MFDCS would operate as a stand-alone processor to minimize the impact on the current Orbiter software. Supervisory crew command of all current functions would be retained through the use of several operating modes in the system. Both the design concept and the processes followed in defining the concept are described.

ATHENA: system design and implementation for a next generation x-ray telescope

Science.gov (United States)

Ayre, M.; Bavdaz, M.; Ferreira, I.; Wille, E.; Lumb, D.; Linder, M.

2015-08-01

ATHENA, Europe's next generation x-ray telescope, has recently been selected for the 'L2' slot in ESA's Cosmic Vision Programme, with a mandate to address the 'Hot and Energetic Universe' Cosmic Vision science theme. The mission is currently in the Assessment/Definition Phase (A/B1), with a view to formal adoption after a successful System Requirements Review in 2019. This paper will describe the reference mission architecture and spacecraft design produced during Phase 0 by the ESA Concurrent Design Facility (CDF), in response to the technical requirements and programmatic boundary conditions. The main technical requirements and their mapping to resulting design choices will be presented, at both mission and spacecraft level. An overview of the spacecraft design down to subsystem level will then be presented (including the telescope and instruments), remarking on the critically-enabling technologies where appropriate. Finally, a programmatic overview will be given of the on-going Assessment Phase, and a snapshot of the prospects for securing the `as-proposed' mission within the cost envelope will be given.

Internal Acoustics of the ISS and Other Spacecraft

Science.gov (United States)

Allen, Christopher S.

2017-01-01

It is important to control the acoustic environment inside spacecraft and space habitats to protect for astronaut communications, alarm audibility, and habitability, and to reduce astronauts' risk for sleep disturbance, and hear-ing loss. But this is not an easy task, given the various design trade-offs, and it has been difficult, historically, to achieve. Over time it has been found that successful control of spacecraft acoustic levels is achieved by levying firm requirements at the system-level, using a systems engineering approach for design and development, and then validating these requirements with acoustic testing. In the systems engineering method, the system-level requirements must be flowed down to sub-systems and component noise sources, using acoustic analysis and acoustic modelling to develop allocated requirements for the sub-systems and components. Noise controls must also be developed, tested, and implemented so the sub-systems and components can achieve their allocated limits. It is also important to have management support for acoustics efforts to maintain their priority against the various trade-offs, including mass, volume, power, cost, and schedule. In this extended abstract and companion presentation, the requirements, approach, and results for controlling acoustic levels in most US spacecraft since Apollo will be briefly discussed. The approach for controlling acoustic levels in the future US space vehicle, Orion Multipurpose Crew Vehicle (MPCV), will also be briefly discussed. These discussions will be limited to the control of continuous noise inside the space vehicles. Other types of noise, such as launch, landing, and abort noise, intermittent noise, Extra-Vehicular Activity (EVA) noise, emergency operations/off-nominal noise, noise exposure, and impulse noise are important, but will not be discussed because of time limitations.

Artificial intelligence costs, benefits, risks for selected spacecraft ground system automation scenarios

Science.gov (United States)

Truszkowski, Walter F.; Silverman, Barry G.; Kahn, Martha; Hexmoor, Henry

1988-01-01

In response to a number of high-level strategy studies in the early 1980s, expert systems and artificial intelligence (AI/ES) efforts for spacecraft ground systems have proliferated in the past several years primarily as individual small to medium scale applications. It is useful to stop and assess the impact of this technology in view of lessons learned to date, and hopefully, to determine if the overall strategies of some of the earlier studies both are being followed and still seem relevant. To achieve that end four idealized ground system automation scenarios and their attendant AI architecture are postulated and benefits, risks, and lessons learned are examined and compared. These architectures encompass: (1) no AI (baseline), (2) standalone expert systems, (3) standardized, reusable knowledge base management systems (KBMS), and (4) a futuristic unattended automation scenario. The resulting artificial intelligence lessons learned, benefits, and risks for spacecraft ground system automation scenarios are described.

Shaded Spacecraft Radiators to Be Used on the Daytime Surface of the Mercury Planet, the Moon, and Asteroids of the Solar System Inner Part

Directory of Open Access Journals (Sweden)

V. A. Igrickii

2016-01-01

Full Text Available During the daytime a surface of the Moon, Mercury planet, and asteroids of the Solar system inner part, significantly heats up, and infrared radiation of the local soil becomes essential. At the same time direct solar radiation and reflected from the surface solar radiation reach the maximum too. These radiation fluxes can significantly decrease the efficiency of spacecraft radiators in the daytime. This effect is especially strong on the Mercury surface where direct solar radiation is 10 times stronger than solar radiation near the Earth. As a result, on the daytime surface of the Mercury the conventional low-temperature radiators become completely disabled.The article describes the development of the special shaded spacecraft radiators to be used in daytime on the Mercury and other atmosphereless bodies of the Solar system inner part. To solve this task are used mirror shades. The shape of these shades is developed to improve operation conditions of the spacecraft radiator through the appropriate scheme of radiation reflection. The task is discussed in 2D and 3D cases. A new design of shaded spacecraft radiators is proposed, and reasonable proportions of radiators are determined. The performance capability of proposed radiators for environments of the Mercury and the Moon is estimated using the zonal method in view of partial mirror reflection. The calculations showed that the developed shaded spacecraft radiators are capable to work on the Mercury surface as the low-temperature radiators even during the daytime. New radiators provide minimum accepted operating temperature of 241К (-32°С, meanwhile radiators of common design have minimum operating temperature of 479К (206°С. Using such radiators on the Moon enables us to increase effectiveness of spacecraft radiators and to decrease their minimum operating temperature from 270К (-3°С to 137К (-136°С.

Design Process of Flight Vehicle Structures for a Common Bulkhead and an MPCV Spacecraft Adapter

Science.gov (United States)

Aggarwal, Pravin; Hull, Patrick V.

2015-01-01

Design and manufacturing space flight vehicle structures is a skillset that has grown considerably at NASA during that last several years. Beginning with the Ares program and followed by the Space Launch System (SLS); in-house designs were produced for both the Upper Stage and the SLS Multipurpose crew vehicle (MPCV) spacecraft adapter. Specifically, critical design review (CDR) level analysis and flight production drawing were produced for the above mentioned hardware. In particular, the experience of this in-house design work led to increased manufacturing infrastructure for both Marshal Space Flight Center (MSFC) and Michoud Assembly Facility (MAF), improved skillsets in both analysis and design, and hands on experience in building and testing (MSA) full scale hardware. The hardware design and development processes from initiation to CDR and finally flight; resulted in many challenges and experiences that produced valuable lessons. This paper builds on these experiences of NASA in recent years on designing and fabricating flight hardware and examines the design/development processes used, as well as the challenges and lessons learned, i.e. from the initial design, loads estimation and mass constraints to structural optimization/affordability to release of production drawing to hardware manufacturing. While there are many documented design processes which a design engineer can follow, these unique experiences can offer insight into designing hardware in current program environments and present solutions to many of the challenges experienced by the engineering team.

Wave-Based Attitude Control of Spacecraft with Fuel Sloshing Dynamics

Directory of Open Access Journals (Sweden)

Thompson Joseph William

2016-06-01

Full Text Available Wave-Based Control has been previously applied successfully to simple under-actuated flexible mechanical systems. Spacecraft and rockets with structural flexibility and sloshing are examples of such systems but have added difficulties due to non-uniform structure, external disturbing forces and non-ideal actuators and sensors. The aim of this paper is to extend the application of WBC to spacecraft systems, to compare the performance of WBC to other popular controllers and to carry out experimental validation of the designed control laws. A mathematical model is developed for an upper stage accelerating rocket moving in a single plane. Fuel sloshing is represented by an equivalent mechanical pendulum model. A wave-based controller is designed for the upper stage AVUM of the European launcher Vega. In numerical simulations the controller successfully suppresses the sloshing motion. A major advantage of the strategy is that no measurement of the pendulum states (sloshing motion is required.

Modeling the angular motion dynamics of spacecraft with a magnetic attitude control system based on experimental studies and dynamic similarity

Science.gov (United States)

Kulkov, V. M.; Medvedskii, A. L.; Terentyev, V. V.; Firsyuk, S. O.; Shemyakov, A. O.

2017-12-01

The problem of spacecraft attitude control using electromagnetic systems interacting with the Earth's magnetic field is considered. A set of dimensionless parameters has been formed to investigate the spacecraft orientation regimes based on dynamically similar models. The results of experimental studies of small spacecraft with a magnetic attitude control system can be extrapolated to the in-orbit spacecraft motion control regimes by using the methods of the dimensional and similarity theory.

Spacecraft Hybrid (Mixed-Actuator) Attitude Control Experiences on NASA Science Missions

Science.gov (United States)

Dennehy, Cornelius J.

2014-01-01

There is a heightened interest within NASA for the design, development, and flight implementation of mixed-actuator hybrid attitude control systems for science spacecraft that have less than three functional reaction wheel actuators. This interest is driven by a number of recent reaction wheel failures on aging, but what could be still scientifically productive, NASA spacecraft if a successful hybrid attitude control mode can be implemented. Over the years, hybrid (mixed-actuator) control has been employed for contingency attitude control purposes on several NASA science mission spacecraft. This paper provides a historical perspective of NASA's previous engineering work on spacecraft mixed-actuator hybrid control approaches. An update of the current situation will also be provided emphasizing why NASA is now so interested in hybrid control. The results of the NASA Spacecraft Hybrid Attitude Control Workshop, held in April of 2013, will be highlighted. In particular, the lessons learned captured from that workshop will be shared in this paper. An update on the most recent experiences with hybrid control on the Kepler spacecraft will also be provided. This paper will close with some future considerations for hybrid spacecraft control.

Global Precipitation Measurement (GPM) Mission Core Spacecraft Systems Engineering Challenges

Science.gov (United States)

Bundas, David J.; ONeill, Deborah; Field, Thomas; Meadows, Gary; Patterson, Peter

2006-01-01

The Global Precipitation Measurement (GPM) Mission is a collaboration between the National Aeronautics and Space Administration (NASA) and the Japanese Aerospace Exploration Agency (JAXA), and other US and international partners, with the goal of monitoring the diurnal and seasonal variations in precipitation over the surface of the earth. These measurements will be used to improve current climate models and weather forecasting, and enable improved storm and flood warnings. This paper gives an overview of the mission architecture and addresses the status of some key trade studies, including the geolocation budgeting, design considerations for spacecraft charging, and design issues related to the mitigation of orbital debris.

Spacecraft Charging: Hazard Causes, Hazard Effects, Hazard Controls

Science.gov (United States)

Koontz, Steve.

2018-01-01

Spacecraft flight environments are characterized both by a wide range of space plasma conditions and by ionizing radiation (IR), solar ultraviolet and X-rays, magnetic fields, micrometeoroids, orbital debris, and other environmental factors, all of which can affect spacecraft performance. Dr. Steven Koontz's lecture will provide a solid foundation in the basic engineering physics of spacecraft charging and charging effects that can be applied to solving practical spacecraft and spacesuit engineering design, verification, and operations problems, with an emphasis on spacecraft operations in low-Earth orbit, Earth's magnetosphere, and cis-Lunar space.

Optical System Design and Integration of the Mercury Laser Altimeter

Science.gov (United States)

Ramos-Izquierdo, Luis; Scott, V. Stanley, III; Schmidt, Stephen; Britt, Jamie; Mamakos, William; Trunzo, Raymond; Cavanaugh, John; Miller, Roger

2005-01-01

The Mercury Laser Altimeter (MLA). developed for the 2004 MESSENGER mission to Mercury, is designed to measure the planet's topography via laser ranging. A description of the MLA optical system and its measured optical performance during instrument-level and spacecraft-level integration and testing are presented.

3D Reconfigurable MPSoC for Unmanned Spacecraft Navigation

Science.gov (United States)

Dekoulis, George

2016-07-01

This paper describes the design of a new lightweight spacecraft navigation system for unmanned space missions. The system addresses the demands for more efficient autonomous navigation in the near-Earth environment or deep space. The proposed instrumentation is directly suitable for unmanned systems operation and testing of new airborne prototypes for remote sensing applications. The system features a new sensor technology and significant improvements over existing solutions. Fluxgate type sensors have been traditionally used in unmanned defense systems such as target drones, guided missiles, rockets and satellites, however, the guidance sensors' configurations exhibit lower specifications than the presented solution. The current implementation is based on a recently developed material in a reengineered optimum sensor configuration for unprecedented low-power consumption. The new sensor's performance characteristics qualify it for spacecraft navigation applications. A major advantage of the system is the efficiency in redundancy reduction achieved in terms of both hardware and software requirements.

Optimal Design of Integrated Systems Health Management (ISHM) Systems for improving safety in NASA's Exploration Vehicles: A Two-Level Multidisciplinary Design Approach

Science.gov (United States)

Mehr, Ali Farhang; Tumer, Irem; Barszcz, Eric

2005-01-01

Integrated Vehicle Health Management (ISHM) systems are used to detect, assess, and isolate functional failures in order to improve safety of space systems such as Orbital Space Planes (OSPs). An ISHM system, as a whole, consists of several subsystems that monitor different components of an OSP including: Spacecraft, Launch Vehicle, Ground Control, and the International Space Station. In this research, therefore, we propose a new methodology to design and optimize ISHM as a distributed system with multiple disciplines (that correspond to different subsystems of OSP safety). A paramount amount of interest has been given in the literature to the multidisciplinary design optimization of problems with such architecture (as will be reviewed in the full paper).

LDEF materials results for spacecraft applications: Executive summary

Science.gov (United States)

Whitaker, A. F.; Dooling, D.

1995-03-01

To address the challenges of space environmental effects, NASA designed the Long Duration Exposure Facility (LDEF) for an 18-month mission to expose thousands of samples of candidate materials that might be used on a space station or other orbital spacecraft. LDEF was launched in April 1984 and was to have been returned to Earth in 1985. Changes in mission schedules postponed retrieval until January 1990, after 69 months in orbit. Analyses of the samples recovered from LDEF have provided spacecraft designers and managers with the most extensive data base on space materials phenomena. Many LDEF samples were greatly changed by extended space exposure. Among even the most radially altered samples, NASA and its science teams are finding a wealth of surprising conclusions and tantalizing clues about the effects of space on materials. Many were discussed at the first two LDEF results conferences and subsequent professional papers. The LDEF Materials Results for Spacecraft Applications Conference was convened in Huntsville to discuss implications for spacecraft design. Already, paint and thermal blanket selections for space station and other spacecraft have been affected by LDEF data. This volume synopsizes those results.

Compact Chemical Monitor for Silver Ions in Spacecraft Water Systems, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — NASA has identified silver ions as the best candidate biocide for use in the potable water system on next-generation spacecraft. Though significant work has been...

Time Frequency Analysis of Spacecraft Propellant Tank Spinning Slosh

Science.gov (United States)

Green, Steven T.; Burkey, Russell C.; Sudermann, James

2010-01-01

Many spacecraft are designed to spin about an axis along the flight path as a means of stabilizing the attitude of the spacecraft via gyroscopic stiffness. Because of the assembly requirements of the spacecraft and the launch vehicle, these spacecraft often spin about an axis corresponding to a minor moment of inertia. In such a case, any perturbation of the spin axis will cause sloshing motions in the liquid propellant tanks that will eventually dissipate enough kinetic energy to cause the spin axis nutation (wobble) to grow further. This spinning slosh and resultant nutation growth is a primary design problem of spinning spacecraft and one that is not easily solved by analysis or simulation only. Testing remains the surest way to address spacecraft nutation growth. This paper describes a test method and data analysis technique that reveal the resonant frequency and damping behavior of liquid motions in a spinning tank. Slosh resonant frequency and damping characteristics are necessary inputs to any accurate numerical dynamic simulation of the spacecraft.

Small Spacecraft for Planetary Science

Science.gov (United States)

Baker, John; Castillo-Rogez, Julie; Bousquet, Pierre-W.; Vane, Gregg; Komarek, Tomas; Klesh, Andrew

2016-07-01

As planetary science continues to explore new and remote regions of the Solar system with comprehensive and more sophisticated payloads, small spacecraft offer the possibility for focused and more affordable science investigations. These small spacecraft or micro spacecraft (attitude control and determination, capable computer and data handling, and navigation are being met by technologies currently under development to be flown on CubeSats within the next five years. This paper will discuss how micro spacecraft offer an attractive alternative to accomplish specific science and technology goals and what relevant technologies are needed for these these types of spacecraft. Acknowledgements: Part of this work is being carried out at the Jet Propulsion Laboratory, California Institute of Technology under contract to NASA. Government sponsorship acknowledged.

Micro-Inspector Spacecraft for Space Exploration Missions

Science.gov (United States)

Mueller, Juergen; Alkalai, Leon; Lewis, Carol

2005-01-01

NASA is seeking to embark on a new set of human and robotic exploration missions back to the Moon, to Mars, and destinations beyond. Key strategic technical challenges will need to be addressed to realize this new vision for space exploration, including improvements in safety and reliability to improve robustness of space operations. Under sponsorship by NASA's Exploration Systems Mission, the Jet Propulsion Laboratory (JPL), together with its partners in government (NASA Johnson Space Center) and industry (Boeing, Vacco Industries, Ashwin-Ushas Inc.) is developing an ultra-low mass (missions. The micro-inspector will provide remote vehicle inspections to ensure safety and reliability, or to provide monitoring of in-space assembly. The micro-inspector spacecraft represents an inherently modular system addition that can improve safety and support multiple host vehicles in multiple applications. On human missions, it may help extend the reach of human explorers, decreasing human EVA time to reduce mission cost and risk. The micro-inspector development is the continuation of an effort begun under NASA's Office of Aerospace Technology Enabling Concepts and Technology (ECT) program. The micro-inspector uses miniaturized celestial sensors; relies on a combination of solar power and batteries (allowing for unlimited operation in the sun and up to 4 hours in the shade); utilizes a low-pressure, low-leakage liquid butane propellant system for added safety; and includes multi-functional structure for high system-level integration and miniaturization. Versions of this system to be designed and developed under the H&RT program will include additional capabilities for on-board, vision-based navigation, spacecraft inspection, and collision avoidance, and will be demonstrated in a ground-based, space-related environment. These features make the micro-inspector design unique in its ability to serve crewed as well as robotic spacecraft, well beyond Earth-orbit and into arenas such

An Overview of the Space Environments and Spacecraft Effects Organization Concept

Science.gov (United States)

Edwards, David L.; Burns, Howard D.; Garrett, Henry B.; Miller, Sharon K.; Peddie, Darilyn; Porter Ron; Spann, James F.; Xapsos, Michael A.

2012-01-01

The National Aeronautics and Space Administration (NASA) is embarking on a course to expand human presence beyond Low Earth Orbit (LEO) while also expanding its mission to explore our Earth, and the solar system. Destinations such as Near Earth Asteroids (NEA), Mars and its moons, and the outer planets are but a few of the mission targets. Each new destination presents an opportunity to increase our knowledge on the solar system and the unique environments for each mission target. NASA has multiple technical and science discipline areas specializing in specific space environments fields that will serve to enable these missions. To complement these existing discipline areas, a concept is presented focusing on the development of a space environment and spacecraft effects (SESE) organization. This SESE organization includes disciplines such as space climate, space weather, natural and induced space environments, effects on spacecraft materials and systems, and the transition of research information into application. This space environment and spacecraft effects organization will be composed of Technical Working Groups (TWG). These technical working groups will survey customers and users, generate products, and provide knowledge supporting four functional areas: design environments, engineering effects, operational support, and programmatic support. The four functional areas align with phases in the program mission lifecycle and are briefly described below. Design environments are used primarily in the mission concept and design phases of a program. Environment effects focuses on the material, component, sub-system, and system-level response to the space environment and include the selection and testing to verify design and operational performance. Operational support provides products based on real time or near real time space weather to mission operators to aid in real time and near-term decision-making. The programmatic support function maintains an interface with

Artificial intelligence costs, benefits, and risks for selected spacecraft ground system automation scenarios

Science.gov (United States)

Truszkowski, Walter F.; Silverman, Barry G.; Kahn, Martha; Hexmoor, Henry

1988-01-01

In response to a number of high-level strategy studies in the early 1980s, expert systems and artificial intelligence (AI/ES) efforts for spacecraft ground systems have proliferated in the past several years primarily as individual small to medium scale applications. It is useful to stop and assess the impact of this technology in view of lessons learned to date, and hopefully, to determine if the overall strategies of some of the earlier studies both are being followed and still seem relevant. To achieve that end four idealized ground system automation scenarios and their attendant AI architecture are postulated and benefits, risks, and lessons learned are examined and compared. These architectures encompass: (1) no AI (baseline); (2) standalone expert systems; (3) standardized, reusable knowledge base management systems (KBMS); and (4) a futuristic unattended automation scenario. The resulting artificial intelligence lessons learned, benefits, and risks for spacecraft ground system automation scenarios are described.

Putting Integrated Systems Health Management Capabilities to Work: Development of an Advanced Caution and Warning System for Next-Generation Crewed Spacecraft Missions

Science.gov (United States)

Mccann, Robert S.; Spirkovska, Lilly; Smith, Irene

2013-01-01

Integrated System Health Management (ISHM) technologies have advanced to the point where they can provide significant automated assistance with real-time fault detection, diagnosis, guided troubleshooting, and failure consequence assessment. To exploit these capabilities in actual operational environments, however, ISHM information must be integrated into operational concepts and associated information displays in ways that enable human operators to process and understand the ISHM system information rapidly and effectively. In this paper, we explore these design issues in the context of an advanced caution and warning system (ACAWS) for next-generation crewed spacecraft missions. User interface concepts for depicting failure diagnoses, failure effects, redundancy loss, "what-if" failure analysis scenarios, and resolution of ambiguity groups are discussed and illustrated.

The Global Precipitation Measurement (GPM Spacecraft Power System Design and Orbital Performance

Directory of Open Access Journals (Sweden)

Dakermanji George

2017-01-01

The paper describes the power system design details, its performance to date and the lithium ion battery model that was developed for use in the energy balance analysis and is being used to predict the on-orbit health of the battery.

Nonlinear model and attitude dynamics of flexible spacecraft with large amplitude slosh

Science.gov (United States)

Deng, Mingle; Yue, Baozeng

2017-04-01

This paper is focused on the nonlinearly modelling and attitude dynamics of spacecraft coupled with large amplitude liquid sloshing dynamics and flexible appendage vibration. The large amplitude fuel slosh dynamics is included by using an improved moving pulsating ball model. The moving pulsating ball model is an equivalent mechanical model that is capable of imitating the whole liquid reorientation process. A modification is introduced in the capillary force computation in order to more precisely estimate the settling location of liquid in microgravity or zero-g environment. The flexible appendage is modelled as a three dimensional Bernoulli-Euler beam and the assumed modal method is employed to derive the nonlinear mechanical model for the overall coupled system of liquid filled spacecraft with appendage. The attitude maneuver is implemented by the momentum transfer technique, and a feedback controller is designed. The simulation results show that the liquid sloshing can always result in nutation behavior, but the effect of flexible deformation of appendage depends on the amplitude and direction of attitude maneuver performed by spacecraft. Moreover, it is found that the liquid sloshing and the vibration of flexible appendage are coupled with each other, and the coupling becomes more significant with more rapid motion of spacecraft. This study reveals that the appendage's flexibility has influence on the liquid's location and settling time in microgravity. The presented nonlinear system model can provide an important reference for the overall design of the modern spacecraft composed of rigid platform, liquid filled tank and flexible appendage.

Robustness and Actuator Bandwidth of MRP-Based Sliding Mode Control for Spacecraft Attitude Control Problems

Science.gov (United States)

Keum, Jung-Hoon; Ra, Sung-Woong

2009-12-01

Nonlinear sliding surface design in variable structure systems for spacecraft attitude control problems is studied. A robustness analysis is performed for regular form of system, and calculation of actuator bandwidth is presented by reviewing sliding surface dynamics. To achieve non-singular attitude description and minimal parameterization, spacecraft attitude control problems are considered based on modified Rodrigues parameters (MRP). It is shown that the derived controller ensures the sliding motion in pre-determined region irrespective of unmodeled effects and disturbances.

Spacecraft Testing Programs: Adding Value to the Systems Engineering Process

Science.gov (United States)

Britton, Keith J.; Schaible, Dawn M.

2011-01-01

Testing has long been recognized as a critical component of spacecraft development activities - yet many major systems failures may have been prevented with more rigorous testing programs. The question is why is more testing not being conducted? Given unlimited resources, more testing would likely be included in a spacecraft development program. Striking the right balance between too much testing and not enough has been a long-term challenge for many industries. The objective of this paper is to discuss some of the barriers, enablers, and best practices for developing and sustaining a strong test program and testing team. This paper will also explore the testing decision factors used by managers; the varying attitudes toward testing; methods to develop strong test engineers; and the influence of behavior, culture and processes on testing programs. KEY WORDS: Risk, Integration and Test, Validation, Verification, Test Program Development

Development and Analysis of a Resource-Aware Power Management System as Applied to Small Spacecraft

Energy Technology Data Exchange (ETDEWEB)

Shriver, Patrick [Univ. of Colorado, Boulder, CO (United States)

2005-01-01

In this thesis, an overall framework and solution method for managing the limited power resources of a small spacecraft is presented. Analogous to mobile computing technology, a primary limiting factor is the available power resources. In spite of the millions of dollars budgeted for research and development over decades, improvements in battery efficiency remains low. This situation is exacerbated by advances in payload technology that lead to increasingly power-hungry and data-intensive instruments. The challenge for the small spacecraft is to maximize capabilities and performance while meeting difficult design requirements and small project budgets.

