System design of the Pioneer Venus spacecraft. Volume 9: Attitude control/mechanisms subsystems studies

NASA Technical Reports Server (NTRS)

Neil, A. L.

1973-01-01

The Pioneer Venus mission study was conducted for a probe spacecraft and an orbiter spacecraft to be launched by either a Thor/Delta or an Atlas/Centaur launch vehicle. Both spacecraft are spin stabilized. The spin speed is controlled by ground commands to as low as 5 rpm for science instrument scanning on the orbiter and as high as 71 rpm for small probes released from the probe bus. A major objective in the design of the attitude control and mechanism subsystem (ACMS) was to provide, in the interest of costs, maximum commonality of the elements between the probe bus and orbiter spacecraft configurations. This design study was made considering the use of either launch vehicle. The basic functional requirements of the ACMS are derived from spin axis pointing and spin speed control requirements implicit in the acquisition, cruise, encounter and orbital phases of the Pioneer Venus missions.

Bioresearch Module Design Definition and Space Shuttle Vehicle Integration Study. Volume 1: Basic Report

NASA Technical Reports Server (NTRS)

Lang, A. L., Jr.

1971-01-01

Preliminary designs of the Bioexplorer spacecraft, developed in an earlier study program, are analyzed and updated to conform to a new specification which includes use of both the Scout and the space shuttle vehicle for launch. The updated spacecraft is referred to as bioresearch module. It is capable of supporting a variety of small biological experiments in near-earth and highly elliptical earth orbits. The baseline spacecraft design is compatible with the Scout launch vehicle. Inboard profile drawings, weight statements, interface drawings, and spacecraft parts and aerospace ground equipment lists are provided to document the design. The baseline design was analyzed to determine the design and cost impact of a set of optional features. These include reduced experiment power and thermal load, addition of an experiment television monitor, and replacement of VHF with S-band communications. The impact of these options on power required, weight change and cost is defined.

Study on light weight design of truss structures of spacecrafts

NASA Astrophysics Data System (ADS)

Zeng, Fuming; Yang, Jianzhong; Wang, Jian

2015-08-01

Truss structure is usually adopted as the main structure form for spacecrafts due to its high efficiency in supporting concentrated loads. Light-weight design is now becoming the primary concern during conceptual design of spacecrafts. Implementation of light-weight design on truss structure always goes through three processes: topology optimization, size optimization and composites optimization. During each optimization process, appropriate algorithm such as the traditional optimality criterion method, mathematical programming method and the intelligent algorithms which simulate the growth and evolution processes in nature will be selected. According to the practical processes and algorithms, combined with engineering practice and commercial software, summary is made for the implementation of light-weight design on truss structure for spacecrafts.

System design of the Pioneer Venus spacecraft. Volume 4: Probe bus and orbiter spacecraft vehicle studies

NASA Technical Reports Server (NTRS)

Bozajian, J. M.

1973-01-01

The requirements, trades, and design descriptions for the probe bus and orbiter spacecraft configurations, structure, thermal control, and harness are defined. Designs are developed for Thor/Delta and Atlas/Centaur launch vehicles with the latter selected as the final baseline. The major issues examined in achieving the baseline design are tabulated. The importance of spin axis orientation because of the effect on science experiments and earth communications is stressed.

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

NASA Technical Reports Server (NTRS)

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.

Study of ballistic mode Mercury Orbiter missions. Volume 1: Summary report

NASA Technical Reports Server (NTRS)

Hollenbeck, G. R.

1973-01-01

A summary is given of the scope, approach, and major results of the study of ballistic mode Mercury orbit missions (the Mariner Venus-Mercury spacecraft). The performance potential of ballistic flight mode is presented along with a study of alternate flight techniques. Orbit selection considerations are discussed in terms of the thermal environment of Mercury. Orbiter science experiments are summarized. Technology assessments were conducted for major subsystems appropriate to spin-stabilized and three-axis-stabilized spacecraft designs. Conclusions from this study are: ballistic mode Mercury orbiter missions offer adequate performance for effective follow-up of the MVM'73 science findings; the existing and programmed technology base is adequate for implementation of Mercury orbit spacecraft design; and when pending MVM flyby has been accomplished and the results analyzed, the data base will be adequate to support detailed orbiter spacecraft design efforts.

Vulnerability of manned spacecraft to crew loss from orbital debris penetration

NASA Technical Reports Server (NTRS)

Williamsen, J. E.

1994-01-01

Orbital debris growth threatens the survival of spacecraft systems from impact-induced failures. Whereas the probability of debris impact and spacecraft penetration may currently be calculated, another parameter of great interest to safety engineers is the probability that debris penetration will cause actual spacecraft or crew loss. Quantifying the likelihood of crew loss following a penetration allows spacecraft designers to identify those design features and crew operational protocols that offer the highest improvement in crew safety for available resources. Within this study, a manned spacecraft crew survivability (MSCSurv) computer model is developed that quantifies the conditional probability of losing one or more crew members, P(sub loss/pen), following the remote likelihood of an orbital debris penetration into an eight module space station. Contributions to P(sub loss/pen) are quantified from three significant penetration-induced hazards: pressure wall rupture (explosive decompression), fragment-induced injury, and 'slow' depressurization. Sensitivity analyses are performed using alternate assumptions for hazard-generating functions, crew vulnerability thresholds, and selected spacecraft design and crew operations parameters. These results are then used to recommend modifications to the spacecraft design and expected crew operations that quantitatively increase crew safety from orbital debris impacts.

Modeling the fundamental characteristics and processes of the spacecraft functioning

NASA Technical Reports Server (NTRS)

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.

Conceptual spacecraft systems design and synthesis

NASA Technical Reports Server (NTRS)

Wright, R. L.; Deryder, D. D.; Ferebee, M. J., Jr.

1984-01-01

An interactive systems design and synthesis is performed on future spacecraft concepts using the Interactive Design and Evaluation of Advanced Systems (IDEAS) computer-aided design and analysis system. The capabilities and advantages of the systems-oriented interactive computer-aided design and analysis system are described. The synthesis of both large antenna and space station concepts, and space station evolutionary growth designs is demonstrated. The IDEAS program provides the user with both an interactive graphics and an interactive computing capability which consists of over 40 multidisciplinary synthesis and analysis modules. Thus, the user can create, analyze, and conduct parametric studies and modify earth-orbiting spacecraft designs (space stations, large antennas or platforms, and technologically advanced spacecraft) at an interactive terminal with relative ease. The IDEAS approach is useful during the conceptual design phase of advanced space missions when a multiplicity of parameters and concepts must be analyzed and evaluated in a cost-effective and timely manner.

Mars Aeronomy Explorer (MAX): Study Employing Distributed Micro-Spacecraft

NASA Technical Reports Server (NTRS)

Shotwell, Robert F.; Gray, Andrew A.; Illsley, Peter M.; Johnson, M.; Sherwood, Robert L.; Vozoff, M.; Ziemer, John K.

2005-01-01

An overview of a Mars Aeronomy Explorer (MAX) mission design study performed at NASA's Jet Propulsion Laboratory is presented herein. The mission design consists of ten micro-spacecraft orbiters launched on a Delta IV to Mars polar orbit to determine the spatial, diurnal and seasonal variation of the constituents of the Martian upper atmosphere and ionosphere over the course of one Martian year. The spacecraft are designed to allow penetration of the upper atmosphere to at least 90 km. This property coupled with orbit precession will yield knowledge of the nature of the solar wind interaction with Mars, the influence of the Mars crustal magnetic field on ionospheric processes, and the measurement of present thermal and nonthermal escape rates of atmospheric constituents. The mission design incorporates alternative design paradigms that are more appropriate for-and in some cases motivate-distributed micro-spacecraft. These design paradigms are not defined by a simple set of rules, but rather a way of thinking about the function of instruments, mission reliability/risk, and cost in a systemic framework.

An Educational Multimedia Presentation on the Introduction to Spacecraft Charging

NASA Technical Reports Server (NTRS)

Lin, E.; dePayrebrune, M.

2004-01-01

Over the last few decades, significant knowledge has been gained in how to protect spacecraft from charging; however, the continuing technical advancement in the design and build of satellites requires on-going effort in the study of spacecraft charging. A situation that we have encountered is that not all satellite designers and builders are familiar with the problem of spacecraft charging. The design of a satellite involves many talented people with diverse backgrounds, ranging from manufacturing and assembly to engineering and program management. The complex design and build of a satellite system requires people with highly specialized skills such that cross-specialization is often not achievable. As a result, designers and builders of satellites are not usually familiar with the problems outside their specialization. This is also true for spacecraft charging. Not everyone is familiar with the definition of spacecraft charging and the damage that spacecraft charging can cause. Understanding the problem is an important first step in getting everyone involved in addressing the appropriate spacecraft charging issues during the satellite design and build phases. To address this important first step, an educational multimedia presentation has been created to inform the general engineering community about the basics of spacecraft charging. The content of this educational presentation is based on relevant published technical papers. The presentation was developed using Macromedia Flash. This software produces a more dynamic learning environment than a typical slide show , resulting in a more effective learning experience. The end result is that the viewer will have learned about the basics of spacecraft charging. This presentation is available to the public through our website, www.dplscience.com, free of charge. Viewers are encouraged to pass this presentation to colleagues within their own work environment. This paper describes the content of the multimedia presentation.

Lean spacecraft avionics trade study

NASA Technical Reports Server (NTRS)

Main, John A.

1994-01-01

Spacecraft design is generally an exercise in design trade-offs: fuel vs. weight, power vs. solar cell area, radiation exposure vs. shield weight, etc. Proper analysis of these trades is critical in the development of lightweight, efficient, 'lean' satellites. The modification of the launch plans for the Magnetosphere Imager (MI) to a Taurus launcher from the much more powerful Delta has forced a reduction in spacecraft weight availability into the mission orbit from 1300 kg to less than 500 kg. With weight now a driving factor it is imperative that the satellite design be extremely efficient and lean. The accuracy of engineering trades now takes on an added importance. An understanding of spacecraft subsystem interactions is critical in the development of a good spacecraft design, yet it is a challenge to define these interactions while the design is immature. This is currently an issue in the development of the preliminary design of the MI. The interaction and interfaces between this spacecraft and the instruments it carries are currently unclear since the mission instruments are still under development. It is imperative, however, to define these interfaces so that avionics requirements ideally suited to the mission's needs can be determined.

An autonomous satellite architecture integrating deliberative reasoning and behavioural intelligence

NASA Technical Reports Server (NTRS)

Lindley, Craig A.

1993-01-01

This paper describes a method for the design of autonomous spacecraft, based upon behavioral approaches to intelligent robotics. First, a number of previous spacecraft automation projects are reviewed. A methodology for the design of autonomous spacecraft is then presented, drawing upon both the European Space Agency technological center (ESTEC) automation and robotics methodology and the subsumption architecture for autonomous robots. A layered competency model for autonomous orbital spacecraft is proposed. A simple example of low level competencies and their interaction is presented in order to illustrate the methodology. Finally, the general principles adopted for the control hardware design of the AUSTRALIS-1 spacecraft are described. This system will provide an orbital experimental platform for spacecraft autonomy studies, supporting the exploration of different logical control models, different computational metaphors within the behavioral control framework, and different mappings from the logical control model to its physical implementation.

Mars Orbiter Study. Volume 2: Mission Design, Science Instrument Accommodation, Spacecraft Design

NASA Technical Reports Server (NTRS)

Drean, R.; Macpherson, D.; Steffy, D.; Vargas, T.; Shuman, B.; Anderson, K.; Richards, B.

1982-01-01

Spacecraft system and subsystem designs were developed at the conceptual level to perform either of two Mars Orbiter Missions, a Climatology Mission and an Aeronomy Mission. The objectives of these missions are to obtain and return data to increase knowledge of Mars.

Preliminary thermal design of the COLD-SAT spacecraft

NASA Technical Reports Server (NTRS)

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.

Interplanetary spacecraft design using solar electric propulsion

NASA Technical Reports Server (NTRS)

Duxbury, J. H.; Paul, G. M.

1974-01-01

Emphasis of the electric propulsion technology program is now on the application of solar electric propulsion to scientific missions. Candidate planetary, cometary, and geosynchronous missions are being studied. The object of this paper is to describe a basic spacecraft design proposed as the means to accomplish (1) a comet Encke slow flyby, (2) a comet Encke rendezvous, and (3) an out-of-the-ecliptic mission. The discussion includes design differences foreseen for the various missions and indicates those areas where spacecraft design commonality is possible. Particular emphasis is placed on a solar electric propulsion module design which permits an attractive degree of design inheritance from mission to mission.

Electrical Grounding Architecture for Unmanned Spacecraft

NASA Technical Reports Server (NTRS)

1998-01-01

This handbook is approved for use by NASA Headquarters and all NASA Centers and is intended to provide a common framework for consistent practices across NASA programs. This handbook was developed to describe electrical grounding design architecture options for unmanned spacecraft. This handbook is written for spacecraft system engineers, power engineers, and electromagnetic compatibility (EMC) engineers. Spacecraft grounding architecture is a system-level decision which must be established at the earliest point in spacecraft design. All other grounding design must be coordinated with and be consistent with the system-level architecture. This handbook assumes that there is no one single 'correct' design for spacecraft grounding architecture. There have been many successful satellite and spacecraft programs from NASA, using a variety of grounding architectures with different levels of complexity. However, some design principles learned over the years apply to all types of spacecraft development. This handbook summarizes those principles to help guide spacecraft grounding architecture design for NASA and others.

Systems design study of the Pioneer Venus spacecraft. Volume 1. Technical analyses and tradeoffs, sections 1-4 (part 1 of 4)

NASA Technical Reports Server (NTRS)

1973-01-01

The results are reported of the Pioneer Venus studies from 2 October 1972 through 30 June 1973. Many missions were considered, involving two launch vehicles (Thor/Delta and Atlas/Centaur), and different launch opportunities and spacecraft configurations to meet varying science requirements, all at minimum cost. The sequence of events is described and the specific studies conducted are summarized. The effects of science payload on mission and spacecraft design are discussed along with the mission analyses.

The Auroral Particles experiment

NASA Technical Reports Server (NTRS)

1981-01-01

An instrument for the detection of particles in the energy range of 0.1 ev to 80 Kev was designed, built, tested, calibrated, and flown onboard the spacecraft ATS-6. Data from this instrument generated the following research: intensive studies of the plasma in the vicinity of the spacecraft; global variations of plasmas; correlative studies using either other spacecraft or ground based measurements; and studies of spacecraft interactions with ambient plasmas including charging, local electric fields due to differential charging, and active control of spacecraft potential. Results from this research are presented.

Study of multiple asteroid flyby missions

NASA Technical Reports Server (NTRS)

1972-01-01

The feasibility, scientific objectives, mission profile characteristics, and implementation of an asteroid belt exploration mission by a spacecraft guided to intercept three or more asteroids at close range are discussed. A principal consideration in planning a multiasteroid mission is to cut cost by adapting an available and flight-proven spacecraft design such as Pioneer F and G, augmenting its propulsion and guidance capabilities and revising the scientific payload complement in accordance with required mission characteristics. Spacecraft modification necessary to meet the objectives and requirements of the mission were studied. A ground rule of the study was to hold design changes to a minimum and to utilize available technology as much as possible. However, with mission dates not projected before the end of this decade, a reasonable technology growth in payload instrument design and some subsystem components is anticipated that can be incorporated in the spacecraft adaptation.

Solar-C Conceptual Spacecraft Design Study: Final Review. Release 2

NASA Technical Reports Server (NTRS)

Hopkins, Randall; Baysinger, Mike; Thomas, Dan; Heaton, Andy; Stough, Rob; Hill, Spencer; Owens, Jerry; Young, Roy; Fabisinski, Leo; Thomas, Scott;

Mars, Phobos, and Deimos Sample Return Enabled by ARRM Alternative Trade Study Spacecraft

NASA Technical Reports Server (NTRS)

Englander, Jacob A.; Vavrina, Matthew; Merrill, Raymond G.; Qu, Min; Naasz, Bo J.

2014-01-01

The Asteroid Robotic Redirect Mission (ARRM) has been the topic of many mission design studies since 2011. The reference ARRM spacecraft uses a powerful solar electric propulsion (SEP) system and a bag device to capture a small asteroid from an Earth-like orbit and redirect it to a distant retrograde orbit (DRO) around the moon. The ARRM Option B spacecraft uses the same propulsion system and multi-Degree of Freedom (DoF) manipulators device to retrieve a very large sample (thousands of kilograms) from a 100+ meter diameter farther-away Near Earth Asteroid (NEA). This study will demonstrate that the ARRM Option B spacecraft design can also be used to return samples from Mars and its moons - either by acquiring a large rock from the surface of Phobos or Deimos, and or by rendezvousing with a sample-return spacecraft launched from the surface of Mars.

Mars, Phobos, and Deimos Sample Return Enabled by ARRM Alternative Trade Study Spacecraft

NASA Technical Reports Server (NTRS)

Englander, Jacob A.; Vavrina, Matthew; Naasz, Bo; Merill, Raymond G.; Qu, Min

2014-01-01

The Asteroid Robotic Redirect Mission (ARRM) has been the topic of many mission design studies since 2011. The reference ARRM spacecraft uses a powerful solar electric propulsion (SEP) system and a bag device to capture a small asteroid from an Earth-like orbit and redirect it to a distant retrograde orbit (DRO) around the moon. The ARRM Option B spacecraft uses the same propulsion system and multi-Degree of Freedom (DoF) manipulators device to retrieve a very large sample (thousands of kilograms) from a 100+ meter diameter farther-away Near Earth Asteroid (NEA). This study will demonstrate that the ARRM Option B spacecraft design can also be used to return samples from Mars and its moons - either by acquiring a large rock from the surface of Phobos or Deimos, and/or by rendezvousing with a sample-return spacecraft launched from the surface of Mars.

Tracking and data relay satellite system configuration and tradeoff study. Volume 4: Spacecraft and subsystem design, part 1

NASA Technical Reports Server (NTRS)

Hill, T. E.

1972-01-01

The design and development of the Tracking and Data Relay satellite are discussed. The subjects covered are: (1) spacecraft mechanical and structural design, (2) attitude stabilization and control subsystem, (3) propulsion system, (4) electrical power subsystem, (5) thermal control, and (6) reliability engineering.