Influence of a new generation of operations support systems on current spacecraft operations philosophy: The users feedback

Science.gov (United States)

Darroy, Jean Michel

1993-01-01

Current trends in the spacecraft mission operations area (spacecraft & mission complexity, project duration, required flexibility are requiring a breakthrough for what concerns philosophy, organization, and support tools. A major evolution is related to space operations 'informationalization', i.e adding to existing operations support & data processing systems a new generation of tools based on advanced information technologies (object-oriented programming, artificial intelligence, data bases, hypertext) that automate, at least partially, operations tasks that used be performed manually (mission & project planning/scheduling, operations procedures elaboration & execution, data analysis & failure diagnosis). All the major facets of this 'informationalization' are addressed at MATRA MARCONI SPACE, operational applications were fielded and generic products are becoming available. These various applications have generated a significant feedback from the users (at ESA, CNES, ARIANESPACE, MATRA MARCONI SPACE), which is now allowing us to precisely measure how the deployment of this new generation of tools, that we called OPSWARE, can 'reengineer' current spacecraft mission operations philosophy, how it can make space operations faster, better, and cheaper. This paper can be considered as an update of the keynote address 'Knowledge-Based Systems for Spacecraft Control' presented during the first 'Ground Data Systems for Spacecraft Control' conference in Darmstadt, June 1990, with a special emphasis on these last two years users feedback.

Reactor/Brayton power systems for nuclear electric spacecraft

Science.gov (United States)

Layton, J. P.

1980-01-01

Studies are currently underway to assess the technological feasibility of a nuclear-reactor-powered spacecraft propelled by electric thrusters. This vehicle would be capable of performing detailed exploration of the outer planets of the solar system during the remainder of this century. The purpose of this study was to provide comparative information on a closed cycle gas turbine power conversion system. The results have shown that the performance is very competitive and that a 400 kWe space power system is dimensionally compatible with a single Space Shuttle launch. Performance parameters of system mass and radiator area were determined for systems from 100 to 1000 kWe. A 400 kWe reference system received primary attention. The components of this system were defined and a conceptual layout was developed with encouraging results. The preliminary mass determination for the complete power system was very close to the desired goal of 20 kg/kWe. Use of more advanced technology (higher turbine inlet temperature) will substantially improve system performance characteristics.

Small Rocket/Spacecraft Technology (SMART) Platform

Science.gov (United States)

Esper, Jaime; Flatley, Thomas P.; Bull, James B.; Buckley, Steven J.

2011-01-01

The NASA Goddard Space Flight Center (GSFC) and the Department of Defense Operationally Responsive Space (ORS) Office are exercising a multi-year collaborative agreement focused on a redefinition of the way space missions are designed and implemented. A much faster, leaner and effective approach to space flight requires the concerted effort of a multi-agency team tasked with developing the building blocks, both programmatically and technologically, to ultimately achieve flights within 7-days from mission call-up. For NASA, rapid mission implementations represent an opportunity to find creative ways for reducing mission life-cycle times with the resulting savings in cost. This in tum enables a class of missions catering to a broader audience of science participants, from universities to private and national laboratory researchers. To that end, the SMART (Small Rocket/Spacecraft Technology) micro-spacecraft prototype demonstrates an advanced avionics system with integrated GPS capability, high-speed plug-and-playable interfaces, legacy interfaces, inertial navigation, a modular reconfigurable structure, tunable thermal technology, and a number of instruments for environmental and optical sensing. Although SMART was first launched inside a sounding rocket, it is designed as a free-flyer.

SSTI- Lewis Spacecraft Nickel-Hydrogen Battery

Science.gov (United States)

Tobias, R. F.

1997-01-01

Topics considered include: NASA-Small Spacecraft Technology Initiative (SSTI) objectives, SSTI-Lewis overview, battery requirement, two cells Common Pressure Vessel (CPV) design summary, CPV electric performance, battery design summary, battery functional description, battery performance.

Relativistic Spacecraft Propelled by Directed Energy

Science.gov (United States)

Kulkarni, Neeraj; Lubin, Philip; Zhang, Qicheng

2018-04-01

Achieving relativistic flight to enable extrasolar exploration is one of the dreams of humanity and the long-term goal of our NASA Starlight program. We derive a relativistic solution for the motion of a spacecraft propelled by radiation pressure from a directed energy (DE) system. Depending on the system parameters, low-mass spacecraft can achieve relativistic speeds, thus enabling interstellar exploration. The diffraction of the DE system plays an important role and limits the maximum speed of the spacecraft. We consider “photon recycling” as a possible method to achieving higher speeds. We also discuss recent claims that our previous work on this topic is incorrect and show that these claims arise from an improper treatment of causality.

Space Transportation System Cargo projects: inertial stage/spacecraft integration plan. Volume 1: Management plan

Science.gov (United States)

1981-01-01

The Kennedy Space Center (KSC) Management System for the Inertial Upper Stage (IUS) - spacecraft processing from KSC arrival through launch is described. The roles and responsibilities of the agencies and test team organizations involved in IUS-S/C processing at KSC for non-Department of Defense missions are described. Working relationships are defined with respect to documentation preparation, coordination and approval, schedule development and maintenance, test conduct and control, configuration management, quality control and safety. The policy regarding the use of spacecraft contractor test procedures, IUS contractor detailed operating procedures and KSC operations and maintenance instructions is defined. Review and approval requirements for each documentation system are described.

Analysis of the Apollo spacecraft operational data management system. Executive summary

Science.gov (United States)

1971-01-01

A study was made of Apollo, Skylab, and several other data management systems to determine those techniques which could be applied to the management of operational data for future manned spacecraft programs. The results of the study are presented and include: (1) an analysis of present data management systems, (2) a list of requirements for future operational data management systems, (3) an evaluation of automated data management techniques, and (4) a plan for data management applicable to future space programs.

The Near Earth Object Scout Spacecraft: A Low Cost Approach to in-situ Characterization of the NEO Population

Science.gov (United States)

Koontz, Steven L.; Condon, Gerald; Graham, Lee; Bevilacqua, Ricardo

2014-01-01

In this paper we describe a micro/nano satellite spacecraft and a supporting mission profile and architecture designed to enable preliminary in-situ characterization of a significant number of Near Earth Objects (NEOs) at reasonable cost. The spacecraft will be referred to as the NEO Scout. NEO Scout spacecraft are to be placed in GTO, GEO, or cis-lunar space as secondary payloads on launch vehicles headed for GTO or beyond and will begin their mission after deployment from the launcher. A distinguishing key feature of the NEO scout system is to design the mission timeline and spacecraft to rendezvous with and land on the target NEOs during close approach to the Earth-Moon system using low-thrust/high- impulse propulsion systems. Mission feasibility and preliminary design analysis are presented along with detailed trajectory calculations. The use of micro/nano satellites in low-cost interplanetary exploration is attracting increasing attention and is the subject of several annual workshops and published design studies (1-4). The NEO population consists of those asteroids and short period comets orbiting the Sun with a perihelion of 1.3 astronomical units or less (5-8). As of July 30, 2013 10065 Near-Earth objects have been discovered. The spin rate, mass, density, surface physical (especially mechanical) properties, composition, and mineralogy of the vast majority of these objects are highly uncertain and the limited available telescopic remote sensing data imply a very diverse population (5-8). In-situ measurements by robotic spacecraft are urgently needed to provide the characterization data needed to support hardware and mission design for more ambitious human and robotic NEO operations. Large numbers of NEOs move into close proximity with the Earth-Moon system every year (9). The JPL Near-Earth Object Human Space Flight Accessible Targets Study (NHATS) (10) has produced detailed mission profile and delta V requirements for various NEO missions ranging from 30

Large-Scale Spacecraft Fire Safety Tests

Science.gov (United States)

Urban, David; Ruff, Gary A.; Ferkul, Paul V.; Olson, Sandra; Fernandez-Pello, A. Carlos; T'ien, James S.; Torero, Jose L.; Cowlard, Adam J.; Rouvreau, Sebastien; Minster, Olivier;

The design and development of a spacecraft appendage tie down mechanism

Science.gov (United States)

Nygren, W. D.; Head, R.

1985-01-01

The design and evolution is described of a spacecraft Appendage Tie Down Mechanism (ATDM). Particular emphasis is paid to the mechanical aspects of using dry lubricants to increase the efficiency of acme threads and worm gearing. The ATDM consists of five major components. These are a dc torque motor, a worm gear speed reducer, the tension bolt (or T-bolt), nut capture and centering jaws and the capture nut. In addition, there are several minor components such as limit switch assemblies and an antibackdrive mechanism which couples the drive motor to the worm shaft. A development model of the ATDM in various configurations was under test for some time. In its latest version, it has successfully completed thermal vacuum testing, vibration testing, and extended life testing.

Galileo spacecraft solid-state imaging system view of Antarctica

Science.gov (United States)

1990-01-01

Galileo spacecraft solid-state imaging system view of Antarctica was taken during its first encounter with the Earth. This color picture of Antarctica is part of a mosaic of pictures covering the entire polar continent showing the Ross Ice Shelf and its border with the sea and mountains poking through the ice near the McMurdo Station. From top to bottom, the frame looks across about half of Antarctica. View provided by the Jet Propulsion Laboratory (JPL) with alternate number P-37297.

Thermal analysis of hybrid single-phase, two-phase and heat pump thermal control system (TCS) for future spacecraft

International Nuclear Information System (INIS)

Lee, S.H.; Mudawar, I.; Hasan, Mohammad M.

2016-01-01

Highlights: • Hybrid Thermal Control System (H-TCS) is proposed for future spacecraft. • Thermodynamic performance of H-TCS is examined for different space missions. • Operational modes including single-phase, two-phase and heat pump are explored. • R134a is deemed most appropriate working fluid. - Abstract: An urgent need presently exists to develop a new class of versatile spacecraft capable of conducting different types of missions and enduring varying gravitational and temperature environments, including Lunar, Martian and Near Earth Object (NEOs). This study concerns the spacecraft's Thermal Control System (TCS), which tackles heat acquisition, especially from crew and avionics, heat transport, and ultimate heat rejection by radiation. The primary goal of the study is to explore the design and thermal performance of a Hybrid Thermal Control System (H-TCS) that would satisfy the diverse thermal requirements of the different space missions. The H-TCS must endure both ‘cold’ and ‘hot’ environments, reduce weight and size, and enhance thermodynamic performance. Four different operational modes are considered: single-phase, two-phase, basic heat pump and heat pump with liquid-side, suction-side heat exchanger. A thermodynamic trade study is conducted for six different working fluids to assess important performance parameters including mass flow rate of the working fluid, maximum pressure, radiator area, compressor/pump work, and coefficient of performance (COP). R134a is determined to be most suitable based on its ability to provide a balanced compromise between reducing flow rate and maintaining low system pressure, and a moderate coefficient of performance (COP); this fluid is also both nontoxic and nonflammable, and features zero ozone depletion potential (ODP) and low global warming potential (GWP). It is shown how specific mission stages dictate which mode of operation is most suitable, and this information is used to size the radiator for the

Revamping Spacecraft Operational Intelligence

Science.gov (United States)

Hwang, Victor

2012-01-01

The EPOXI flight mission has been testing a new commercial system, Splunk, which employs data mining techniques to organize and present spacecraft telemetry data in a high-level manner. By abstracting away data-source specific details, Splunk unifies arbitrary data formats into one uniform system. This not only reduces the time and effort for retrieving relevant data, but it also increases operational visibility by allowing a spacecraft team to correlate data across many different sources. Splunk's scalable architecture coupled with its graphing modules also provide a solid toolset for generating data visualizations and building real-time applications such as browser-based telemetry displays.

Development of an advanced spacecraft tandem mass spectrometer

Science.gov (United States)

Drew, Russell C.

1992-03-01

The purpose of this research was to apply current advanced technology in electronics and materials to the development of a miniaturized Tandem Mass Spectrometer that would have the potential for future development into a package suitable for spacecraft use. The mass spectrometer to be used as a basis for the tandem instrument would be a magnetic sector instrument, of Nier-Johnson configuration, as used on the Viking Mars Lander mission. This instrument configuration would then be matched with a suitable second stage MS to provide the benefits of tandem MS operation for rapid identification of unknown organic compounds. This tandem instrument is configured with a newly designed GC system to aid in separation of complex mixtures prior to MS analysis. A number of important results were achieved in the course of this project. Among them were the development of a miniaturized GC subsystem, with a unique desorber-injector, fully temperature feedback controlled oven with powered cooling for rapid reset to ambient conditions, a unique combination inlet system to the MS that provides for both membrane sampling and direct capillary column sample transfer, a compact and ruggedized alignment configuration for the MS, an improved ion source design for increased sensitivity, and a simple, rugged tandem MS configuration that is particularly adaptable to spacecraft use because of its low power and low vacuum pumping requirements. The potential applications of this research include use in manned spacecraft like the space station as a real-time detection and warning device for the presence of potentially harmful trace contaminants of the spacecraft atmosphere, use as an analytical device for evaluating samples collected on the Moon or a planetary surface, or even use in connection with monitoring potentially hazardous conditions that may exist in terrestrial locations such as launch pads, environmental test chambers or other sensitive areas. Commercial development of the technology

Small Spacecraft Constellation Concept for Mars Atmospheric Radio Occultations

Science.gov (United States)

Asmar, S. W.; Mannucci, A. J.; Ao, C. O.; Kobayashi, M. M.; Lazio, J.; Marinan, A.; Massone, G.; McCandless, S. E.; Preston, R. A.; Seubert, J.; Williamson, W.

2017-12-01

First demonstrated in 1965 when Mariner IV flew by Mars and determined the salient features of its atmosphere, radio occultation experiments have been carried out on numerous planetary missions with great discoveries. These experiments utilize the now classic configuration of a signal from a single planetary spacecraft to Earth receiving stations, where the science data are acquired. The Earth science community advanced the technique to utilizing a constellation of spacecraft with the radio occultation links between the spacecraft, enabled by the infrastructure of the Global Positioning System. With the advent of small and less costly spacecraft, such as planetary CubeSats and other variations, such as the anticipated innovative Mars Cube One mission, crosslinks among small spacecraft can be used to study other planets in the near future. Advantages of this type of experiment include significantly greater geographical coverage, which could reach global coverage over a few weeks with a small number of spacecraft. Repeatability of the global coverage can lead to examining temperature-pressure profiles and ionospheric electron density profiles, on daily, seasonal, annual, or other time scales of interest. The higher signal-to-noise ratio for inter-satellite links, compared to a link to Earth, decreases the design demands on the instrumentation (smaller antennas and transmitters, etc.). After an actual Mars crosslink demonstration, this concept has been in development using Mars as a possible target. Scientific objectives, delivery methods, operational scenarios and end-to-end configuration have been documented. Science objectives include determining the state and variability of the lower Martian atmosphere, which has been an identified as a high priority objective by the Mars Exploration Program Analysis Group, particularly as it relates to entry, descent, and landing and ascent for future crewed and robotic missions. This paper will present the latest research on the

Integrating standard operating procedures with spacecraft automation, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Spacecraft automation has the potential to assist crew members and spacecraft operators in managing spacecraft systems during extended space missions. Automation can...

The spacecraft encounters of Comet Halley

Science.gov (United States)

Asoka Mendis, D.; Tsurutani, Bruce T.

1986-01-01

The characteristics of the Comet Halley spacecraft 'fleet' (VEGA 1 and VEGA 2, Giotto, Suisei, and Sakigake) are presented. The major aims of these missions were (1) to discover and characterize the nucleus, (2) to characterize the atmosphere and ionosphere, (3) to characterize the dust, and (4) to characterize the nature of the large-scale comet-solar wind interaction. While the VEGA and Giotto missions were designed to study all four areas, Suisei addressed the second and fourth. Sakigake was designed to study the solar wind conditions upstream of the comet. It is noted that NASA's Deep Space Network played an important role in spacecraft tracking.

Space Evaporator Absorber Radiator (SEAR) for Thermal Storage on Manned Spacecraft

Science.gov (United States)

Izenson, Michael G.; Chen, Weibo; Chepko, Ariane; Bue, Grant; Quinn, Gregory

2015-01-01

Future manned exploration spacecraft will need to operate in challenging thermal environments. State-of-the-art technology for active thermal control relies on sublimating water ice and venting the vapor overboard in very hot environments, and or heavy phase change material heat exchangers for thermal storage. These approaches can lead to large loss of water and a significant mass penalties for the spacecraft. This paper describes an innovative thermal control system that uses a Space Evaporator Absorber Radiator (SEAR) to control spacecraft temperatures in highly variable environments without venting water. SEAR uses heat pumping and energy storage by LiCl/water absorption to enable effective cooling during hot periods and regeneration during cool periods. The LiCl absorber technology has the potential to absorb over 800 kJ per kg of system mass, compared to phase change heat sink systems that typically achieve approx. 50 kJ/kg. This paper describes analysis models to predict performance and optimize the size of the SEAR system, estimated size and mass of key components, and an assessment of potential mass savings compared with alternative thermal management approaches. We also describe a concept design for an ISS test package to demonstrate operation of a subscale system in zero gravity.

Temperature control of the Mariner class spacecraft - A seven mission summary.

Science.gov (United States)

Dumas, L. N.

1973-01-01

Mariner spacecraft have completed five missions of scientific investigation of the planets. Two additional missions are planned. A description of the thermal design of these seven spacecraft is given herein. The factors which have influenced the thermal design include the mission requirements and constraints, the flight environment, certain programmatic considerations and the experience gained as each mission is completed. These factors are reviewed and the impact of each on thermal design and developmental techniques is assessed. It is concluded that the flight success of these spacecraft indicates that adequate temperature control has been obtained, but that improvements in design data, hardware performance and analytical techniques are needed.

Cooperative control of multi-agent systems optimal and adaptive design approaches

CERN Document Server

Lewis, Frank L; Hengster-Movric, Kristian; Das, Abhijit

2014-01-01

Task complexity, communication constraints, flexibility and energy-saving concerns are all factors that may require a group of autonomous agents to work together in a cooperative manner. Applications involving such complications include mobile robots, wireless sensor networks, unmanned aerial vehicles (UAVs), spacecraft, and so on. In such networked multi-agent scenarios, the restrictions imposed by the communication graph topology can pose severe problems in the design of cooperative feedback control systems.  Cooperative control of multi-agent systems is a challenging topic for both control theorists and practitioners and has been the subject of significant recent research. Cooperative Control of Multi-Agent Systems extends optimal control and adaptive control design methods to multi-agent systems on communication graphs.  It develops Riccati design techniques for general linear dynamics for cooperative state feedback design, cooperative observer design, and cooperative dynamic output feedback design.  B...

Theoretical analysis of infrared radiation shields of spacecraft

Science.gov (United States)

Shealy, D. L.

1984-01-01

For a system of N diffuse, gray body radiation shields which view only adjacent surfaces and space, the net radiation method for enclosures has been used to formulate a system of linear, nonhomogeneous equations in terms of the temperatures to the fourth power of each surface in the coupled system of enclosures. The coefficients of the unknown temperatures in the system of equations are expressed in terms of configuration factors between adjacent surfaces and the emissivities. As an application, a system of four conical radiation shields for a spin stabilized STARPROBE spacecraft has been designed and analyzed with respect to variations of the cone half angles, the intershield spacings, and emissivities.

Training for spacecraft technical analysts

Science.gov (United States)

Ayres, Thomas J.; Bryant, Larry

1989-01-01

Deep space missions such as Voyager rely upon a large team of expert analysts who monitor activity in the various engineering subsystems of the spacecraft and plan operations. Senior teammembers generally come from the spacecraft designers, and new analysts receive on-the-job training. Neither of these methods will suffice for the creation of a new team in the middle of a mission, which may be the situation during the Magellan mission. New approaches are recommended, including electronic documentation, explicit cognitive modeling, and coached practice with archived data.

Historical Mass, Power, Schedule, and Cost Growth for NASA Spacecraft

Science.gov (United States)

Hayhurst, Marc R.; Bitten, Robert E.; Shinn, Stephen A.; Judnick, Daniel C.; Hallgrimson, Ingrid E.; Youngs, Megan A.

2016-01-01

Although spacecraft developers have been moving towards standardized product lines as the aerospace industry has matured, NASA's continual need to push the cutting edge of science to accomplish unique, challenging missions can still lead to spacecraft resource growth over time. This paper assesses historical mass, power, cost, and schedule growth for multiple NASA spacecraft from the last twenty years and compares to industry reserve guidelines to understand where the guidelines may fall short. Growth is assessed from project start to launch, from the time of the preliminary design review (PDR) to launch and from the time of the critical design review (CDR) to launch. Data is also assessed not just at the spacecraft bus level, but also at the subsystem level wherever possible, to help obtain further insight into possible drivers of growth. Potential recommendations to minimize spacecraft mass, power, cost, and schedule growth for future missions are also discussed.

Internet Technology on Spacecraft

Science.gov (United States)

Rash, James; Parise, Ron; Hogie, Keith; Criscuolo, Ed; Langston, Jim; Powers, Edward I. (Technical Monitor)

2000-01-01

The Operating Missions as Nodes on the Internet (OMNI) project has shown that Internet technology works in space missions through a demonstration using the UoSAT-12 spacecraft. An Internet Protocol (IP) stack was installed on the orbiting UoSAT-12 spacecraft and tests were run to demonstrate Internet connectivity and measure performance. This also forms the basis for demonstrating subsequent scenarios. This approach provides capabilities heretofore either too expensive or simply not feasible such as reconfiguration on orbit. The OMNI project recognized the need to reduce the risk perceived by mission managers and did this with a multi-phase strategy. In the initial phase, the concepts were implemented in a prototype system that includes space similar components communicating over the TDRS (space network) and the terrestrial Internet. The demonstration system includes a simulated spacecraft with sample instruments. Over 25 demonstrations have been given to mission and project managers, National Aeronautics and Space Administration (NASA), Department of Defense (DoD), contractor technologists and other decisions makers, This initial phase reached a high point with an OMNI demonstration given from a booth at the Johnson Space Center (JSC) Inspection Day 99 exhibition. The proof to mission managers is provided during this second phase with year 2000 accomplishments: testing the use of Internet technologies onboard an actual spacecraft. This was done with a series of tests performed using the UoSAT-12 spacecraft. This spacecraft was reconfigured on orbit at very low cost. The total period between concept and the first tests was only 6 months! On board software was modified to add an IP stack to support basic IP communications. Also added was support for ping, traceroute and network timing protocol (NTP) tests. These tests show that basic Internet functionality can be used onboard spacecraft. The performance of data was measured to show no degradation from current

Architectural and Behavioral Systems Design Methodology and Analysis for Optimal Habitation in a Volume-Limited Spacecraft for Long Duration Flights

Science.gov (United States)

Kennedy, Kriss J.; Lewis, Ruthan; Toups, Larry; Howard, Robert; Whitmire, Alexandra; Smitherman, David; Howe, Scott

2016-01-01

As our human spaceflight missions change as we reach towards Mars, the risk of an adverse behavioral outcome increases, and requirements for crew health, safety, and performance, and the internal architecture, will need to change to accommodate unprecedented mission demands. Evidence shows that architectural arrangement and habitability elements impact behavior. Net habitable volume is the volume available to the crew after accounting for elements that decrease the functional volume of the spacecraft. Determination of minimum acceptable net habitable volume and associated architectural design elements, as mission duration and environment varies, is key to enabling, maintaining, andor enhancing human performance and psychological and behavioral health. Current NASA efforts to derive minimum acceptable net habitable volumes and study the interaction of covariates and stressors, such as sensory stimulation, communication, autonomy, and privacy, and application to internal architecture design layouts, attributes, and use of advanced accommodations will be presented. Furthermore, implications of crew adaptation to available volume as they transfer from Earth accommodations, to deep space travel, to planetary surface habitats, and return, will be discussed.

Power requirements for commercial communications spacecraft

Science.gov (United States)

Billerbeck, W. J.

1985-01-01

Historical data on commercial spacecraft power systems are presented and their power requirements to the growth of satellite communications channel usage are related. Some approaches for estimating future power requirements of this class of spacecraft through the year 2000 are proposed. The key technology drivers in satellite power systems are addressed. Several technological trends in such systems are described, focusing on the most useful areas for research and development of major subsystems, including solar arrays, energy storage, and power electronics equipment.

Autonomous Navigation of the SSTI/Lewis Spacecraft Using the Global Positioning System (GPS)

Science.gov (United States)

Hart, R. C.; Long, A. C.; Lee, T.

1997-01-01

The National Aeronautics and Space Administration (NASA) Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) is pursuing the application of Global Positioning System (GPS) technology to improve the accuracy and economy of spacecraft navigation. High-accuracy autonomous navigation algorithms are being flight qualified in conjunction with GSFC's GPS Attitude Determination Flyer (GADFLY) experiment on the Small Satellite Technology Initiative (SSTI) Lewis spacecraft, which is scheduled for launch in 1997. Preflight performance assessments indicate that these algorithms can provide a real-time total position accuracy of better than 10 meters (1 sigma) and velocity accuracy of better than 0.01 meter per second (1 sigma), with selective availability at typical levels. This accuracy is projected to improve to the 2-meter level if corrections to be provided by the GPS Wide Area Augmentation System (WAAS) are included.

Imaging Flash Lidar for Safe Landing on Solar System Bodies and Spacecraft Rendezvous and Docking

Science.gov (United States)

Amzajerdian, Farzin; Roback, Vincent E.; Bulyshev, Alexander E.; Brewster, Paul F.; Carrion, William A; Pierrottet, Diego F.; Hines, Glenn D.; Petway, Larry B.; Barnes, Bruce W.; Noe, Anna M.

2015-01-01

NASA has been pursuing flash lidar technology for autonomous, safe landing on solar system bodies and for automated rendezvous and docking. During the final stages of the landing from about 1 kilometer to 500 meters above the ground, the flash lidar can generate 3-Dimensional images of the terrain to identify hazardous features such as craters, rocks, and steep slopes. The onboard flight computer can then use the 3-D map of terrain to guide the vehicle to a safe location. As an automated rendezvous and docking sensor, the flash lidar can provide relative range, velocity, and bearing from an approaching spacecraft to another spacecraft or a space station. NASA Langley Research Center has developed and demonstrated a flash lidar sensor system capable of generating 16,000 pixels range images with 7 centimeters precision, at 20 Hertz frame rate, from a maximum slant range of 1800 m from the target area. This paper describes the lidar instrument and presents the results of recent flight tests onboard a rocket-propelled free-flyer vehicle (Morpheus) built by NASA Johnson Space Center. The flights were conducted at a simulated lunar terrain site, consisting of realistic hazard features and designated landing areas, built at NASA Kennedy Space Center specifically for this demonstration test. This paper also provides an overview of the plan for continued advancement of the flash lidar technology aimed at enhancing its performance to meet both landing and automated rendezvous and docking applications.

Performance evaluation and design of flight vehicle control systems

CERN Document Server

Falangas, Eric T

2015-01-01

This book will help students, control engineers and flight dynamics analysts to model and conduct sophisticated and systemic analyses of early flight vehicle designs controlled with multiple types of effectors and to design and evaluate new vehicle concepts in terms of satisfying mission and performance goals. Performance Evaluation and Design of Flight Vehicle Control Systems begins by creating a dynamic model of a generic flight vehicle that includes a range of elements from airplanes and launch vehicles to re-entry vehicles and spacecraft. The models may include dynamic effects dealing with structural flexibility, as well as dynamic coupling between structures and actuators, propellant sloshing, and aeroelasticity, and they are typically used for control analysis and design. The book shows how to efficiently combine different types of effectors together, such as aero-surfaces, TVC, throttling engines and RCS, to operate as a system by developing a mixing logic atrix. Methods of trimming a vehicle controll...

Navigation Operations with Prototype Components of an Automated Real-Time Spacecraft Navigation System

Science.gov (United States)

Cangahuala, L.; Drain, T. R.

1999-01-01

At present, ground navigation support for interplanetary spacecraft requires human intervention for data pre-processing, filtering, and post-processing activities; these actions must be repeated each time a new batch of data is collected by the ground data system.

The Heliopause Electrostatic Rapid Transit System (HERTS) Design, Trades, and Analyses Performed in a Two Year NASA Investigation of Electric Sail Propulsion Systems

Science.gov (United States)

Wiegmann, Bruce M.

2017-01-01

The Heliopause Electrostatic Rapid Transit System (HERTS) was one of the seven total Phase II NASA Innovative Advanced Concepts (NIAC) that was down-selected in 2015 for continued funding and research. In Phase I our team learned that a spacecraft propelled by an Electric Sail (E-Sail) can travel great astronomical distances, such as to the Heliopause region of the solar system (approx. 100 to 120 AU) in approximately one quarter of the time (10 years) versus the time it took the Voyager spacecraft launched in 1977 (36 years). The completed work within the Phase II NIAC funded effort builds upon the work that was done in the Phase I NIAC and is focused on: 1) Testing of plasma interaction with a charged wire in a MSFC simulated solar environment vacuum test chamber. 2) Development of a Particle-in-Cell (PIC) models that are validated in the plasma testing and used to extrapolate to the E-Sail propulsion system design. 3) Conceptual design of a Technology Demonstration Mission (TDM) spacecraft developed to showcase E-Sail propulsion systems. 4) Down selection of both: a) Materials for a multi km length conductor and, b) Best configuration of the proposed conductor deployment subsystem. This paper will document the findings to date (June, 2017) of the above focused areas.

Spacecraft computer technology at Southwest Research Institute

Science.gov (United States)

Shirley, D. J.

1993-01-01

Southwest Research Institute (SwRI) has developed and delivered spacecraft computers for a number of different near-Earth-orbit spacecraft including shuttle experiments and SDIO free-flyer experiments. We describe the evolution of the basic SwRI spacecraft computer design from those weighing in at 20 to 25 lb and using 20 to 30 W to newer models weighing less than 5 lb and using only about 5 W, yet delivering twice the processing throughput. Because of their reduced size, weight, and power, these newer designs are especially applicable to planetary instrument requirements. The basis of our design evolution has been the availability of more powerful processor chip sets and the development of higher density packaging technology, coupled with more aggressive design strategies in incorporating high-density FPGA technology and use of high-density memory chips. In addition to reductions in size, weight, and power, the newer designs also address the necessity of survival in the harsh radiation environment of space. Spurred by participation in such programs as MSTI, LACE, RME, Delta 181, Delta Star, and RADARSAT, our designs have evolved in response to program demands to be small, low-powered units, radiation tolerant enough to be suitable for both Earth-orbit microsats and for planetary instruments. Present designs already include MIL-STD-1750 and Multi-Chip Module (MCM) technology with near-term plans to include RISC processors and higher-density MCM's. Long term plans include development of whole-core processors on one or two MCM's.

Research-Based Monitoring, Prediction, and Analysis Tools of the Spacecraft Charging Environment for Spacecraft Users

Science.gov (United States)

Zheng, Yihua; Kuznetsova, Maria M.; Pulkkinen, Antti A.; Maddox, Marlo M.; Mays, Mona Leila

2015-01-01

The Space Weather Research Center (http://swrc. gsfc.nasa.gov) at NASA Goddard, part of the Community Coordinated Modeling Center (http://ccmc.gsfc.nasa.gov), is committed to providing research-based forecasts and notifications to address NASA's space weather needs, in addition to its critical role in space weather education. It provides a host of services including spacecraft anomaly resolution, historical impact analysis, real-time monitoring and forecasting, tailored space weather alerts and products, and weekly summaries and reports. In this paper, we focus on how (near) real-time data (both in space and on ground), in combination with modeling capabilities and an innovative dissemination system called the integrated Space Weather Analysis system (http://iswa.gsfc.nasa.gov), enable monitoring, analyzing, and predicting the spacecraft charging environment for spacecraft users. Relevant tools and resources are discussed.

On-orbit supervisor for controlling spacecraft

Science.gov (United States)

Vandervoort, Richard J.

1992-07-01

Spacecraft systems of the 1990's and beyond will be substantially more complex than their predecessors. They will have demanding performance requirements and will be expected to operate more autonomously. This underscores the need for innovative approaches to Fault Detection, Isolation and Recovery (FDIR). A hierarchical expert system is presented that provides on-orbit supervision using intelligent FDIR techniques. Each expert system in the hierarchy supervises the operation of a local set of spacecraft functions. Spacecraft operational goals flow top down while responses flow bottom up. The expert system supervisors have a fairly high degree of autonomy. Bureaucratic responsibilities are minimized to conserve bandwidth and maximize response time. Data for FDIR can be acquired local to an expert and from other experts. By using a blackboard architecture for each supervisor, the system provides a great degree of flexibility in implementing the problem solvers for each problem domain. In addition, it provides for a clear separation between facts and knowledge, leading to an efficient system capable of real time response.

Application of advanced electronics to a future spacecraft computer design

Science.gov (United States)

Carney, P. C.

1980-01-01

Advancements in hardware and software technology are summarized with specific emphasis on spacecraft computer capabilities. Available state of the art technology is reviewed and candidate architectures are defined.

Design and Stability of an On-Orbit Attitude Control System Using Reaction Control Thrusters

Science.gov (United States)

Hall, Robert A.; Hough, Steven; Orphee, Carolina; Clements, Keith

2016-01-01

Basic principles for the design and stability of a spacecraft on-orbit attitude control system employing on-off Reaction Control System (RCS) thrusters are presented. Both vehicle dynamics and the control system actuators are inherently nonlinear, hence traditional linear control system design approaches are not directly applicable. This paper has two main aspects: It summarizes key RCS design principles from earlier NASA vehicles, notably the Space Shuttle and Space Station programs, and introduces advances in the linear modelling and analyses of a phase plane control system derived in the initial development of the NASA's next upper stage vehicle, the Exploration Upper Stage (EUS). Topics include thruster hardware specifications, phase plane design and stability, jet selection approaches, filter design metrics, and RCS rotational maneuver logic.

Radiation Effects on Spacecraft Structural Materials

International Nuclear Information System (INIS)

Wang, Jy-An J.; Ellis, Ronald J.; Hunter, Hamilton T.; Singleterry, Robert C. Jr.

2002-01-01

Research is being conducted to develop an integrated technology for the prediction of aging behavior for space structural materials during service. This research will utilize state-of-the-art radiation experimental apparatus and analysis, updated codes and databases, and integrated mechanical and radiation testing techniques to investigate the suitability of numerous current and potential spacecraft structural materials. Also included are the effects on structural materials in surface modules and planetary landing craft, with or without fission power supplies. Spacecraft structural materials would also be in hostile radiation environments on the surface of the moon and planets without appreciable atmospheres and moons around planets with large intense magnetic and radiation fields (such as the Jovian moons). The effects of extreme temperature cycles in such locations compounds the effects of radiation on structural materials. This paper describes the integrated methodology in detail and shows that it will provide a significant technological advance for designing advanced spacecraft. This methodology will also allow for the development of advanced spacecraft materials through the understanding of the underlying mechanisms of material degradation in the space radiation environment. Thus, this technology holds a promise for revolutionary advances in material damage prediction and protection of space structural components as, for example, in the development of guidelines for managing surveillance programs regarding the integrity of spacecraft components, and the safety of the aging spacecraft. (authors)

Structural Design of Glass and Ceramic Components for Space System Safety

Science.gov (United States)

Bernstein, Karen S.

2007-01-01

Manned space flight programs will always have windows as part of the structural shell of the crew compartment. Astronauts and cosmonauts need to and enjoy looking out of the spacecraft windows at Earth, at approaching vehicles, at scientific objectives and at the stars. With few exceptions spacecraft windows have been made of glass, and the lessons learned over forty years of manned space flight have resulted in a well-defined approach for using this brittle, unforgiving material in NASA's vehicles, in windows and other structural applications. This chapter will outline the best practices that have developed at NASA for designing, verifying and accepting glass (and ceramic) windows and other components for safe and reliable use in any space system.

Worldwide Spacecraft Crew Hatch History

Science.gov (United States)

Johnson, Gary

2009-01-01

The JSC Flight Safety Office has developed this compilation of historical information on spacecraft crew hatches to assist the Safety Tech Authority in the evaluation and analysis of worldwide spacecraft crew hatch design and performance. The document is prepared by SAIC s Gary Johnson, former NASA JSC S&MA Associate Director for Technical. Mr. Johnson s previous experience brings expert knowledge to assess the relevancy of data presented. He has experience with six (6) of the NASA spacecraft programs that are covered in this document: Apollo; Skylab; Apollo Soyuz Test Project (ASTP), Space Shuttle, ISS and the Shuttle/Mir Program. Mr. Johnson is also intimately familiar with the JSC Design and Procedures Standard, JPR 8080.5, having been one of its original developers. The observations and findings are presented first by country and organized within each country section by program in chronological order of emergence. A host of reference sources used to augment the personal observations and comments of the author are named within the text and/or listed in the reference section of this document. Careful attention to the selection and inclusion of photos, drawings and diagrams is used to give visual association and clarity to the topic areas examined.

A Sustainable Spacecraft Component Database Solution, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Numerous spacecraft component databases have been developed to support NASA, DoD, and contractor design centers and design tools. Despite the clear utility of...

Testing of an End-Point Control Unit Designed to Enable Precision Control of Manipulator-Coupled Spacecraft

Science.gov (United States)

Montgomery, Raymond C.; Ghosh, Dave; Tobbe, Patrick A.; Weathers, John M.; Manouchehri, Davoud; Lindsay, Thomas S.

1994-01-01

This paper presents an end-point control concept designed to enable precision telerobotic control of manipulator-coupled spacecraft. The concept employs a hardware unit (end-point control unit EPCU) that is positioned between the end-effector of the Space Shuttle Remote Manipulator System and the payload. Features of the unit are active compliance (control of the displacement between the end-effector and the payload), to allow precision control of payload motions, and inertial load relief, to prevent the transmission of loads between the end-effector and the payload. This paper presents the concept and studies the active compliance feature using a simulation and hardware. Results of the simulation show the effectiveness of the EPCU in smoothing the motion of the payload. Results are presented from initial, limited tests of a laboratory hardware unit on a robotic arm testbed at the l Space Flight Center. Tracking performance of the arm in a constant speed automated retraction and extension maneuver of a heavy payload with and without the unit active is compared for the design speed and higher speeds. Simultaneous load reduction and tracking performance are demonstrated using the EPCU.

High-Fidelity Dynamic Modeling of Spacecraft in the Continuum--Rarefied Transition Regime

Science.gov (United States)

Turansky, Craig P.

The state of the art of spacecraft rarefied aerodynamics seldom accounts for detailed rigid-body dynamics. In part because of computational constraints, simpler models based upon the ballistic and drag coefficients are employed. Of particular interest is the continuum-rarefied transition regime of Earth's thermosphere where gas dynamic simulation is difficult yet wherein many spacecraft operate. The feasibility of increasing the fidelity of modeling spacecraft dynamics is explored by coupling rarefied aerodynamics with rigid-body dynamics modeling similar to that traditionally used for aircraft in atmospheric flight. Presented is a framework of analysis and guiding principles which capitalize on the availability of increasing computational methods and resources. Aerodynamic force inputs for modeling spacecraft in two dimensions in a rarefied flow are provided by analytical equations in the free-molecular regime, and the direct simulation Monte Carlo method in the transition regime. The application of the direct simulation Monte Carlo method to this class of problems is examined in detail with a new code specifically designed for engineering-level rarefied aerodynamic analysis. Time-accurate simulations of two distinct geometries in low thermospheric flight and atmospheric entry are performed, demonstrating non-linear dynamics that cannot be predicted using simpler approaches. The results of this straightforward approach to the aero-orbital coupled-field problem highlight the possibilities for future improvements in drag prediction, control system design, and atmospheric science. Furthermore, a number of challenges for future work are identified in the hope of stimulating the development of a new subfield of spacecraft dynamics.

Spacecraft Cabin Particulate Monitor, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — We propose to design, build and test an optical extinction monitor for the detection of spacecraft cabin particulates. This monitor will be sensitive to particle...

The Effect of Golden Pothos in Reducing the Level of Volatile Organic Compounds in a Simulated Spacecraft Cabin

Science.gov (United States)

Ursprung, Matthew; Amiri, Azita; Kayatin, Matthew; Perry, Jay

2016-01-01

The impact of Golden Pothos on indoor air quality was studied against a simulated spacecraft trace contaminant load model, consistent with the International Space Station (ISS), containing volatile organic compounds (VOCs) and formaldehyde. Previous research provides inconclusive results on the efficacy of plant VOC removal which this projects seeks to rectify through a better experimental design. This work develops a passive system for removing common VOC's from spacecraft and household indoor air and decreasing the necessity for active cabin trace contaminant removal systems.

High-Performance Fire Detector for Spacecraft, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — The danger from fire aboard spacecraft is immediate with only moments for detection and suppression. Spacecraft are unique high-value systems where the cost of...

A Reconfigurable Testbed Environment for Spacecraft Autonomy

Science.gov (United States)

Biesiadecki, Jeffrey; Jain, Abhinandan

1996-01-01

A key goal of NASA's New Millennium Program is the development of technology for increased spacecraft on-board autonomy. Achievement of this objective requires the development of a new class of ground-based automony testbeds that can enable the low-cost and rapid design, test, and integration of the spacecraft autonomy software. This paper describes the development of an Autonomy Testbed Environment (ATBE) for the NMP Deep Space I comet/asteroid rendezvous mission.

Technology for organization of the onboard system for processing and storage of ERS data for ultrasmall spacecraft

Science.gov (United States)

Strotov, Valery V.; Taganov, Alexander I.; Konkin, Yuriy V.; Kolesenkov, Aleksandr N.

2017-10-01

Task of processing and analysis of obtained Earth remote sensing data on ultra-small spacecraft board is actual taking into consideration significant expenditures of energy for data transfer and low productivity of computers. Thereby, there is an issue of effective and reliable storage of the general information flow obtained from onboard systems of information collection, including Earth remote sensing data, into a specialized data base. The paper has considered peculiarities of database management system operation with the multilevel memory structure. For storage of data in data base the format has been developed that describes a data base physical structure which contains required parameters for information loading. Such structure allows reducing a memory size occupied by data base because it is not necessary to store values of keys separately. The paper has shown architecture of the relational database management system oriented into embedment into the onboard ultra-small spacecraft software. Data base for storage of different information, including Earth remote sensing data, can be developed by means of such database management system for its following processing. Suggested database management system architecture has low requirements to power of the computer systems and memory resources on the ultra-small spacecraft board. Data integrity is ensured under input and change of the structured information.

Adaptive nonlinear robust relative pose control of spacecraft autonomous rendezvous and proximity operations.

Science.gov (United States)

Sun, Liang; Huo, Wei; Jiao, Zongxia

2017-03-01

This paper studies relative pose control for a rigid spacecraft with parametric uncertainties approaching to an unknown tumbling target in disturbed space environment. State feedback controllers for relative translation and relative rotation are designed in an adaptive nonlinear robust control framework. The element-wise and norm-wise adaptive laws are utilized to compensate the parametric uncertainties of chaser and target spacecraft, respectively. External disturbances acting on two spacecraft are treated as a lumped and bounded perturbation input for system. To achieve the prescribed disturbance attenuation performance index, feedback gains of controllers are designed by solving linear matrix inequality problems so that lumped disturbance attenuation with respect to the controlled output is ensured in the L 2 -gain sense. Moreover, in the absence of lumped disturbance input, asymptotical convergence of relative pose are proved by using the Lyapunov method. Numerical simulations are performed to show that position tracking and attitude synchronization are accomplished in spite of the presence of couplings and uncertainties. Copyright © 2016 ISA. Published by Elsevier Ltd. All rights reserved.

Study of reactor Brayton power systems for nuclear electric spacecraft

Science.gov (United States)

1979-01-01

The feasibility of using Brayton power systems for nuclear electric spacecraft was investigated. The primary performance parameters of systems mass and radiator area were determined for systems from 100 to 1000 kW sub e. Mathematical models of all system components were used to determine masses and volumes. Two completely independent systems provide propulsion power so that no single-point failure can jeopardize a mission. The waste heat radiators utilize armored heat pipes to limit meteorite puncture. The armor thickness was statistically determined to achieve the required probability of survival. A 400 kW sub e reference system received primary attention as required by the contract. The components of this system were defined and a conceptual layout was developed with encouraging results. An arrangement with redundant Brayton power systems having a 1500 K (2240 F) turbine inlet temperature was shown to be compatible with the dimensions of the space shuttle orbiter payload bay.

Flight Plasma Diagnostics for High-Power, Solar-Electric Deep-Space Spacecraft

Science.gov (United States)

Johnson, Lee; De Soria-Santacruz Pich, Maria; Conroy, David; Lobbia, Robert; Huang, Wensheng; Choi, Maria; Sekerak, Michael J.

2018-01-01

NASA's Asteroid Redirect Robotic Mission (ARRM) project plans included a set of plasma and space environment instruments, the Plasma Diagnostic Package (PDP), to fulfill ARRM requirements for technology extensibility to future missions. The PDP objectives were divided into the classes of 1) Plasma thruster dynamics, 2) Solar array-specific environmental effects, 3) Plasma environmental spacecraft effects, and 4) Energetic particle spacecraft environment. A reference design approach and interface requirements for ARRM's PDP was generated by the PDP team at JPL and GRC. The reference design consisted of redundant single-string avionics located on the ARRM spacecraft bus as well as solar array, driving and processing signals from multiple copies of several types of plasma, effects, and environments sensors distributed over the spacecraft and array. The reference design sensor types were derived in part from sensors previously developed for USAF Research Laboratory (AFRL) plasma effects campaigns such as those aboard TacSat-2 in 2007 and AEHF-2 in 2012.

Addressing EO-1 Spacecraft Pulsed Plasma Thruster EMI Concerns

Science.gov (United States)

Zakrzwski, C. M.; Davis, Mitch; Sarmiento, Charles; Bauer, Frank H. (Technical Monitor)

2001-01-01

The Pulsed Plasma Thruster (PPT) Experiment on the Earth Observing One (EO-1) spacecraft has been designed to demonstrate the capability of a new generation PPT to perform spacecraft attitude control. Results from PPT unit level radiated electromagnetic interference (EMI) tests led to concerns about potential interference problems with other spacecraft subsystems. Initial plans to address these concerns included firing the PPT at the spacecraft level both in atmosphere, with special ground support equipment. and in vacuum. During the spacecraft level tests, additional concerns where raised about potential harm to the Advanced Land Imager (ALI). The inadequacy of standard radiated emission test protocol to address pulsed electromagnetic discharges and the lack of resources required to perform compatibility tests between the PPT and an ALI test unit led to changes in the spacecraft level validation plan. An EMI shield box for the PPT was constructed and validated for spacecraft level ambient testing. Spacecraft level vacuum tests of the PPT were deleted. Implementation of the shield box allowed for successful spacecraft level testing of the PPT while eliminating any risk to the ALI. The ALI demonstration will precede the PPT demonstration to eliminate any possible risk of damage of ALI from PPT operation.

Integrated orbital servicing study follow-on. Volume 2: Technical analysis and system design

Science.gov (United States)

1978-01-01

In-orbit service functional and physical requirements to support both low and high Earth orbit servicing/maintenance operations were defined, an optimum servicing system configuration was developed and mockups and early prototype hardware were fabricated to demonstrate and validate the concepts selected. Significant issues addressed include criteria for concept selection; representative mission equipment and approaches to their design for serviceability; significant serviceable spacecraft design aspects; servicer mechanism operation in one-g; approaches for the demonstration/simulation; and service mechanism structure design approach.

Trajectory Control of Rendezvous with Maneuver Target Spacecraft

Science.gov (United States)

Zhou, Zhinqiang

2012-01-01

In this paper, a nonlinear trajectory control algorithm of rendezvous with maneuvering target spacecraft is presented. The disturbance forces on the chaser and target spacecraft and the thrust forces on the chaser spacecraft are considered in the analysis. The control algorithm developed in this paper uses the relative distance and relative velocity between the target and chaser spacecraft as the inputs. A general formula of reference relative trajectory of the chaser spacecraft to the target spacecraft is developed and applied to four different proximity maneuvers, which are in-track circling, cross-track circling, in-track spiral rendezvous and cross-track spiral rendezvous. The closed-loop differential equations of the proximity relative motion with the control algorithm are derived. It is proven in the paper that the tracking errors between the commanded relative trajectory and the actual relative trajectory are bounded within a constant region determined by the control gains. The prediction of the tracking errors is obtained. Design examples are provided to show the implementation of the control algorithm. The simulation results show that the actual relative trajectory tracks the commanded relative trajectory tightly. The predicted tracking errors match those calculated in the simulation results. The control algorithm developed in this paper can also be applied to interception of maneuver target spacecraft and relative trajectory control of spacecraft formation flying.

Air dehumidification by membrane with cold water for manned spacecraft environmental control

Directory of Open Access Journals (Sweden)

Shang Yonghong

2017-01-01

Full Text Available The traditional condensation dehumidification method requires additional gas-liquid separation and water recovery process in the manned spacecraft humidity control system, which would increase weight and complexity of systems. A new membrane dehumidification with cold water is proposed, which uses water vapor partial pressure difference to promote water vapor transmembrane mass transfer for dehumidification. The permeability of the membrane was measured and the experimental results agree well with the theoretical calculations. Based on the simulation of dehumidification process of cold water-membrane, the influence of module structure and working condition on dehumidification performance was analyzed, which provided reference for the design of membrane module construct. It can be seen from the simulation and experiments that the cold water-membrane dehumidification can effectively reduce the thermal load of the manned spacecraft.

Ascent performance feasibility for next-generation spacecraft

Science.gov (United States)

Mancuso, Salvatore Massimo

This thesis deals with the optimization of the ascent trajectories for single-stage suborbital (SSSO), single-stage-to-orbit (SSTO), and two-stage-to-orbit (TSTO) rocket-powered spacecraft. The maximum payload weight problem has been solved using the sequential gradient-restoration algorithm. For the TSTO case, some modifications to the original version of the algorithm have been necessary in order to deal with discontinuities due to staging and the fact that the functional being minimized depends on interface conditions. The optimization problem is studied for different values of the initial thrust-to-weight ratio in the range 1.3 to 1.6, engine specific impulse in the range 400 to 500 sec, and spacecraft structural factor in the range 0.08 to 0.12. For the TSTO configuration, two subproblems are studied: uniform structural factor between stages and nonuniform structural factor between stages. Due to the regular behavior of the results obtained, engineering approximations have been developed which connect the maximum payload weight to the engine specific impulse and spacecraft structural factor; in turn, this leads to useful design considerations. Also, performance sensitivity to the scale of the aerodynamic drag is studied, and it is shown that its effect on payload weight is relatively small, even for drag changes approaching ± 50%. The main conclusions are that: the design of a SSSO configuration appears to be feasible; the design of a SSTO configuration might be comfortably feasible, marginally feasible, or unfeasible, depending on the parameter values assumed; the design of a TSTO configuration is not only feasible, but its payload appears to be considerably larger than that of a SSTO configuration. Improvements in engine specific impulse and spacecraft structural factor are desirable and crucial for SSTO feasibility; indeed, it appears that aerodynamic improvements do not yield significant improvements in payload weight.

An IP-Based Software System for Real-time, Closed Loop, Multi-Spacecraft Mission Simulations

Science.gov (United States)

Cary, Everett; Davis, George; Higinbotham, John; Burns, Richard; Hogie, Keith; Hallahan, Francis

2003-01-01

This viewgraph presentation provides information on the architecture of a computerized testbest for simulating Distributed Space Systems (DSS) for controlling spacecraft flying in formation. The presentation also discusses and diagrams the Distributed Synthesis Environment (DSE) for simulating and planning DSS missions.