Spacecraft design project: Low Earth orbit communications satellite

NASA Technical Reports Server (NTRS)

Moroney, Dave; Lashbrook, Dave; Mckibben, Barry; Gardener, Nigel; Rivers, Thane; Nottingham, Greg; Golden, Bill; Barfield, Bill; Bruening, Joe; Wood, Dave

1991-01-01

This is the final product of the spacecraft design project completed to fulfill the academic requirements of the Spacecraft Design and Integration 2 course (AE-4871) taught at the U.S. Naval Postgraduate School. The Spacecraft Design and Integration 2 course is intended to provide students detailed design experience in selection and design of both satellite system and subsystem components, and their location and integration into a final spacecraft configuration. The design team pursued a design to support a Low Earth Orbiting (LEO) communications system (GLOBALSTAR) currently under development by the Loral Cellular Systems Corporation. Each of the 14 team members was assigned both primary and secondary duties in program management or system design. Hardware selection, spacecraft component design, analysis, and integration were accomplished within the constraints imposed by the 11 week academic schedule and the available design facilities.

Parametric Design within an Atomic Design Process (ADP) applied to Spacecraft Design

NASA Astrophysics Data System (ADS)

Ramos Alarcon, Rafael

This thesis describes research investigating the development of a model for the initial design of complex systems, with application to spacecraft design. The design model is called an atomic design process (ADP) and contains four fundamental stages (specifications, configurations, trade studies and drivers) that constitute the minimum steps of an iterative process that helps designers find a feasible solution. Representative design models from the aerospace industry are reviewed and are compared with the proposed model. The design model's relevance, adaptability and scalability features are evaluated through a focused design task exercise with two undergraduate teams and a long-term design exercise performed by a spacecraft payload team. The implementation of the design model is explained in the context in which the model has been researched. This context includes the organization (a student-run research laboratory at the University of Michigan), its culture (academically oriented), members that have used the design model and the description of the information technology elements meant to provide support while using the model. This support includes a custom-built information management system that consolidates relevant information that is currently being used in the organization. The information is divided in three domains: personnel development history, technical knowledge base and laboratory operations. The focused study with teams making use of the design model to complete an engineering design exercise consists of the conceptual design of an autonomous system, including a carrier and a deployable lander that form the payload of a rocket with an altitude range of over 1000 meters. Detailed results from each of the stages of the design process while implementing the model are presented, and an increase in awareness of good design practices in the teams while using the model are explained. A long-term investigation using the design model consisting of the successful characterization of an imaging system for a spacecraft is presented. The spacecraft is designed to take digital color images from low Earth orbit. The dominant drivers from each stage of the design process are indicated as they were identified, with the accompanying hardware development leading to the final configuration that comprises the flight spacecraft.

Cryogenic On-Orbit Liquid Depot-Storage, Acquisition and Transfer (COLD-SAT) Experiment Conceptual Design and Feasibility Study

NASA Technical Reports Server (NTRS)

Kramer, Edward (Editor)

1998-01-01

The cryogenic fluid management technologies required for the exploration of the solar system can only be fully developed via space-based experiments. A dedicated spacecraft is the most efficient way to perform these experiments. This report documents the extended conceptual design of the COLD-SAT spacecraft, capable of meeting these experimental requirements. All elements, including the spacecraft, ground segment, launch site modifications and launch vehicle operations, and flight operations are included. Greatly expanded coverage is provided for those areas unique to this cryogenic spacecraft, such as the experiment system, attitude control system, and spacecraft operations. Supporting analyses are included as are testing requirements, facilities surveys, and proposed project timelines.

Design of a laboratory study of contaminant film darkening in space

NASA Astrophysics Data System (ADS)

Judeikis, H. S.; Arnold, G. S.; Hill, M.; Young Owl, R. C.; Hall, D. F.

This paper reports the philosophy, design, and initial results of a program aimed at improving control of the optical effects of contamination in the design of a spacecraft. The types of basic data needed to produce criteria for the selection of spacecraft material based on the effects of the contaminant films they produce are discussed. The results of a spacecraft nonmetallic materials list analysis and radiation sensitivity estimates are presented. A rationale is given for simulation analysis of the potential effects of the geosynchronous environment on organic contaminant films.

Saturn Uranus atmospheric entry probe mission spacecraft system definition study

NASA Technical Reports Server (NTRS)

1973-01-01

The modifications required of the Pioneer F/G spacecraft design for it to deliver an atmospheric entry probe to the planets Saturn and Uranus are investigated. It is concluded that it is feasible to conduct such a mission within the constraints and interfaces defined. The spacecraft required to perform the mission is derived from the Pioneer F/G design, and the modifications required are generally routinely conceived and executed. The entry probe is necessarily a new design, although it draws on the technology of past, present, and imminent programs of planetary atmospheric investigations.

Adaptation of G-TAG Software for Validating Touch-and-Go Comet Surface Sampling Design Methodology

NASA Technical Reports Server (NTRS)

Mandic, Milan; Acikmese, Behcet; Blackmore, Lars

2011-01-01

The G-TAG software tool was developed under the R&TD on Integrated Autonomous Guidance, Navigation, and Control for Comet Sample Return, and represents a novel, multi-body dynamics simulation software tool for studying TAG sampling. The G-TAG multi-body simulation tool provides a simulation environment in which a Touch-and-Go (TAG) sampling event can be extensively tested. TAG sampling requires the spacecraft to descend to the surface, contact the surface with a sampling collection device, and then to ascend to a safe altitude. The TAG event lasts only a few seconds but is mission-critical with potentially high risk. Consequently, there is a need for the TAG event to be well characterized and studied by simulation and analysis in order for the proposal teams to converge on a reliable spacecraft design. This adaptation of the G-TAG tool was developed to support the Comet Odyssey proposal effort, and is specifically focused to address comet sample return missions. In this application, the spacecraft descends to and samples from the surface of a comet. Performance of the spacecraft during TAG is assessed based on survivability and sample collection performance. For the adaptation of the G-TAG simulation tool to comet scenarios, models are developed that accurately describe the properties of the spacecraft, approach trajectories, and descent velocities, as well as the models of the external forces and torques acting on the spacecraft. The adapted models of the spacecraft, descent profiles, and external sampling forces/torques were more sophisticated and customized for comets than those available in the basic G-TAG simulation tool. Scenarios implemented include the study of variations in requirements, spacecraft design (size, locations, etc. of the spacecraft components), and the environment (surface properties, slope, disturbances, etc.). The simulations, along with their visual representations using G-View, contributed to the Comet Odyssey New Frontiers proposal effort by indicating problems and/or benefits of different approaches and designs.

Automating Structural Analysis of Spacecraft Vehicles

NASA Technical Reports Server (NTRS)

Hrinda, Glenn A.

2004-01-01

A major effort within NASA's vehicle analysis discipline has been to automate structural analysis and sizing optimization during conceptual design studies of advanced spacecraft. Traditional spacecraft structural sizing has involved detailed finite element analysis (FEA) requiring large degree-of-freedom (DOF) finite element models (FEM). Creation and analysis of these models can be time consuming and limit model size during conceptual designs. The goal is to find an optimal design that meets the mission requirements but produces the lightest structure. A structural sizing tool called HyperSizer has been successfully used in the conceptual design phase of a reusable launch vehicle and planetary exploration spacecraft. The program couples with FEA to enable system level performance assessments and weight predictions including design optimization of material selections and sizing of spacecraft members. The software's analysis capabilities are based on established aerospace structural methods for strength, stability and stiffness that produce adequately sized members and reliable structural weight estimates. The software also helps to identify potential structural deficiencies early in the conceptual design so changes can be made without wasted time. HyperSizer's automated analysis and sizing optimization increases productivity and brings standardization to a systems study. These benefits will be illustrated in examining two different types of conceptual spacecraft designed using the software. A hypersonic air breathing, single stage to orbit (SSTO), reusable launch vehicle (RLV) will be highlighted as well as an aeroshell for a planetary exploration vehicle used for aerocapture at Mars. By showing the two different types of vehicles, the software's flexibility will be demonstrated with an emphasis on reducing aeroshell structural weight. Member sizes, concepts and material selections will be discussed as well as analysis methods used in optimizing the structure. Analysis based on the HyperSizer structural sizing software will be discussed. Design trades required to optimize structural weight will be presented.

Systems design and comparative analysis of large antenna concepts

NASA Technical Reports Server (NTRS)

Garrett, L. B.; Ferebee, M. J., Jr.

1983-01-01

Conceptual designs are evaluated and comparative analyses conducted for several large antenna spacecraft for Land Mobile Satellite System (LMSS) communications missions. Structural configurations include trusses, hoop and column and radial rib. The study was conducted using the Interactive Design and Evaluation of Advanced Spacecraft (IDEAS) system. The current capabilities, development status, and near-term plans for the IDEAS system are reviewed. Overall capabilities are highlighted. IDEAS is an integrated system of computer-aided design and analysis software used to rapidly evaluate system concepts and technology needs for future advanced spacecraft such as large antennas, platforms, and space stations. The system was developed at Langley to meet a need for rapid, cost-effective, labor-saving approaches to the design and analysis of numerous missions and total spacecraft system options under consideration. IDEAS consists of about 40 technical modules efficient executive, data-base and file management software, and interactive graphics display capabilities.

Spacecraft radiator systems

NASA Technical Reports Server (NTRS)

Anderson, Grant A. (Inventor)

2012-01-01

A spacecraft radiator system designed to provide structural support to the spacecraft. Structural support is provided by the geometric "crescent" form of the panels of the spacecraft radiator. This integration of radiator and structural support provides spacecraft with a semi-monocoque design.

Data base architecture for instrument characteristics critical to spacecraft conceptual design

NASA Technical Reports Server (NTRS)

Rowell, Lawrence F.; Allen, Cheryl L.

1990-01-01

Spacecraft designs are driven by the payloads and mission requirements that they support. Many of the payload characteristics, such as mass, power requirements, communication requirements, moving parts, and so forth directly affect the choices for the spacecraft structural configuration and its subsystem design and component selection. The conceptual design process, which translates mission requirements into early spacecraft concepts, must be tolerant of frequent changes in the payload complement and resource requirements. A computer data base was designed and implemented for the purposes of containing the payload characteristics pertinent for spacecraft conceptual design, tracking the evolution of these payloads over time, and enabling the integration of the payload data with engineering analysis programs for improving the efficiency in producing spacecraft designs. In-house tools were used for constructing the data base and for performing the actual integration with an existing program for optimizing payload mass locations on the spacecraft.

NASA STD-4005: The LEO Spacecraft Charging Design Standard

NASA Technical Reports Server (NTRS)

Ferguson, Dale C.

2006-01-01

Power systems with voltages higher than about 55 volts may charge in Low Earth Orbit (LEO) enough to cause destructive arcing. The NASA STD-4005 LEO Spacecraft Charging Design Standard will help spacecraft designers prevent arcing and other deleterious effects on LEO spacecraft. The Appendices, an Information Handbook based on the popular LEO Spacecraft Charging Design Guidelines by Ferguson and Hillard, serve as a useful explanation and accompaniment to the Standard.

A thermal scale modeling study for Apollo and Apollo applications, volume 1

NASA Technical Reports Server (NTRS)

Shannon, R. L.

1972-01-01

The program is reported for developing and demonstrating the capabilities of thermal scale modeling as a thermal design and verification tool for Apollo and Apollo Applications Projects. The work performed for thermal scale modeling of STB; cabin atmosphere/spacecraft cabin wall thermal interface; closed loop heat rejection radiator; and docked module/spacecraft thermal interface are discussed along with the test facility requirements for thermal scale model testing of AAP spacecraft. It is concluded that thermal scale modeling can be used as an effective thermal design and verification tool to provide data early in a spacecraft development program.

A study of spacecraft technology and design concepts. Volume 2: Appendices

NASA Technical Reports Server (NTRS)

Zylius, F. A.

1985-01-01

Electrical, mechanical, and software subsystem needs in the Post 1990 space operations environment are considered as well as the effect of radiation environment on spacecraft configuration. Criteria are given for selecting a specific design or technology concept from among the alternatives available.

Atmosphere explorer missions C, D, and E. Spacecraft experiment interface definition study

NASA Technical Reports Server (NTRS)

1972-01-01

The Atmosphere Explorer Missions C, D, & E Spacecraft/Experiment Interface Definition Study is discussed. The objectives of the study included an analysis of the accommodation requirements of the experiments for the three missions, an assessment of the overall effect of these requirements on the spacecraft system design and performance, and the detailed definition of all experiment/spacecraft electrical, mechanical, and environmental interfaces. In addition, the study included the identification and definition of system characteristics required to ensure compatibility with the consolidated STADAN and MSFN communications networks.

Concurrent engineering: Spacecraft and mission operations system design

NASA Technical Reports Server (NTRS)

Landshof, J. A.; Harvey, R. J.; Marshall, M. H.

1994-01-01

Despite our awareness of the mission design process, spacecraft historically have been designed and developed by one team and then turned over as a system to the Mission Operations organization to operate on-orbit. By applying concurrent engineering techniques and envisioning operability as an essential characteristic of spacecraft design, tradeoffs can be made in the overall mission design to minimize mission lifetime cost. Lessons learned from previous spacecraft missions will be described, as well as the implementation of concurrent mission operations and spacecraft engineering for the Near Earth Asteroid Rendezvous (NEAR) program.

Interplanetary propulsion using inertial fusion

NASA Technical Reports Server (NTRS)

Orth, C. D.; Hogan, W. J.; Hoffman, N.; Murray, K.; Klein, G.; Diaz, F. C.

1987-01-01

Inertial fusion can be used to power spacecraft within the solar system and beyond. Such spacecraft have the potential for short-duration manned-mission performance exceeding other technologies. We are conducting a study to assess the systems aspects of inertial fusion as applied to such missions, based on the conceptual engine design of Hyde (1983) we describe the required systems for an entirely new spacecraft design called VISTA that is based on the use of DT fuel. We give preliminary design details for the power conversion and power conditioning systems for manned missions to Mars of total duration of about 100 days. Specific mission performance results will be published elsewhere, after the study has been completed.

Limits on Interconnection Bandwidth for On-Board Processing

NASA Technical Reports Server (NTRS)

Lux, James P.

2006-01-01

This viewgraph presentation reviews the constraints, and concerns of spacecraft design, in particular spacecraft instrumentation design and the issues concerning space communication. The advantages and disadvantages of several communication options are reviewed. Ultimately there will be spacecraft communication not between boxes on spacecraft, but between spacecraft. The future of spacecraft communication is interplanetary networks.

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

NASA Technical Reports Server (NTRS)

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.

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

NASA Technical Reports Server (NTRS)

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".

Spacecraft (Mobile Satellite) configuration design study

NASA Technical Reports Server (NTRS)

1985-01-01

The relative costs to procure and operate a two-satellite mobile satellite system designed to operate either in the UHF band of the L Band, and with several antenna diameter options in each frequency band was investigated. As configured, the size of the spacecraft is limited to the current RCA Series 4000 Geosynchronous Communications Spacecraft bus, which spans the range from 4000 to 5800 pounds in the transfer orbit. The Series 4000 bus forms the basis around which the Mobile Satellite transponder and associated antennas were appended. Although the resultant configuration has little outward resemblance to the present Series 4000 microwave communications spacecraft, the structure, attitude control, thermal, power, and command and control subsystems of the Series 4000 spacecraft are all adapted to support the Mobile Satellite mission.

Interplanetary propulsion using inertial fusion

NASA Technical Reports Server (NTRS)

Orth, Charles D.; Hoffman, Nate; Murray, Kathy; Klein, Gail; Diaz, Franklin Chang

1987-01-01

Inertial fusion can be used to power spacecraft within the solar system and beyond. Such spacecraft have the potential for short duration manned mission performance exceeding other technologies. A study was conducted to assess the systems aspects of inertial as applied to such missions, based on the conceptual engine design of Hyde (1983). The required systems for an entirely new spacecraft design called VISTA that is based on the use of DT fuel is described. Preliminary design details are given for the power conversion and power conditioning systems for manned missions to Mars of total duration of about 100 days.

Flexible Multi-Body Spacecraft Simulator: Design, Construction, and Experiments

DTIC Science & Technology

2017-12-01

BODY SPACECRAFT SIMULATOR: DESIGN , CONSTRUCTION, AND EXPERIMENTS by Adam L. Atwood December 2017 Thesis Advisor: Mark Karpenko Second...TYPE AND DATES COVERED Master’s thesis 4. TITLE AND SUBTITLE FLEXIBLE MULTI-BODY SPACECRAFT SIMULATOR: DESIGN , CONSTRUCTION, AND EXPERIMENTS 5...spacecraft simulator for use in testing optimal control-based slew and maneuver designs . The simulator is modified from an earlier prototype, which

Performance Evaluation of the Gravity Probe B Design

NASA Technical Reports Server (NTRS)

Francis, Ronnie; Wells, Eugene M.

1996-01-01

This report documents the simulation of the Lockheed Martin designed Gravity Probe B (GPB) spacecraft developed tool by bd Systems Inc using the TREETOPS simulation. This study quantifies the effects of flexibility and liquid helium slosh on GPB spacecraft control performance. The TREETOPS simulation tool permits the simulation of flexible structures given that a flexible body model of the structure is available. For purposes of this study, a flexible model of the GPB spacecraft was obtained from Lockheed Martin. To model the liquid helium slosh effects, computational fluid dynamics (CFD) results' were obtained, and used to develop a dynamic model of the slosh effects. The flexible body and slosh effects were incorporated separately into the TREETOPS simulation, which places the vehicle in a 650 km circular polar orbit and subjects the spacecraft to realistic environmental disturbances and sensor error quantities. In all of the analysis conducted in this study the spacecraft is pointed at an inertially fixed guide star (GS) and is rotating at a constant rate about this line of sight.

Architecture Study on Telemetry Coverage for Immediate Post-Separation Phase

NASA Technical Reports Server (NTRS)

Cheung, Kar-Ming; Lee, Charles H.; Kellogg, Kent H.; Stocklin, Frank J.; Zillig, David J.; Fielhauer, Karl B.

2008-01-01

This paper presents the preliminary results of an architecture study that provides continuous telemetry coverage for NASA missions for immediate post-separation phase. This study is a collaboration effort between Jet Propulsion Laboratory (JPL), Goddard Space Flight Center (GSFC), and Applied Physics Laboratory (APL). After launch when the spacecraft separated from the upper stage, the spacecraft typically executes a number of mission-critical operations prior to the deployment of solar panels and the activation of the primary communication subsystem. JPL, GSFC, and APL have similar design principle statements that require continuous coverage of mission-critical telemetry during the immediate post-separation phase. To conform to these design principles, an architecture that consists of a separate spacecraft transmitter and a robust communication network capable of tracking the spacecraft signals is needed.This paper presents the preliminary results of an architecture study that provides continuous telemetry coverage for NASA missions for immediate post-separation phase. This study is a collaboration effort between Jet Propulsion Laboratory (JPL), Goddard Space Flight Center (GSFC), and Applied Physics Laboratory (APL). After launch when the spacecraft separated from the upper stage, the spacecraft typically executes a number of mission-critical operations prior to the deployment of solar panels and the activation of the primary communication subsystem. JPL, GSFC, and APL have similar design principle statements that require continuous coverage of mission-critical telemetry during the immediate post-separation phase. To conform to these design principles, an architecture that consists of a separate spacecraft transmitter and a robust communication network capable of tracking the spacecraft signals is needed. The main results of this study are as follows: 1) At low altitude (< 10000 km) when most post-separation critical operations are executed, Earth-based network (e.g. Deep Space Network (DSN)) can only provide limited coverage, whereas space-based network (e.g. Space Network (SN)) can provide continuous coverage. 2) Commercial-off-the-shelf SN compatible transmitters are available for small satellite applications. In this paper we present the detailed coverage analysis of Earth-based and Space-based networks. We identify the key functional and performance requirements of the architecture, and describe the proposed selection criteria of the spacecraft transmitter. We conclude the paper with a proposed forward plan.