Post-capture vibration suppression of spacecraft via a bio-inspired isolation system

Science.gov (United States)

Dai, Honghua; Jing, Xingjian; Wang, Yu; Yue, Xiaokui; Yuan, Jianping

2018-05-01

Inspired by the smooth motions of a running kangaroo, a bio-inspired quadrilateral shape (BIQS) structure is proposed to suppress the vibrations of a free-floating spacecraft subject to periodic or impulsive forces, which may be encountered during on-orbit servicing missions. In particular, the BIQS structure is installed between the satellite platform and the capture mechanism. The dynamical model of the BIQS isolation system, i.e. a BIQS structure connecting the platform and the capture mechanism at each side, is established by Lagrange's equations to simulate the post-capture dynamical responses. The BIQS system suffering an impulsive force is dealt with by means of a modified version of Lagrange's equations. Furthermore, the classical harmonic balance method is used to solve the nonlinear dynamical system subject to periodic forces, while for the case under impulsive forces the numerical integration method is adopted. Due to the weightless environment in space, the present BIQS system is essentially an under-constrained dynamical system with one of its natural frequencies being identical to zero. The effects of system parameters, such as the number of layers in BIQS, stiffness, assembly angle, rod length, damping coefficient, masses of satellite platform and capture mechanism, on the isolation performance of the present system are thoroughly investigated. In addition, comparisons between the isolation performances of the presently proposed BIQS isolator and the conventional spring-mass-damper (SMD) isolator are conducted to demonstrate the advantages of the present isolator. Numerical simulations show that the BIQS system has a much better performance than the SMD system under either periodic or impulsive forces. Overall, the present BIQS isolator offers a highly efficient passive way for vibration suppressions of free-floating spacecraft.

Heat pipe applications for future Air Force spacecraft

International Nuclear Information System (INIS)

Mahefkey, T.; Barthelemy, R.R.

1980-01-01

This paper summarizes the envisioned, future usage of high and low temperature heat pipes in advanced Air Force spacecraft. Thermal control requirements for a variety of communications, surveillance, and space defense missions are forecast. Thermal design constraints implied by survivability to potential weapons effects are outlined. Applications of heat pipes to meet potential low and high power spacecraft mission requirements and envisioned design constraints are suggested. A brief summary of past Air Force sponsored heat pipe development efforts is presented and directions for future development outlined, including those applicable to advanced photovoltaic and nuclear power subsystem applications of heat pipes

Microgravity Flammability Experiments for Spacecraft Fire Safety

DEFF Research Database (Denmark)

Legros, Guillaume; Minster, Olivier; Tóth, Balazs

2012-01-01

As fire behaviour in manned spacecraft still remains poorly understood, an international topical team has been created to design a validation experiment that has an unprecedented large scale for a microgravity flammability experiment. While the validation experiment is being designed for a re-sup...

Space Launch System Spacecraft and Payload Elements: Progress Toward Crewed Launch and Beyond

Science.gov (United States)

Schorr, Andrew A.; Smith, David Alan; Holcomb, Shawn; Hitt, David

2017-01-01

While significant and substantial progress continues to be accomplished toward readying the Space Launch System (SLS) rocket for its first test flight, work is already underway on preparations for the second flight - using an upgraded version of the vehicle - and beyond. Designed to support human missions into deep space, SLS is the most powerful human-rated launch vehicle the United States has ever undertaken, and is one of three programs being managed by the National Aeronautics and Space Administration's (NASA's) Exploration Systems Development division. The Orion spacecraft program is developing a new crew vehicle that will support human missions beyond low Earth orbit (LEO), and the Ground Systems Development and Operations (GSDO) program is transforming Kennedy Space Center (KSC) into a next-generation spaceport capable of supporting not only SLS but also multiple commercial users. Together, these systems will support human exploration missions into the proving ground of cislunar space and ultimately to Mars. For its first flight, SLS will deliver a near-term heavy-lift capability for the nation with its 70-metric-ton (t) Block 1 configuration. Each element of the vehicle now has flight hardware in production in support of the initial flight of the SLS, which will propel Orion around the moon and back. Encompassing hardware qualification, structural testing to validate hardware compliance and analytical modeling, progress is on track to meet the initial targeted launch date. In Utah and Mississippi, booster and engine testing are verifying upgrades made to proven shuttle hardware. At Michoud Assembly Facility (MAF) in Louisiana, the world's largest spacecraft welding tool is producing tanks for the SLS core stage. Providing the Orion crew capsule/launch vehicle interface and in-space propulsion via a cryogenic upper stage, the Spacecraft/Payload Integration and Evolution (SPIE) element serves a key role in achieving SLS goals and objectives. The SPIE element

Modeling and Analysis of Realistic Fire Scenarios in Spacecraft

Science.gov (United States)

Brooker, J. E.; Dietrich, D. L.; Gokoglu, S. A.; Urban, D. L.; Ruff, G. A.

2015-01-01

An accidental fire inside a spacecraft is an unlikely, but very real emergency situation that can easily have dire consequences. While much has been learned over the past 25+ years of dedicated research on flame behavior in microgravity, a quantitative understanding of the initiation, spread, detection and extinguishment of a realistic fire aboard a spacecraft is lacking. Virtually all combustion experiments in microgravity have been small-scale, by necessity (hardware limitations in ground-based facilities and safety concerns in space-based facilities). Large-scale, realistic fire experiments are unlikely for the foreseeable future (unlike in terrestrial situations). Therefore, NASA will have to rely on scale modeling, extrapolation of small-scale experiments and detailed numerical modeling to provide the data necessary for vehicle and safety system design. This paper presents the results of parallel efforts to better model the initiation, spread, detection and extinguishment of fires aboard spacecraft. The first is a detailed numerical model using the freely available Fire Dynamics Simulator (FDS). FDS is a CFD code that numerically solves a large eddy simulation form of the Navier-Stokes equations. FDS provides a detailed treatment of the smoke and energy transport from a fire. The simulations provide a wealth of information, but are computationally intensive and not suitable for parametric studies where the detailed treatment of the mass and energy transport are unnecessary. The second path extends a model previously documented at ICES meetings that attempted to predict maximum survivable fires aboard space-craft. This one-dimensional model implies the heat and mass transfer as well as toxic species production from a fire. These simplifications result in a code that is faster and more suitable for parametric studies (having already been used to help in the hatch design of the Multi-Purpose Crew Vehicle, MPCV).

Orion: Design of a system for assured low-cost human access to space

Science.gov (United States)

Elvander, Josh; Heifetz, Andy; Hunt, Teresa; Zhu, Martin

1994-01-01

In recent years, Congress and the American people have begun to seriously question the role and importance of future manned spaceflight. This is mainly due to two factors: a decline in technical competition caused by the collapse of communism, and the high costs associated with the Space Shuttle transportation system. With these factors in mind, the ORION system was designed to enable manned spaceflight at a low cost, while maintaining the ability to carry out diverse missions, each with a high degree of flexibility. It is capable of performing satellite servicing missions, supporting a space station via crew rotation and resupply, and delivering satellites into geosynchronous orbit. The components of the system are a primary launch module, an upper stage, and a manned spacecraft capable of dynamic reentry. For satellite servicing and space station resupply missions, the ORION system utilizes three primary modules, an upper stage, and the spacecraft, which is delivered to low earth orbit and used to rendezvous, transfer materials, and make repairs. For launching a geosynchronous satellite, one primary module and an upper stage are used to deliver the satellite, along with an apogee kick motor, into orbit. The system is designed with reusability and modularity in mind in an attempt to lower cost.

The Implementation of Satellite Attitude Control System Software Using Object Oriented Design

Science.gov (United States)

Reid, W. Mark; Hansell, William; Phillips, Tom; Anderson, Mark O.; Drury, Derek

1998-01-01

NASA established the Small Explorer (SNMX) program in 1988 to provide frequent opportunities for highly focused and relatively inexpensive space science missions. The SMEX program has produced five satellites, three of which have been successfully launched. The remaining two spacecraft are scheduled for launch within the coming year. NASA has recently developed a prototype for the next generation Small Explorer spacecraft (SMEX-Lite). This paper describes the object-oriented design (OOD) of the SMEX-Lite Attitude Control System (ACS) software. The SMEX-Lite ACS is three-axis controlled and is capable of performing sub-arc-minute pointing. This paper first describes high level requirements governing the SMEX-Lite ACS software architecture. Next, the context in which the software resides is explained. The paper describes the principles of encapsulation, inheritance, and polymorphism with respect to the implementation of an ACS software system. This paper will also discuss the design of several ACS software components. Specifically, object-oriented designs are presented for sensor data processing, attitude determination, attitude control, and failure detection. Finally, this paper will address the establishment of the ACS Foundation Class (AFC) Library. The AFC is a large software repository, requiring a minimal amount of code modifications to produce ACS software for future projects.

LQG/LTR optimal attitude control of small flexible spacecraft using free-free boundary conditions

Science.gov (United States)

Fulton, Joseph M.

Due to the volume and power limitations of a small satellite, careful consideration must be taken while designing an attitude control system for 3-axis stabilization. Placing redundancy in the system proves difficult and utilizing power hungry, high accuracy, active actuators is not a viable option. Thus, it is customary to find dependable, passive actuators used in conjunction with small scale active control components. This document describes the application of Elastic Memory Composite materials in the construction of a flexible spacecraft appendage, such as a gravity gradient boom. Assumed modes methods are used with Finite Element Modeling information to obtain the equations of motion for the system while assuming free-free boundary conditions. A discussion is provided to illustrate how cantilever mode shapes are not always the best assumption when modeling small flexible spacecraft. A key point of interest is first resonant modes may be needed in the system design plant in spite of these modes being greater than one order of magnitude in frequency when compared to the crossover frequency of the controller. LQG/LTR optimal control techniques are implemented to compute attitude control gains while controller robustness considerations determine appropriate reduced order controllers and which flexible modes to include in the design model. Key satellite designer concerns in the areas of computer processor sizing, material uncertainty impacts on the system model, and system performance variations resulting from appendage length modifications are addressed.

Attitude dynamics and control of a spacecraft using shifting mass distribution

Science.gov (United States)

Ahn, Young Tae

conditions are satisfied. In order to increase the accuracy, generally, more than two control systems are installed in a satellite. Combination with another actuator will be examined to fulfill the full attitude control maneuver. Future work can also include more realistic spacecraft design and operational considerations on the behavior of this type of control system.

Autonomous Spacecraft Communication Interface for Load Planning

Science.gov (United States)

Dever, Timothy P.; May, Ryan D.; Morris, Paul H.

2014-01-01

Ground-based controllers can remain in continuous communication with spacecraft in low Earth orbit (LEO) with near-instantaneous communication speeds. This permits near real-time control of all of the core spacecraft systems by ground personnel. However, as NASA missions move beyond LEO, light-time communication delay issues, such as time lag and low bandwidth, will prohibit this type of operation. As missions become more distant, autonomous control of manned spacecraft will be required. The focus of this paper is the power subsystem. For present missions, controllers on the ground develop a complete schedule of power usage for all spacecraft components. This paper presents work currently underway at NASA to develop an architecture for an autonomous spacecraft, and focuses on the development of communication between the Mission Manager and the Autonomous Power Controller. These two systems must work together in order to plan future load use and respond to unanticipated plan deviations. Using a nominal spacecraft architecture and prototype versions of these two key components, a number of simulations are run under a variety of operational conditions, enabling development of content and format of the messages necessary to achieve the desired goals. The goals include negotiation of a load schedule that meets the global requirements (contained in the Mission Manager) and local power system requirements (contained in the Autonomous Power Controller), and communication of off-plan disturbances that arise while executing a negotiated plan. The message content is developed in two steps: first, a set of rapid-prototyping "paper" simulations are preformed; then the resultant optimized messages are codified for computer communication for use in automated testing.

Environmentally-induced discharge transient coupling to spacecraft

Science.gov (United States)

Viswanathan, R.; Barbay, G.; Stevens, N. J.

1985-01-01

The Hughes SCREENS (Space Craft Response to Environments of Space) technique was applied to generic spin and 3-axis stabilized spacecraft models. It involved the NASCAP modeling for surface charging and lumped element modeling for transients coupling into a spacecraft. A differential voltage between antenna and spun shelf of approx. 400 V and current of 12 A resulted from discharge at antenna for the spinner and approx. 3 kv and 0.3 A from a discharge at solar panels for the 3-axis stabilized Spacecraft. A typical interface circuit response was analyzed to show that the transients would couple into the Spacecraft System through ground points, which are most vulnerable. A compilation and review was performed on 15 years of available data from electron and ion current collection phenomena. Empirical models were developed to match data and compared with flight data of Pix-1 and Pix-2 mission. It was found that large space power systems would float negative and discharge if operated at or above 300 V. Several recommendations are given to improve the models and to apply them to large space systems.

Spacecraft operations

CERN Document Server

Sellmaier, Florian; Schmidhuber, Michael

2015-01-01

The book describes the basic concepts of spaceflight operations, for both, human and unmanned missions. The basic subsystems of a space vehicle are explained in dedicated chapters, the relationship of spacecraft design and the very unique space environment are laid out. Flight dynamics are taught as well as ground segment requirements. Mission operations are divided into preparation including management aspects, execution and planning. Deep space missions and space robotic operations are included as special cases. The book is based on a course held at the German Space Operation Center (GSOC).

Particle Morphology and Elemental Composition of Smoke Generated by Overheating Common Spacecraft Materials

Science.gov (United States)

Meyer, Marit E.

2015-01-01

Fire safety in the indoor spacecraft environment is concerned with a unique set of fuels which are designed to not combust. Unlike terrestrial flaming fires, which often can consume an abundance of wood, paper and cloth, spacecraft fires are expected to be generated from overheating electronics consisting of flame resistant materials. Therefore, NASA prioritizes fire characterization research for these fuels undergoing oxidative pyrolysis in order to improve spacecraft fire detector design. A thermal precipitator designed and built for spacecraft fire safety test campaigns at the NASA White Sands Test Facility (WSTF) successfully collected an abundance of smoke particles from oxidative pyrolysis. A thorough microscopic characterization has been performed for ten types of smoke from common spacecraft materials or mixed materials heated at multiple temperatures using the following techniques: SEM, TEM, high resolution TEM, high resolution STEM and EDS. Resulting smoke particle morphologies and elemental compositions have been observed which are consistent with known thermal decomposition mechanisms in the literature and chemical make-up of the spacecraft fuels. Some conclusions about particle formation mechanisms are explored based on images of the microstructure of Teflon smoke particles and tar ball-like particles from Nomex fabric smoke.

Robust Spacecraft Component Detection in Point Clouds

Directory of Open Access Journals (Sweden)

Quanmao Wei

2018-03-01

Full Text Available Automatic component detection of spacecraft can assist in on-orbit operation and space situational awareness. Spacecraft are generally composed of solar panels and cuboidal or cylindrical modules. These components can be simply represented by geometric primitives like plane, cuboid and cylinder. Based on this prior, we propose a robust automatic detection scheme to automatically detect such basic components of spacecraft in three-dimensional (3D point clouds. In the proposed scheme, cylinders are first detected in the iteration of the energy-based geometric model fitting and cylinder parameter estimation. Then, planes are detected by Hough transform and further described as bounded patches with their minimum bounding rectangles. Finally, the cuboids are detected with pair-wise geometry relations from the detected patches. After successive detection of cylinders, planar patches and cuboids, a mid-level geometry representation of the spacecraft can be delivered. We tested the proposed component detection scheme on spacecraft 3D point clouds synthesized by computer-aided design (CAD models and those recovered by image-based reconstruction, respectively. Experimental results illustrate that the proposed scheme can detect the basic geometric components effectively and has fine robustness against noise and point distribution density.

Robust Spacecraft Component Detection in Point Clouds.

Science.gov (United States)

Wei, Quanmao; Jiang, Zhiguo; Zhang, Haopeng

2018-03-21

Automatic component detection of spacecraft can assist in on-orbit operation and space situational awareness. Spacecraft are generally composed of solar panels and cuboidal or cylindrical modules. These components can be simply represented by geometric primitives like plane, cuboid and cylinder. Based on this prior, we propose a robust automatic detection scheme to automatically detect such basic components of spacecraft in three-dimensional (3D) point clouds. In the proposed scheme, cylinders are first detected in the iteration of the energy-based geometric model fitting and cylinder parameter estimation. Then, planes are detected by Hough transform and further described as bounded patches with their minimum bounding rectangles. Finally, the cuboids are detected with pair-wise geometry relations from the detected patches. After successive detection of cylinders, planar patches and cuboids, a mid-level geometry representation of the spacecraft can be delivered. We tested the proposed component detection scheme on spacecraft 3D point clouds synthesized by computer-aided design (CAD) models and those recovered by image-based reconstruction, respectively. Experimental results illustrate that the proposed scheme can detect the basic geometric components effectively and has fine robustness against noise and point distribution density.

Metal hydride hydrogen and heat storage systems as enabling technology for spacecraft applications

Energy Technology Data Exchange (ETDEWEB)

Reissner, Alexander, E-mail: reissner@fotec.at [FOTEC Forschungs- und Technologietransfer GmbH, Viktor Kaplan Straße 2, 2700 Wiener Neustadt (Austria); University of Applied Sciences Wiener Neustadt, Johannes Gutenberg-Straße 3, 2700 Wiener Neustadt (Austria); Pawelke, Roland H.; Hummel, Stefan; Cabelka, Dusan [FOTEC Forschungs- und Technologietransfer GmbH, Viktor Kaplan Straße 2, 2700 Wiener Neustadt (Austria); Gerger, Joachim [University of Applied Sciences Wiener Neustadt, Johannes Gutenberg-Straße 3, 2700 Wiener Neustadt (Austria); Farnes, Jarle, E-mail: Jarle.farnes@prototech.no [CMR Prototech AS, Fantoftvegen 38, PO Box 6034, 5892 Bergen (Norway); Vik, Arild; Wernhus, Ivar; Svendsen, Tjalve [CMR Prototech AS, Fantoftvegen 38, PO Box 6034, 5892 Bergen (Norway); Schautz, Max, E-mail: max.schautz@esa.int [European Space Agency, ESTEC – Keplerlaan 1, 2201 AZ Noordwijk Zh (Netherlands); Geneste, Xavier, E-mail: xavier.geneste@esa.int [European Space Agency, ESTEC – Keplerlaan 1, 2201 AZ Noordwijk Zh (Netherlands)

2015-10-05

Highlights: • A metal hydride tank concept for heat and hydrogen storage is presented. • The tank is part of a closed-loop reversible fuel cell system for space application. • For several engineering issues specific to the spacecraft application, solutions have been developed. • The effect of water contamination has been approximated for Ti-doped NaAlH{sub 4}. • A novel heat exchanger design has been realized by Selective Laser Melting. - Abstract: The next generation of telecommunication satellites will demand a platform payload performance in the range of 30+ kW within the next 10 years. At this high power output, a Regenerative Fuel Cell Systems (RFCS) offers an efficiency advantage in specific energy density over lithium ion batteries. However, a RFCS creates a substantial amount of heat (60–70 kJ per mol H{sub 2}) during fuel cell operation. This requires a thermal hardware that accounts for up to 50% of RFCS mass budget. Thus the initial advantage in specific energy density is reduced. A metal hydride tank for combined storage of heat and hydrogen in a RFCS may overcome this constraint. Being part of a consortium in an ongoing European Space Agency project, FOTEC is building a technology demonstrator for such a combined hydrogen and heat storage system.

Design of a Thermal Precipitator for the Characterization of Smoke Particles from Common Spacecraft Materials

Science.gov (United States)

Meyer, Marit Elisabeth

2015-01-01

A thermal precipitator (TP) was designed to collect smoke aerosol particles for microscopic analysis in fire characterization research. Information on particle morphology, size and agglomerate structure obtained from these tests supplements additional aerosol data collected. Modeling of the thermal precipitator throughout the design process was performed with the COMSOL Multiphysics finite element software package, including the Eulerian flow field and thermal gradients in the fluid. The COMSOL Particle Tracing Module was subsequently used to determine particle deposition. Modeling provided optimized design parameters such as geometry, flow rate and temperatures. The thermal precipitator was built and testing verified the performance of the first iteration of the device. The thermal precipitator was successfully operated and provided quality particle samples for microscopic analysis, which furthered the body of knowledge on smoke particulates. This information is a key element of smoke characterization and will be useful for future spacecraft fire detection research.

Multi-Objective Optimization for Solid Amine CO2 Removal Assembly in Manned Spacecraft

Directory of Open Access Journals (Sweden)

Rong A

2017-07-01

Full Text Available Carbon Dioxide Removal Assembly (CDRA is one of the most important systems in the Environmental Control and Life Support System (ECLSS for a manned spacecraft. With the development of adsorbent and CDRA technology, solid amine is increasingly paid attention due to its obvious advantages. However, a manned spacecraft is launched far from the Earth, and its resources and energy are restricted seriously. These limitations increase the design difficulty of solid amine CDRA. The purpose of this paper is to seek optimal design parameters for the solid amine CDRA. Based on a preliminary structure of solid amine CDRA, its heat and mass transfer models are built to reflect some features of the special solid amine adsorbent, Polyethylenepolyamine adsorbent. A multi-objective optimization for the design of solid amine CDRA is discussed further in this paper. In this study, the cabin CO2 concentration, system power consumption and entropy production are chosen as the optimization objectives. The optimization variables consist of adsorption cycle time, solid amine loading mass, adsorption bed length, power consumption and system entropy production. The Improved Non-dominated Sorting Genetic Algorithm (NSGA-II is used to solve this multi-objective optimization and to obtain optimal solution set. A design example of solid amine CDRA in a manned space station is used to show the optimal procedure. The optimal combinations of design parameters can be located on the Pareto Optimal Front (POF. Finally, Design 971 is selected as the best combination of design parameters. The optimal results indicate that the multi-objective optimization plays a significant role in the design of solid amine CDRA. The final optimal design parameters for the solid amine CDRA can guarantee the cabin CO2 concentration within the specified range, and also satisfy the requirements of lightweight and minimum energy consumption.

Probing interferometric parallax with interplanetary spacecraft

Science.gov (United States)

Rodeghiero, G.; Gini, F.; Marchili, N.; Jain, P.; Ralston, J. P.; Dallacasa, D.; Naletto, G.; Possenti, A.; Barbieri, C.; Franceschini, A.; Zampieri, L.

2017-07-01

We describe an experimental scenario for testing a novel method to measure distance and proper motion of astronomical sources. The method is based on multi-epoch observations of amplitude or intensity correlations between separate receiving systems. This technique is called Interferometric Parallax, and efficiently exploits phase information that has traditionally been overlooked. The test case we discuss combines amplitude correlations of signals from deep space interplanetary spacecraft with those from distant galactic and extragalactic radio sources with the goal of estimating the interplanetary spacecraft distance. Interferometric parallax relies on the detection of wavefront curvature effects in signals collected by pairs of separate receiving systems. The method shows promising potentialities over current techniques when the target is unresolved from the background reference sources. Developments in this field might lead to the construction of an independent, geometrical cosmic distance ladder using a dedicated project and future generation instruments. We present a conceptual overview supported by numerical estimates of its performances applied to a spacecraft orbiting the Solar System. Simulations support the feasibility of measurements with a simple and time-saving observational scheme using current facilities.

An Application of the "Virtual Spacecraft" Concept in Evaluation of the Mars Pathfinder Lander Low Gain Antenna

Science.gov (United States)

Pogorzelski, R. J.; Beckon, R. J.

1997-01-01

The virtual spacecraft concept is embodied in a set of subsystems, either in the form of hardware or computational models, which together represent all, or a portion of, a spacecraft. For example, the telecommunications transponder may be a hardware prototype while the propulsion system may exist only as a simulation. As the various subsystems are realized in hardware, the spacecraft becomes progressively less virtual. This concept is enabled by JPL's Mission System Testbed which is a set of networked workstations running a message passing operating system called "TRAMEL" which stands for Task Remote Asynchronous Message Exchange Layer. Each simulation on the workstations, which may in fact be hardware controlled by the workstation, "publishes" its operating parameters on TRAMEL and other simulations requiring those parameters as input may "subscribe" to them. In this manner, the whole simulation operates as a single virtual system. This paper describes a simulation designed to evaluate a communications link between the earth and the Mars Pathfinder Lander module as it descends under a parachute through the Martian atmosphere toward the planet's surface. This link includes a transmitter and a low gain antenna on the spacecraft and a receiving antenna and receiver on the earth as well as a simulation of the dynamics of the spacecraft. The transmitter, the ground station antenna, the receiver and the dynamics are all simulated computationally while the spacecraft antenna is implemented in hardware on a very simple spacecraft mockup. The dynamics simulation is a record of one output of the ensemble of outputs of a Monte Carlo simulation of the descent. Additionally, the antenna/spacecraft mock-up system was simulated using APATCH, a shooting and bouncing ray code developed by Demaco, Inc. The antenna simulation, the antenna hardware, and the link simulation are all physically located in different facilities at JPL separated by several hundred meters and are linked via

Artist concept of Galileo spacecraft

Science.gov (United States)

1988-01-01

Galileo spacecraft is illustrated in artist concept. Gallileo, named for the Italian astronomer, physicist and mathematician who is credited with construction of the first complete, practical telescope in 1620, will make detailed studies of Jupiter. A cooperative program with the Federal Republic of Germany the Galileo mission will amplify information acquired by two Voyager spacecraft in their brief flybys. Galileo is a two-element system that includes a Jupiter-orbiting observatory and an entry probe. Jet Propulsion Laboratory (JPL) is Galileo project manager and builder of the main spacecraft. Ames Research Center (ARC) has responsibility for the entry probe, which was built by Hughes Aircraft Company and General Electric. Galileo will be deployed from the payload bay (PLB) of Atlantis, Orbiter Vehicle (OV) 104, during mission STS-34.