Brane Craft

NASA Technical Reports Server (NTRS)

Janson, Siegfried

2017-01-01

A Brane Craft is a membrane spacecraft with solar cells, command and control electronics, communications systems, antennas, propulsion systems, attitude and proximity sensors, and shape control actuators as thin film structures manufactured on 10 micron thick plastic sheets. This revolutionary spacecraft design can have a thickness of tens of microns with a surface area of square meters to maximize area-to-mass ratios for exceptionally low-mass spacecraft. Communications satellites, solar power satellites, solar electric propulsion stages, and solar sails can benefit from Brane Craft design. It also enables new missions that require low-mass spacecraft with exceptionally high delta-V. Active removal of orbital debris from Earth orbit is the target application for this study.

Preliminary Structural Sensitivity Study of Hypersonic Inflatable Aerodynamic Decelerator Using Probabilistic Methods

NASA Technical Reports Server (NTRS)

Lyle, Karen H.

2014-01-01

Acceptance of new spacecraft structural architectures and concepts requires validated design methods to minimize the expense involved with technology validation via flighttesting. This paper explores the implementation of probabilistic methods in the sensitivity analysis of the structural response of a Hypersonic Inflatable Aerodynamic Decelerator (HIAD). HIAD architectures are attractive for spacecraft deceleration because they are lightweight, store compactly, and utilize the atmosphere to decelerate a spacecraft during re-entry. However, designers are hesitant to include these inflatable approaches for large payloads or spacecraft because of the lack of flight validation. In the example presented here, the structural parameters of an existing HIAD model have been varied to illustrate the design approach utilizing uncertainty-based methods. Surrogate models have been used to reduce computational expense several orders of magnitude. The suitability of the design is based on assessing variation in the resulting cone angle. The acceptable cone angle variation would rely on the aerodynamic requirements.

Description of the Spacecraft Control Laboratory Experiment (SCOLE) facility

NASA Technical Reports Server (NTRS)

Williams, Jeffrey P.; Rallo, Rosemary A.

1987-01-01

A laboratory facility for the study of control laws for large flexible spacecraft is described. The facility fulfills the requirements of the Spacecraft Control Laboratory Experiment (SCOLE) design challenge for a laboratory experiment, which will allow slew maneuvers and pointing operations. The structural apparatus is described in detail sufficient for modelling purposes. The sensor and actuator types and characteristics are described so that identification and control algorithms may be designed. The control implementation computer and real-time subroutines are also described.

Description of the Spacecraft Control Laboratory Experiment (SCOLE) facility

NASA Technical Reports Server (NTRS)

Williams, Jeffrey P.; Rallo, Rosemary A.

1987-01-01

A laboratory facility for the study of control laws for large flexible spacecraft is described. The facility fulfills the requirements of the Spacecraft Control Laboratory Experiment (SCOLE) design challenge for laboratory experiments, which will allow slew maneuvers and pointing operations. The structural apparatus is described in detail sufficient for modelling purposes. The sensor and actuator types and characteristics are described so that identification and control algorithms may be designed. The control implementation computer and real-time subroutines are also described.

Spacecraft design project: High temperature superconducting infrared imaging satellite

NASA Technical Reports Server (NTRS)

1991-01-01

The High Temperature Superconductor Infrared Imaging Satellite (HTSCIRIS) is designed to perform the space based infrared imaging and surveillance mission. The design of the satellite follows the black box approach. The payload is a stand alone unit, with the spacecraft bus designed to meet the requirements of the payload as listed in the statement of work. Specifications influencing the design of the spacecraft bus were originated by the Naval Research Lab. A description of the following systems is included: spacecraft configuration, orbital dynamics, radio frequency communication subsystem, electrical power system, propulsion, attitude control system, thermal control, and structural design. The issues of testing and cost analysis are also addressed. This design project was part of the course Advanced Spacecraft Design taught at the Naval Postgraduate School.

Near Earth asteroid rendezvous

NASA Technical Reports Server (NTRS)

1992-01-01

The Spacecraft Design Course is the capstone design class for the M.S. in astronautics at the Naval Postgraduate School. The Fall 92 class designed a spacecraft for the Near Earth Asteroid Rendezvous Mission (NEAR). The NEAR mission uses a robotic spacecraft to conduct up-close reconnaissance of a near-earth asteroid. Such a mission will provide information on Solar System formation and possible space resources. The spacecraft is intended to complete a NEAR mission as a relatively low-budget program while striving to gather as much information about the target asteroid as possible. A complete mission analysis and detailed spacecraft design were completed. Mission analysis includes orbit comparison and selection, payload and telemetry requirements, spacecraft configuration, and launch vehicle selection. Spacecraft design includes all major subsystems: structure, electrical power, attitude control, propulsion, payload integration, and thermal control. The resulting spacecraft demonstrates the possibility to meet the NEAR mission requirements using existing technology, 'off-the-shelf' components, and a relatively low-cost launch vehicle.

Design Study for a Global Magnetospheric Dynamics Mission

NASA Technical Reports Server (NTRS)

Russell, C. T.

1999-01-01

A successful design was developed, one with many advantages over the original mission. The time spent in orbit was more evenly spread over the region being investigated. The radiation close was significantly lower and the mission did not rely on gravity assist at the moon and thus did not have to make measurements that far out in the tail. A spacecraft design was developed that keeps interference from the engines to a minimum. The design however was quite specific for four spacecraft. It could not be easily scaled to five spacecraft for example. One problem was discovered that is a concern for all similar missions. Inter- spacecraft communication can determine the spacing of the vehicles easily and to the accuracy required. However, the orientation of the polyhedron with the spacecraft at its vertices is not well known for small separations. Ground station range measurements give the line of sight location well but not the angle around that vector. This is a problem any such mission needs to solve. Neither the navigation teams at Goddard nor at Lewis were willing to attempt to solve this problem. At the completion of the study a report was made to the AGU meeting in San Francisco and a paper published in the volume "Science Closure and Enabling Technologies for Constellation Class Missions". This paper is attached.

Preentry communications study. Outer planets atmospheric entry probe

NASA Technical Reports Server (NTRS)

Hinrichs, C. A.

1976-01-01

A pre-entry communications study is presented for a relay link between a Jupiter entry probe and a spacecraft in hyperbolic orbit. Two generic communications links of interest are described: a pre-entry link to a spun spacecraft antenna, and a pre-entry link to a despun spacecraft antenna. The propagation environment of Jupiter is defined. Although this is one of the least well known features of Jupiter, enough information exists to reasonably establish bounds on the performance of a communications link. Within these bounds, optimal carrier frequencies are defined. The next step is to identify optimal relative geometries between the probe and the spacecraft. Optimal trajectories are established for both spun and despun spacecraft antennas. Given the optimal carrier frequencies, and the optimal trajectories, the data carrying capacities of the pre-entry links are defined. The impact of incorporating pre-entry communications into a basic post entry probe is then assessed. This assessment covers the disciplines of thermal control, power source, mass properties and design layout. A conceptual design is developed of an electronically despun antenna for use on a Pioneer class of spacecraft.

Standard spacecraft procurement analysis: A case study in NASA-DoD coordination in space programs. Ph.D. Thesis - Rand Graduate Inst.

NASA Technical Reports Server (NTRS)

Harris, E. D.

1980-01-01

The Space Test Program Standard Satellite (STPSS), a design proposed by the Air Force, and two NASA candidates, the Applications Explorer Mission spacecraft (AEM) and the Multimission Modular Spacecraft (MMS), were considered during the first phase. During the second phase, a fourth candidate was introduced, a larger, more capable AEM (L-AEM), configured by the Boeing Company under NASA sponsorship to meet the specifications jointly agreed upon by NASA and the Air Force. Total program costs for a variety of procurement options, each of which is capable of performing all of the Air Force Space Test Program missions during the 1980-1990 time period, were used as the principal measure for distinguishing among procurement options. Program cost does not provide a basis for choosing among the AEM, STPSS, and MMS spacecraft, given their present designs. The availability of the L-AEM spacecraft, or some very similar design, would provide a basis for minimizing the cost of the Air Force Space Test Program.

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

NASA Technical Reports Server (NTRS)

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.

System design of the Pioneer Venus spacecraft. Volume 5: Probe vehicle studies

NASA Technical Reports Server (NTRS)

Nolte, L. J.; Stephenson, D. S.

1973-01-01

A summary of the key issues and studies conducted for the Pioneer Venus spacecraft and the resulting probe designs are presented. The key deceleration module issues are aerodynamic configuration and heat shield material selection. The design and development of the pressure vessel module are explained. Thermal control and science integration of the pressure vessel module are explained. The deceleration module heat shield, parachute and separation/despin are reported. The Thor/Delta and Atlas/Centaur baseline descriptions are provided.

Structures and materials technology needs for communications and remote sensing spacecraft

NASA Technical Reports Server (NTRS)

Gronet, M. J.; Jensen, G. A.; Hoskins, J. W.

1995-01-01

This report documents trade studies conducted from the perspective of a small spacecraft developer to determine and quantify the structures and structural materials technology development needs for future commercial and NASA small spacecraft to be launched in the period 1999 to 2005. Emphasis is placed on small satellites weighing less than 1800 pounds for two focus low-Earth orbit missions: commercial communications and remote sensing. The focus missions are characterized in terms of orbit, spacecraft size, performance, and design drivers. Small spacecraft program personnel were interviewed to determine their technology needs, and the results are summarized. A systems-analysis approach for quantifying the benefits of inserting advanced state-of-the-art technologies into a current reference, state-of-the-practice small spacecraft design is developed and presented. This approach is employed in a set of abbreviated trade studies to quantify the payoffs of using a subset of 11 advanced technologies selected from the interview results The 11 technology development opportunities are then ranked based on their relative payoff. Based on the strong potential for significant benefits, recommendations are made to pursue development of 8 and the 11 technologies. Other important technology development areas identified are recommended for further study.

Magellan aerobraking periapse corridor design

NASA Technical Reports Server (NTRS)

Cook, Richard A.; Lyons, Daniel T.

1992-01-01

One extended mission idea for the Magellan project uses aerobraking techniques to circularize the current orbit. A major technical issue in this proposal is the design of the periapse altitude corridor. Aerobraking would cause a number of significant side effects on both the spacecraft and ground system. Heating and aerodynamic torques on the spacecraft are key issues, as are the corridor control maneuver frequency and aerobrake duration. Spacecraft and ground systems operational limits have been identified in an attempt to constrain the corridor design. A simulation program has been developed to model the aerobraking corridor control process. This paper presents study results using this program which relate to the feasibility of this aerobraking concept.

Spacecraft design project multipurpose satellite bus MPS

NASA Technical Reports Server (NTRS)

Kellman, Lyle; Riley, John; Szostak, Michael; Watkins, Joseph; Willhelm, Joseph; Yale, Gary

1990-01-01

The thrust of this project was to design not a single spacecraft, but to design a multimission bus capable of supporting several current payloads and unnamed, unspecified future payloads. Spiraling costs of spacecraft and shrinking defense budgets necessitated a fresh look at the feasibility of a multimission spacecraft bus. The design team chose two very diverse and different payloads, and along with them two vastly different orbits, to show that multimission spacecraft buses are an area where indeed more research and effort needs to be made. Tradeoffs, of course, were made throughout the design, but optimization of subsystem components limited weight and volume penalties, performance degradation, and reliability concerns. Simplicity was chosen over more complex, sophisticated and usually more efficient designs. Cost of individual subsystem components was not a primary concern in the design phase, but every effort was made to chose flight tested and flight proven hardware. Significant cost savings could be realized if a standard spacecraft bus was indeed designed and purchased in finite quantities.

Radiation Effects and Protection for Moon and Mars Missions

NASA Technical Reports Server (NTRS)

Parnell, Thomas A.; Watts, John W., Jr.; Armstrong, Tony W.

1998-01-01

Manned and robotic missions to the Earth's moon and Mars are exposed to a continuous flux of Galactic Cosmic Rays (GCR) and occasional, but intense, fluxes of Solar Energetic Particles (SEP). These natural radiations impose hazards to manned exploration, but also present some constraints to the design of robotic missions. The hazards to interplanetary flight crews and their uncertainties have been studied recently by a National Research Council Committee (Space Studies Board 1996). Considering the present uncertainty estimates, thick spacecraft shielding would be needed for manned missions, some of which could be accomplished with onboard equipment and expendables. For manned and robotic missions, the effects of radiation on electronics, sensors, and controls require special consideration in spacecraft design. This paper describes the GCR and SEP particle fluxes, secondary particles behind shielding, uncertainties in radiobiological effects and their impact on manned spacecraft design, as well as the major effects on spacecraft equipment. The principal calculational tools and considerations to mitigate the radiation effects are discussed, and work in progress to reduce uncertainties is included.

Incorporating CCSDS telemetry standards and philosophy on Cassini

NASA Technical Reports Server (NTRS)

Day, John C.; Elson, Anne B.

1995-01-01

The Cassini project at the Jet Propulsion Laboratory (JPL) is implementing a spacecraft telemetry system based on the Consultative Committee for Space Data Systems (CCSDS) packet telemetry standards. Resolving the CCSDS concepts with a Ground Data System designed to handle time-division-multiplexed telemetry and also handling constraints unique to a deep-space planetary spacecraft (such as fixed downlink opportunities, small downlink rates and requirements for on-board data storage) have resulted in spacecraft and ground system design challenges. Solving these design challenges involved adapting and extending the CCSDS telemetry standards as well as changes to the spacecraft and ground system designs. The resulting spacecraft/ground system design is an example of how new ideas and philosophies can be incorporated into existing systems and design approaches without requiring significant rework. In addition, it shows that the CCSDS telemetry standards can be successfully applied to deep-space planetary spacecraft.

The spacecraft encounters of Comet Halley

NASA Technical Reports Server (NTRS)

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.

Spacecraft configuration study for second generation mobile satellite system

NASA Technical Reports Server (NTRS)

Louie, M.; Vonstentzsch, W.; Zanella, F.; Hayes, R.; Mcgovern, F.; Tyner, R.

1985-01-01

A high power, high performance communicatons satellite bus being developed is designed to satisfy a broad range of multimission payload requirements in a cost effective manner and is compatible with both STS and expendable launchers. Results are presented of tradeoff studies conducted to optimize the second generation mobile satellite system for its mass, power, and physical size. Investigations of the 20-meter antenna configuration, transponder linearization techniques, needed spacecraft modifications, and spacecraft power, dissipation, mass, and physical size indicate that the advanced spacecraft bus is capable of supporting the required payload for the satellite.

Design study LANDSAT follow-on mission unique communications system

NASA Technical Reports Server (NTRS)

1976-01-01

Spacecraft subsystem design, performance evaluation, and system tradeoffs are presented for the LANDSAT follow-on mission (LF/O) spacecraft to TDRSS link for the transmission of thematic mapper (TM) and multispectral scanner (MSS) data and for the LF/O spacecraft to STDN and other direct users link for the transmission of TM data. Included are requirements definition, link analysis, subsystem and hardware tradeoffs, conceptual selection, hardware definition, and identification of required new technology. Cost estimates of the recommended communication system including both recurring and non recurring costs are discussed.

Simulation of Attitude and Trajectory Dynamics and Control of Multiple Spacecraft

NASA Technical Reports Server (NTRS)

Stoneking, Eric T.

2009-01-01

Agora software is a simulation of spacecraft attitude and orbit dynamics. It supports spacecraft models composed of multiple rigid bodies or flexible structural models. Agora simulates multiple spacecraft simultaneously, supporting rendezvous, proximity operations, and precision formation flying studies. The Agora environment includes ephemerides for all planets and major moons in the solar system, supporting design studies for deep space as well as geocentric missions. The environment also contains standard models for gravity, atmospheric density, and magnetic fields. Disturbance force and torque models include aerodynamic, gravity-gradient, solar radiation pressure, and third-body gravitation. In addition to the dynamic and environmental models, Agora supports geometrical visualization through an OpenGL interface. Prototype models are provided for common sensors, actuators, and control laws. A clean interface accommodates linking in actual flight code in place of the prototype control laws. The same simulation may be used for rapid feasibility studies, and then used for flight software validation as the design matures. Agora is open-source and portable across computing platforms, making it customizable and extensible. It is written to support the entire GNC (guidance, navigation, and control) design cycle, from rapid prototyping and design analysis, to high-fidelity flight code verification. As a top-down design, Agora is intended to accommodate a large range of missions, anywhere in the solar system. Both two-body and three-body flight regimes are supported, as well as seamless transition between them. Multiple spacecraft may be simultaneously simulated, enabling simulation of rendezvous scenarios, as well as formation flying. Built-in reference frames and orbit perturbation dynamics provide accurate modeling of precision formation control.

Solar Power System Options for the Radiation and Technology Demonstration Spacecraft

NASA Technical Reports Server (NTRS)

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.

Star field attitude sensor study for the Pioneer Venus spacecraft

NASA Technical Reports Server (NTRS)

Rudolf, W. P.; Reed, D. R.

1972-01-01

The characteristics of a star field attitude sensor for use with the Pioneer Venus spacecraft are presented. The aspects of technical feasibility, system interface considerations, and cost of flight hardware development are discussed. The tradeoffs which relate to performance, design, cost, and reliability are analyzed. The configuration of the system for installation in the spacecraft is described.

Research study: Space vehicle control systems

NASA Technical Reports Server (NTRS)

Likins, P. W.; Longman, R. W.

1979-01-01

From the control point of view, spacecraft are classified into two main groups: those for which the spacecraft is fully defined before the control system is designed; and those for which the control system must be specified before certain interchangeable parts of a multi-purpose spacecraft are selected for future missions. Consideration is given to both classes of problems.

Operating in the space plasma environment: A spacecraft charging study of the Solar X-ray Imager

NASA Technical Reports Server (NTRS)

Herr, Joel L.; Mccollum, Matthew B.; James, Bonnie F.

1994-01-01

This study presents the results of a spacecraft charging effects protection study conducted on the Solar X-ray Imager (SXI). The SXI is being developed by NASA Marshall Space Flight Center for NOAA's Space Environment Laboratory, and will be used to aid in forecasting energetic particle events and geomagnetic storms. Images will provide information on the intensity and location of solar flares, coronal mass ejections, and high speed solar streams. The SXI will be flown on a next-generation GOES sometime in the mid to late 1990's. Charging due to the encounter with a worst-case magnetic substorm environment is modeled using the NASCAP/GEO computer code. Charging levels of exterior surfaces and the floating potential of the spacecraft relative to plasma are determined as a function of spacecraft design, operational configuration, and orbital conditions. Areas where large surface voltage gradients exist on or near the SXI are identified as possible arc-discharge sites. Results of the charging analysis are then used to develop design recommendations that will limit the effects of spacecraft charging on the SXI operation.

The Human as a System - Monitoring Spacecraft Net Habitable Volume throughout the Design Lifecycle

NASA Technical Reports Server (NTRS)

Szabo, Richard; Kallay, Anna; Twyford, Evan; Maida, Jim

2007-01-01

Spacecraft design has historically allocated specific volume and mass "not to exceed" requirements upon individual systems and their accompanying hardware (e.g., life support, avionics) early in their conceptual design in an effort to align the spacecraft with propulsion capabilities. If the spacecraft is too heavy or too wide for the launch stack - it does not get off the ground. This approach has predictably ended with the crew being allocated whatever open, pressurized volume remains. With the recent inauguration of a new human-rated spacecraft - NASA human factors personnel have found themselves in the unique position to redefine the human as a system from the very foundation of design. They seek to develop and monitor a "not to fall below" requirement for crew net habitable volume (NHV) - balanced against the "not to exceed" system volume requirements, with the spacecraft fitting the crew versus the crew having to fit inside the spacecraft.

Low Earth orbit communications satellite

NASA Technical Reports Server (NTRS)

Moroney, D.; Lashbrook, D.; Mckibben, B.; Gardener, N.; Rivers, T.; Nottingham, G.; Golden, B.; Barfield, B.; Bruening, J.; Wood, D.

1992-01-01

A current thrust in satellite communication systems considers a low-Earth orbiting constellations of satellites for continuous global coverage. Conceptual design studies have been done at the time of this design project by LORAL Aerospace Corporation under the program name GLOBALSTAR and by Motorola under their IRIDIUM program. This design project concentrates on the spacecraft design of the GLOBALSTAR low-Earth orbiting communication system. Overview information on the program was gained through the Federal Communications Commission licensing request. The GLOBALSTAR system consists of 48 operational satellites positioned in a Walker Delta pattern providing global coverage and redundancy. The operational orbit is 1389 km (750 nmi) altitude with eight planes of six satellites each. The orbital planes are spaced 45 deg., and the spacecraft are separated by 60 deg. within the plane. A Delta 2 launch vehicle is used to carry six spacecraft for orbit establishment. Once in orbit, the spacecraft will utilize code-division multiple access (spread spectrum modulation) for digital relay, voice, and radio determination satellite services (RDSS) yielding position determination with accuracy up to 200 meters.

The Voyager spacecraft /James Watt International Gold Medal Lecture/

NASA Technical Reports Server (NTRS)

Heacock, R. L.

1980-01-01

The Voyager Project background is reviewed with emphasis on selected features of the Voyager spacecraft. Investigations by the Thermo-electric Outer Planets Spacecraft Project are discussed, including trajectories, design requirements, and the development of a Self Test and Repair computer, and a Computer Accessed Telemetry System. The design and configuration of the spacecraft are described, including long range communications, attitude control, solar independent power, sequencing and control data handling, and spacecraft propulsion. The development program, maintained by JPL, experienced a variety of problems such as design deficiencies, and process control and manufacturing problems. Finally, the spacecraft encounter with Jupiter is discussed, and expectations for the Saturn encounter are expressed.

Thermal re-design of the Galileo spacecraft for a Venus-earth-earth-gravity assist (VEEGA) trajectory

NASA Technical Reports Server (NTRS)

Reeve, R.

1989-01-01

The cancellation of the Centaur upper stage program in the aftermath of the Challenger tragedy forced a redesign of the flight trajectory of the Galileo spacecraft to Jupiter, i.e., from a direct trajectory to the Venus-earth-earth-gravity-assist (VEEGA) trajectory on the lower energy two-stage inertial upper stage (IUS), with the result that the spacecraft would be exposed to more than twofold increase in peak solar irradiance. This paper describes the general system-level thermal redesign effort for the Galileo spacecraft, from the start of feasibility studies to its final implementation. Results indicate that the addition of sunshades and the generous utilization of second-surface aluminized Kapton surface material for reflecting high percentages of incident solar irradiation would 'harden' the spacecraft's existing thermal protection system adequately, provided that sun-pointing at the relatively higher solar irradiance levels could be maintained. The final miximum flight temperature predictions for the spacecraft's subsystem thermal designs are given.

General Methodology for Designing Spacecraft Trajectories

NASA Technical Reports Server (NTRS)

Condon, Gerald; Ocampo, Cesar; Mathur, Ravishankar; Morcos, Fady; Senent, Juan; Williams, Jacob; Davis, Elizabeth C.

2012-01-01

A methodology for designing spacecraft trajectories in any gravitational environment within the solar system has been developed. The methodology facilitates modeling and optimization for problems ranging from that of a single spacecraft orbiting a single celestial body to that of a mission involving multiple spacecraft and multiple propulsion systems operating in gravitational fields of multiple celestial bodies. The methodology consolidates almost all spacecraft trajectory design and optimization problems into a single conceptual framework requiring solution of either a system of nonlinear equations or a parameter-optimization problem with equality and/or inequality constraints.

Six degree of freedom simulation system for evaluating automated rendezvous and docking spacecraft

NASA Technical Reports Server (NTRS)

Rourke, Kenneth H.; Tsugawa, Roy K.

1991-01-01

Future logistics supply and servicing vehicles such as cargo transfer vehicles (CTV) must have full 6 degree of freedom (6DOF) capability in order to perform requisite rendezvous, proximity operations, and capture operations. The design and performance issues encountered when developing a 6DOF maneuvering spacecraft are very complex with subtle interactions which are not immediately obvious or easily anticipated. In order to deal with these complexities and develop robust maneuvering spacecraft designs, a simulation system and associated family of tools are used at TRW for generating and validating spacecraft performance requirements and guidance algorithms. An overview of the simulator and tools is provided. These are used by TRW for autonomous rendezvous and docking research projects including CTV studies.

Improving spacecraft design using a multidisciplinary design optimization methodology

NASA Astrophysics Data System (ADS)

Mosher, Todd Jon

2000-10-01

Spacecraft design has gone from maximizing performance under technology constraints to minimizing cost under performance constraints. This is characteristic of the "faster, better, cheaper" movement that has emerged within NASA. Currently spacecraft are "optimized" manually through a tool-assisted evaluation of a limited set of design alternatives. With this approach there is no guarantee that a systems-level focus will be taken and "feasibility" rather than "optimality" is commonly all that is achieved. To improve spacecraft design in the "faster, better, cheaper" era, a new approach using multidisciplinary design optimization (MDO) is proposed. Using MDO methods brings structure to conceptual spacecraft design by casting a spacecraft design problem into an optimization framework. Then, through the construction of a model that captures design and cost, this approach facilitates a quicker and more straightforward option synthesis. The final step is to automatically search the design space. As computer processor speed continues to increase, enumeration of all combinations, while not elegant, is one method that is straightforward to perform. As an alternative to enumeration, genetic algorithms are used and find solutions by reviewing fewer possible solutions with some limitations. Both methods increase the likelihood of finding an optimal design, or at least the most promising area of the design space. This spacecraft design methodology using MDO is demonstrated on three examples. A retrospective test for validation is performed using the Near Earth Asteroid Rendezvous (NEAR) spacecraft design. For the second example, the premise that aerobraking was needed to minimize mission cost and was mission enabling for the Mars Global Surveyor (MGS) mission is challenged. While one might expect no feasible design space for an MGS without aerobraking mission, a counterintuitive result is discovered. Several design options that don't use aerobraking are feasible and cost effective. The third example is an original commercial lunar mission entitled Eagle-eye. This example shows how an MDO approach is applied to an original mission with a larger feasible design space. It also incorporates a simplified business case analysis.

Advanced space power requirements and techniques. Task 1: Mission projections and requirements. Volume 3: Appendices. [cost estimates and computer programs

NASA Technical Reports Server (NTRS)

Wolfe, M. G.

1978-01-01

Contents: (1) general study guidelines and assumptions; (2) launch vehicle performance and cost assumptions; (3) satellite programs 1959 to 1979; (4) initiative mission and design characteristics; (5) satellite listing; (6) spacecraft design model; (7) spacecraft cost model; (8) mission cost model; and (9) nominal and optimistic budget program cost summaries.

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

NASA Technical Reports Server (NTRS)

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.

Spacecraft formation flying for Earth-crossing object deflections using a power limited laser ablating

NASA Astrophysics Data System (ADS)

Yoo, Sung-Moon; Song, Young-Joo; Park, Sang-Young; Choi, Kyu-Hong

2009-06-01

A formation flying strategy with an Earth-crossing object (ECO) is proposed to avoid the Earth collision. Assuming that a future conceptual spacecraft equipped with a powerful laser ablation tool already rendezvoused with a fictitious Earth collision object, the optimal required laser operating duration and direction histories are accurately derived to miss the Earth. Based on these results, the concept of formation flying between the object and the spacecraft is applied and analyzed as to establish the spacecraft's orbital motion design strategy. A fictitious "Apophis"-like object is established to impact with the Earth and two major deflection scenarios are designed and analyzed. These scenarios include the cases for the both short and long laser operating duration to avoid the Earth impact. Also, requirement of onboard laser tool's for both cases are discussed. As a result, the optimal initial conditions for the spacecraft to maintain its relative trajectory to the object are discovered. Additionally, the discovered optimal initial conditions also satisfied the optimal required laser operating conditions with no additional spacecraft's own fuel expenditure to achieve the spacecraft formation flying with the ECO. The initial conditions founded in the current research can be used as a spacecraft's initial rendezvous points with the ECO when designing the future deflection missions with laser ablation tools. The results with proposed strategy are expected to make more advances in the fields of the conceptual studies, especially for the future deflection missions using powerful laser ablation tools.

TROPIX plasma interactions group report

NASA Astrophysics Data System (ADS)

Herr, Joel L.; Chock, Ricaurte

1993-10-01

The purpose is to summarize the spacecraft charging analysis conducted by the plasma interactions group during the period from April 1993 to July 1993, on the proposed TROPIX spacecraft, and to make design recommendations which will limit the detrimental effects introduced by spacecraft charging. The recommendations were presented to the TROPIX study team at a Technical Review meeting held on 15 July 1993.

TROPIX plasma interactions group report

NASA Technical Reports Server (NTRS)

Herr, Joel L.; Chock, Ricaurte

1993-01-01

The purpose is to summarize the spacecraft charging analysis conducted by the plasma interactions group during the period from April 1993 to July 1993, on the proposed TROPIX spacecraft, and to make design recommendations which will limit the detrimental effects introduced by spacecraft charging. The recommendations were presented to the TROPIX study team at a Technical Review meeting held on 15 July 1993.

Suitable configurations for triangular formation flying about collinear libration points under the circular and elliptic restricted three-body problems

NASA Astrophysics Data System (ADS)

Ferrari, Fabio; Lavagna, Michèle

2018-06-01

The design of formations of spacecraft in a three-body environment represents one of the most promising challenges for future space missions. Two or more cooperating spacecraft can greatly answer some very complex mission goals, not achievable by a single spacecraft. The dynamical properties of a low acceleration environment such as the vicinity of libration points associated to a three-body system, can be effectively exploited to design spacecraft configurations able of satisfying tight relative position and velocity requirements. This work studies the evolution of an uncontrolled formation orbiting in the proximity of periodic orbits about collinear libration points under the Circular and Elliptic Restricted Three-Body Problems. A three spacecraft triangularly-shaped formation is assumed as a representative geometry to be investigated. The study identifies initial configurations that provide good performance in terms of formation keeping, and investigates key parameters that control the relative dynamics between the spacecraft within the three-body system. Formation keeping performance is quantified by monitoring shape and size changes of the triangular formation. The analysis has been performed under five degrees of freedom to define the geometry, the orientation and the location of the triangle in the synodic rotating frame.

BioSentinel: Enabling CubeSat-Scale Biogical Research Beyond Low Earth Orbit

NASA Technical Reports Server (NTRS)

Sorgenfrei, Matt; Lewis, Brian S.

2014-01-01

The introduction of the Space Launch System will provide NASA with a new means of access to space beyond low Earth orbit (LEO), creating opportunities for scientific research in a range of spacecraft sizes. This presentation describes the preliminary design of the BioSentinel spacecraft, a CubeSat measuring 10cm x 20cm x 30cm, which has been manifested for launch on the maiden voyage of the Space Launch System in 2017. BioSentinel will provide the first direct experimental data from a biological study conducted beyond LEO in over forty years, which in turn will help to pave the way for future human exploration missions. The combination of an advanced biology payload with standard spacecraft bus components required for operation in deep space within a CubeSat form factor poses a unique challenge, and this paper will describe the early design trades under consideration. The baseline spacecraft design calls for the biology payload to occupy four cube-units of volume (denoted 4U), with all spacecraft bus components occupying the remaining 2U.

Rapid Spacecraft Development: Results and Lessons Learned

NASA Technical Reports Server (NTRS)

Watson, William A.

2002-01-01

The Rapid Spacecraft Development Office (RSDO) at NASA's Goddard Space Flight Center is responsible for the management and direction of a dynamic and versatile program for the definition, competition, and acquisition of multiple indefinite delivery and indefinite quantity contracts - resulting in a catalog of spacecraft buses. Five spacecraft delivery orders have been placed by the RSDO and one spacecraft has been launched. Numerous concept and design studies have been performed, most with the intent of leading to a future spacecraft acquisition. A collection of results and lessons learned is recorded to highlight management techniques, methods and processes employed in the conduct of spacecraft acquisition. Topics include working relationships under fixed price delivery orders, price and value, risk management, contingency reserves, and information restrictions.

Taurus lightweight manned spacecraft Earth orbiting vehicle

NASA Technical Reports Server (NTRS)

Chase, Kevin A.; Vandersall, Eric J.; Plotkin, Jennifer; Travisano, Jeffrey J.; Loveless, Dennis; Kaczmarek, Michael; White, Anthony G.; Est, Andy; Bulla, Gregory; Henry, Chris

1991-01-01

The Taurus Lightweight Manned Spacecraft (LMS) was developed by students of the University of Maryland's Aerospace Engineering course in Space Vehicle Design. That course required students to design an Alternative Manned Spacecraft (AMS) to augment or replace the Space Transportation System and meet the following design requirements: (1) launch on the Taurus Booster being developed by Orbital Sciences Corporation; (2) 99.9 percent assured crew survival rate; (3) technology cutoff data of 1 Jan. 1991; (4) compatibility with current space administration infrastructure; and (5) first flight by May 1995. The Taurus LMS design meets the above requirements and represents an initial step towards larger and more complex spacecraft. The Taurus LMS has a very limited application when compared to the Space Shuttle, but it demonstrates that the U.S. can have a safe, reliable, and low cost space system. The Taurus LMS is a short mission duration spacecraft designed to place one man into low earth orbit (LEO). The driving factor for this design was the low payload carrying capabilities of the Taurus Booster--1300 kg to a 300 km orbit. The Taurus LMS design is divided into six major design sections. The human factors system deals with the problems of life support and spacecraft cooling. The propulsion section contains the abort system, the Orbital Maneuvering System (OMS), the Reaction Control System (RCS), and power generation. The thermal protection systems and spacecraft structure are contained in the structures section. The avionics section includes navigation, attitude determination, data processing, communication systems, and sensors. The mission analysis section was responsible for ground processing and spacecraft astrodynamics. The systems integration section pulled the above sections together into one spacecraft and addressed costing and reliability.

Taurus Lightweight Manned Spacecraft Earth orbiting vehicle

NASA Technical Reports Server (NTRS)

Bosset, M.

1991-01-01

The Taurus Lightweight Manned Spacecraft (LMS) was developed by students of the University of Maryland's Aerospace Engineering course in Space Vehicle Design. That course required students to design an Alternative Manned Spacecraft (AMS) to augment or replace the Space Transportation System and meet the following design requirements: (1) launch on the Taurus Booster being developed by Orbital Sciences Corporation; (2) 99.9 percent assured crew survival rate; (3) technology cutoff date of 1 Jan. 1991; (4) compatibility with current space administration infrastructure; and (5) first flight by May 1995. The Taurus LMS design meets the above requirements and represents an initial step toward larger and more complex spacecraft. The Taurus LMS has a very limited application when compared to the space shuttle, but it demonstrates that the U.S. can have a safe, reliable, and low-cost space system. The Taurus LMS is a short mission duration spacecraft designed to place one man into low Earth orbit (LEO). The driving factor for this design was the low payload carrying capabilities of the Taurus Booster - 1300 kg to a 300-km orbit. The Taurus LMS design is divided into six major design sections. The Human Factors section deals with the problems of life support and spacecraft cooling. The Propulsion section contains the Abort System, the Orbital Maneuvering System (OMS), the Reaction Control System (RCS), and Power Generation. The thermal protection systems and spacecraft structure are contained in the Structures section. The Avionics section includes Navigation, Attitude Determination, Data Processing, Communication systems, and Sensors. The Mission Analysis section was responsible for ground processing and spacecraft astrodynamics. The Systems Integration Section pulled the above sections together into one spacecraft, and addressed costing and reliability.

Design and Development of an Equipotential Voltage Reference (Grounding) System for a Low-Cost Rapid-Development Modular Spacecraft Architecture

NASA Technical Reports Server (NTRS)

Lukash, James A.; Daley, Earl

2011-01-01

This work describes the design and development effort to adapt rapid-development space hardware by creating a ground system using solutions of low complexity, mass, & cost. The Lunar Atmosphere and Dust Environment Explorer (LADEE) spacecraft is based on the modular common spacecraft bus architecture developed at NASA Ames Research Center. The challenge was building upon the existing modular common bus design and development work and improving the LADEE spacecraft design by adding an Equipotential Voltage Reference (EVeR) system, commonly referred to as a ground system. This would aid LADEE in meeting Electromagnetic Environmental Effects (E3) requirements, thereby making the spacecraft more compatible with itself and its space environment. The methods used to adapt existing hardware are presented, including provisions which may be used on future spacecraft.

Statistical Rick Estimation for Communication System Design --- A Preliminary Look

NASA Astrophysics Data System (ADS)

Babuscia, A.; Cheung, K.-M.