Systems aspects of a space nuclear reactor power system

International Nuclear Information System (INIS)

Jaffe, L.; Fujita, T.; Beatty, R.

1988-01-01

Selected systems aspects of a 300 kW nuclear reactor power system for spacecraft have been studied. The approach included examination of two candidate missions and their associated spacecraft, and a number of special topics dealing with the power system design and operation. The missions considered were a reusable orbital transfer vehicle and a space-based radar. The special topics included: Power system configuration and scaling, launch vehicle integration, operating altitude, orbital storage, start-up, thawing, control, load following, procedures in case of malfunction, restart, thermal and nuclear radiation to other portions of the spacecraft, thermal stresses between subsystems, boom and cable designs, vibration modes, attitude control, reliability, and survivability. Among the findings are that the stowed length of the power system is important to mission design and that orbital storage for months to years may be needed for missions involving orbital assembly

Hybrid spacecraft attitude control system

OpenAIRE

Renuganth Varatharajoo; Ramly Ajir; Tamizi Ahmad

2016-01-01

The hybrid subsystem design could be an attractive approach for futurespacecraft to cope with their demands. The idea of combining theconventional Attitude Control System and the Electrical Power System ispresented in this article. The Combined Energy and Attitude ControlSystem (CEACS) consisting of a double counter rotating flywheel assemblyis investigated for small satellites in this article. Another hybrid systemincorporating the conventional Attitude Control System into the ThermalControl...

Intelligent Data Visualization for Cross-Checking Spacecraft System Diagnosis

Science.gov (United States)

Ong, James C.; Remolina, Emilio; Breeden, David; Stroozas, Brett A.; Mohammed, John L.

2012-01-01

Any reasoning system is fallible, so crew members and flight controllers must be able to cross-check automated diagnoses of spacecraft or habitat problems by considering alternate diagnoses and analyzing related evidence. Cross-checking improves diagnostic accuracy because people can apply information processing heuristics, pattern recognition techniques, and reasoning methods that the automated diagnostic system may not possess. Over time, cross-checking also enables crew members to become comfortable with how the diagnostic reasoning system performs, so the system can earn the crew s trust. We developed intelligent data visualization software that helps users cross-check automated diagnoses of system faults more effectively. The user interface displays scrollable arrays of timelines and time-series graphs, which are tightly integrated with an interactive, color-coded system schematic to show important spatial-temporal data patterns. Signal processing and rule-based diagnostic reasoning automatically identify alternate hypotheses and data patterns that support or rebut the original and alternate diagnoses. A color-coded matrix display summarizes the supporting or rebutting evidence for each diagnosis, and a drill-down capability enables crew members to quickly view graphs and timelines of the underlying data. This system demonstrates that modest amounts of diagnostic reasoning, combined with interactive, information-dense data visualizations, can accelerate system diagnosis and cross-checking.

Spacecraft attitude and velocity control system

Science.gov (United States)

Paluszek, Michael A. (Inventor); Piper, Jr., George E. (Inventor)

1992-01-01

A spacecraft attitude and/or velocity control system includes a controller which responds to at least attitude errors to produce command signals representing a force vector F and a torque vector T, each having three orthogonal components, which represent the forces and torques which are to be generated by the thrusters. The thrusters may include magnetic torquer or reaction wheels. Six difference equations are generated, three having the form ##EQU1## where a.sub.j is the maximum torque which the j.sup.th thruster can produce, b.sub.j is the maximum force which the j.sup.th thruster can produce, and .alpha..sub.j is a variable representing the throttling factor of the j.sup.th thruster, which may range from zero to unity. The six equations are summed to produce a single scalar equation relating variables .alpha..sub.j to a performance index Z: ##EQU2## Those values of .alpha. which maximize the value of Z are determined by a method for solving linear equations, such as a linear programming method. The Simplex method may be used. The values of .alpha..sub.j are applied to control the corresponding thrusters.

Advanced engineering software for in-space assembly and manned planetary spacecraft

Science.gov (United States)

Delaquil, Donald; Mah, Robert

1990-01-01

Meeting the objectives of the Lunar/Mars initiative to establish safe and cost-effective extraterrestrial bases requires an integrated software/hardware approach to operational definitions and systems implementation. This paper begins this process by taking a 'software-first' approach to systems design, for implementing specific mission scenarios in the domains of in-space assembly and operations of the manned Mars spacecraft. The technological barriers facing implementation of robust operational systems within these two domains are discussed, and preliminary software requirements and architectures that resolve these barriers are provided.

A Comparison of Learning Technologies for Teaching Spacecraft Software Development

Science.gov (United States)

Straub, Jeremy

2014-01-01

The development of software for spacecraft represents a particular challenge and is, in many ways, a worst case scenario from a design perspective. Spacecraft software must be "bulletproof" and operate for extended periods of time without user intervention. If the software fails, it cannot be manually serviced. Software failure may…

Applications of different design methodologies in navigation systems and development at JPL

Science.gov (United States)

Thurman, S. W.

1990-01-01

The NASA/JPL deep space navigation system consists of a complex array of measurement systems, data processing systems, and support facilities, with components located both on the ground and on-board interplanetary spacecraft. From its beginings nearly 30 years ago, this system has steadily evolved and grown to meet the demands for ever-increasing navigation accuracy placed on it by a succession of unmanned planetary missions. Principal characteristics of this system are its capabilities and great complexity. Three examples in the design and development of interplanetary space navigation systems are examined in order to make a brief assessment of the usefulness of three basic design theories, known as normative, rational, and heuristic. Evaluation of the examples indicates that a heuristic approach, coupled with rational-based mathematical and computational analysis methods, is used most often in problems such as orbit determination strategy development and mission navigation system design, while normative methods have seen only limited use is such applications as the development of large software systems and in the design of certain operational navigation subsystems.

Spacecraft attitude maneuver control using two parallel mounted 3-DOF spherical actuators

Directory of Open Access Journals (Sweden)

Guidan Li

2017-02-01

Full Text Available A parallel configuration using two 3-degree-of-freedom (3-DOF spherical electromagnetic momentum exchange actuators is investigated for large angle spacecraft attitude maneuvers. First, the full dynamic equations of motion for the spacecraft system are derived by the Newton-Euler method. To facilitate computation, virtual gimbal coordinate frames are established. Second, a nonlinear control law in terms of quaternions is developed via backstepping method. The proposed control law compensates the coupling torques arising from the spacecraft rotation, and is robust against the external disturbances. Then, the singularity problem is analyzed. To avoid singularities, a modified weighed Moore-Pseudo inverse velocity steering law based on null motion is proposed. The weighted matrices are carefully designed to switch the actuators and redistribute the control torques. The null motion is used to reorient the rotor away from the tilt angle saturation state. Finally, numerical simulations of rest-to-rest maneuvers are performed to validate the effectiveness of the proposed method.

Spacecraft Charging and the Microwave Anisotropy Probe Spacecraft

Science.gov (United States)

Timothy, VanSant J.; Neergaard, Linda F.

1998-01-01

The Microwave Anisotropy Probe (MAP), a MIDEX mission built in partnership between Princeton University and the NASA Goddard Space Flight Center (GSFC), will study the cosmic microwave background. It will be inserted into a highly elliptical earth orbit for several weeks and then use a lunar gravity assist to orbit around the second Lagrangian point (L2), 1.5 million kilometers, anti-sunward from the earth. The charging environment for the phasing loops and at L2 was evaluated. There is a limited set of data for L2; the GEOTAIL spacecraft measured relatively low spacecraft potentials (approx. 50 V maximum) near L2. The main area of concern for charging on the MAP spacecraft is the well-established threat posed by the "geosynchronous region" between 6-10 Re. The launch in the autumn of 2000 will coincide with the falling of the solar maximum, a period when the likelihood of a substorm is higher than usual. The likelihood of a substorm at that time has been roughly estimated to be on the order of 20% for a typical MAP mission profile. Because of the possibility of spacecraft charging, a requirement for conductive spacecraft surfaces was established early in the program. Subsequent NASCAP/GEO analyses for the MAP spacecraft demonstrated that a significant portion of the sunlit surface (solar cell cover glass and sunshade) could have nonconductive surfaces without significantly raising differential charging. The need for conductive materials on surfaces continually in eclipse has also been reinforced by NASCAP analyses.

Manifold dynamics in the Earth-Moon system via isomorphic mapping with application to spacecraft end-of-life strategies

Science.gov (United States)

Pontani, Mauro; Giancotti, Marco; Teofilatto, Paolo

2014-12-01

Recently, manifold dynamics has assumed an increasing relevance for analysis and design of low-energy missions, both in the Earth-Moon system and in alternative multibody environments. With regard to lunar missions, exterior and interior transfers, based on the transit through the regions where the collinear libration points L1 and L2 are located, have been studied for a long time and some space missions have already taken advantage of the results of these studies. This paper is focused on the definition and use of a special isomorphic mapping for low-energy mission analysis. A convenient set of cylindrical coordinates is employed to describe the spacecraft dynamics (i.e. position and velocity), in the context of the circular restricted three-body problem, used to model the spacecraft motion in the Earth-Moon system. This isomorphic mapping of trajectories allows the identification and intuitive representation of periodic orbits and of the related invariant manifolds, which correspond to tubes that emanate from the curve associated with the periodic orbit. Heteroclinic connections, i.e. the trajectories that belong to both the stable and the unstable manifolds of two distinct periodic orbits, can be easily detected by means of this representation. This paper illustrates the use of isomorphic mapping for finding (a) periodic orbits, (b) heteroclinic connections between trajectories emanating from two Lyapunov orbits, the first at L1, and the second at L2, and (c) heteroclinic connections between trajectories emanating from the Lyapunov orbit at L1 and from a particular unstable lunar orbit. Heteroclinic trajectories are asymptotic trajectories that travels at zero-propellant cost. In practical situations, a modest delta-v budget is required to perform transfers along the manifolds. This circumstance implies the possibility of performing complex missions, by combining different types of trajectory arcs belonging to the manifolds. This work studies also the possible

Antimicrobial Materials for Advanced Microbial Control in Spacecraft Water Systems

Science.gov (United States)

Birmele, Michele; Caro, Janicce; Newsham, Gerard; Roberts, Michael; Morford, Megan; Wheeler, Ray

2012-01-01

Microbial detection, identification, and control are essential for the maintenance and preservation of spacecraft water systems. Requirements set by NASA put limitations on the energy, mass, materials, noise, cost, and crew time that can be devoted to microbial control. Efforts are being made to attain real-time detection and identification of microbial contamination in microgravity environments. Research for evaluating technologies for capability enhancement on-orbit is currently focused on the use of adenosine triphosphate (ATP) analysis for detection purposes and polymerase chain reaction (peR) for microbial identification. Additional research is being conducted on how to control for microbial contamination on a continual basis. Existing microbial control methods in spacecraft utilize iodine or ionic silver biocides, physical disinfection, and point-of-use sterilization filters. Although these methods are effective, they require re-dosing due to loss of efficacy, have low human toxicity thresholds, produce poor taste, and consume valuable mass and crew time. Thus, alternative methods for microbial control are needed. This project also explores ultraviolet light-emitting diodes (UV-LEDs), surface passivation methods for maintaining residual biocide levels, and several antimicrobial materials aimed at improving current microbial control techniques, as well as addressing other materials presently under analysis and future directions to be pursued.

Comprehensive Fault Tolerance and Science-Optimal Attitude Planning for Spacecraft Applications

Science.gov (United States)

Nasir, Ali

Spacecraft operate in a harsh environment, are costly to launch, and experience unavoidable communication delay and bandwidth constraints. These factors motivate the need for effective onboard mission and fault management. This dissertation presents an integrated framework to optimize science goal achievement while identifying and managing encountered faults. Goal-related tasks are defined by pointing the spacecraft instrumentation toward distant targets of scientific interest. The relative value of science data collection is traded with risk of failures to determine an optimal policy for mission execution. Our major innovation in fault detection and reconfiguration is to incorporate fault information obtained from two types of spacecraft models: one based on the dynamics of the spacecraft and the second based on the internal composition of the spacecraft. For fault reconfiguration, we consider possible changes in both dynamics-based control law configuration and the composition-based switching configuration. We formulate our problem as a stochastic sequential decision problem or Markov Decision Process (MDP). To avoid the computational complexity involved in a fully-integrated MDP, we decompose our problem into multiple MDPs. These MDPs include planning MDPs for different fault scenarios, a fault detection MDP based on a logic-based model of spacecraft component and system functionality, an MDP for resolving conflicts between fault information from the logic-based model and the dynamics-based spacecraft models" and the reconfiguration MDP that generates a policy optimized over the relative importance of the mission objectives versus spacecraft safety. Approximate Dynamic Programming (ADP) methods for the decomposition of the planning and fault detection MDPs are applied. To show the performance of the MDP-based frameworks and ADP methods, a suite of spacecraft attitude planning case studies are described. These case studies are used to analyze the content and

An Evaluation of Ultra-High Pressure Regulator for Robotic Lunar Landing Spacecraft

Science.gov (United States)

Burnside, Christopher; Trinh, Huu; Pedersen, Kevin

2011-01-01

The Robotic Lunar Lander Development (RLLD) Project Office at NASA Marshall Space Flight Center (MSFC) has studied several lunar surface science mission concepts. These missions focus on spacecraft carrying multiple science instruments and power systems that will allow extended operations on the lunar surface. Initial trade studies of launch vehicle options for these mission concepts indicate that the spacecraft design will be significantly mass-constrained. To minimize mass and facilitate efficient packaging, the notional propulsion system for these landers has a baseline of an ultra-high pressure (10,000 psig) helium pressurization system that has been used on Defense missiles. The qualified regulator is capable of short duration use; however, the hardware has not been previously tested at NASA spacecraft requirements with longer duration. Hence, technical risks exist in using this missile-based propulsion component for spacecraft applications. A 10,000-psig helium pressure regulator test activity is being carried out as part of risk reduction testing for MSFC RLLD project. The goal of the test activity is to assess the feasibility of commercial off-the-shelf ultra-high pressure regulator by testing with a representative flight mission profile. Slam-start, gas blowdown, water expulsion, lock-up, and leak tests are also performed on the regulator to assess performance under various operating conditions. The preliminary test results indicated that the regulator can regulate helium to a stable outlet pressure of 740 psig within the +/- 5% tolerance band and maintain a lock-up pressure less than +5% for all tests conducted. Numerous leak tests demonstrated leakage less than 10-3 standard cubic centimeters per second (SCCS) for internal seat leakage at lock-up and less than10-5 SCCS for external leakage through the regulator ambient reference cavity. The successful tests have shown the potential for 10,000 psig helium systems in NASA spacecraft and have reduced risk

VIP: A knowledge-based design aid for the engineering of space systems

Science.gov (United States)

Lewis, Steven M.; Bellman, Kirstie L.

1990-01-01

The Vehicles Implementation Project (VIP), a knowledge-based design aid for the engineering of space systems is described. VIP combines qualitative knowledge in the form of rules, quantitative knowledge in the form of equations, and other mathematical modeling tools. The system allows users rapidly to develop and experiment with models of spacecraft system designs. As information becomes available to the system, appropriate equations are solved symbolically and the results are displayed. Users may browse through the system, observing dependencies and the effects of altering specific parameters. The system can also suggest approaches to the derivation of specific parameter values. In addition to providing a tool for the development of specific designs, VIP aims at increasing the user's understanding of the design process. Users may rapidly examine the sensitivity of a given parameter to others in the system and perform tradeoffs or optimizations of specific parameters. A second major goal of VIP is to integrate the existing corporate knowledge base of models and rules into a central, symbolic form.

Guidance and Navigation for Rendezvous and Proximity Operations with a Non-Cooperative Spacecraft at Geosynchronous Orbit

Science.gov (United States)

Barbee, Brent William; Carpenter, J. Russell; Heatwole, Scott; Markley, F. Landis; Moreau, Michael; Naasz, Bo J.; VanEepoel, John

2010-01-01

The feasibility and benefits of various spacecraft servicing concepts are currently being assessed, and all require that the servicer spacecraft perform rendezvous, proximity, and capture operations with the target spacecraft to be serviced. Many high-value spacecraft, which would be logical targets for servicing from an economic point of view, are located in geosynchronous orbit, a regime in which autonomous rendezvous and capture operations are not commonplace. Furthermore, existing GEO spacecraft were not designed to be serviced. Most do not have cooperative relative navigation sensors or docking features, and some servicing applications, such as de-orbiting of a non-functional spacecraft, entail rendezvous and capture with a spacecraft that may be non-functional or un-controlled. Several of these challenges have been explored via the design of a notional mission in which a nonfunctional satellite in geosynchronous orbit is captured by a servicer spacecraft and boosted into super-synchronous orbit for safe disposal. A strategy for autonomous rendezvous, proximity operations, and capture is developed, and the Orbit Determination Toolbox (ODTBX) is used to perform a relative navigation simulation to assess the feasibility of performing the rendezvous using a combination of angles-only and range measurements. Additionally, a method for designing efficient orbital rendezvous sequences for multiple target spacecraft is utilized to examine the capabilities of a servicer spacecraft to service multiple targets during the course of a single mission.

DET/MPS - THE GSFC ENERGY BALANCE PROGRAM, DIRECT ENERGY TRANSFER/MULTIMISSION SPACECRAFT MODULAR POWER SYSTEM (UNIX VERSION)

Science.gov (United States)

Jagielski, J. M.

1994-01-01

The DET/MPS programs model and simulate the Direct Energy Transfer and Multimission Spacecraft Modular Power System in order to aid both in design and in analysis of orbital energy balance. Typically, the DET power system has the solar array directly to the spacecraft bus, and the central building block of MPS is the Standard Power Regulator Unit. DET/MPS allows a minute-by-minute simulation of the power system's performance as it responds to various orbital parameters, focusing its output on solar array output and battery characteristics. While this package is limited in terms of orbital mechanics, it is sufficient to calculate eclipse and solar array data for circular or non-circular orbits. DET/MPS can be adjusted to run one or sequential orbits up to about one week, simulated time. These programs have been used on a variety of Goddard Space Flight Center spacecraft projects. DET/MPS is written in FORTRAN 77 with some VAX-type extensions. Any FORTRAN 77 compiler that includes VAX extensions should be able to compile and run the program with little or no modifications. The compiler must at least support free-form (or tab-delineated) source format and 'do do-while end-do' control structures. DET/MPS is available for three platforms: GSC-13374, for DEC VAX series computers running VMS, is available in DEC VAX Backup format on a 9-track 1600 BPI tape (standard distribution) or TK50 tape cartridge; GSC-13443, for UNIX-based computers, is available on a .25 inch streaming magnetic tape cartridge in UNIX tar format; and GSC-13444, for Macintosh computers running AU/X with either the NKR FORTRAN or AbSoft MacFORTRAN II compilers, is available on a 3.5 inch 800K Macintosh format diskette. Source code and test data are supplied. The UNIX version of DET requires 90K of main memory for execution. DET/MPS was developed in 1990. A/UX and Macintosh are registered trademarks of Apple Computer, Inc. VMS, DEC VAX and TK50 are trademarks of Digital Equipment Corporation. UNIX is a

Distributed Autonomous Control of Multiple Spacecraft During Close Proximity Operations

National Research Council Canada - National Science Library

McCamish, Shawn B

2007-01-01

This research contributes to multiple spacecraft control by developing an autonomous distributed control algorithm for close proximity operations of multiple spacecraft systems, including rendezvous...

Spacecraft Charge Monitor

Science.gov (United States)

Goembel, L.

2003-12-01

We are currently developing a flight prototype Spacecraft Charge Monitor (SCM) with support from NASA's Small Business Innovation Research (SBIR) program. The device will use a recently proposed high energy-resolution electron spectroscopic technique to determine spacecraft floating potential. The inspiration for the technique came from data collected by the Atmosphere Explorer (AE) satellites in the 1970s. The data available from the AE satellites indicate that the SCM may be able to determine spacecraft floating potential to within 0.1 V under certain conditions. Such accurate measurement of spacecraft charge could be used to correct biases in space plasma measurements. The device may also be able to measure spacecraft floating potential in the solar wind and in orbit around other planets.

The Heliopause Electrostatic Rapid Transit System (HERTS) - Design, Trades, and Analyses Performed in a Two Year NASA Investigation of Electric Sail Propulsion Systems

Science.gov (United States)

Wiegmann, Bruce M.; Scheider, Todd; Heaton, Andrew; Vaughn, Jason; Stone, Nobie; Wright, Ken

2017-01-01

Personnel from NASA's MSFC have been investigating the feasibility of an advanced propulsion system known as the Electric Sail (E-Sail) for future scientific exploration missions. This team initially won a NASA Space Technology Mission Directorate (STMD) Phase I NASA Innovative Advanced Concept (NIAC) award and then a two-year follow-on Phase II NIAC award in October 2015. This paper documents the findings from this three-year investigation. An Electric sail, a propellant-less propulsion system, uses solar wind ions to rapidly travel either to deep space or the inner solar system. Scientific spacecraft could reach Pluto in 5 years, or the boundary of the solar system in ten to twelve years compared to the thirty-five plus years the Voyager spacecraft took. The team's recent focuses have been: 1) Developing a Particle in Cell (PIC) numeric engineering model from MSFC's experimental data on the interaction between simulated solar wind and a charged bare wire that can be applied to a variety of missions, 2) Determining what missions could benefit from this revolutionary propulsion system, 3) Conceptualizing spacecraft designs for various tasks: to reach the solar system's edge, to orbit the sun as Heliophysics sentinels, or to examine a multitude of asteroids.

Practical Applications of Cosmic Ray Science: Spacecraft, Aircraft, Ground-Based Computation and Control Systems, and Human Health and Safety

Science.gov (United States)

Atwell, William; Koontz, Steve; Normand, Eugene

2012-01-01

Three twentieth century technological developments, 1) high altitude commercial and military aircraft; 2) manned and unmanned spacecraft; and 3) increasingly complex and sensitive solid state micro-electronics systems, have driven an ongoing evolution of basic cosmic ray science into a set of practical engineering tools needed to design, test, and verify the safety and reliability of modern complex technological systems. The effects of primary cosmic ray particles and secondary particle showers produced by nuclear reactions with the atmosphere, can determine the design and verification processes (as well as the total dollar cost) for manned and unmanned spacecraft avionics systems. Similar considerations apply to commercial and military aircraft operating at high latitudes and altitudes near the atmospheric Pfotzer maximum. Even ground based computational and controls systems can be negatively affected by secondary particle showers at the Earth s surface, especially if the net target area of the sensitive electronic system components is large. Finally, accumulation of both primary cosmic ray and secondary cosmic ray induced particle shower radiation dose is an important health and safety consideration for commercial or military air crews operating at high altitude/latitude and is also one of the most important factors presently limiting manned space flight operations beyond low-Earth orbit (LEO). In this paper we review the discovery of cosmic ray effects on the performance and reliability of microelectronic systems as well as human health and the development of the engineering and health science tools used to evaluate and mitigate cosmic ray effects in ground-based atmospheric flight, and space flight environments. Ground test methods applied to microelectronic components and systems are used in combinations with radiation transport and reaction codes to predict the performance of microelectronic systems in their operating environments. Similar radiation transport

Iodine Plasma (Electric Propulsion) Interaction with Spacecraft Materials

Science.gov (United States)

2016-12-28

Teflon (AGT5, Ag-FEP) Thermal control surface (radiator) Spacecraft Exposure Soda-lime glass (74% SiO2 , 13% Na2O, 8% CaO, 4% MgO, 1% other oxide... Glass Solar panel cover Spacecraft Exposure Buna-N (acrylonitrile butadiene rubber) Seals Iodine Feed System Carbon fiber composite (epoxy resin...Fe Propellant isolator Spacecraft Exposure Lanthanum Hexaboride, LaB6 Cathode emitter Inside Cathode Yes MACOR (46% SiO2 , 17% MgO, 16% Al2O3, 10

Spacecraft Environmental Interactions Technology, 1983

Science.gov (United States)

1985-01-01

State of the art of environment interactions dealing with low-Earth-orbit plasmas; high-voltage systems; spacecraft charging; materials effects; and direction of future programs are contained in over 50 papers.