2012-02-01

Spacecraft are complex systems that involve different subsystems with multiple relationships among them. For these reasons, the design of a spacecraft is a time-evolving process that starts from requirements and evolves over time across different design phases. During this process, a lot of changes can happen. They can affect mass and power at the component level, at the subsystem level, and even at the system level. Each spacecraft has to respect the overall constraints in terms of mass and power: for this reason, it is important to be sure that the design does not exceed these limitations. Current practice in system models primarily deals with this problem, allocating margins on individual components and on individual subsystems. However, a statistical characterization of the fluctuations in mass and power of the overall system (i.e., the spacecraft) is missing. This lack of adequate statistical characterization would result in a risky spacecraft design that might not fit the mission constraints and requirements, or in a conservative design that might not fully utilize the available resources. Due to the complexity of the problem and to the different expertise and knowledge required to develop a complete risk model for a spacecraft design, this article is focused on risk estimation for a specific spacecraft subsystem: the communication subsystem. The current research aims to be a proof of concept of a risk-based design optimization approach, which can then be further expanded to the design of other subsystems as well as to the whole spacecraft. The objective of this research is to develop a mathematical approach to quantify the likelihood that the major design drivers of mass and power of a space communication system would meet the spacecraft and mission requirements and constraints through the mission design lifecycle. Using this approach, the communication system designers will be able to evaluate and to compare different communication architectures in a risk trade-off perspective. The results described in this article include a baseline communication system design tool and a statistical characterization of the design risks through a combination of historical mission data and expert opinion contributions. An application example of the communication system of a university spacecraft is presented. IPNPR Volume 42-189 Tagged File.txt

Design Challenges of Power Systems for Instrumented Spacecraft with Very Low Perigees in the Earth's Ionosphere

NASA Technical Reports Server (NTRS)

Moran, Vickie Eakin; Manzer, Dominic D.; Pfaff, Robert E.; Grebowsky, Joseph M.; Gervin, Jan C.

1999-01-01

Designing a solar array to power a spacecraft bus supporting a set of instruments making in situ plasma and neutral atmosphere measurements in the ionosphere at altitudes of 120km or lower poses several challenges. The driving scientific requirements are the field-of-view constraints of the instruments resulting in a three-axis stabilized spacecraft, the need for an electromagnetically unperturbed environment accomplished by designing an electrostatically conducting solar array surface to avoid large potentials, making the spacecraft body as small and as symmetric as possible, and body-mounting the solar array. Furthermore, the life and thermal constraints, in the midst of the effects of the dense atmosphere at low altitude, drive the cross-sectional area of the spacecraft to be small particularly normal to the ram direction. Widely varying sun angles and eclipse durations add further complications, as does the growing desire for multiple spacecraft to resolve spatial and temporal variations packaged into a single launch vehicle. Novel approaches to insure adequate orbit-averaged power levels of approximately 250W include an oval-shaped cross section to increase the solar array collecting area during noon-midnight orbits and the use of a flywheel energy storage system. The flywheel could also be used to help maintain the spacecraft's attitude, particularly during excursions to the lowest perigee altitudes. This paper discusses the approaches used in conceptual power designs for both the proposed Dipper and the Global Electrodynamics Connections (GEC) Mission currently being studied at the NASA/Goddard Space Flight Center.

Using SFOC to fly the Magellan Venus mapping mission

NASA Technical Reports Server (NTRS)

Bucher, Allen W.; Leonard, Robert E., Jr.; Short, Owen G.

1993-01-01

Traditionally, spacecraft flight operations at the Jet Propulsion Laboratory (JPL) have been performed by teams of spacecraft experts utilizing ground software designed specifically for the current mission. The Jet Propulsion Laboratory set out to reduce the cost of spacecraft mission operations by designing ground data processing software that could be used by multiple spacecraft missions, either sequentially or concurrently. The Space Flight Operations Center (SFOC) System was developed to provide the ground data system capabilities needed to monitor several spacecraft simultaneously and provide enough flexibility to meet the specific needs of individual projects. The Magellan Spacecraft Team utilizes the SFOC hardware and software designed for engineering telemetry analysis, both real-time and non-real-time. The flexibility of the SFOC System has allowed the spacecraft team to integrate their own tools with SFOC tools to perform the tasks required to operate a spacecraft mission. This paper describes how the Magellan Spacecraft Team is utilizing the SFOC System in conjunction with their own software tools to perform the required tasks of spacecraft event monitoring as well as engineering data analysis and trending.

Aerothermal Analysis and Design of the Gravity Recovery and Climate Experiment (GRACE) Spacecraft

NASA Technical Reports Server (NTRS)

Mazanek, Daniel D.; Kumar, Renjith R.; Qu, Min; Seywald, Hans

2000-01-01

The Gravity Recovery and Climate Experiment (GRACE) primary mission will be performed by making measurements of the inter-satellite range change between two co-planar, low altitude near-polar orbiting satellites. Understanding the uncertainties in the disturbance environment, particularly the aerodynamic drag and torques, is critical in several mission areas. These include an accurate estimate of the spacecraft orbital lifetime, evaluation of spacecraft attitude control requirements, and estimation of the orbital maintenance maneuver frequency necessitated by differences in the drag forces acting on both satellites. The FREEMOL simulation software has been developed and utilized to analyze and suggest design modifications to the GRACE spacecraft. Aerodynamic accommodation bounding analyses were performed and worst-case envelopes were obtained for the aerodynamic torques and the differential ballistic coefficients between the leading and trailing GRACE spacecraft. These analyses demonstrate how spacecraft aerodynamic design and analysis can benefit from a better understanding of spacecraft surface accommodation properties, and the implications for mission design constraints such as formation spacing control.

Extensive Radiation Shielding Analysis for Different Spacecraft Orbits

NASA Astrophysics Data System (ADS)

Çay, Yiǧit; Kaymaz, Zerefsan

2016-07-01

Radiation environment around Earth poses a great danger for spacecraft and causes immature de-orbiting or loss of the spacecraft in near Earth space environment. In this study, a student project has been designed to build a CubeSat, PolarBeeSail (PBS), with an orbit having inclination of 80°, 4 Re in perigee and 20 Re in apogee to study the polar magnetospheric environment. An extensive radiation dose analyses were carried out for PBS orbit, and integral and differential fluxes were calculated using SPENVIS tools. A shielding analysis was performed and an optimum Aluminum thickness, 3 mm, was obtained. These results for PBS were then compared for other orbits at different altitudes both for polar and equatorial orbits. For this purpose, orbital characteristics of POES-19 and GOES-15 were used. The resulting proton flux analyses, TID analyses, and further shielding studies were conducted; comparisons and recommendations were made for future design of spacecraft that will use these environments.

Making Human Spaceflight Practical and Affordable: Spacecraft Designs and their Degree of Operability

NASA Technical Reports Server (NTRS)

Crocker, Alan R.

2011-01-01

As we push toward new and diverse space transportation capabilities, reduction in operations cost becomes increasingly important. Achieving affordable and safe human spaceflight capabilities will be the mark of success for new programs and new providers. The ability to perceive the operational implications of design decisions is crucial in developing safe yet cost competitive space transportation systems. Any human spaceflight program - government or commercial - must make countless decisions either to implement spacecraft system capabilities or adopt operational constraints or workarounds to account for the lack of such spacecraft capabilities. These decisions can benefit from the collective experience that NASA has accumulated in building and operating crewed spacecraft over the last five decades. This paper reviews NASA s history in developing and operating human rated spacecraft, reviewing the key aspects of spacecraft design and their resultant impacts on operations phase complexity and cost. Specific examples from current and past programs - including the Space Shuttle and International Space Station - are provided to illustrate design traits that either increase or increase cost and complexity associated with spacecraft operations. These examples address factors such as overall design performance margins, levels of redundancy, degree of automated failure response, type and quantity of command and telemetry interfaces, and the definition of reference scenarios for analysis and test. Each example - from early program requirements, design implementation and resulting real-time operations experience - to tell the end-to-end "story" Based on these experiences, specific techniques are recommended to enable earlier and more effective assessment of operations concerns during the design process. A formal method for the assessment of spacecraft operability is defined and results of such operability assessments for recent spacecraft designs are provided. Recent experience in applying these techniques to Orion spacecraft development is reviewed to highlight the direct benefits of early operational assessment and collaborative development efforts.

Anomaly Trends for Missions to Mars: Mars Global Surveyor and Mars Odyssey

NASA Technical Reports Server (NTRS)

Green, Nelson W.; Hoffman, Alan R.

2008-01-01

Conducted as a part of NASA Ultra-Reliability effort: Goal is to design for increased reliability in all NASA missions. Desire is to increase reliability by a factor of 10. Study provides a baseline for current technology. Analyzed anomalies for spacecraft orbiting Mars. Long lived spacecraft. Comparison with current rover missions and past orbiters. Looked for trends to assist design of future missions.

Cluster: A fleet of four spacecraft to study plasma structures in three dimensions

NASA Technical Reports Server (NTRS)

Schmidt, R.; Goldstein, M. L.

1988-01-01

The four Cluster spacecraft are spin stabilized spacecraft which are designed and built under stringent requirements as far as electromagnetic cleanliness is concerned. Conductive surfaces and low electromagnetic background noise are mandatory for accurate electric field and cold plasma measurements. The mission is implemented in collaboration between ESA and NASA. A Russian mission will be closely coordinated with Cluster.

NIR Color vs Launch Date: A 20-year Analysis of Space Weathering Effects on the Boeing 376 Spacecraft

NASA Astrophysics Data System (ADS)

Frith, J.; Anz-Meador, P.; Lederer, S.; Cowardin, H.; Buckalew, B.

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 2002. 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.

NIR Color vs Launch Date: A 20-Year Analysis of Space Weathering Effects on the Boeing 376 Spacecraft

NASA Technical Reports Server (NTRS)

Frith, James; Anz-Meador, Philip; Lederer, Sue; Cowardin, Heather; Buckalew, Brent

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 2002. 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.

Spacecraft attitude control for a solar electric geosynchronous transfer mission

NASA Technical Reports Server (NTRS)

Leroy, B. E.; Regetz, J. D., Jr.

1975-01-01

A study of the Attitude Control System (ACS) is made for a solar electric propulsion geosynchronous transfer mission. The basic mission considered is spacecraft injection into a low altitude, inclined orbit followed by low thrust orbit changing to achieve geosynchronous orbit. Because of the extended thrusting time, the mission performance is a strong function of the attitude control system. Two attitude control system design options for an example mission evolve from consideration of the spacecraft configuration, the environmental disturbances, and the probable ACS modes of operation. The impact of these design options on other spacecraft subsystems is discussed. The factors which must be considered in determining the ACS actuation and sensing subsystems are discussed. The effects of the actuation and sensing subsystems on the mission performance are also considered.

Spacecraft Systems Engineering, 3rd Edition

NASA Astrophysics Data System (ADS)

Fortescue, Peter; Stark, John; Swinerd, Graham

2003-03-01

Following on from the hugely successful previous editions, the third edition of Spacecraft Systems Engineering incorporates the most recent technological advances in spacecraft and satellite engineering. With emphasis on recent developments in space activities, this new edition has been completely revised. Every chapter has been updated and rewritten by an expert engineer in the field, with emphasis on the bus rather than the payload. Encompassing the fundamentals of spacecraft engineering, the book begins with front-end system-level issues, such as environment, mission analysis and system engineering, and progresses to a detailed examination of subsystem elements which represent the core of spacecraft design - mechanical, electrical, propulsion, thermal, control etc. This quantitative treatment is supplemented by an appreciation of the interactions between the elements, which deeply influence the process of spacecraft systems design. In particular the revised text includes * A new chapter on small satellites engineering and applications which has been contributed by two internationally-recognised experts, with insights into small satellite systems engineering. * Additions to the mission analysis chapter, treating issues of aero-manouevring, constellation design and small body missions. In summary, this is an outstanding textbook for aerospace engineering and design students, and offers essential reading for spacecraft engineers, designers and research scientists. The comprehensive approach provides an invaluable resource to spacecraft manufacturers and agencies across the world.

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

NASA Technical Reports Server (NTRS)

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 to 420 days in duration and assuming chemical propulsion. Similar studies have been reported assuming high power electric propulsion for manned NEO rendezvous missions (11). The delta V requirement breakdown and mission profile data from references 10 and 11 are used as a basis for sizing the NEO Scout spacecraft and for conducting preliminary feasibility assessments using the Tsiokolvsky rocket equation, a (worst-case) delta V requirement of 10 km/sec, and a maximum spacecraft dry mass of 20 kg. Using chemical propellant for a 10 km/sec delta V drives spacecraft wet mass well above 300 kg so that chemical propulsion is a non-starter for the proposed mission profile and spacecraft wet mass limits. In contrast, a solar electric propulsion system needs only 8 kg of Xe propellant to accelerate the spacecraft to 10 km/sec in 163 days with 0.02 N of thrust and 500 W of power from1.6 sq m of 29% efficient solar panels. In a second example, accelerating a 4 kg payload to 7 km/sec over 180 days requires about 6.7 kg of propellant and 1.2 kg of solar panels (12 kg total spacecraft wet mass).

Thermal Design, Test and Analysis of PharmaSat, a Small Class D Spacecraft with a Biological Experiment

NASA Technical Reports Server (NTRS)

Diaz-Aguado, Millan F.; VanOutryve, Cassandra; Ghassemiah, Shakib; Beasley, Christopher; Schooley, Aaron

2009-01-01

Small spacecraft have been increasing in popularity because of their low cost, short turnaround and relative efficiency. In the past, small spacecraft have been primarily used for technology demonstrations, but advances in technology have made the miniaturization of space science possible [1,2]. PharmaSat is a low cost, small three cube size spacecraft, with a biological experiment on board, built at NASA (National Aeronautics and Space Administration) Ames Research Center. The thermal design of small spacecraft presents challenges as their smaller surface areas translate into power and thermal constraints. The spacecraft is thermally designed to run colder in the Low Earth Orbit space environment, and heated to reach the temperatures required by the science payload. The limited power supply obtained from the solar panels on small surfaces creates a constraint in the power used to heat the payload to required temperatures. The pressurized payload is isolated with low thermally conductance paths from the large ambient temperature changes. The thermal design consists of different optical properties of section surfaces, Multi Layer Insulation (MLI), low thermal conductance materials, flexible heaters and thermal spreaders. The payload temperature is controlled with temperature sensors and flexible heaters. Finite Element Analysis (FEA) and testing were used to aid the thermal design of the spacecraft. Various tests were conducted to verify the thermal design. An infrared imager was used on the electronic boards to find large heat sources and eliminate any possible temperature runaways. The spacecraft was tested in a thermal vacuum chamber to optimize the thermal and power analysis and qualify the thermal design of the spacecraft for the mission.

Laser communication experiment. Volume 1: Design study report: Spacecraft transceiver. Part 3: LCE design specifications

NASA Technical Reports Server (NTRS)

1970-01-01

The requirements for the design, fabrication, performance, and testing of a 10.6 micron optical heterodyne receiver subsystem for use in a laser communication system are presented. The receiver subsystem, as a part of the laser communication experiment operates in the ATS 6 satellite and in a transportable ground station establishing two-way laser communications between the spacecraft and the transportable ground station. The conditions under which environmental tests are conducted are reported.

Technology requirements for a generic aerocapture system. [for atmospheric entry

NASA Technical Reports Server (NTRS)

Cruz, M. I.

1980-01-01

The technology requirements for the design of a generic aerocapture vehicle system are summarized. These spacecraft have the capability of completely eliminating fuel-costly retropropulsion for planetary orbit capture through a single aerodynamically controlled atmospheric braking pass from a hyperbolic trajectory into a near circular orbit. This generic system has application at both the inner and outer planets. Spacecraft design integration, navigation, communications, and aerothermal protection system design problems were assessed in the technology requirements study and are discussed in this paper.

Inner structural vibration isolation method for a single control moment gyroscope

NASA Astrophysics Data System (ADS)

Zhang, Jingrui; Guo, Zixi; Zhang, Yao; Tang, Liang; Guan, Xin

2016-01-01

Assembling and manufacturing errors of control moment gyros (CMG) often generate high frequency vibrations which are detrimental to spacecrafts with high precision pointing requirement. In this paper, some design methods of vibration isolation between CMG and spacecraft is dealt with. As a first step, the dynamic model of the CMG with and without supporting isolation structures is studied and analyzed. Subsequently, the frequency domain analysis of CMG with isolation system is performed and the effectiveness of the designed system is ascertained. Based on the above studies, an adaptive design suitable with appropriate design parameters are carried out. A numerical analysis is also performed to understand the effectiveness of the system and the comparison made. The simulation results clearly indicate that when the ideal isolation structure was implemented in the spacecraft, the vibrations generated by the rotor were found to be greatly reduced, while the capacity of the output torque was not lost, which means that the isolation system will not affect the performance of attitude control.

Overview of SDCM - The Spacecraft Design and Cost Model

NASA Technical Reports Server (NTRS)

Ferebee, Melvin J.; Farmer, Jeffery T.; Andersen, Gregory C.; Flamm, Jeffery D.; Badi, Deborah M.

1988-01-01

The Spacecraft Design and Cost Model (SDCM) is a computer-aided design and analysis tool for synthesizing spacecraft configurations, integrating their subsystems, and generating information concerning on-orbit servicing and costs. SDCM uses a bottom-up method in which the cost and performance parameters for subsystem components are first calculated; the model then sums the contributions from individual components in order to obtain an estimate of sizes and costs for each candidate configuration within a selected spacecraft system. An optimum spacraft configuration can then be selected.

Geopotential research mission, science, engineering and program summary

NASA Technical Reports Server (NTRS)

Keating, T. (Editor); Taylor, P. (Editor); Kahn, W. (Editor); Lerch, F. (Editor)

1986-01-01

This report is based upon the accumulated scientific and engineering studies pertaining to the Geopotential Research Mission (GRM). The scientific need and justification for the measurement of the Earth's gravity and magnetic fields are discussed. Emphasis is placed upon the studies and conclusions of scientific organizations and NASA advisory groups. The engineering design and investigations performed over the last 4 years are described, and a spacecraft design capable of fulfilling all scientific objectives is presented. In addition, critical features of the scientific requirements and state-of-the-art limitations of spacecraft design, mission flight performance, and data processing are discussed.

Mars Polar Lander undergoes testing in SAEF-2

NASA Technical Reports Server (NTRS)

1998-01-01

In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), KSC technicians lower the Mars Polar Lander onto a workstand. The spacecraft is undergoing testing of science instruments and basic spacecraft subsystems. The solar-powered spacecraft, targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999, is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The Lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere.