MW-Class Electric Propulsion System Designs

Science.gov (United States)

LaPointe, Michael R.; Oleson, Steven; Pencil, Eric; Mercer, Carolyn; Distefano, Salvador

2011-01-01

Electric propulsion systems are well developed and have been in commercial use for several years. Ion and Hall thrusters have propelled robotic spacecraft to encounters with asteroids, the Moon, and minor planetary bodies within the solar system, while higher power systems are being considered to support even more demanding future space science and exploration missions. Such missions may include orbit raising and station-keeping for large platforms, robotic and human missions to near earth asteroids, cargo transport for sustained lunar or Mars exploration, and at very high-power, fast piloted missions to Mars and the outer planets. The Advanced In-Space Propulsion Project, High Efficiency Space Power Systems Project, and High Power Electric Propulsion Demonstration Project were established within the NASA Exploration Technology Development and Demonstration Program to develop and advance the fundamental technologies required for these long-range, future exploration missions. Under the auspices of the High Efficiency Space Power Systems Project, and supported by the Advanced In-Space Propulsion and High Power Electric Propulsion Projects, the COMPASS design team at the NASA Glenn Research Center performed multiple parametric design analyses to determine solar and nuclear electric power technology requirements for representative 300-kW class and pulsed and steady-state MW-class electric propulsion systems. This paper describes the results of the MW-class electric power and propulsion design analysis. Starting with the representative MW-class vehicle configurations, and using design reference missions bounded by launch dates, several power system technology improvements were introduced into the parametric COMPASS simulations to determine the potential system level benefits such technologies might provide. Those technologies providing quantitative system level benefits were then assessed for technical feasibility, cost, and time to develop. Key assumptions and primary

Attitude dynamics and control of spacecraft with a partially filled liquid tank and flexible panels

Science.gov (United States)

Liu, Feng; Yue, Baozeng; Zhao, Liangyu

2018-02-01

A liquid-filled flexible spacecraft is essentially a time-variant fully-coupled system, whose dynamics characteristics are closely associated with its motion features. This paper focuses on the mathematical modelling and attitude control of the spacecraft coupled with fuel sloshing dynamics and flexible solar panels vibration. The slosh motion is represented by a spherical pendulum, whose motion description method is improved by using split variable operation. Benefiting from this improvement, the nonlinear lateral sloshing and the rotary sloshing as well as the rigid motion of a liquid respect to the spacecraft can be approximately described. The assumed modes discretization method has been adopted to approximate the elastic displacements of the attached panels, and the coupled dynamics is derived by using the Lagrangian formulation. A variable substitution method is proposed to obtain the apparently-uncoupled mathematical model of the rigid-flexible-liquid spacecraft. After linearization, this model can be directly used for designing Lyapunov output-feedback attitude controller (OFAC). With only torque actuators, and attitude and rate sensors installed, this kind of attitude controller, as simulation results show, is capable of not only bringing the spacecraft to the desired orientation, but also suppressing the effect of flex and slosh on the attitude motion of the spacecraft.

Integration of a Fire Detector into a Spacecraft

Science.gov (United States)

Linford, R. M. F.

1972-01-01

A detector sensitive to only the ultraviolet radiation emitted by flames has been selected as the basic element of the NASA Skylab fire detection system. It is sensitive to approximately 10(exp -12)W of radiation and will detect small flames at distances in excess of 3m. The performance of the detector was verified by experiments in an aircraft flying zero-gravity parabolas to simulate the characteristics of a fire which the detector must sense. Extensive investigation and exacting design was necessary to exclude all possible sources of false alarms. Optical measurements were made on all the spacecraft windows to determine the amount of solar radiation transmitted. The lighting systems and the onboard experiments also were appraised for ultraviolet emissions. Proton-accelerator tests were performed to determine the interaction of the Earth's trapped radiation belts with the detectors and the design of the instrument was modified to negate these effects.

Dream missions space colonies, nuclear spacecraft and other possibilities

CERN Document Server

van Pelt, Michel

2017-01-01

This book takes the reader on a journey through the history of extremely ambitious, large and complex space missions that never happened. What were the dreams and expectations of the visionaries behind these plans, and why were they not successful in bringing their projects to reality thus far? As spaceflight development progressed, new technologies and ideas led to pushing the boundaries of engineering and technology though still grounded in real scientific possibilities. Examples are space colonies, nuclear-propelled interplanetary spacecraft, space telescopes consisting of multiple satellites and canon launch systems. Each project described in this book says something about the dreams and expectations of their time, and their demise was often linked to an important change in the cultural, political and social state of the world. For each mission or spacecraft concept, the following will be covered: • Description of the design. • Overview of the history of the concept and the people involved. • Why it...

Nutation instability of spinning solid rocket motor spacecraft

Directory of Open Access Journals (Sweden)

Dan YANG

2017-08-01

Full Text Available The variation of mass, and moment of inertia of a spin-stabilized spacecraft leads to concern about the nutation instability. Here a careful analysis on the nutation instability is performed on a spacecraft propelled by solid rocket booster (SRB. The influences of specific solid propellant designs on transversal angular velocity are discussed. The results show that the typical SRB of End Burn suppresses the non-principal axial angular velocity. On the contrary, the frequently used SRB of Radial Burn could amplify the transversal angular velocity. The nutation instability caused by a design of Radial Burn could be remedied by the addition of End Burn at the same time based on the study of the combination design of both End Burn and Radial Burn. The analysis of the results proposes the design conception of how to control the nutation motion. The method is suitable to resolve the nutation instability of solid rocket motor with complex propellant patterns.

A Quantized State Approach to On-line Simulation for Spacecraft Autonomy

DEFF Research Database (Denmark)

Alminde, Lars; Stoustrup, Jakob; Bendtsen, Jan Dimon

2006-01-01

Future space applications will require an increased level of operational autonomy. This calls for declarative methods for spacecraft state estimation and control, so that the spacecraft engineer can focus on modeling the spacecraft rather than implementing all details of the on-line system. Celeb...

Navigating the MESSENGER Spacecraft through End of Mission

Science.gov (United States)

Bryan, C. G.; Williams, B. G.; Williams, K. E.; Taylor, A. H.; Carranza, E.; Page, B. R.; Stanbridge, D. R.; Mazarico, E.; Neumann, G. A.; O'Shaughnessy, D. J.; McAdams, J. V.; Calloway, A. B.

2015-12-01

The MErcury Surface, Space ENvironment, GEochemistry, and Ranging (MESSENGER) spacecraft orbited the planet Mercury from March 2011 until the end of April 2015, when it impacted the planetary surface after propellant reserves used to maintain the orbit were depleted. This highly successful mission was led by the principal investigator, Sean C. Solomon, of Columbia University. The Johns Hopkins University Applied Physics Laboratory (JHU/APL) designed and assembled the spacecraft and served as the home for spacecraft operations. Spacecraft navigation for the entirety of the mission was provided by the Space Navigation and Flight Dynamics Practice (SNAFD) of KinetX Aerospace. Orbit determination (OD) solutions were generated through processing of radiometric tracking data provided by NASA's Deep Space Network (DSN) using the MIRAGE suite of orbital analysis tools. The MESSENGER orbit was highly eccentric, with periapsis at a high northern latitude and periapsis altitude in the range 200-500 km for most of the orbital mission phase. In a low-altitude "hover campaign" during the final two months of the mission, periapsis altitudes were maintained within a narrow range between about 35 km and 5 km. Navigating a spacecraft so near a planetary surface presented special challenges. Tasks required to meet those challenges included the modeling and estimation of Mercury's gravity field and of solar and planetary radiation pressure, and the design of frequent orbit-correction maneuvers. Superior solar conjunction also presented observational modeling issues. One key to the overall success of the low-altitude hover campaign was a strategy to utilize data from an onboard laser altimeter as a cross-check on the navigation team's reconstructed and predicted estimates of periapsis altitude. Data obtained from the Mercury Laser Altimeter (MLA) on a daily basis provided near-real-time feedback that proved invaluable in evaluating alternative orbit estimation strategies, and

Passive radiative cooling of a HTS coil for attitude orbit control in micro-spacecraft

Science.gov (United States)

Inamori, Takaya; Ozaki, Naoya; Saisutjarit, Phongsatorn; Ohsaki, Hiroyuki

2015-02-01

This paper proposes a novel radiative cooling system for a high temperature superconducting (HTS) coil for an attitude orbit control system in nano- and micro-spacecraft missions. These days, nano-spacecraft (1-10 kg) and micro-spacecraft (10-100 kg) provide space access to a broader range of spacecraft developers and attract interest as space development applications. In planetary and high earth orbits, most previous standard-size spacecraft used thrusters for their attitude and orbit control, which are not available for nano- and micro-spacecraft missions because of the strict power consumption, space, and weight constraints. This paper considers orbit and attitude control methods that use a superconducting coil, which interacts with on-orbit space plasmas and creates a propulsion force. Because these spacecraft cannot use an active cooling system for the superconducting coil because of their mass and power consumption constraints, this paper proposes the utilization of a passive radiative cooling system, in which the superconducting coil is thermally connected to the 3 K cosmic background radiation of deep space, insulated from the heat generation using magnetic holders, and shielded from the sun. With this proposed cooling system, the HTS coil is cooled to 60 K in interplanetary orbits. Because the system does not use refrigerators for its cooling system, the spacecraft can achieve an HTS coil with low power consumption, small mass, and low cost.

Precise Relative Positioning of Formation Flying Spacecraft using GPS

NARCIS (Netherlands)

Kroes, R.

2006-01-01

Spacecraft formation flying is considered as a key technology for advanced space missions. Compared to large individual spacecraft, the distribution of sensor systems amongst multiple platforms offers improved flexibility, shorter times to mission, and the prospect of being more cost effective.

Concept Assessment of a Fission Fragment Rocket Engine (FFRE) Propelled Spacecraft

Science.gov (United States)

Werka, Robert; Clark, Rod; Sheldon, Rob; Percy, Tom

2012-01-01

The March, 2012 issue of Aerospace America stated that ?the near-to-medium prospects for applying advanced propulsion to create a new era of space exploration are not very good. In the current world, we operate to the Moon by climbing aboard a Carnival Cruise Lines vessel (Saturn 5), sail from the harbor (liftoff) shedding whole decks of the ship (staging) along the way and, having reached the return leg of the journey, sink the ship (burnout) and return home in a lifeboat (Apollo capsule). Clearly this is an illogical way to travel, but forced on Explorers by today's propulsion technology. However, the article neglected to consider the one propulsion technology, using today's physical principles that offer continuous, substantial thrust at a theoretical specific impulse of 1,000,000 sec. This engine unequivocally can create a new era of space exploration that changes the way spacecraft operate. Today's space Explorers could travel in Cruise Liner fashion using the technology not considered by Aerospace America, the novel Dusty Plasma Fission Fragment Rocket Engine (FFRE). This NIAC study addresses the FFRE as well as its impact on Exploration Spacecraft design and operation. It uses common physics of the relativistic speed of fission fragments to produce thrust. It radiatively cools the fissioning dusty core and magnetically controls the fragments direction to practically implement previously patented, but unworkable designs. The spacecraft hosting this engine is no more complex nor more massive than the International Space Station (ISS) and would employ the successful ISS technology for assembly and check-out. The elements can be lifted in "chunks" by a Heavy Lift Launcher. This Exploration Spacecraft would require the resupply of small amounts of nuclear fuel for each journey and would be an in-space asset for decades just as any Cruise Liner on Earth. This study has synthesized versions of the FFRE, integrated one concept onto a host spacecraft designed for

LO2/LH2 propulsion for outer planet orbiter spacecraft

Science.gov (United States)

Garrison, P. W.; Sigurdson, K. B.

1983-01-01

Galileo class orbiter missions (750-1500 kg) to the outer planets require a large postinjection delta-V for improved propulsion performance. The present investigation shows that a pump-fed low thrust LO2/LH2 propulsion system can provide a significantly larger net on-orbit mass for a given delta-V than a state-of-the-art earth storable, N2O4/monomethylhydrazine pressure-fed propulsion system. A description is given of a conceptual design for a LO2/LH2 pump-fed propulsion system developed for a Galileo class mission to the outer planets. Attention is given to spacecraft configuration, details regarding the propulsion system, the thermal control of the cryogenic propellants, and aspects of mission performance.

A passively controlled appendage deployment system for the San Marco D/L spacecraft

Science.gov (United States)

Lang, W. E.; Frisch, H. P.; Schwartz, D. A.

1984-01-01

The analytical simulation of deployment dynamics of these two axis concepts as well as the evolution of practical designs for the add on deployable inertia boom units is described. With the boom free to swing back in response to Coriolis forces as well as outwards in response to centrifugal forces, the kinematics of motion are complex but admit the possibility of absorbing deployment energy in frictional or other damping devices about the radial axis, where large amplitude motions can occur and where the design envelope allows more available volume. An acceptable range is defined for frictional damping for any given spin rate. Inadequate damping allows boom motions which strike the spacecraft; excessive damping causes the boom to swing out and latch with damaging violence. The acceptable range is a design parameter and must accommodate spin rate tolerance and also the tolerance and repeatability of the damping mechanisms.

Pi-Sat: A Low Cost Small Satellite and Distributed Spacecraft Mission System Test Platform

Science.gov (United States)

Cudmore, Alan

2015-01-01

Current technology and budget trends indicate a shift in satellite architectures from large, expensive single satellite missions, to small, low cost distributed spacecraft missions. At the center of this shift is the SmallSatCubesat architecture. The primary goal of the Pi-Sat project is to create a low cost, and easy to use Distributed Spacecraft Mission (DSM) test bed to facilitate the research and development of next-generation DSM technologies and concepts. This test bed also serves as a realistic software development platform for Small Satellite and Cubesat architectures. The Pi-Sat is based on the popular $35 Raspberry Pi single board computer featuring a 700Mhz ARM processor, 512MB of RAM, a flash memory card, and a wealth of IO options. The Raspberry Pi runs the Linux operating system and can easily run Code 582s Core Flight System flight software architecture. The low cost and high availability of the Raspberry Pi make it an ideal platform for a Distributed Spacecraft Mission and Cubesat software development. The Pi-Sat models currently include a Pi-Sat 1U Cube, a Pi-Sat Wireless Node, and a Pi-Sat Cubesat processor card.The Pi-Sat project takes advantage of many popular trends in the Maker community including low cost electronics, 3d printing, and rapid prototyping in order to provide a realistic platform for flight software testing, training, and technology development. The Pi-Sat has also provided fantastic hands on training opportunities for NASA summer interns and Pathways students.

Relativity time-delay experiments utilizing 'Mariner' spacecraft

Science.gov (United States)

Esposito, P. B.; Anderson, J. D.

1974-01-01

Relativity predicts that the transit time of a signal propagated from the earth to a spacecraft and retransmitted back to earth ought to exhibit an additional, variable time delay. The present work describes some of the analytical techniques employed in experiments using Mariner spacecraft designed to test the accuracy of this prediction. Two types of data are analyzed in these relativity experiments; these include phase-coherent, two-way Doppler shift and round-trip, transit-time measurements. Results of Mariner 6 and 7 relativistic time-delay experiments are in agreement with Einstein's theory of general relativity with an uncertainty of 3%.

The SPOT-HRV instrument - An overview of design and performance

Science.gov (United States)

Midan, J. P.

1983-10-01

The SPOT spacecraft's High Visible Resolution (HVR) earth resources sensor performance requirements, system and subsystem design features, and technology development considerations, are discussed. Attention is given to such problem areas involving extensive design tradeoff analyses and testing as those uncovered by mechanical design and thermal distortion studies and SNR analysis and calibration considerations. The SPOT spacecraft will be placed in orbit in 1985.

Spacecraft Health Automated Reasoning Prototype (SHARP): The fiscal year 1989 SHARP portability evaluations task for NASA Solar System Exploration Division's Voyager project

Science.gov (United States)

Atkinson, David J.; Doyle, Richard J.; James, Mark L.; Kaufman, Tim; Martin, R. Gaius

1990-01-01

A Spacecraft Health Automated Reasoning Prototype (SHARP) portability study is presented. Some specific progress is described on the portability studies, plans for technology transfer, and potential applications of SHARP and related artificial intelligence technology to telescience operations. The application of SHARP to Voyager telecommunications was a proof-of-capability demonstration of artificial intelligence as applied to the problem of real time monitoring functions in planetary mission operations. An overview of the design and functional description of the SHARP system is also presented as it was applied to Voyager.

Novel Design Aspects of the Space Technology 5 Mechanical Subsystem

Science.gov (United States)

Rossoni, Peter; McGill, William

2003-01-01

This paper describes several novel design elements of the Space Technology 5 (ST5) spacecraft mechanical subsystem. The spacecraft structure itself takes a significant step in integrating electronics into the primary structure. The deployment system restrains the spacecraft during launch and imparts a predetermined spin rate upon release from its secondary payload accommodations. The deployable instrument boom incorporates some traditional as well as new techniques for lightweight and stiffness. Analysis and test techniques used to validate these technologies are described. Numerous design choices were necessitated due to the compact spacecraft size and strict mechanical subsystem requirements.

Ultra Long-Life Spacecraft for Long Duration Space Exploration Missions

Science.gov (United States)

Chau, Savio

2002-01-01

After decades of Solar System exploration, NASA has almost completed the initial reconnaissance, and has been planning for landing and sample return missions on many planets, satellites, comets, and asteroids. The next logical step of space exploration is to expand the frontier into other missions within and outside the solar system. These missions can easily last for more than 30 to 50 years. Most of the current technologies and spacecraft design techniques are not adequate to support such long life missions. Many breakthrough technologies and non-conventional system architecture have to develop in order to sustain such long life missions.Some of these technologies are being developed by the NASA Exploration Team (neXt). Based on the projected requirements for ultra long life missions, the costs and benefits of the required technologies can be quantified. The ultra long-life space system should have four attributes: long-term survivability, administration of consumable resources, evolvability and adaptability, and low-cost long-term operations of the spacecraft. The discussion of survivability is the focus of this paper. Conventional fault tolerant system design has to tolerate only random failures, which can be handled effectively by dual or triple redundancy for a relatively short time. In contrast, the predominant failure mode in an ultra long-life system is the wear-out of components. All active components in the system are destined to fail before the end of the mission. Therefore, an ultra long-life system would require a large number of redundant components. This would be impractical in conventional fault tolerant systems because their fault tolerance techniques are very inefficient. For instance, a conventional dual-string avionics system duplicates the all the components including the processor, memory, and I/O controllers on a spacecraft. However, when the same component in both strings fail (e.g., the processor), the system will fail although all other

Fuzzy based attitude controller for flexible spacecraft with on/off thrusters

Science.gov (United States)

Knapp, Roger Glenn

1993-05-01

A fuzzy-based attitude controller is designed for attitude control of a generic spacecraft with on/off thrusters. The controller is comprised of packages of rules dedicated to addressing different objectives (e.g., disturbance rejection, low fuel consumption, avoiding the excitation of flexible appendages, etc.). These rule packages can be inserted or removed depending on the requirements of the particular spacecraft and are parameterized based on vehicle parameters such as inertia or operational parameters such as the maneuvering rate. Individual rule packages can be 'weighted' relative to each other to emphasize the importance of one objective relative to another. Finally, the fuzzy controller and rule packages are demonstrated using the high-fidelity Space Shuttle Interactive On-Orbit Simulator (IOS) while performing typical on-orbit operations and are subsequently compared with the existing shuttle flight control system performance.

Relative Status Determination for Spacecraft Relative Motion Based on Dual Quaternion

Directory of Open Access Journals (Sweden)

Jun Sun

2014-01-01

Full Text Available For the two-satellite formation, the relative motion and attitude determination algorithm is a key component that affects the flight quality and mission efficiency. The relative status determination algorithm is proposed based on the Extended Kalman Filter (EKF and the system state optimal estimate linearization. Aiming at the relative motion of the spacecraft formation navigation problem, the spacecraft relative kinematics and dynamics model are derived from the dual quaternion in the algorithm. Then taking advantage of EKF technique, combining with the dual quaternion integrated dynamic models, considering the navigation algorithm using the fusion measurement by the gyroscope and star sensors, the relative status determination algorithm is designed. At last the simulation is done to verify the feasibility of the algorithm. The simulation results show that the EKF algorithm has faster convergence speed and higher accuracy.

High-precision relative position and attitude measurement for on-orbit maintenance of spacecraft

Science.gov (United States)

Zhu, Bing; Chen, Feng; Li, Dongdong; Wang, Ying

2018-02-01

In order to realize long-term on-orbit running of satellites, space stations, etc spacecrafts, in addition to the long life design of devices, The life of the spacecraft can also be extended by the on-orbit servicing and maintenance. Therefore, it is necessary to keep precise and detailed maintenance of key components. In this paper, a high-precision relative position and attitude measurement method used in the maintenance of key components is given. This method mainly considers the design of the passive cooperative marker, light-emitting device and high resolution camera in the presence of spatial stray light and noise. By using a series of algorithms, such as background elimination, feature extraction, position and attitude calculation, and so on, the high precision relative pose parameters as the input to the control system between key operation parts and maintenance equipment are obtained. The simulation results show that the algorithm is accurate and effective, satisfying the requirements of the precision operation technique.

Using neuromorphic optical sensors for spacecraft absolute and relative navigation

Science.gov (United States)

Shake, Christopher M.

We develop a novel attitude determination system (ADS) for use on nano spacecraft using neuromorphic optical sensors. The ADS intends to support nano-satellite operations by providing low-cost, low-mass, low-volume, low-power, and redundant attitude determination capabilities with quick and straightforward onboard programmability for real time spacecraft operations. The ADS is experimentally validated with commercial-off-the-shelf optical devices that perform sensing and image processing on the same circuit board and are biologically inspired by insects' vision systems, which measure optical flow while navigating in the environment. The firmware on the devices is modified to both perform the additional biologically inspired task of tracking objects and communicate with a PC/104 form-factor embedded computer running Real Time Application Interface Linux used on a spacecraft simulator. Algorithms are developed for operations using optical flow, point tracking, and hybrid modes with the sensors, and the performance of the system in all three modes is assessed using a spacecraft simulator in the Advanced Autonomous Multiple Spacecraft (ADAMUS) laboratory at Rensselaer. An existing relative state determination method is identified to be combined with the novel ADS to create a self-contained navigation system for nano spacecraft. The performance of the method is assessed in simulation and found not to match the results from its authors using only conditions and equations already published. An improved target inertia tensor method is proposed as an update to the existing relative state method, but found not to perform as expected, but is presented for others to build upon.

Practical Applications of Cosmic Ray Science: Spacecraft, Aircraft, Ground-Based Computation and Control Systems, Exploration, and Human Health and Safety

Science.gov (United States)

Koontz, Steve

2015-01-01

In this presentation a review of galactic cosmic ray (GCR) effects on microelectronic systems and human health and safety is given. The methods used to evaluate and mitigate unwanted cosmic ray effects in ground-based, atmospheric flight, and space flight environments are also reviewed. However not all GCR effects are undesirable. We will also briefly review how observation and analysis of GCR interactions with planetary atmospheres and surfaces and reveal important compositional and geophysical data on earth and elsewhere. About 1000 GCR particles enter every square meter of Earth’s upper atmosphere every second, roughly the same number striking every square meter of the International Space Station (ISS) and every other low- Earth orbit spacecraft. GCR particles are high energy ionized atomic nuclei (90% protons, 9% alpha particles, 1% heavier nuclei) traveling very close to the speed of light. The GCR particle flux is even higher in interplanetary space because the geomagnetic field provides some limited magnetic shielding. Collisions of GCR particles with atomic nuclei in planetary atmospheres and/or regolith as well as spacecraft materials produce nuclear reactions and energetic/highly penetrating secondary particle showers. Three twentieth century technology developments have driven an ongoing evolution of basic cosmic ray science into a set of practical engineering tools needed to design, test, and verify the safety and reliability of modern complex technological systems and assess effects on human health and safety effects. The key technology developments are: 1) high altitude commercial and military aircraft; 2) manned and unmanned spacecraft; and 3) increasingly complex and sensitive solid state micro-electronics systems. Space and geophysical exploration needs drove the development of the instruments and analytical tools needed to recover compositional and structural data from GCR induced nuclear reactions and secondary particle showers. Finally, the

Dynamics and control of Lorentz-augmented spacecraft relative motion

CERN Document Server

Yan, Ye; Yang, Yueneng

2017-01-01

This book develops a dynamical model of the orbital motion of Lorentz spacecraft in both unperturbed and J2-perturbed environments. It explicitly discusses three kinds of typical space missions involving relative orbital control: spacecraft hovering, rendezvous, and formation flying. Subsequently, it puts forward designs for both open-loop and closed-loop control schemes propelled or augmented by the geomagnetic Lorentz force. These control schemes are entirely novel and represent a significantly departure from previous approaches.

SpaceX's Dragon America's next generation spacecraft

CERN Document Server

Seedhouse, Erik

2016-01-01

This book describes Dragon V2, a futuristic vehicle that not only provides a means for NASA to transport its astronauts to the orbiting outpost but also advances SpaceX’s core objective of reusability. A direct descendant of Dragon, Dragon V2 can be retrieved, refurbished and re-launched. It is a spacecraft with the potential to completely revolutionize the economics of an industry where equipment costing hundreds of millions of dollars is routinely discarded after a single use. It was presented by SpaceX CEO Elon Musk in May 2014 as the spaceship that will carry NASA astronauts to the International Space Station as soon as 2016 SpaceX’s Dragon – America’s Next Generation Spacecraft describes the extraordinary feats of engineering and human achievement that have placed this revolutionary spacecraft at the forefront of the launch industry and positioned it as the precursor for ultimately transporting humans to Mars. It describes the design and development of Dragon, provides mission highlights of the f...

Design and Simulation of a MEMS Control Moment Gyroscope for the Sub-Kilogram Spacecraft

Directory of Open Access Journals (Sweden)

Weizheng Yuan

2010-04-01

Full Text Available A novel design of a microelectromechanical systems (MEMS control moment gyroscope (MCMG was proposed in this paper in order to generate a torque output with a magnitude of 10-6 N∙m. The MCMG consists of two orthogonal angular vibration systems, i.e., the rotor and gimbal; the coupling between which is based on the Coriolis effect and will cause a torque output in the direction perpendicular to the two vibrations. The angular rotor vibration was excited by the in-plane electrostatic rotary comb actuators, while the angular gimbal vibration was driven by an out-of-plane electrostatic parallel plate actuator. A possible process flow to fabricate the structure was proposed and discussed step by step. Furthermore, an array configuration using four MCMGs as an effective element, in which the torque was generated with a phase difference of 90 degrees between every two MCMGs, was proposed to smooth the inherent fluctuation of the torque output for a vibrational MCMG. The parasitic torque was cancelled by two opposite MCMGs with a phase difference of 180 degrees. The designed MCMG was about 1.1 cm × 1.1 cm × 0.04 cm in size and 0.1 g in weight. The simulation results showed that the maximum torque output of a MCMG, the resonant frequency of which was approximately 1,000 Hz, was about 2.5 × 10-8 N∙m. The element with four MCMGs could generate a torque of 5 × 10-8 N∙m. The torque output could reach a magnitude of 10-6 N∙m when the frequency was improved from 1,000 Hz to 10,000 Hz. Using arrays of 4 × 4 effective elements on a 1 kg spacecraft with a standard form factor of 10 cm × 10 cm × 10 cm, a 10 degrees attitude change could be achieved in 26.96s.