Mars Polar Lander undergoes testing in SAEF-2

NASA Technical Reports Server (NTRS)

1998-01-01

In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), KSC technicians look over the Mars Polar Lander. The spacecraft is undergoing testing of science instruments and basic spacecraft subsystems. Targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999, the solar-powered spacecraft is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The Lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere.

KSC-98pc1374

NASA Image and Video Library

1998-10-22

KENNEDY SPACE CENTER, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), workers check out the solar panel on the Mars Polar Lander. The spacecraft is undergoing testing of science instruments and basic spacecraft subsystems. The solar-powered spacecraft, targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999, is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

KSC-98pc1373

NASA Image and Video Library

1998-10-22

KENNEDY SPACE CENTER, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), workers adjust the Mars Polar Lander on its workstand. The spacecraft is undergoing testing of science instruments and basic spacecraft subsystems. The solar-powered spacecraft, targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999, is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

KSC-98pc1599

NASA Image and Video Library

1998-10-23

In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), KSC technicians begin to lift the Mars Polar Lander to move it to a workstand. The spacecraft is undergoing testing of science instruments and basic spacecraft subsystems. The solar-powered spacecraft, targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999, is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

KSC-98pc1372

NASA Image and Video Library

1998-10-22

KENNEDY SPACE CENTER, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), a technician checks out the Mars Polar Lander on its workstand. The spacecraft is undergoing testing of science instruments and basic spacecraft subsystems. The solar-powered spacecraft, targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999, is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

KSC-98pc1600

NASA Image and Video Library

1998-10-23

In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), KSC technicians guide the raised Mars Polar Lander to another site. The spacecraft is undergoing testing of science instruments and basic spacecraft subsystems. The solar-powered spacecraft, targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999, is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

KSC-98pc1601

NASA Image and Video Library

1998-10-23

In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), KSC technicians lower the Mars Polar Lander onto a workstand. The spacecraft is undergoing testing of science instruments and basic spacecraft subsystems. The solar-powered spacecraft, targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999, is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The Lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

KSC-98pc1605

NASA Image and Video Library

1998-10-29

KENNEDY SPACE CENTER, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), KSC technicians look over the Mars Polar Lander. The spacecraft is undergoing testing of science instruments and basic spacecraft subsystems. Targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999, the solar-powered spacecraft is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The Lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

Systems design and analysis of the microwave radiometer spacecraft

NASA Technical Reports Server (NTRS)

Garrett, L. B.

1981-01-01

Systems design and analysis data were generated for microwave radiometer spacecraft concept using the Large Advanced Space Systems (LASS) computer aided design and analysis program. Parametric analyses were conducted for perturbations off the nominal-orbital-altitude/antenna-reflector-size and for control/propulsion system options. Optimized spacecraft mass, structural element design, and on-orbit loading data are presented. Propulsion and rigid-body control systems sensitivities to current and advanced technology are established. Spacecraft-induced and environmental effects on antenna performance (surface accuracy, defocus, and boresight off-set) are quantified and structured material frequencies and modal shapes are defined.

Benefits of Application of Advanced Technologies for a Neptune Orbiter, Atmospheric Probes and Triton Lander

NASA Technical Reports Server (NTRS)

Somers, Alan; Celano, Luigi; Kauffman, Jeffrey; Rogers, Laura; Peterson, Craig

2005-01-01

Missions with planned launch dates several years from today pose significant design challenges in properly accounting for technology advances that may occur in the time leading up to actual spacecraft design, build, test and launch. Conceptual mission and spacecraft designs that rely solely on off the shelf technology will result in conservative estimates that may not be attractive or truly representative of the mission as it actually will be designed and built. This past summer, as part of one of NASA s Vision Mission Studies, a group of students at the Laboratory for Spacecraft and Mission Design (LSMD) have developed and analyzed different Neptune mission baselines, and determined the benefits of various assumed technology improvements. The baseline mission uses either a chemical propulsion system or a solar-electric system. Insertion into orbit around Neptune is achieved by means of aerocapture. Neptune s large moon Triton is used as a tour engine. With these technologies a comprehensive Cassini-class investigation of the Neptune system is possible. Technologies under investigation include the aerocapture heat shield and thermal protection system, both chemical and solar electric propulsion systems, spacecraft power, and energy storage systems.

Proceedings of the Spacecraft Charging Technology Conference

NASA Technical Reports Server (NTRS)

Pike, C. P. (Editor); Lovell, R. R. (Editor)

1977-01-01

Over 50 papers from the spacecraft charging conference are included on subjects such as: (1) geosynchronous plasma environment, (2) spacecraft modeling, (3) spacecraft materials characterization, (4) spacecraft materials development, and (5) satellite design and test.

Principles of Timekeeping for the NEAR and STEREO Spacecraft

NASA Technical Reports Server (NTRS)

Cooper, Stanley B.; Wolff, J. (Technical Monitor)

2001-01-01

This paper discusses the details of the inherently different timekeeping systems for two interplanetary missions, the NEAR Shoemaker mission to orbit the near-Earth asteroid 433 Eros and the STEREO (Solar Terrestrial Relations Observatory) mission to study and characterize solar coronal mass ejections. It also reveals the surprising dichotomy between two major categories of spacecraft timekeeping systems with respect to the relationship between spacecraft clock resolution and accuracy. The paper is written in a tutorial style so that it can be easily used as a reference for designing or analyzing spacecraft timekeeping systems.

Planetary quarantine. Space research and technology

NASA Technical Reports Server (NTRS)

1973-01-01

The impact of satisfying satellite quarantine constraints on outer planet missions and spacecraft design are studied by considering the effects of planetary radiation belts, solar wind radiation, and space vacuum on microorganism survival. Post launch recontamination studies evaluate the effects of mission environments on particle distributions on spacecraft surfaces and effective cleaning and decontamination techniques.

Pioneer spacecraft operation at low and high spin rates

NASA Technical Reports Server (NTRS)

1973-01-01

The feasibility of executing major changes upward or downward from the nominal spin rate for which the Pioneer F&G spacecraft was designed was investigated along with the extent of system and subsystem modifications required to implement these mode changes in future spacecraft evolving from the baseline Pioneer F and G. Results of a previous study are re-examined and updated for an extended range of spin rate variations for missions that include outer planet orbiters, outer planet flyby and outer planet probe delivery. However, in the interest of design simplicity and cost economy, major modifications of the baseline Pioneer system and subsystem concept were avoided.

Advanced X-Ray Timing Array Mission: Conceptual Spacecraft Design Study

NASA Technical Reports Server (NTRS)

Hopkins, R. C.; Johnson, L.; Thomas, H. D.; Wilson-Hodge, C. A.; Baysinger, M.; Maples, C. D.; Fabisinski, L.L.; Hornsby, L.; Thompson, K. S.; Miernik, J. H.

2011-01-01

The Advanced X-Ray Timing Array (AXTAR) is a mission concept for submillisecond timing of bright galactic x-ray sources. The two science instruments are the Large Area Timing Array (LATA) (a collimated instrument with 2-50-keV coverage and over 3 square meters of effective area) and a Sky Monitor (SM), which acts as a trigger for pointed observations of x-ray transients. The spacecraft conceptual design team developed two spacecraft concepts that will enable the AXTAR mission: A minimal configuration to be launched on a Taurus II and a larger configuration to be launched on a Falcon 9 or similar vehicle.

Spacecraft level impacts of integrating concentrator solar arrays

DOE Office of Scientific and Technical Information (OSTI.GOV)

Allen, D.M.; Piszczor, M.F. Jr.

1994-12-31

The paper describes the results of a study to determine the impacts of integrating concentrator solar arrays on spacecraft design and performance. First, concentrator array performance is summarized for the AEC-Able/Entech SCARLET array, the Ioffe refractive and reflective concepts being developed in Russia, the Martin Marietta SLATS system, and other concentrator concepts that have been designed or developed. Concentrator array performance is compared to rigid and flex blanket planar array technologies at the array level. Then other impacts on the spacecraft are quantified. Conclusions highlight the most important results as they relate to recommended approaches in developing concentrator arrays formore » satellites.« less

A Study of Fluid Interface Configurations in Exploration Vehicle Propellant Tanks

NASA Technical Reports Server (NTRS)

Zimmerli, Gregory A.; Asipauskas, Marius; Chen, Yongkang; Weislogel, Mark M.

2010-01-01

The equilibrium shape and location of fluid interfaces in spacecraft propellant tanks while in low-gravity is of interest to system designers, but can be challenging to predict. The propellant position can affect many aspects of the spacecraft such as the spacecraft center of mass, response to thruster firing due to sloshing, liquid acquisition, propellant mass gauging, and thermal control systems. We use Surface Evolver, a fluid interface energy minimizing algorithm, to investigate theoretical equilibrium liquid-vapor interfaces for spacecraft propellant tanks similar to those that have been considered for NASA's new class of Exploration vehicles. The choice of tank design parameters we consider are derived from the NASA Exploration Systems Architecture Study report. The local acceleration vector employed in the computations is determined by estimating low-Earth orbit (LEO) atmospheric drag effects and centrifugal forces due to a fixed spacecraft orientation with respect to the Earth or Moon, and rotisserie-type spacecraft rotation. Propellant/vapor interface positions are computed for the Earth Departure Stage and Altair lunar lander descent and ascent stage tanks for propellant loads applicable to LEO and low-lunar orbit. In some of the cases investigated the vapor ullage bubble is located at the drain end of the tank, where propellant management device hardware is often located.

Spacecraft Charging Technology, 1980

NASA Technical Reports Server (NTRS)

1981-01-01

The third Spacecraft Charging Technology Conference proceedings contain 66 papers on the geosynchronous plasma environment, spacecraft modeling, charged particle environment interactions with spacecraft, spacecraft materials characterization, and satellite design and testing. The proceedings is a compilation of the state of the art of spacecraft charging and environmental interaction phenomena.

Earth Observatory Satellite system definition study. Report 7: EOS system definition report

NASA Technical Reports Server (NTRS)

1974-01-01

The Earth Observatory Satellite (EOS) study is summarized to show the modular design of a general purpose spacecraft, a mission peculiar segment which performs the EOS-A mission, an Operations Control Center, a Data Processing Facility, and a design for Low Cost Readout Stations. The study verified the practicality and feasibility of the modularized spacecraft with the capability of supporting many missions in the Earth Observation spectrum. The various subjects considered in the summary are: (1) orbit/launch vehicle tradeoff studies and recommendations, (2) instrument constraints and interfaces, (3) design/cost tradeoff and recommendations, (4) low cost management approach and recommendations, (5) baseline system description and specifications, and (6) space shuttle utilization and interfaces.

Influence of Natural Environments in Spacecraft Design, Development, and Operation

NASA Technical Reports Server (NTRS)

Edwards, Dave

2013-01-01

Spacecraft are growing in complexity and sensitivity to environmental effects. The spacecraft engineer must understand and take these effects into account in building reliable, survivable, and affordable spacecraft. Too much protections, however, means unnecessary expense while too little will potentially lead to early mission loss. The ability to balance cost and risk necessitates an understanding of how the environment impacts the spacecraft and is a critical factor in its design. This presentation is intended to address both the space environment and its effects with the intent of introducing the influence of the environment on spacecraft performance.

Influence of Natural Environments in Spacecraft Design, Development, and Operation

NASA Technical Reports Server (NTRS)

Edwards, Dave

2012-01-01

Spacecraft are growing in complexity and sensitivity to environmental effects. The spacecraft engineer must understand and take these effects into account in building reliable, survivable, and affordable spacecraft. Too much protections, however, means unnecessary expense while too little will potentially lead to early mission loss. The ability to balance cost and risk necessitates an understanding of how the environment impacts the spacecraft and is a critical factor in its design. This presentation is intended to address both the space environment and its effects with the intent of introducing the influence of the environment on spacecraft performance.

Preventing Spacecraft Failures Due to Tribological Problems

NASA Technical Reports Server (NTRS)

Fusaro, Robert L.

2001-01-01

Many mechanical failures that occur on spacecraft are caused by tribological problems. This publication presents a study that was conducted by the author on various preventatives, analyses, controls and tests (PACTs) that could be used to prevent spacecraft mechanical system failure. A matrix is presented in the paper that plots tribology failure modes versus various PACTs that should be performed before a spacecraft is launched in order to insure success. A strawman matrix was constructed by the author and then was sent out to industry and government spacecraft designers, scientists and builders of spacecraft for their input. The final matrix is the result of their input. In addition to the matrix, this publication describes the various PACTs that can be performed and some fundamental knowledge on the correct usage of lubricants for spacecraft applications. Even though the work was done specifically to prevent spacecraft failures the basic methodology can be applied to other mechanical system areas.

Mars Observer trajectory and orbit design

NASA Technical Reports Server (NTRS)

Beerer, Joseph G.; Roncoli, Ralph B.

1991-01-01

The Mars Observer launch, interplanetary, Mars orbit insertion, and mapping orbit designs are described. The design objective is to enable a near-maximum spacecraft mass to be placed in orbit about Mars. This is accomplished by keeping spacecraft propellant requirements to a minimum, selecting a minimum acceptable launch period, equalizing the spacecraft velocity change requirement at the beginning and end of the launch period, and constraining the orbit insertion maneuvers to be coplanar. The mapping orbit design objective is to provide the opportunity for global observation of the planet by the science instruments while facilitating the spacecraft design. This is realized with a sun-synchronous near-polar orbit whose ground-track pattern covers the planet at progressively finer resolution.

Tracking and data relay satellite system configuration and tradeoff study, part 1. Volume 1: Summary volume

NASA Technical Reports Server (NTRS)

1972-01-01

Study efforts directed at defining all TDRS system elements are summarized. Emphasis was placed on synthesis of a space segment design optimized to support low and medium data rate user spacecraft and launched with Delta 2914. A preliminary design of the satellite was developed and conceptual designs of the user spacecraft terminal and TDRS ground station were defined. As a result of the analyses and design effort it was determined that (1) a 3-axis-stabilized tracking and data relay satellite launched on a Delta 2914 provides telecommunications services considerably in excess of that required by the study statement; and (2) the design concept supports the needs of the space shuttle and has sufficient growth potential and flexibility to provide telecommunications services to high data rate users. Recommendations for further study are included.

LDEF Materials Results for Spacecraft Applications

NASA Technical Reports Server (NTRS)

Whitaker, Ann F. (Compiler); Gregory, John (Compiler)

1993-01-01

These proceedings describe the application of LDEF data to spacecraft and payload design, and emphasize where space environmental effects on materials research and development is needed as defined by LDEF data. The LDEF six years of exposure of materials has proven to be by far the most comprehensive source of information ever obtained on the long-term performance of materials in the space environment. The conference provided a forum for materials scientists and engineers to review and critically assess the LDEF results from the standpoint of their relevance, significance, and impact on spacecraft design practice. The impact of the LDEF findings on materials selection and qualification, and the needs and plans for further study, were addressed from several perspectives. Many timely and needed changes and modifications in external spacecraft materials selection have occurred as a result of LDEF investigations.

POWOW: A Modular, High Power Spacecraft Concept

NASA Technical Reports Server (NTRS)

Brandhorst, Henry W., Jr.

2000-01-01

A robust space infrastructure encompasses a broad range of mission needs along with an imperative to reduce costs of satellites meeting those needs. A critical commodity for science, commercial and civil satellites is power at an affordable cost. The POWOW (POwer WithOut Wires) spacecraft concept was created to provide, at one end of the scale, multi-megawatts of power yet also be composed of modules that can meet spacecraft needs in the kilowatt range. With support from the NASA-sponsored Space Solar Power Exploratory Research and Technology Program, the POWOW spacecraft concept was designed to meet Mars mission needs - while at the same time having elements applicable to a range of other missions. At Mars, the vehicle would reside in an aerosynchronous orbit and beam power to a variety of locations on the surface. It is the purpose of this paper to present the latest concept design results. The Space Power Institute along with four companies: Able Engineering, Inc., Entech, Inc., Primex Aerospace Co., and TECSTAR have produced a modular, power-rich electrically propelled spacecraft design that meets these requirements. In addition, it also meets a range of civil and commercial needs. The spacecraft design is based on multijunction Ill-V solar cells, the new Stretched Lens Aurora (SLA) module, a lightweight array design based on a multiplicity of 8 kW end-of-life subarrays and electric thrusters. The solar cells have excellent radiation resistance and efficiencies above 30%. The SLA has a concentration ratio up to 15x while maintaining an operating temperature of 80 C. The design of the 8 kW array building block will be presented and its applicability to commercial and government missions will be discussed. Electric propulsion options include Hall, MPD and ion thrusters of various power levels and trade studies have been conducted to define the most advantageous options. The present baseline spacecraft design providing 900 kW using technologies expected to be available in 2003 will be described. Areal power densities of nearly 400 W/meters squared at 80 C operating temperatures and wing level specific powers of over 400 W/kg are projected. Details of trip times and payloads to Mars will be presented as well as trade studies of various electric propulsion options. Trip times compare favorably with chemical propulsion options. Because the design is modular, learning curve methodology can be applied to determine expected cost reductions. These results will also be included. This paper has not been presented at a previous meeting.

Design/cost tradeoff studies. Earth Observatory Satellite system definition study (EOS)

NASA Technical Reports Server (NTRS)

1974-01-01

The results of design/cost tradeoff studies conducted during the Earth Observatory Satellite system definition studies are presented. The studies are concerned with the definition of a basic modular spacecraft capable of supporting a variety of operational and/or research and development missions, with the deployment either by conventional launch vehicles or by means of the space shuttle. The three levels investigated during the study are: (1) subsystem tradeoffs, (2) spacecraft tradeoffs, and (3) system tradeoffs. The range of requirements which the modular concept must span is discussed. The mechanical, thermal, power, data and electromagnetic compatibility aspects of modularity are analyzed. Other data are provided for the observatory design concept, the payloads, integration and test, the ground support equipment, and ground data management systems.

Parallel Estimation and Control Architectures for Deep-Space Formation Flying Spacecraft

NASA Technical Reports Server (NTRS)

Hadaegh, Fred Y.; Smith, Roy S.

2006-01-01

The formation flying of precisely controlled spacecraft in deep space can be used to implement optical instruments capable of imaging planets in other solar systems. The distance of the formation from Earth necessitates a significant level of autonomy and each spacecraft must base its actions on its estimates of the location and velocity of the other spacecraft. Precise coordination and control is the key requirement in such missions and the flow of information between spacecraft must be carefully designed. Doing this in an efficient and optimal manner requires novel techniques for the design of the on-board estimators. The use of standard Kalman filter-based designs can lead to unanticipated dynamics--which we refer to as disagreement dynamics--in the estimators' errors. We show how communication amongst the spacecraft can be designed in order to control all of the dynamics within the formation. We present several results relating the topology of the communication network to the resulting closed-loop control dynamics of the formation. The consequences for the design of the control, communication and coordination are discussed.