User interface design principles for the SSM/PMAD automated power system

International Nuclear Information System (INIS)

Jakstas, L.M.; Myers, C.J.

1991-01-01

Computer-human interfaces are an integral part of developing software for spacecraft power systems. A well designed and efficient user interface enables an engineer to effectively operate the system, while it concurrently prevents the user from entering data which is beyond boundary conditions or performing operations which are out of context. A user interface should also be designed to ensure that the engineer easily obtains all useful and critical data for operating the system and is aware of all faults and states in the system. Martin Marietta, under contract to NASA George C. Marshall Space Flight Center, has developed a user interface for the Space Station Module Power Management and Distribution (SSM/PMAD) automated power system testbed which provides human access to the functionality of the power system, as well as exemplifying current techniques in user interface design. The testbed user interface was designed to enable an engineer to operate the system easily without having significant knowledge of computer systems, as well as provide an environment in which the engineer can monitor and interact with the SSM/PMAD system hardware. The design of the interface supports a global view of the most important data form the various hardware and software components, as well as enabling the user to obtain additional or more detailed data when needed. The components and representations of the SSM/PMAD testbed user interface are examined in this paper. An engineer's interactions with the system are also described

Manned spacecraft electrical power systems

Science.gov (United States)

Simon, William E.; Nored, Donald L.

1987-01-01

A brief history of the development of electrical power systems from the earliest manned space flights illustrates a natural trend toward a growth of electrical power requirements and operational lifetimes with each succeeding space program. A review of the design philosophy and development experience associated with the Space Shuttle Orbiter electrical power system is presented, beginning with the state of technology at the conclusion of the Apollo Program. A discussion of prototype, verification, and qualification hardware is included, and several design improvements following the first Orbiter flight are described. The problems encountered, the scientific and engineering approaches used to meet the technological challenges, and the results obtained are stressed. Major technology barriers and their solutions are discussed, and a brief Orbiter flight experience summary of early Space Shuttle missions is included. A description of projected Space Station power requirements and candidate system concepts which could satisfy these anticipated needs is presented. Significant challenges different from Space Shuttle, innovative concepts and ideas, and station growth considerations are discussed. The Phase B Advanced Development hardware program is summarized and a status of Phase B preliminary tradeoff studies is presented.

Performance verification and system parameter identification of spacecraft tape recorder control servo

Science.gov (United States)

Mukhopadhyay, A. K.

1979-01-01

Design adequacy of the lead-lag compensator of the frequency loop, accuracy checking of the analytical expression for the electrical motor transfer function, and performance evaluation of the speed control servo of the digital tape recorder used on-board the 1976 Viking Mars Orbiters and Voyager 1977 Jupiter-Saturn flyby spacecraft are analyzed. The transfer functions of the most important parts of a simplified frequency loop used for test simulation are described and ten simulation cases are reported. The first four of these cases illustrate the method of selecting the most suitable transfer function for the hysteresis synchronous motor, while the rest verify and determine the servo performance parameters and alternative servo compensation schemes. It is concluded that the linear methods provide a starting point for the final verification/refinement of servo design by nonlinear time response simulation and that the variation of the parameters of the static/dynamic Coulomb friction is as expected in a long-life space mission environment.

Using the Design for Demise Philosophy to Reduce Casualty Risk Due to Reentering Spacecraft

Science.gov (United States)

Kelley, R. L.

2012-01-01

Recently the reentry of a number of vehicles has garnered public attention due to their risk of human casualty due to fragments surviving reentry. In order to minimize this risk for their vehicles, a number of NASA programs have actively sought to minimize the number of components likely to survive reentry at the end of their spacecraft's life in order to meet and/or exceed NASA safety standards for controlled and uncontrolled reentering vehicles. This philosophy, referred to as "Design for Demise" or D4D, has steadily been adopted, to at least some degree, by numerous programs. The result is that many programs are requesting evaluations of components at the early stages of vehicle design, as they strive to find ways to reduce the number surviving components while ensuring that the components meet the performance requirements of their mission. This paper will discuss some of the methods that have been employed to ensure that the consequences of the vehicle s end-of-life are considered at the beginning of the design process. In addition this paper will discuss the technical challenges overcome, as well as some of the more creative solutions which have been utilized to reduce casualty risk.

The Glory Program: Global Science from a Unique Spacecraft Integration

Science.gov (United States)

Bajpayee Jaya; Durham, Darcie; Ichkawich, Thomas

2006-01-01

The Glory program is an Earth and Solar science mission designed to broaden science community knowledge of the environment. The causes and effects of global warming have become a concern in recent years and Glory aims to contribute to the knowledge base of the science community. Glory is designed for two functions: one is solar viewing to monitor the total solar irradiance and the other is observing the Earth s atmosphere for aerosol composition. The former is done with an active cavity radiometer, while the latter is accomplished with an aerosol polarimeter sensor to discern atmospheric particles. The Glory program is managed by NASA Goddard Space Flight Center (GSFC) with Orbital Sciences in Dulles, VA as the prime contractor for the spacecraft bus, mission operations, and ground system. This paper will describe some of the more unique features of the Glory program including the integration and testing of the satellite and instruments as well as the science data processing. The spacecraft integration and test approach requires extensive analysis and additional planning to ensure existing components are successfully functioning with the new Glory components. The science mission data analysis requires development of mission unique processing systems and algorithms. Science data analysis and distribution will utilize our national assets at the Goddard Institute for Space Studies (GISS) and the University of Colorado's Laboratory for Atmospheric and Space Physics (LASP). The Satellite was originally designed and built for the Vegetation Canopy Lidar (VCL) mission, which was terminated in the middle of integration and testing due to payload development issues. The bus was then placed in secure storage in 2001 and removed from an environmentally controlled container in late 2003 to be refurbished to meet the Glory program requirements. Functional testing of all the components was done as a system at the start of the program, very different from a traditional program

Design and Integration of an All-Magnetic Attitude Control System for FASTSAT-HSV01's Multiple Pointing Objectives

Science.gov (United States)

DeKock, Brandon; Sanders, Devon; Vanzwieten, Tannen; Capo-Lugo, Pedro

2011-01-01

The FASTSAT-HSV01 spacecraft is a microsatellite with magnetic torque rods as it sole attitude control actuator. FASTSAT s multiple payloads and mission functions require the Attitude Control System (ACS) to maintain Local Vertical Local Horizontal (LVLH)-referenced attitudes without spin-stabilization, while the pointing errors for some attitudes be significantly smaller than the previous best-demonstrated for this type of control system. The mission requires the ACS to hold multiple stable, unstable, and non-equilibrium attitudes, as well as eject a 3U CubeSat from an onboard P-POD and recover from the ensuing tumble. This paper describes the Attitude Control System, the reasons for design choices, how the ACS integrates with the rest of the spacecraft, and gives recommendations for potential future applications of the work.

SMART-1: the first spacecraft of the future

Science.gov (United States)

2003-09-01

This is the first of a series of missions designed to test key technologies for future spacecraft —SMART stands for 'Small Missions for Advanced Research and Technology'. In the case of SMART-1, the two main new technologies to be tested are a new 'solar-electric propulsion' system and miniaturised spacecraft and instrumentation. Together, these technologies make up a spacecraft with revolutionary qualities: smaller, lighter, capable of carrying more scientific instruments, greater fuel efficiency. All of which also considerably reduces the cost of the mission. So, the idea behind SMART-1 is to pioneer a futuristic philosophy, the motto of which could be: 'more science for less money'. Even though it is the first of a kind, SMART-1 has been developed in less than four years, and at about a fifth of the cost of a major science mission for ESA: only 110 million euros. That includes the launch, the operations and a dozen scientific experiments. This was achieved partly by using new management methods — such as working with smaller teams both within ESA and in the industry — and partly because of some of the new features inherent in SMART-1, such as the miniaturisation and novel design. Giuseppe Racca, SMART-1 Project Manager, explains: "What has been our trick? First, a short development period in itself means less money. But also, with its small size — which was a requirement of the mission because we are testing miniaturised hardware — the spacecraft is able to 'share' a commercial Ariane flight with two other passengers. Besides, since we were not constrained by any existing design or heritage, we could be more innovative and elegant in our architecture. For example, the new SMART-1 electrical architecture has enabled us to simplify the system tests considerably." SMART-1 could almost be a toy spacecraft — it weighs only 367 kilograms and fits into a cube just one metre across (the solar panel wings extend about 14 metres) — although one able to

Large-Scale Spacecraft Fire Safety Experiments in ISS Resupply Vehicles

Science.gov (United States)

Ruff, Gary A.; Urban, David

2013-01-01

Our understanding of the fire safety risk in manned spacecraft has been limited by the small scale of the testing we have been able to conduct in low-gravity. Fire growth and spread cannot be expected to scale linearly with sample size so we cannot make accurate predictions of the behavior of realistic scale fires in spacecraft based on the limited low-g testing to date. As a result, spacecraft fire safety protocols are necessarily very conservative and costly. Future crewed missions are expected to be longer in duration than previous exploration missions outside of low-earth orbit and accordingly, more complex in terms of operations, logistics, and safety. This will increase the challenge of ensuring a fire-safe environment for the crew throughout the mission. Based on our fundamental uncertainty of the behavior of fires in low-gravity, the need for realistic scale testing at reduced gravity has been demonstrated. To address this concern, a spacecraft fire safety research project is underway to reduce the uncertainty and risk in the design of spacecraft fire safety systems by testing at nearly full scale in low-gravity. This project is supported by the NASA Advanced Exploration Systems Program Office in the Human Exploration and Operations Mission Directorate. The activity of this project is supported by an international topical team of fire experts from other space agencies to maximize the utility of the data and to ensure the widest possible scrutiny of the concept. The large-scale space flight experiment will be conducted on three missions; each in an Orbital Sciences Corporation Cygnus vehicle after it has deberthed from the ISS. Although the experiment will need to meet rigorous safety requirements to ensure the carrier vehicle does not sustain damage, the absence of a crew allows the fire products to be released into the cabin. The tests will be fully automated with the data downlinked at the conclusion of the test before the Cygnus vehicle reenters the

DET/MPS - THE GSFC ENERGY BALANCE PROGRAM, DIRECT ENERGY TRANSFER/MULTIMISSION SPACECRAFT MODULAR POWER SYSTEM (MACINTOSH A/UX VERSION)

Science.gov (United States)

Jagielski, J. M.

1994-01-01

The DET/MPS programs model and simulate the Direct Energy Transfer and Multimission Spacecraft Modular Power System in order to aid both in design and in analysis of orbital energy balance. Typically, the DET power system has the solar array directly to the spacecraft bus, and the central building block of MPS is the Standard Power Regulator Unit. DET/MPS allows a minute-by-minute simulation of the power system's performance as it responds to various orbital parameters, focusing its output on solar array output and battery characteristics. While this package is limited in terms of orbital mechanics, it is sufficient to calculate eclipse and solar array data for circular or non-circular orbits. DET/MPS can be adjusted to run one or sequential orbits up to about one week, simulated time. These programs have been used on a variety of Goddard Space Flight Center spacecraft projects. DET/MPS is written in FORTRAN 77 with some VAX-type extensions. Any FORTRAN 77 compiler that includes VAX extensions should be able to compile and run the program with little or no modifications. The compiler must at least support free-form (or tab-delineated) source format and 'do do-while end-do' control structures. DET/MPS is available for three platforms: GSC-13374, for DEC VAX series computers running VMS, is available in DEC VAX Backup format on a 9-track 1600 BPI tape (standard distribution) or TK50 tape cartridge; GSC-13443, for UNIX-based computers, is available on a .25 inch streaming magnetic tape cartridge in UNIX tar format; and GSC-13444, for Macintosh computers running AU/X with either the NKR FORTRAN or AbSoft MacFORTRAN II compilers, is available on a 3.5 inch 800K Macintosh format diskette. Source code and test data are supplied. The UNIX version of DET requires 90K of main memory for execution. DET/MPS was developed in 1990. A/UX and Macintosh are registered trademarks of Apple Computer, Inc. VMS, DEC VAX and TK50 are trademarks of Digital Equipment Corporation. UNIX is a

DET/MPS - THE GSFC ENERGY BALANCE PROGRAM, DIRECT ENERGY TRANSFER/MULTIMISSION SPACECRAFT MODULAR POWER SYSTEM (DEC VAX VMS VERSION)

Science.gov (United States)

Jagielski, J. M.

1994-01-01

The DET/MPS programs model and simulate the Direct Energy Transfer and Multimission Spacecraft Modular Power System in order to aid both in design and in analysis of orbital energy balance. Typically, the DET power system has the solar array directly to the spacecraft bus, and the central building block of MPS is the Standard Power Regulator Unit. DET/MPS allows a minute-by-minute simulation of the power system's performance as it responds to various orbital parameters, focusing its output on solar array output and battery characteristics. While this package is limited in terms of orbital mechanics, it is sufficient to calculate eclipse and solar array data for circular or non-circular orbits. DET/MPS can be adjusted to run one or sequential orbits up to about one week, simulated time. These programs have been used on a variety of Goddard Space Flight Center spacecraft projects. DET/MPS is written in FORTRAN 77 with some VAX-type extensions. Any FORTRAN 77 compiler that includes VAX extensions should be able to compile and run the program with little or no modifications. The compiler must at least support free-form (or tab-delineated) source format and 'do do-while end-do' control structures. DET/MPS is available for three platforms: GSC-13374, for DEC VAX series computers running VMS, is available in DEC VAX Backup format on a 9-track 1600 BPI tape (standard distribution) or TK50 tape cartridge; GSC-13443, for UNIX-based computers, is available on a .25 inch streaming magnetic tape cartridge in UNIX tar format; and GSC-13444, for Macintosh computers running AU/X with either the NKR FORTRAN or AbSoft MacFORTRAN II compilers, is available on a 3.5 inch 800K Macintosh format diskette. Source code and test data are supplied. The UNIX version of DET requires 90K of main memory for execution. DET/MPS was developed in 1990. A/UX and Macintosh are registered trademarks of Apple Computer, Inc. VMS, DEC VAX and TK50 are trademarks of Digital Equipment Corporation. UNIX is a

The Physics and Technology of Solar Sail Spacecraft.

Science.gov (United States)

Dwivedi, B. N.; McInnes, C. R.

1991-01-01

Various aspects of the solar sail spacecraft such as solar sailing, solar sail design, navigation with solar sails, solar sail mission applications and future prospects for solar sailing are described. Several possible student projects are suggested. (KR)

The Software Design for the Wide-Field Infrared Explorer Attitude Control System

Science.gov (United States)

Anderson, Mark O.; Barnes, Kenneth C.; Melhorn, Charles M.; Phillips, Tom

1998-01-01

The Wide-Field Infrared Explorer (WIRE), currently scheduled for launch in September 1998, is the fifth of five spacecraft in the NASA/Goddard Small Explorer (SMEX) series. This paper presents the design of WIRE's Attitude Control System flight software (ACS FSW). WIRE is a momentum-biased, three-axis stabilized stellar pointer which provides high-accuracy pointing and autonomous acquisition for eight to ten stellar targets per orbit. WIRE's short mission life and limited cryogen supply motivate requirements for Sun and Earth avoidance constraints which are designed to prevent catastrophic instrument damage and to minimize the heat load on the cryostat. The FSW implements autonomous fault detection and handling (FDH) to enforce these instrument constraints and to perform several other checks which insure the safety of the spacecraft. The ACS FSW implements modules for sensor data processing, attitude determination, attitude control, guide star acquisition, actuator command generation, command/telemetry processing, and FDH. These software components are integrated with a hierarchical control mode managing module that dictates which software components are currently active. The lowest mode in the hierarchy is the 'safest' one, in the sense that it utilizes a minimal complement of sensors and actuators to keep the spacecraft in a stable configuration (power and pointing constraints are maintained). As higher modes in the hierarchy are achieved, the various software functions are activated by the mode manager, and an increasing level of attitude control accuracy is provided. If FDH detects a constraint violation or other anomaly, it triggers a safing transition to a lower control mode. The WIRE ACS FSW satisfies all target acquisition and pointing accuracy requirements, enforces all pointing constraints, provides the ground with a simple means for reconfiguring the system via table load, and meets all the demands of its real-time embedded environment (16 MHz Intel

An Analysis of 20 Years of Space Weathering Effects on the Boeing 376 Spacecraft

Science.gov (United States)

Frith, James; Anz-Meador, Phillip; Cowardin, Heather; Buckalew, Brent; Lederer, Susan

2015-01-01

The Boeing HS-376 spin stabilized spacecraft was a popular design that was launched continuously into geosynchronous orbit starting in 1980, with the last launch occurring in 2003. Over 50 of the HS-376 buses were produced to fulfill a variety of different communication missions for countries all over the world. The design of the bus is easily approximated as a telescoping cylinder that is covered with solar cells and an Earth-facing antenna that is despun at the top of the cylinder. The similarity in design and the number of spacecraft launched over a long period of time make the HS-376 a prime target for studying the effects of solar weathering on solar panels as a function of time. A selection of primarily non-operational HS-376 spacecraft launched over a 20-year time period were observed using the United Kingdom Infrared Telescope on Mauna Kea and multi-band, near-infrared photometry produced. Each spacecraft was observed for an entire night cycling through ZYJHK filters and time-varying colors produced to compare near-infrared color as a function of launch date. The resulting analysis shown here may help in the future to set launch date constraints on the parent object of unidentified debris objects or other unknown spacecraft.

Experiments study on attitude coupling control method for flexible spacecraft

Science.gov (United States)

Wang, Jie; Li, Dongxu

2018-06-01

High pointing accuracy and stabilization are significant for spacecrafts to carry out Earth observing, laser communication and space exploration missions. However, when a spacecraft undergoes large angle maneuver, the excited elastic oscillation of flexible appendages, for instance, solar wing and onboard antenna, would downgrade the performance of the spacecraft platform. This paper proposes a coupling control method, which synthesizes the adaptive sliding mode controller and the positive position feedback (PPF) controller, to control the attitude and suppress the elastic vibration simultaneously. Because of its prominent performance for attitude tracking and stabilization, the proposed method is capable of slewing the flexible spacecraft with a large angle. Also, the method is robust to parametric uncertainties of the spacecraft model. Numerical simulations are carried out with a hub-plate system which undergoes a single-axis attitude maneuver. An attitude control testbed for the flexible spacecraft is established and experiments are conducted to validate the coupling control method. Both numerical and experimental results demonstrate that the method discussed above can effectively decrease the stabilization time and improve the attitude accuracy of the flexible spacecraft.

Precision Closed-Loop Orbital Maneuvering System Design and Performance for the Magnetospheric Multi-Scale Mission (MMS) Formation

Science.gov (United States)

Chai, Dean; Queen, Steve; Placanica, Sam

2015-01-01

NASA's Magnetospheric Multi-Scale (MMS) mission successfully launched on March 13, 2015 (UTC) consists of four identically instrumented spin-stabilized observatories that function as a constellation to study magnetic reconnection in space. The need to maintain sufficiently accurate spatial and temporal formation resolution of the observatories must be balanced against the logistical constraints of executing overly-frequent maneuvers on a small fleet of spacecraft. These two considerations make for an extremely challenging maneuver design problem. This paper focuses on the design elements of a 6-DOF spacecraft attitude control and maneuvering system capable of delivering the high-precision adjustments required by the constellation designers---specifically, the design, implementation, and on-orbit performance of the closed-loop formation-class maneuvers that include initialization, maintenance, and re-sizing. The maneuvering control system flown on MMS utilizes a micro-gravity resolution accelerometer sampled at a high rate in order to achieve closed-loop velocity tracking of an inertial target with arc-minute directional and millimeter-per-second magnitude accuracy. This paper summarizes the techniques used for correcting bias drift, sensor-head offsets, and centripetal aliasing in the acceleration measurements. It also discusses the on-board pre-maneuver calibration and compensation algorithms as well as the implementation of the post-maneuver attitude adjustments.

Flexible spacecraft dynamics, control and guidance technologies by giovanni campolo

CERN Document Server

Mazzini, Leonardo

2016-01-01

This book is an up-to-date compendium on spacecraft attitude and orbit control (AOC) that offers a systematic and complete treatment of the subject with the aim of imparting the theoretical and practical knowledge that is required by designers, engineers, and researchers. After an introduction on the kinematics of the flexible and agile space vehicles, the modern architecture and functions of an AOC system are described and the main AOC modes reviewed with possible design solutions and examples. The dynamics of the flexible body in space are then considered using an original Lagrangian approach suitable for the control applications of large space flexible structures. Subsequent chapters address optimal control theory, attitude control methods, and orbit control applications, including the optimal orbital transfer with finite and infinite thrust. The theory is integrated with a description of current propulsion systems, with the focus especially on the new electric propulsion systems and state of the art senso...

Proposed gravity-gradient dynamics experiments in lunar orbit using the RAE-B spacecraft

Science.gov (United States)

Blanchard, D. L.; Walden, H.

1973-01-01

A series of seven gravity-gradient dynamics experiments is proposed utilizing the Radio Astronomy Explorer (RAE-B) spacecraft in lunar orbit. It is believed that none of the experiments will impair the spacecraft structure or adversely affect the continuation of the scientific mission of the satellite. The first experiment is designed to investigate the spacecraft dynamical behavior in the absence of libration damper action and inertia. It requires stable gravity-gradient capture of the spacecraft in lunar orbit with small amplitude attitude librations as a prerequisite. Four subsequent experiments involve partial retraction, ultimately followed by full redeployment, of one or two of the 230-meter booms forming the lunar-directed Vee-antenna. These boom length change operations will induce moderate amplitude angular librations of the spacecraft.

Space Power Program, Instrumentation and Control System Architecture, Preconceptual Design, for Information

International Nuclear Information System (INIS)

JM Ross

2005-01-01

The purpose of this letter is to forward the Prometheus preconceptual Instrumentation and Control (I and C) system architecture (Enclosure (1)) to NR for information as part of the Prometheus closeout work. The preconceptual 1 and C system architecture was considered a key planning document for development of the I and C system for Project Prometheus. This architecture was intended to set the technical approach for the entire I and C system. It defines interfaces to other spacecraft systems, defines hardware blocks for future development, and provides a basis for accurate cost and schedule estimates. Since the system requirements are not known at this time, it was anticipated that the architecture would evolve as the design of the reactor module was matured

An Empirical Comparison between Two Recursive Filters for Attitude and Rate Estimation of Spinning Spacecraft

Science.gov (United States)

Harman, Richard R.

2006-01-01

The advantages of inducing a constant spin rate on a spacecraft are well known. A variety of science missions have used this technique as a relatively low cost method for conducting science. Starting in the late 1970s, NASA focused on building spacecraft using 3-axis control as opposed to the single-axis control mentioned above. Considerable effort was expended toward sensor and control system development, as well as the development of ground systems to independently process the data. As a result, spinning spacecraft development and their resulting ground system development stagnated. In the 1990s, shrinking budgets made spinning spacecraft an attractive option for science. The attitude requirements for recent spinning spacecraft are more stringent and the ground systems must be enhanced in order to provide the necessary attitude estimation accuracy. Since spinning spacecraft (SC) typically have no gyroscopes for measuring attitude rate, any new estimator would need to rely on the spacecraft dynamics equations. One estimation technique that utilized the SC dynamics and has been used successfully in 3-axis gyro-less spacecraft ground systems is the pseudo-linear Kalman filter algorithm. Consequently, a pseudo-linear Kalman filter has been developed which directly estimates the spacecraft attitude quaternion and rate for a spinning SC. Recently, a filter using Markley variables was developed specifically for spinning spacecraft. The pseudo-linear Kalman filter has the advantage of being easier to implement but estimates the quaternion which, due to the relatively high spinning rate, changes rapidly for a spinning spacecraft. The Markley variable filter is more complicated to implement but, being based on the SC angular momentum, estimates parameters which vary slowly. This paper presents a comparison of the performance of these two filters. Monte-Carlo simulation runs will be presented which demonstrate the advantages and disadvantages of both filters.

Event-triggered attitude control of spacecraft

Science.gov (United States)

Wu, Baolin; Shen, Qiang; Cao, Xibin

2018-02-01

The problem of spacecraft attitude stabilization control system with limited communication and external disturbances is investigated based on an event-triggered control scheme. In the proposed scheme, information of attitude and control torque only need to be transmitted at some discrete triggered times when a defined measurement error exceeds a state-dependent threshold. The proposed control scheme not only guarantees that spacecraft attitude control errors converge toward a small invariant set containing the origin, but also ensures that there is no accumulation of triggering instants. The performance of the proposed control scheme is demonstrated through numerical simulation.

A report on SHARP (Spacecraft Health Automated Reasoning Prototype) and the Voyager Neptune encounter

Science.gov (United States)

Martin, R. G. (Editor); Atkinson, D. J.; James, M. L.; Lawson, D. L.; Porta, H. J.

1990-01-01

The development and application of the Spacecraft Health Automated Reasoning Prototype (SHARP) for the operations of the telecommunications systems and link analysis functions in Voyager mission operations are presented. An overview is provided of the design and functional description of the SHARP system as it was applied to Voyager. Some of the current problems and motivations for automation in real-time mission operations are discussed, as are the specific solutions that SHARP provides. The application of SHARP to Voyager telecommunications had the goal of being a proof-of-capability demonstration of artificial intelligence as applied to the problem of real-time monitoring functions in planetary mission operations. AS part of achieving this central goal, the SHARP application effort was also required to address the issue of the design of an appropriate software system architecture for a ground-based, highly automated spacecraft monitoring system for mission operations, including methods for: (1) embedding a knowledge-based expert system for fault detection, isolation, and recovery within this architecture; (2) acquiring, managing, and fusing the multiple sources of information used by operations personnel; and (3) providing information-rich displays to human operators who need to exercise the capabilities of the automated system. In this regard, SHARP has provided an excellent example of how advanced artificial intelligence techniques can be smoothly integrated with a variety of conventionally programmed software modules, as well as guidance and solutions for many questions about automation in mission operations.

Control system design for the constellation acquisition phase of the LISA mission

Energy Technology Data Exchange (ETDEWEB)

Cirillo, Francesca; Gath, Peter F, E-mail: francesca.cirillo@astrium.eads.ne, E-mail: peter.gath@astrium.eads.ne [Astrium GmbH Satellites, 88039 Friedrichshafen (Germany)

2009-03-01

The objective of the constellation acquisition phase for the LISA mission is to establish the three laser links between the three spacecraft of the LISA constellation so that the interferometric measurements for the science experiment can commence. The laser beam acquisition for LISA is extremely challenging given the 5 million km distance between the spacecraft, the inherent limits of the attitude sensors accuracy, the orbit determination accuracy issues and the time required to phase-lock the incoming and outgoing laser signals. This paper presents the design of the control system for the acquisition phase of the LISA constellation: the acquisition operational procedure is outlined, guidance laws are defined together with the Gyro Mode attitude control principle, which implements a Kalman filter for disturbances rejection purposes. Constellation-wide non-linear simulations demonstrate that the LISA constellation acquisition phase is feasible by means of the proposed control strategy.

Control system design for the constellation acquisition phase of the LISA mission

International Nuclear Information System (INIS)

Cirillo, Francesca; Gath, Peter F

2009-01-01

The objective of the constellation acquisition phase for the LISA mission is to establish the three laser links between the three spacecraft of the LISA constellation so that the interferometric measurements for the science experiment can commence. The laser beam acquisition for LISA is extremely challenging given the 5 million km distance between the spacecraft, the inherent limits of the attitude sensors accuracy, the orbit determination accuracy issues and the time required to phase-lock the incoming and outgoing laser signals. This paper presents the design of the control system for the acquisition phase of the LISA constellation: the acquisition operational procedure is outlined, guidance laws are defined together with the Gyro Mode attitude control principle, which implements a Kalman filter for disturbances rejection purposes. Constellation-wide non-linear simulations demonstrate that the LISA constellation acquisition phase is feasible by means of the proposed control strategy.

Space Power Program, Instrumentation and Control System Architecture, Pre-conceptual Design, for Information

Energy Technology Data Exchange (ETDEWEB)

JM Ross

2005-10-20

The purpose of this letter is to forward the Prometheus preconceptual Instrumentation and Control (I&C) system architecture (Enclosure (1)) to NR for information as part of the Prometheus closeout work. The preconceptual 1&C system architecture was considered a key planning document for development of the I&C system for Project Prometheus. This architecture was intended to set the technical approach for the entire I&C system. It defines interfaces to other spacecraft systems, defines hardware blocks for future development, and provides a basis for accurate cost and schedule estimates. Since the system requirements are not known at this time, it was anticipated that the architecture would evolve as the design of the reactor module was matured.

A Research on the Electrical Test Fault Diagnostic and Data Mining of a Manned Spacecraft

Directory of Open Access Journals (Sweden)

Yang Feng

2017-01-01

Full Text Available The paper introduces the modeling method and modeling tool for the fault diagnosis of manned spacecraft, the multi-signal flow graph model of a manned space equipment was established using this method; the framework of the fault detection and diagnosis system of manned spacecraft is proposed, the function of ground system and function of the spacecraft are clearly defined. The structure of the functional module is given separately; finally, the tool builds the fault detection and diagnosis system, the application of fault diagnosis method for manned spacecraft is used for reference.

Determination of Realistic Fire Scenarios in Spacecraft

Science.gov (United States)

Dietrich, Daniel L.; Ruff, Gary A.; Urban, David

2013-01-01

This paper expands on previous work that examined how large a fire a crew member could successfully survive and extinguish in the confines of a spacecraft. The hazards to the crew and equipment during an accidental fire include excessive pressure rise resulting in a catastrophic rupture of the vehicle skin, excessive temperatures that burn or incapacitate the crew (due to hyperthermia), carbon dioxide build-up or accumulation of other combustion products (e.g. carbon monoxide). The previous work introduced a simplified model that treated the fire primarily as a source of heat and combustion products and sink for oxygen prescribed (input to the model) based on terrestrial standards. The model further treated the spacecraft as a closed system with no capability to vent to the vacuum of space. The model in the present work extends this analysis to more realistically treat the pressure relief system(s) of the spacecraft, include more combustion products (e.g. HF) in the analysis and attempt to predict the fire spread and limiting fire size (based on knowledge of terrestrial fires and the known characteristics of microgravity fires) rather than prescribe them in the analysis. Including the characteristics of vehicle pressure relief systems has a dramatic mitigating effect by eliminating vehicle overpressure for all but very large fires and reducing average gas-phase temperatures.

Concurrent System Engineering and Risk Reduction for Dual-Band (RF/optical) Spacecraft Communications

Science.gov (United States)

Fielhauer, Karl, B.; Boone, Bradley, G.; Raible, Daniel, E.

2012-01-01

This paper describes a system engineering approach to examining the potential for combining elements of a deep-space RF and optical communications payload, for the purpose of reducing the size, weight and power burden on the spacecraft and the mission. Figures of merit and analytical methodologies are discussed to conduct trade studies, and several potential technology integration strategies are presented. Finally, the NASA Integrated Radio and Optical Communications (iROC) project is described, which directly addresses the combined RF and optical approach.

Rotational Motion Control of a Spacecraft

DEFF Research Database (Denmark)

Wisniewski, Rafal; Kulczycki, P.

2001-01-01

The paper adopts the energy shaping method to control of rotational motion. A global representation of the rigid body motion is given in the canonical form by a quaternion and its conjugate momenta. A general method for motion control on a cotangent bundle to the 3-sphere is suggested. The design...... algorithm is validated for three-axis spacecraft attitude control...

Rotational motion control of a spacecraft

DEFF Research Database (Denmark)

Wisniewski, Rafal; Kulczycki, P.

2003-01-01

The paper adopts the energy shaping method to control of rotational motion. A global representation of the rigid body motion is given in the canonical form by a quaternion and its conjugate momenta. A general method for motion control on a cotangent bundle to the 3-sphere is suggested. The design...... algorithm is validated for three-axis spacecraft attitude control. Udgivelsesdato: APR...

A Study of Learning Curve Impact on Three Identical Small Spacecraft

Science.gov (United States)

Chen, Guangming; McLennan, Douglas D.

2003-01-01

With an eye to the future strategic needs of NASA, the New Millennium Program is funding the Space Technology 5 (ST-5) project to address the future needs in the area of small satellites in constellation missions. The ST-5 project, being developed at Goddard Space Flight Center, involves the development and simultaneous launch of three small, 20-kilogram-class spacecraft. ST-5 is only a test drive and future NASA science missions may call for fleets of spacecraft containing tens of smart and capable satellites in an intelligent constellation. The objective of ST-5 project is to develop three such pioneering small spacecraft for flight validation of several critical new technologies. The ST-5 project team at Goddard Space Flight Center has completed the spacecraft design, is now building and testing the three flight units. The launch readiness date (LRD) is in December 2005. A critical part of ST-5 mission is to prove that it is possible to build these small but capable spacecraft with recurring cost low enough to make future NASA s multi- spacecraft constellation missions viable from a cost standpoint.

Video-Game-Like Engine for Depicting Spacecraft Trajectories

Science.gov (United States)

Upchurch, Paul R.

2009-01-01

GoView is a video-game-like software engine, written in the C and C++ computing languages, that enables real-time, three-dimensional (3D)-appearing visual representation of spacecraft and trajectories (1) from any perspective; (2) at any spatial scale from spacecraft to Solar-system dimensions; (3) in user-selectable time scales; (4) in the past, present, and/or future; (5) with varying speeds; and (6) forward or backward in time. GoView constructs an interactive 3D world by use of spacecraft-mission data from pre-existing engineering software tools. GoView can also be used to produce distributable application programs for depicting NASA orbital missions on personal computers running the Windows XP, Mac OsX, and Linux operating systems. GoView enables seamless rendering of Cartesian coordinate spaces with programmable graphics hardware, whereas prior programs for depicting spacecraft trajectories variously require non-Cartesian coordinates and/or are not compatible with programmable hardware. GoView incorporates an algorithm for nonlinear interpolation between arbitrary reference frames, whereas the prior programs are restricted to special classes of inertial and non-inertial reference frames. Finally, whereas the prior programs present complex user interfaces requiring hours of training, the GoView interface provides guidance, enabling use without any training.

TTEthernet for Integrated Spacecraft Networks

Science.gov (United States)

Loveless, Andrew

2015-01-01

Aerospace projects have traditionally employed federated avionics architectures, in which each computer system is designed to perform one specific function (e.g. navigation). There are obvious downsides to this approach, including excessive weight (from so much computing hardware), and inefficient processor utilization (since modern processors are capable of performing multiple tasks). There has therefore been a push for integrated modular avionics (IMA), in which common computing platforms can be leveraged for different purposes. This consolidation of multiple vehicle functions to shared computing platforms can significantly reduce spacecraft cost, weight, and design complexity. However, the application of IMA principles introduces significant challenges, as the data network must accommodate traffic of mixed criticality and performance levels - potentially all related to the same shared computer hardware. Because individual network technologies are rarely so competent, the development of truly integrated network architectures often proves unreasonable. Several different types of networks are utilized - each suited to support a specific vehicle function. Critical functions are typically driven by precise timing loops, requiring networks with strict guarantees regarding message latency (i.e. determinism) and fault-tolerance. Alternatively, non-critical systems generally employ data networks prioritizing flexibility and high performance over reliable operation. Switched Ethernet has seen widespread success filling this role in terrestrial applications. Its high speed, flexibility, and the availability of inexpensive commercial off-the-shelf (COTS) components make it desirable for inclusion in spacecraft platforms. Basic Ethernet configurations have been incorporated into several preexisting aerospace projects, including both the Space Shuttle and International Space Station (ISS). However, classical switched Ethernet cannot provide the high level of network

Design and Development of the Solar Dynamics Observatory (SDO) Electrical Power System

Science.gov (United States)

Denney, Keys; Burns, Michael; Kercheval, Bradford

2009-01-01

The SDO spacecraft was designed to help us understand the Sun's influence on Earth and Near-Earth space by studying the solar atmosphere on small scales of space and time and in many wavelengths simultaneously. It will perform its operations in a geosynchronous orbit of the earth. This paper will present background on the SDO mission, an overview of the design and development activities associated specifically with the SDO electrical power system (EPS), as well as the major driving requirements behind the mission design. The primary coverage of the paper will be devoted to some of the challenges faced during the design and development phase. This will include the challenges associated with development of a compatible CompactPCI (cPCI) interface within the Power System Electronics (PSE) in order to utilize a "common" processor card, implementation of new solid state power controllers (SSPC) for primary load distribution switching and over current protection in the PSE, and the design approach adopted to meet single fault tolerance requirements for all of the SDO EPS functions.

Resilience Engineering in Critical Long Term Aerospace Software Systems: A New Approach to Spacecraft Software Safety

Science.gov (United States)

Dulo, D. A.

Safety critical software systems permeate spacecraft, and in a long term venture like a starship would be pervasive in every system of the spacecraft. Yet software failure today continues to plague both the systems and the organizations that develop them resulting in the loss of life, time, money, and valuable system platforms. A starship cannot afford this type of software failure in long journeys away from home. A single software failure could have catastrophic results for the spaceship and the crew onboard. This paper will offer a new approach to developing safe reliable software systems through focusing not on the traditional safety/reliability engineering paradigms but rather by focusing on a new paradigm: Resilience and Failure Obviation Engineering. The foremost objective of this approach is the obviation of failure, coupled with the ability of a software system to prevent or adapt to complex changing conditions in real time as a safety valve should failure occur to ensure safe system continuity. Through this approach, safety is ensured through foresight to anticipate failure and to adapt to risk in real time before failure occurs. In a starship, this type of software engineering is vital. Through software developed in a resilient manner, a starship would have reduced or eliminated software failure, and would have the ability to rapidly adapt should a software system become unstable or unsafe. As a result, long term software safety, reliability, and resilience would be present for a successful long term starship mission.

Low-Frequency Gravitational Wave Searches Using Spacecraft Doppler Tracking

Directory of Open Access Journals (Sweden)

Armstrong J. W.

2006-01-01

Full Text Available This paper discusses spacecraft Doppler tracking, the current-generation detector technology used in the low-frequency (~millihertz gravitational wave band. In the Doppler method the earth and a distant spacecraft act as free test masses with a ground-based precision Doppler tracking system continuously monitoring the earth-spacecraft relative dimensionless velocity $2 Delta v/c = Delta u/ u_0$, where $Delta u$ is the Doppler shift and $ u_0$ is the radio link carrier frequency. A gravitational wave having strain amplitude $h$ incident on the earth-spacecraft system causes perturbations of order $h$ in the time series of $Delta u/ u_0$. Unlike other detectors, the ~1-10 AU earth-spacecraft separation makes the detector large compared with millihertz-band gravitational wavelengths, and thus times-of-flight of signals and radio waves through the apparatus are important. A burst signal, for example, is time-resolved into a characteristic signature: three discrete events in the Doppler time series. I discuss here the principles of operation of this detector (emphasizing transfer functions of gravitational wave signals and the principal noises to the Doppler time series, some data analysis techniques, experiments to date, and illustrations of sensitivity and current detector performance. I conclude with a discussion of how gravitational wave sensitivity can be improved in the low-frequency band.

ADRC for spacecraft attitude and position synchronization in libration point orbits

Science.gov (United States)

Gao, Chen; Yuan, Jianping; Zhao, Yakun

2018-04-01

This paper addresses the problem of spacecraft attitude and position synchronization in libration point orbits between a leader and a follower. Using dual quaternion, the dimensionless relative coupled dynamical model is derived considering computation efficiency and accuracy. Then a model-independent dimensionless cascade pose-feedback active disturbance rejection controller is designed to spacecraft attitude and position tracking control problems considering parameter uncertainties and external disturbances. Numerical simulations for the final approach phase in spacecraft rendezvous and docking and formation flying are done, and the results show high-precision tracking errors and satisfactory convergent rates under bounded control torque and force which validate the proposed approach.

Astronaut L. Gordon Cooper is assisted into his spacecraft for tests

Science.gov (United States)

1963-01-01

NASA and McDonnell Aircraft Corp. spacecraft technicians assist Astronaut L. Gordon Cooper into his spacecraft prior to undergoing tests in the altitude chamber. These tests are used to determine the operating characteristcs of the overall environmental control system.

Preliminary results on ocean dynamics from Skylab and their implications for future spacecraft

Science.gov (United States)

Hayes, J.; Pierson, W. J.; Cardone, V. J.

1975-01-01

The instrument aboard Skylab designated S193 - a combined passive and active microwave radar system acting as a radiometer, scatterometer, and altimeter - is used to measure the surface vector wind speeds in the planetary boundary layer over the oceans. Preliminary results corroborate the hypothesis that sea surface winds in the planetary boundary layer can be determined from satellite data. Future spacecraft plans for measuring a geoid with an accuracy up to 10 cm are discussed.

Towards a standardized grasping and refuelling on-orbit servicing for geo spacecraft

Science.gov (United States)

Medina, Alberto; Tomassini, Angelo; Suatoni, Matteo; Avilés, Marcos; Solway, Nick; Coxhill, Ian; Paraskevas, Iosif S.; Rekleitis, Georgios; Papadopoulos, Evangelos; Krenn, Rainer; Brito, André; Sabbatinelli, Beatrice; Wollenhaupt, Birk; Vidal, Christian; Aziz, Sarmad; Visentin, Gianfranco

2017-05-01

Exploitation of space must benefit from the latest advances in robotics. On-orbit servicing is a clear candidate for the application of autonomous rendezvous and docking mechanisms. However, during the last three decades most of the trials took place combining extravehicular activities (EVAs) with telemanipulated robotic arms. The European Space Agency (ESA) considers that grasping and refuelling are promising near-mid-term capabilities that could be performed by servicing spacecraft. Minimal add-ons on spacecraft to enhance their serviceability may protect them for a changing future in which satellite servicing may become mainstream. ESA aims to conceive and promote standard refuelling provisions that can be installed in present and future European commercial geostationary orbit (GEO) satellite platforms and scientific spacecraft. For this purpose ESA has started the ASSIST activity addressing the analysis, design and validation of internal provisions (such as modifications to fuel, gas, electrical and data architecture to allow servicing) and external provisions (such as integrated berthing fixtures with peripheral electrical, gas, liquid connectors, leak check systems and corresponding optical and radio markers for cooperative rendezvous and docking). This refuelling approach is being agreed with European industry (OHB, Thales Alenia Space) and expected to be consolidated with European commercial operators as a first step to become an international standard; this approach is also being considered for on-orbit servicing spacecraft, such as the SpaceTug, by Airbus DS. This paper describes in detail the operational means, structure, geometry and accommodation of the system. Internal and external provisions will be designed with the minimum possible impact on the current architecture of GEO satellites without introducing additional risks in the development and commissioning of the satellite. End-effector and berthing fixtures are being designed in the range of few

Nonlinear Slewing Spacecraft Control Based on Exergy, Power Flow, and Static and Dynamic Stability

Science.gov (United States)

Robinett, Rush D.; Wilson, David G.

2009-10-01

This paper presents a new nonlinear control methodology for slewing spacecraft, which provides both necessary and sufficient conditions for stability by identifying the stability boundaries, rigid body modes, and limit cycles. Conservative Hamiltonian system concepts, which are equivalent to static stability of airplanes, are used to find and deal with the static stability boundaries: rigid body modes. The application of exergy and entropy thermodynamic concepts to the work-rate principle provides a natural partitioning through the second law of thermodynamics of power flows into exergy generator, dissipator, and storage for Hamiltonian systems that is employed to find the dynamic stability boundaries: limit cycles. This partitioning process enables the control system designer to directly evaluate and enhance the stability and performance of the system by balancing the power flowing into versus the power dissipated within the system subject to the Hamiltonian surface (power storage). Relationships are developed between exergy, power flow, static and dynamic stability, and Lyapunov analysis. The methodology is demonstrated with two illustrative examples: (1) a nonlinear oscillator with sinusoidal damping and (2) a multi-input-multi-output three-axis slewing spacecraft that employs proportional-integral-derivative tracking control with numerical simulation results.

Spacecraft contamination programs within the Air Force Systems Command Laboratories

Science.gov (United States)

Murad, Edmond

1990-01-01

Spacecraft contamination programs exist in five independent AFSC organizations: Geophysics Laboratory (GL), Arnold Engineering and Development Center (AEDC), Rome Air Development Center (RADC/OSCE), Wright Research and Development Center (MLBT), Armament Laboratory (ATL/SAI), and Space Systems Division (SSD/OL-AW). In addition, a sizable program exists at Aerospace Corp. These programs are complementary, each effort addressing a specific area of expertise: GL's effort is aimed at addressing the effects of on-orbit contamination; AEDC's effort is aimed at ground simulation and measurement of optical contamination; RADC's effort addresses the accumulation, measurement, and removal of contamination on large optics; MLBT's effort is aimed at understanding the effect of contamination on materials; ATL's effort is aimed at understanding the effect of plume contamination on systems; SSD's effort is confined to the integration of some contamination experiments sponsored by SSD/CLT; and Aerospace Corp.'s effort is aimed at supporting the needs of the using System Program Offices (SPO) in specific areas, such as contamination during ground handling, ascent phase, laboratory measurements aimed at understanding on-orbit contamination, and mass loss and mass gain in on-orbit operations. These programs are described in some detail, with emphasis on GL's program.

Simulator Facility for Attitude Control and Energy Storage of Spacecraft

National Research Council Canada - National Science Library

Tsiotras, Panagiotis

2002-01-01

This report concerns a designed and built experimental facility that will allow the conduction of experiments for validating advanced attitude control algorithms for spacecraft in a weightless environment...

Human factors issues for interstellar spacecraft

Science.gov (United States)

Cohen, Marc M.; Brody, Adam R.

1991-01-01

Developments in research on space human factors are reviewed in the context of a self-sustaining interstellar spacecraft based on the notion of traveling space settlements. Assumptions about interstellar travel are set forth addressing costs, mission durations, and the need for multigenerational space colonies. The model of human motivation by Maslow (1970) is examined and directly related to the design of space habitat architecture. Human-factors technology issues encompass the human-machine interface, crew selection and training, and the development of spaceship infrastructure during transtellar flight. A scenario for feasible instellar travel is based on a speed of 0.5c, a timeframe of about 100 yr, and an expandable multigenerational crew of about 100 members. Crew training is identified as a critical human-factors issue requiring the development of perceptual and cognitive aids such as expert systems and virtual reality.

Power Subsystem Design for Tiangong-1 Target Spacecraft%天宫一号目标飞行器电源分系统设计

Institute of Scientific and Technical Information of China (English)

陈启忠; 马季军; 王娜; 黄应春; 黄峥; 王振绪

2011-01-01

The sketch and main performances of the power subsystem of Tiangong-1 target spacecraft were introduced in this paper. The key technologies in domestic such as the bus with the voltage 100 V applied on low orbit spacecraft, large-scale nickel-metal hybrid batteries, triple-junction gallium arsenide solar cells and semi-rigid solar wings were given out. The main job of the high voltage device system establishing, the semi-rigid solar dynamics and space environment design and verification, the life time and reliability of the nickel-metal hybrid battery, and the reliability and safety of the high voltage power system were reviewed. The operation on orbit was given out. The research of the power subsystem of Tiangong-1 target spacecraft would establish the foundation for the China＇s next space technology.%介绍了天宫一号（TG1）目标飞行器电源分系统的组成和主要技术指标。分析了国内在低轨飞行器上采用100V高压母线、大批量使用国产氢镍电池、三结砷化镓太阳电池片和半刚性基板等关键技术。回顾了电源分系统研制过程中高电压元器件体系建立、半刚性帆板力学及空间环境设计与验证、氢镍电池在轨寿命和可靠性研究，以及高压电源系统可靠性及安全性研究等主要工作。给出了在轨运行情况。TG-1目标飞行器电源分系统的研制为我国后续空间技术的发展打下了基础。

Designing Spacecraft and Mission Operations Plans to Meet Flight Crew Radiation Dose Requirements: Why is this an "Epic Challenge" for Long-Term Manned Interplanetary Flight

Science.gov (United States)

Koontz, Steven

2012-01-01

Outline of presentation: (1) Radiation Shielding Concepts and Performance - Galactic Cosmic Rays (GCRs) (1a) Some general considerations (1b) Galactic Cosmic Rays (2)GCR Shielding I: What material should I use and how much do I need? (2a) GCR shielding materials design and verification (2b) Spacecraft materials point dose cosmic ray shielding performance - hydrogen content and atomic number (2c) Accelerator point dose materials testing (2d) Material ranking and selection guidelines (2e) Development directions and return on investment (point dose metric) (2f) Secondary particle showers in the human body (2f-1) limited return of investment for low-Z, high-hydrogen content materials (3) GCR shielding II: How much will it cost? (3a) Spacecraft design and verification for mission radiation dose to the crew (3b) Habitat volume, shielding areal density, total weight, and launch cost for two habitat volumes (3c) It's All about the Money - Historical NASA budgets and budget limits (4) So, what can I do about all this? (4a) Program Design Architecture Trade Space (4b) The Vehicle Design Trade Space (4c) Some Near Term Recommendations

Mechanical design of SERT 2 thruster system

Science.gov (United States)

Zavesky, R. J.; Hurst, E. B.

1972-01-01

The mechanical design of the mercury bombardment thruster that was tested on SERT is described. The report shows how the structural, thermal, electrical, material compatibility, and neutral mercury coating considerations affected the design and integration of the subsystems and components. The SERT 2 spacecraft with two thrusters was launched on February 3, 1970. One thruster operated for 3782 hours and the other for 2011 hours. A high voltage short resulting from buildup of loose eroded material was believed to be the cause of failure.

Towards the Development of a Unified Distributed Date System for L1 Spacecraft

Science.gov (United States)

Lazarus, Alan J.; Kasper, Justin C.

2005-01-01

The purpose of this grant, 'Towards the Development of a Unified Distributed Data System for L1 Spacecraft', is to take the initial steps towards the development of a data distribution mechanism for making in-situ measurements more easily accessible to the scientific community. Our obligations as subcontractors to this grant are to add our Faraday Cup plasma data to this initial study and to contribute to the design of a general data distribution system. The year 1 objectives of the overall project as stated in the GSFC proposal are: 1) Both the rsync and Perl based data exchange tools will be fully developed and tested in our mixed, Unix, VMS, Windows and Mac OS X data service environment. Based on the performance comparisons, one will be selected and fully deployed. Continuous data exchange between all L1 solar wind monitors initiated. 2) Data version metadata will be agreed upon, fully documented, and deployed on our data sites. 3) The first version of the data description rules, encoded in a XML Schema, will be finalized. 4) Preliminary set of library routines will be collected, documentation standards and formats agreed on, and desirable routines that have not been implemented identified and assigned. 5) ViSBARD test site implemented to independently validate data mirroring procedures. The specific MIT tasks over the duration of this project are the following: a) implement mirroring service for WIND plasma data b) participate in XML Schema development c) contribute toward routine library.