System design of the Pioneer Venus spacecraft. Volume 11: Launch vehicle utilization

NASA Technical Reports Server (NTRS)

Varga, R. J.

1973-01-01

A summary of the spacecraft descriptions; the probe bus, large probe, small probe, and orbiter is presented. The highlights on the designs of the Atlas/Centaur spacecraft as compared to the corresponding Thor/Delta spacecraft designs are contained. A comparison is made of the two Atlas/Centaur spacecraft for reference. The major differences are the replacement of the probes of the forward end of the probe bus with the mechanically despun antenna of the orbiter and the replacement of the bicone antenna on the aft end with the orbit insertion motor. The cross sections of the large and small probes are compared. The major features of each probe are described. The Thor/Delta and Atlas/Centaur designs for the probe bus and orbiter are analyzed. The usable spacecraft mass for the Atlas/Centaur is roughly twice that for the Thor/Delta if the Type I trajectory is assumed. It is somewhat less for the Type II trajectory in the designated launch years. This additional mass capability leads to cost savings in many areas which are described.

Using a Genetic Algorithm to Design Nuclear Electric Spacecraft

NASA Technical Reports Server (NTRS)

Pannell, William P.

2003-01-01

The basic approach to to design nuclear electric spacecraft is to generate a group of candidate designs, see how "fit" the design are, and carry best design forward to the next generation. Some designs eliminated, some randomly modified and carried forward.

Systems design study of the Pioneer Venus spacecraft. Volume 1. Technical analyses and tradeoffs, sections 5-6 (part 2 of 4)

NASA Technical Reports Server (NTRS)

1973-01-01

System configuration concepts, and tradeoffs are presented for the Atlas/Centaur, and the Thor/Delta probes. Spacecraft system definition, and the probe system definition are discussed along with the mission reliability.

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

NASA Technical Reports Server (NTRS)

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.

Ascent performance feasibility for next-generation spacecraft

NASA Astrophysics Data System (ADS)

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.

Design guidelines for assessing and controlling spacecraft charging effects

NASA Technical Reports Server (NTRS)

Purvis, C. K.; Garrett, H. B.; Whittlesey, A. C.; Stevens, N. J.

1984-01-01

The need for uniform criteria, or guidelines, to be used in all phases of spacecraft design is discussed. Guidelines were developed for the control of absolute and differential charging of spacecraft surfaces by the lower energy space charged particle environment. Interior charging due to higher energy particles is not considered. A guide to good design practices for assessing and controlling charging effects is presented. Uniform design practices for all space vehicles are outlined.

Design guidelines for assessing and controlling spacecraft charging effects

NASA Technical Reports Server (NTRS)

Purvis, C. K.; Garrett, H. B.; Whittlesey, A.; Stevens, N. J.

1985-01-01

The need for uniform criteria, or guidelines, to be used in all phases of spacecraft design is discussed. Guidelines were developed for the control of absolute and differential charging of spacecraft surfaces by the lower energy space charged particle environment. Interior charging due to higher energy particles is not considered. A guide to good design practices for assessing and controlling charging effects is presented. Uniform design practices for all space vehicles are outlined.

GRACE Mission Design: Impact of Uncertainties in Disturbance Environment and Satellite Force Models

NASA Technical Reports Server (NTRS)

Mazanek, Daniel D.; Kumar, Renjith R.; Seywald, Hans; Qu, Min

2000-01-01

The Gravity Recovery and Climate Experiment (GRACE) primary mission will be performed by making measurements of the inter-satellite range change between two co-planar, low altitude, near-polar orbiting satellites. Understanding the uncertainties in the disturbance environment, particularly the aerodynamic drag and torques, is critical in several mission areas. These include an accurate estimate of the spacecraft orbital lifetime, evaluation of spacecraft attitude control requirements, and estimation of the orbital maintenance maneuver frequency necessitated by differences in the drag forces acting on both satellites. The FREEMOL simulation software has been developed and utilized to analyze and suggest design modifications to the GRACE spacecraft. Aerodynamic accommodation bounding analyses were performed and worst-case envelopes were obtained for the aerodynamic torques and the differential ballistic coefficients between the leading and trailing GRACE spacecraft. These analyses demonstrate how spacecraft aerodynamic design and analysis can benefit from a better understanding of spacecraft surface accommodation properties, and the implications for mission design constraints such as formation spacing control.

Human factors in spacecraft design

NASA Technical Reports Server (NTRS)

Harrison, Albert A.; Connors, Mary M.

1990-01-01

This paper describes some of the salient implications of evolving mission parameters for spacecraft design. Among the requirements for future spacecraft are new, higher standards of living, increased support of human productivity, and greater accommodation of physical and cultural variability. Design issues include volumetric allowances, architecture and layouts, closed life support systems, health maintenance systems, recreational facilities, automation, privacy, and decor. An understanding of behavioral responses to design elements is a precondition for critical design decisions. Human factors research results must be taken into account early in the course of the design process.

Conceptual design of a two stage to orbit spacecraft

NASA Technical Reports Server (NTRS)

Armiger, Scott C.; Kwarta, Jennifer S.; Horsley, Kevin B.; Snow, Glenn A.; Koe, Eric C.; Single, Thomas G.

1993-01-01

This project, undertaken through the Advanced Space Design Program, developed a 'Conceptual Design of a Two Stage To Orbit Spacecraft (TSTO).' The design developed utilizes a combination of air breathing and rocket propulsion systems and is fully reusable, with horizontal takeoff and landing capability. The orbiter is carried in an aerodynamically designed bay in the aft section of the booster vehicle to the staging altitude. This TSTO Spacecraft design meets the requirements of replacing the aging Space Shuttle system with a more easily maintained vehicle with more flexible mission capability.

High voltage cabling for high power spacecraft

NASA Technical Reports Server (NTRS)

Dunbar, W. G.

1981-01-01

Studies by NASA have shown that many of the space missions proposed for the time period 1980 to 2000 will require large spacecraft structures to be assembled in orbit. Large antennas and power systems up to 2.5 MW size are predicted to supply the electrical/electronic subsystems, solar electric subsystems, solar electric propulsion, and space processing for the near-term programs. Platforms of 100 meters/length for stable foundations, utility stations, and supports for these multi-antenna and electronic powered mechanisms are also being considered. This paper includes the findings of an analytic and conceptual design study for large spacecraft power distribution, and electrical loads and their influence on the cable and connector requirements for these proposed large spacecraft.

Mars Polar Lander undergoes testing in SAEF-2

NASA Technical Reports Server (NTRS)

1998-01-01

In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), KSC technicians check underneath the Mars Polar Lander as it sits on a workstand. The spacecraft is undergoing testing of science instruments and basic spacecraft subsystems. The solar-powered spacecraft, targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999, is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere.

KSC-98pc1606

NASA Image and Video Library

1998-10-29

KENNEDY SPACE CENTER, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), KSC technicians check underneath the Mars Polar Lander as it sits on a workstand. The spacecraft is undergoing testing of science instruments and basic spacecraft subsystems. The solar-powered spacecraft, targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999, is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

High-power S-band experiment study: Objectives, description, planning and operation of experiment for ATS-G spacecraft

NASA Technical Reports Server (NTRS)

Walp, R. M.

1972-01-01

The results of a study to develop and define requirements for the high power S-band experiment for the ATS-G are summarized. The objectives of the experiment are: (1) to demonstrate high power technology at S-band frequencies in orbiting spacecraft, (2) to employ high power carrier from the spacecraft for conducting interference measurements with Instructional Television Fixed Service systems, and (3) to provide means for performing educationally oriented applications experiments. Experiment organization and operation, and hardware for flight on the ATS-G spacecraft are described. Earth stations designed for the experiment as well as other special ground equipment are also described.

Spacecraft Mission Design for the Mitigation of the 2017 PDC Hypothetical Asteroid Threat

NASA Technical Reports Server (NTRS)

Barbee, Brent W.; Sarli, Bruno V.; Lyzhoft, Josh; Chodas, Paul W.; Englander, Jacob A.

2017-01-01

This paper presents a detailed mission design analysis results for the 2017 Planetary Defense Conference (PDC) Hypothetical Asteroid Impact Scenario, documented at https:cneos.jpl.nasa.govpdcspdc17. The mission design includes campaigns for both reconnaissance (flyby or rendezvous) of the asteroid (to characterize it and the nature of the threat it poses to Earth) and mitigation of the asteroid, via kinetic impactor deflection, nuclear explosive device (NED) deflection, or NED disruption. Relevant scenario parameters are varied to assess the sensitivity of the design outcome, such as asteroid bulk density, asteroid diameter, momentum enhancement factor, spacecraft launch vehicle, and mitigation system type. Different trajectory types are evaluated in the mission design process from purely ballistic to those involving optimal midcourse maneuvers, planetary gravity assists, and/or low-thrust solar electric propulsion. The trajectory optimization is targeted around peak deflection points that were found through a novel linear numerical technique method. The optimization process includes constrain parameters, such as Earth departure date, launch declination, spacecraft, asteroid relative velocity and solar phase angle, spacecraft dry mass, minimum/maximum spacecraft distances from Sun and Earth, and Earth-spacecraft communications line of sight. Results show that one of the best options for the 2017 PDC deflection is solar electric propelled rendezvous mission with a single spacecraft using NED for the deflection.

Polarized-interferometer feasibility study

NASA Technical Reports Server (NTRS)

Raab, F. H.

1983-01-01

The feasibility of using a polarized-interferometer system as a rendezvous and docking sensor for two cooperating spacecraft was studied. The polarized interferometer is a radio frequency system for long range, real time determination of relative position and attitude. Range is determined by round trip signal timing. Direction is determined by radio interferometry. Relative roll is determined from signal polarization. Each spacecraft is equipped with a transponder and an antenna array. The antenna arrays consist of four crossed dipoles that can transmit or receive either circularly or linearly polarized signals. The active spacecraft is equipped with a sophisticated transponder and makes all measurements. The transponder on the passive spacecraft is a relatively simple repeater. An initialization algorithm is developed to estimate position and attitude without any a priori information. A tracking algorithm based upon minimum variance linear estimators is also developed. Techniques to simplify the transponder on the passive spacecraft are investigated and a suitable configuration is determined. A multiple carrier CW signal format is selected. The dependence of range accuracy and ambiguity resolution error probability are derived and used to design a candidate system. The validity of the design and the feasibility of the polarized interferometer concept are verified by simulation.

Numerical Simulations of Spacecraft Charging: Selected Applications

NASA Astrophysics Data System (ADS)

Moulton, J. D.; Delzanno, G. L.; Meierbachtol, C.; Svyatskiy, D.; Vernon, L.; Borovsky, J.; Thomsen, M. F.

2016-12-01

The electrical charging of spacecraft due to bombarding charged particles affects their performance and operation. We study this charging using CPIC, a particle-in-cell code specifically designed for studying plasma-material interactions. CPIC is based on multi-block curvilinear meshes, resulting in near-optimal computational performance while maintaining geometric accuracy. It is interfaced to a mesh generator that creates a computational mesh conforming to complex objects like a spacecraft. Relevant plasma parameters can be imported from the SHIELDS framework (currently under development at LANL), which simulates geomagnetic storms and substorms in the Earth's magnetosphere. Selected physics results will be presented, together with an overview of the code. The physics results include spacecraft-charging simulations with geometry representative of the Van Allen Probes spacecraft, focusing on the conditions that can lead to significant spacecraft charging events. Second, results from a recent study that investigates the conditions for which a high-power (>keV) electron beam could be emitted from a magnetospheric spacecraft will be presented. The latter study proposes a spacecraft-charging mitigation strategy based on the plasma contactor technology that might allow beam experiments to operate in the low-density magnetosphere. High-power electron beams could be used for instance to establish magnetic-field-line connectivity between ionosphere and magnetosphere and help solving long-standing questions in ionospheric/magnetospheric physics.

Extreme Spacecraft Charging in Polar Low Earth Orbit

NASA Technical Reports Server (NTRS)

Colson, Andrew D.; Minow, Joseph I.; NeergaardParker, Linda

2012-01-01

Spacecraft in low altitude, high inclination (including sun-synchronous) orbits are widely used for remote sensing of the Earth's land surface and oceans, monitoring weather and climate, communications, scientific studies of the upper atmosphere and ionosphere, and a variety of other scientific, commercial, and military applications. These systems episodically charge to frame potentials in the kilovolt range when exposed to space weather environments characterized by a high flux of energetic (10 s kilovolt) electrons in regions of low background plasma density which is similar in some ways to the space weather conditions in geostationary orbit responsible for spacecraft charging to kilovolt levels. We first review the physics of space environment interactions with spacecraft materials that control auroral charging rates and the anticipated maximum potentials that should be observed on spacecraft surfaces during disturbed space weather conditions. We then describe how the theoretical values compare to the observational history of extreme charging in auroral environments. Finally, a set of extreme DMSP charging events are described varying in maximum negative frame potential from 0.6 kV to 2 kV, focusing on the characteristics of the charging events that are of importance both to the space system designer and to spacecraft operators. The goal of the presentation is to bridge the gap between scientific studies of auroral charging and the need for engineering teams to understand how space weather impacts both spacecraft design and operations for vehicles on orbital trajectories that traverse auroral charging environments.

Extreme Spacecraft Charging in Polar Low Earth Orbit

NASA Technical Reports Server (NTRS)

Colson, Andrew D.; Minow, Joseph I.; Parker, L. Neergaard

2012-01-01

Spacecraft in low altitude, high inclination (including sun -synchronous) orbits are widely used for remote sensing of the Earth fs land surface and oceans, monitoring weather and climate, communications, scientific studies of the upper atmosphere and ionosphere, and a variety of other scientific, commercial, and military applications. These systems episodically charge to frame potentials in the kilovolt range when exposed to space weather environments characterized by a high flux of energetic (approx.10 fs kilovolt) electrons in regions of low background plasma density. Auroral charging conditions are similar in some ways to the space weather conditions in geostationary orbit responsible for spacecraft charging to kilovolt levels. We first review the physics of space environment interactions with spacecraft materials that control auroral charging rates and the anticipated maximum potentials that should be observed on spacecraft surfaces during disturbed space weather conditions. We then describe how the theoretical values compare to the observational history of extreme charging in auroral environments. Finally, a set of extreme DMSP charging events are described varying in maximum negative frame potential from approx.0.6 kV to approx.2 kV, focusing on the characteristics of the charging events that are of importance both to the space system designer and to spacecraft operators. The goal of the presentation is to bridge the gap between scientific studies of auroral charging and the need for engineering teams to understand how space weather impacts both spacecraft design and operations for vehicles on orbital trajectories that traverse auroral charging environments.

Analytical study of electrical disconnect system for use on manned and unmanned missions

NASA Technical Reports Server (NTRS)

Rosener, A. A.; Lenda, J. A.; Trummer, R. O.; Jonkoniec, T. G.

1977-01-01

The program to survey existing electrical connector availability, and establish an optimum connector design for maintainable spacecraft substation interfaces is reported. Functional and operational requirements are given along with the results of the documentation survey, which disclosed that the MSFC series connectors have the preferred features of current connector technology. Optimum design concepts for EVA tasks, modules serviced by manipulators, and for manipulators independent of other servicing units are presented. It is concluded that separate connector designs are required for spacecraft replaceable modules, and for crewman EVA.

Analysis and design of the Multimission Modular Spacecraft hydrazine propulsion module

NASA Technical Reports Server (NTRS)

Etheridge, F. G.; Woodruff, W. L.

1978-01-01

The translational velocity increment, stabilization and control requirements, vehicle weight, and geometric considerations of the Multimission Modular Spacecraft (MMS) provided the basic data on which to initiate the analysis and design of the hydrazine propulsion modules. The Landsat D was used as the mission model. Tradeoff studies were conducted on thrust level, thruster location, and clustering arrangement together with tankage volume and location. The impact of the use of single and dual seat thruster valves on plumbing configuration, reliability, and overall system cost was studied in detail. Conceptual designs of a recommended propulsion module configuration for both the Delta 3910 and Shuttle were prepared.

Space Weather Impacts on Spacecraft Design and Operations in Auroral Charging Environments

NASA Technical Reports Server (NTRS)

Minow, Joseph I.; Parker, Linda N.

2012-01-01

Spacecraft in low altitude, high inclination (including sun-synchronous) orbits are widely used for remote sensing of the Earth s land surface and oceans, monitoring weather and climate, communications, scientific studies of the upper atmosphere and ionosphere, and a variety of other scientific, commercial, and military applications. These systems are episodically exposed to environments characterized by a high flux of energetic (approx.1 to 10 s kilovolt) electrons in regions of very low background plasma density which is similar in some ways to the space weather conditions in geostationary orbit responsible for spacecraft charging to kilovolt levels. While it is well established that charging conditions in geostationary orbit are responsible for many anomalies and even spacecraft failures, to date there have been relatively few such reports due to charging in auroral environments. This presentation first reviews the physics of the space environment and its interactions with spacecraft materials that control auroral charging rates and the anticipated maximum potentials that should be observed on spacecraft surfaces during disturbed space weather conditions. We then describe how the theoretical values compare to the observational history of extreme charging in auroral environments and discuss how space weather impacts both spacecraft design and operations for vehicles on orbital trajectories that traverse auroral charging environments.

Integrated Controls-Structures Design Methodology for Flexible Spacecraft

NASA Technical Reports Server (NTRS)

Maghami, P. G.; Joshi, S. M.; Price, D. B.

1995-01-01

This paper proposes an approach for the design of flexible spacecraft, wherein the structural design and the control system design are performed simultaneously. The integrated design problem is posed as an optimization problem in which both the structural parameters and the control system parameters constitute the design variables, which are used to optimize a common objective function, thereby resulting in an optimal overall design. The approach is demonstrated by application to the integrated design of a geostationary platform, and to a ground-based flexible structure experiment. The numerical results obtained indicate that the integrated design approach generally yields spacecraft designs that are substantially superior to the conventional approach, wherein the structural design and control design are performed sequentially.

Spacecraft high-voltage power supply construction

NASA Technical Reports Server (NTRS)

Sutton, J. F.; Stern, J. E.

1975-01-01

The design techniques, circuit components, fabrication techniques, and past experience used in successful high-voltage power supplies for spacecraft flight systems are described. A discussion of the basic physics of electrical discharges in gases is included and a design rationale for the prevention of electrical discharges is provided. Also included are typical examples of proven spacecraft high-voltage power supplies with typical specifications for design, fabrication, and testing.

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

NASA Technical Reports Server (NTRS)

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.

Wide-field high-performance geosynchronous imaging

NASA Astrophysics Data System (ADS)

Wood, H. John; Jenstrom, Del; Wilson, Mark; Hinkal, Sanford; Kirchman, Frank

1998-01-01

The NASA Mission to Planet Earth (MTPE) Program and the National Oceanographic and Atmospheric Administration (NOAA) are sponsoring the Advanced Geosynchronous Studies (AGS) to develop technologies and system concepts for Earth observation from geosynchronous orbit. This series of studies is intended to benefit both MTPE science and the NOAA GOES Program. Within the AGS program, advanced imager trade studies have investigated two candidate concepts for near-term advanced geosynchronous imagers. One concept uses a scan mirror to direct the line of sight from a 3-axis stabilized platform. Another eliminates the need for a scan mirror by using an agile spacecraft bus to scan the entire instrument. The purpose of this paper is to discuss the optical design trades and system issues encountered in evaluating the two scanning approaches. The imager design started with a look at first principles: what is the most efficient way to image the Earth in those numerous spectral bands of interest to MTPE scientists and NOAA weather forecasters. Optical design trades included rotating filter wheels and dispersive grating instruments. The design converged on a bandpass filter instrument using four focal planes to cover the spectral range 0.45 to 13.0 micrometers. The first imager design uses a small agile spacecraft supporting an afocal optical telescope. Dichroic beamsplitters feed refractive objectives to four focal planes. The detectors are a series of long linear and rectangular arrays which are scanned in a raster fashion over the 17 degree Earth image. The use of the spacecraft attitude control system to raster the imager field-of-view (FOV) back and forth over the Earth eliminates the need for a scan mirror. However, the price paid is significant energy and time required to reverse the spacecraft slew motions at the end of each scan line. Hence, it is desired to minimize the number of scan lines needed to cover the full Earth disk. This desire, coupled with the ground coverage requirements, drives the telescope design to a 1.6 degree square FOV to provide full Earth disk coverage in less than 12 swaths. The telescope design to accommodate the FOV and image quality requirements is a 30 cm aperture three-element off-axis anastigmat. The size and mass of the imager instrument that result from this optical configuration are larger than desired. But spacecraft reaction wheel torque and power requirements to raster the imager FOV are achievable using existing spacecraft technology. However, launch mass and cost are higher than desired. In the second high-level trade study, the AGS imager team is looking at incorporating a scan mirror and having the satellite three-axis stabilized. The use of the scan mirror eliminates the long turn-around times of the spacecraft scanning approach, allowing for faster Earth coverage. Thus the field of view of the afocal telescope can be reduced by half while still satisfying ground coverage requirements. The optical design of the reduced field afocal telescope is being studied to shrink its size and improve its performance. Both a three-mirror Cassegrain afocal and a two-mirror pair of confocal paraboloids are being considered. With either telescope, the size, mass, and power requirements of this imager are significantly less than those of the first imager design. Both imager designs appear to be feasible and both meet envisioned MTPE and NOAA geosynchronous imaging needs. The AGS imager team is continuing to explore the optical trade space to further optimize imager designs.

Methodology for conceptual remote sensing spacecraft technology: insertion analysis balancing performance, cost, and risk

NASA Astrophysics Data System (ADS)

Bearden, David A.; Duclos, Donald P.; Barrera, Mark J.; Mosher, Todd J.; Lao, Norman Y.

1997-12-01

Emerging technologies and micro-instrumentation are changing the way remote sensing spacecraft missions are developed and implemented. Government agencies responsible for procuring space systems are increasingly requesting analyses to estimate cost, performance and design impacts of advanced technology insertion for both state-of-the-art systems as well as systems to be built 5 to 10 years in the future. Numerous spacecraft technology development programs are being sponsored by Department of Defense (DoD) and National Aeronautics and Space Administration (NASA) agencies with the goal of enhancing spacecraft performance, reducing mass, and reducing cost. However, it is often the case that technology studies, in the interest of maximizing subsystem-level performance and/or mass reduction, do not anticipate synergistic system-level effects. Furthermore, even though technical risks are often identified as one of the largest cost drivers for space systems, many cost/design processes and models ignore effects of cost risk in the interest of quick estimates. To address these issues, the Aerospace Corporation developed a concept analysis methodology and associated software tools. These tools, collectively referred to as the concept analysis and design evaluation toolkit (CADET), facilitate system architecture studies and space system conceptual designs focusing on design heritage, technology selection, and associated effects on cost, risk and performance at the system and subsystem level. CADET allows: (1) quick response to technical design and cost questions; (2) assessment of the cost and performance impacts of existing and new designs/technologies; and (3) estimation of cost uncertainties and risks. These capabilities aid mission designers in determining the configuration of remote sensing missions that meet essential requirements in a cost- effective manner. This paper discuses the development of CADET modules and their application to several remote sensing satellite mission concepts.

Design guide for low cost standardized payloads, volume 2

NASA Technical Reports Server (NTRS)

1972-01-01

Sixteen engineering approaches to low cost standardized payloads in spacecraft are presented. Standard earth observatory satellite, standard U.S. domestic communication satellite, planetary spacecraft subsystems, standard spacecraft, and cluster spacecraft are reviewed.

Magnetic suspension options for spacecraft inertia-wheel applications

NASA Technical Reports Server (NTRS)

Downer, J. R.

1984-01-01

Design criteria for spacecraft inertia-wheel suspensions are listed. The advantages of magnetic suspensions over other suspension types for spacecraft inertia-wheel applications are cited along with the functions performed by magnetic suspension. The common designs for magnetic suspensions are enumerated. Materials selection of permanent magnets and core materials is considered.

Proximity operations considerations affecting spacecraft design

NASA Technical Reports Server (NTRS)

Staas, Steven K.

1991-01-01

Experience from several recent spacecraft development programs, such as Space Station Freedom (SSF) and the Orbital Maneuvering Vehicle (OMV) has shown the need for factoring proximity operations considerations into the vehicle design process. Proximity operations, those orbital maneuvers and procedures which involve operation of two or more spacecraft at ranges of less than one nautical mile, are essential to the construction, servicing, and operation of complex spacecraft. Typical proximity operations considerations which drive spacecraft design may be broken into two broad categories; flight profile characteristics and concerns, and use of various spacecraft systems during proximity operations. Proximity operations flight profile concerns include the following: (1) relative approach/separation line; (2) relative orientation of the vehicles; (3) relative translational and rotational rates; (4) vehicle interaction, in the form of thruster plume impingement, mating or demating operations, or uncontrolled contact/collision; and (5) active vehicle piloting. Spacecraft systems used during proximity operations include the following: (1) sensors, such as radar, laser ranging devices, or optical ranging systems; (2) effector hardware, such as thrusters; (3) flight control software; and (4) mating hardware, needed for docking or berthing operations. A discussion of how these factors affect vehicle design follows, addressing both active and passive/cooperative vehicles.

Spacecraft fault tolerance: The Magellan experience

NASA Technical Reports Server (NTRS)

Kasuda, Rick; Packard, Donna Sexton

1993-01-01

Interplanetary and earth orbiting missions are now imposing unique fault tolerant requirements upon spacecraft design. Mission success is the prime motivator for building spacecraft with fault tolerant systems. The Magellan spacecraft had many such requirements imposed upon its design. Magellan met these requirements by building redundancy into all the major subsystem components and designing the onboard hardware and software with the capability to detect a fault, isolate it to a component, and issue commands to achieve a back-up configuration. This discussion is limited to fault protection, which is the autonomous capability to respond to a fault. The Magellan fault protection design is discussed, as well as the developmental and flight experiences and a summary of the lessons learned.

Spacecraft attitude impacts on COLD-SAT non-vacuum jacketed LH2 supply tank thermal performance

NASA Technical Reports Server (NTRS)

Arif, Hugh

1990-01-01

The Cryogenic On-Orbit Liquid Depot - Storage, Acquisition and Transfer (COLD-SAT) spacecraft will be launched into low earth orbit to perform fluid management experiments on the behavior of subcritical liquid hydrogen (LH2). For determining the optimum on-orbit attitude for the COLD-SAT satellite, a comparative analytical study was performed to determine the thermal impacts of spacecraft attitude on the performance of the COLD-SAT non-vacuum jacketed LH2 supply tank. Tank thermal performance was quantitied by total conductive and radiative heat leakage into the pressure vessel due to the absorbed solar, earth albedo and infra-red on-orbit fluxes, and also by the uniformity of the variation of this leakage on the vessel surface area. Geometric and thermal analysis math models were developed for the spacecraft and the tank as part of this analysis, based on their individual thermal/structural designs. Two quasi-inertial spacecraft attitudes were investigated and their effects on the tank performance compared. The results are one of the criteria by which the spacecraft orientation in orbit was selected for the in-house NASA Lewis Research Center design.

Spacecraft attitude impacts on COLD-SAT non-vacuum jacketed LH2 supply tank thermal performance

NASA Technical Reports Server (NTRS)

Arif, Hugh

1990-01-01

The Cryogenic On-Orbit Liquid Depot - Storage, Acquisition and Transfer (COLD-SAT) spacecraft will be launched into low earth orbit to perform fluid management experiments on the behavior of subcritical liquid hydrogen (LH2). For determining the optimum on-orbit attitude for the COLD-SAT satellite, a comparative analytical study was performed to determine the thermal impacts of spacecraft attitude on the performance of the COLD-SAT non-vacuum jacketed LH2 supply tank. Tank thermal performance was quantified by total conductive and radiative heat leakage into the pressure vessel due to the absorbed solar, earth albedo and infra-red on-orbit fluxes, and also by the uniformity of the variation of this leakage on the vessel surface area. Geometric and thermal analysis math models were developed for the spacecraft and the tank as part of this analysis, based on their individual thermal/structural designs. Two quasi-inertial spacecraft attitudes were investigated and their effects on the tank performance compared. The results are one of the criteria by which the spacecraft orientation in orbit was selected for the in-house NASA Lewis Research Center design.

Voyager spacecraft electrostatic discharge testing

NASA Technical Reports Server (NTRS)

Whittlesey, A.; Inouye, G.

1980-01-01

The program of environmental testing undergone by the Voyager spacecraft in order to simulate the transient voltage effects of electrostatic discharges expected in the energetic plasma environment of Jupiter is reported. The testing consists of studies of the electrostatic discharge characteristics of spacecraft dielectrics in a vacuum-chamber-electron beam facility, brief piece part sensitivity tests on such items as a MOSFET multiplexer and the grounding of the thermal blanket, and assembly tests of the magnetometer boom and the science boom. In addition, testing of a complete spacecraft was performed using two arc sources to simulate long and short duration discharge sources for successive spacecraft shielding and grounding improvements. Due to the testing program, both Voyager 1 and Voyager 2 experienced tolerable electrostatic discharge-caused transient anomalies in science and engineering subsystems, however, a closer duplication of the spacecraft environment is necessary to predict and design actual spacecraft responses more accurately.

Swarms: Optimum aggregations of spacecraft

NASA Technical Reports Server (NTRS)

Mayer, H. L.

1980-01-01

Swarms are aggregations of spacecraft or elements of a space system which are cooperative in function, but physically isolated or only loosely connected. For some missions the swarm configuration may be optimum compared to a group of completely independent spacecraft or a complex rigidly integrated spacecraft or space platform. General features of swarms are induced by considering an ensemble of 26 swarms, examples ranging from Earth centered swarms for commercial application to swarms for exploring minor planets. A concept for a low altitude swarm as a substitute for a space platform is proposed and a preliminary design studied. The salient design feature is the web of tethers holding the 30 km swarm in a rigid two dimensional array in the orbital plane. A mathematical discussion and tutorial in tether technology and in some aspects of the distribution of services (mass, energy, and information to swarm elements) are included.

Contamination of planets by nonsterile flight hardware.

NASA Technical Reports Server (NTRS)

Wolfson, R. P.; Craven, C. W.

1971-01-01

The various factors about space missions and spacecraft involved in the study of nonsterile space flight hardware with respect to their effects on planetary quarantine are reviewed. It is shown that methodology currently exists to evaluate the various potential contamination sources and to take appropriate steps in the design of spacecraft ha rdware and mission parameters so that quarantine constraints are met. This work should be done for each program so that the latest knowledge pertaining to various biological questions is utilized, and so that the specific hardware designs of the program can be assessed. The general trend of specific recommendations include: (1) biasing the launch trajectory away from planet to assure against accidental impact of the spacecraft; (2) selecting planetary orbits that meet quarantine requirements - both for accidental impact and for minimizing contamination probabilities due to ejecta; and (3) manufacturing and handling spacecraft under cleanliness conditions assuring minimum bioload.

Technology Combination Analysis Tool (TCAT) for Active Debris Removal

NASA Astrophysics Data System (ADS)

Chamot, B.; Richard, M.; Salmon, T.; Pisseloup, A.; Cougnet, C.; Axthelm, R.; Saunder, C.; Dupont, C.; Lequette, L.

2013-08-01

This paper present the work of the Swiss Space Center EPFL within the CNES-funded OTV-2 study. In order to find the most performant Active Debris Removal (ADR) mission architectures and technologies, a tool was developed in order to design and compare ADR spacecraft, and to plan ADR campaigns to remove large debris. Two types of architectures are considered to be efficient: the Chaser (single-debris spacecraft), the Mothership/ Kits (multiple-debris spacecraft). Both are able to perform controlled re-entry. The tool includes modules to optimise the launch dates and the order of capture, to design missions and spacecraft, and to select launch vehicles. The propulsion, power and structure subsystems are sized by the tool thanks to high-level parametric models whilst the other ones are defined by their mass and power consumption. Final results are still under investigation by the consortium but two concrete examples of the tool's outputs are presented in the paper.

Impact of Space Transportation System on planetary spacecraft and missions design

NASA Technical Reports Server (NTRS)

Barnett, P. M.

1975-01-01

Results of Jet Propulsion Laboratory (JPL) activities to define and understand alternatives for planetary spacecraft operations with the Space Transportation System (STS) are summarized. The STS presents a set of interfaces, operational alternatives, and constraints in the prelaunch, launch, and near-earth flight phases of a mission. Shuttle-unique features are defined and coupled with JPL's existing program experience to begin development of operationally efficient alternatives, concepts, and methods for STS-launched missions. The time frame considered begins with the arrival of the planetary spacecraft at Kennedy Space Center and includes prelaunch ground operations, Shuttle-powered flight, and near-earth operations, up to acquisition of the spacecraft signal by the Deep Space Network. The areas selected for study within this time frame were generally chosen because they represent the 'driving conditions' on planetary-mission as well as system design and operations.

The Impact of the Space Environment on Space Systems

DTIC Science & Technology

1999-07-20

databases were designed to determine the extent of spacecraft problems from the standpoint of the spacecraft designer. One of their main uses has...the cause of an anomaly, the spacecraft generally lack sensors to determine the state of the environment at the location of the spacecraft at the...from PRADS to basic attitude mode Diagnosis: SEU Sure: 2 Impact: Loss of pointing accuracy. Important for imaging sensors Duration 1 hr to 1 day

Spacecraft and mission design for the SP-100 flight experiment

NASA Technical Reports Server (NTRS)

Deininger, William D.; Vondra, Robert J.

1988-01-01

The design and performance of a spacecraft employing arcjet nuclear electric propulsion, suitable for use in the SP-100 Space Reactor Power System (SRPS) Flight Experiment, are outlined. The vehicle design is based on a 93 kW(e) ammonia arcjet system operating at an experimentally measured specific impulse of 1031 s and an efficiency of 42.3 percent. The arcjet/gimbal assemblies, power conditioning subsystem, propellant feed system, propulsion system thermal control, spacecraft diagnostic instrumentation, and the telemetry requirements are described. A 100 kW(e) SRPS is assumed. The spacecraft mass is baselined at 5675 kg excluding the propellant and propellant feed system. Four mission scenarios are described which are capable of demonstrating the full capability of the SRPS. The missions considered include spacecraft deployment to possible surveillance platform orbits, a spacecraft storage mission, and an orbit raising round trip corresponding to possible orbit transfer vehicle (OTV) missions.

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

NASA Technical Reports Server (NTRS)

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.

STS ancillary equipment study. [user reference book for multimission modular spacecraft missions

NASA Technical Reports Server (NTRS)

Plough, J. A.

1977-01-01

Spaceborne and ground ancillary equipment for multimission module spacecraft are listed to provide documentation for potential users interested in utilizing existing equipment rather than developing payload unique designs. The format of the data form contained in the Ancillary Equipment user reference book is discussed.

Grooved Fuel Rings for Nuclear Thermal Rocket Engines

NASA Technical Reports Server (NTRS)

Emrich, William

2009-01-01

An alternative design concept for nuclear thermal rocket engines for interplanetary spacecraft calls for the use of grooved-ring fuel elements. Beyond spacecraft rocket engines, this concept also has potential for the design of terrestrial and spacecraft nuclear electric-power plants. The grooved ring fuel design attempts to retain the best features of the particle bed fuel element while eliminating most of its design deficiencies. In the grooved ring design, the hydrogen propellant enters the fuel element in a manner similar to that of the Particle Bed Reactor (PBR) fuel element.

A study of the selection of microcomputer architectures to automate planetary spacecraft power systems

NASA Technical Reports Server (NTRS)

Nauda, A.

1982-01-01

Performance and reliability models of alternate microcomputer architectures as a methodology for optimizing system design were examined. A methodology for selecting an optimum microcomputer architecture for autonomous operation of planetary spacecraft power systems was developed. Various microcomputer system architectures are analyzed to determine their application to spacecraft power systems. It is suggested that no standardization formula or common set of guidelines exists which provides an optimum configuration for a given set of specifications.

Optimization techniques applied to passive measures for in-orbit spacecraft survivability

NASA Technical Reports Server (NTRS)

Mog, Robert A.; Price, D. Marvin

1991-01-01

Spacecraft designers have always been concerned about the effects of meteoroid impacts on mission safety. The engineering solution to this problem has generally been to erect a bumper or shield placed outboard from the spacecraft wall to disrupt/deflect the incoming projectiles. Spacecraft designers have a number of tools at their disposal to aid in the design process. These include hypervelocity impact testing, analytic impact predictors, and hydrodynamic codes. Analytic impact predictors generally provide the best quick-look estimate of design tradeoffs. The most complete way to determine the characteristics of an analytic impact predictor is through optimization of the protective structures design problem formulated with the predictor of interest. Space Station Freedom protective structures design insight is provided through the coupling of design/material requirements, hypervelocity impact phenomenology, meteoroid and space debris environment sensitivities, optimization techniques and operations research strategies, and mission scenarios. Major results are presented.

Benefits of Operational Consideration into the Guidance, Navigation, and Control Design of Spacecraft

NASA Technical Reports Server (NTRS)

Morris, Mark C.; Holt, Greg N.

2011-01-01

The following paper points out historical examples where operational consideration into the GN&C design could have helped avoid operational complexity, reduce costs, ensure the ability for a GN&C system to be able to adapt to failures, and in some cases might have helped save mission objectives. A costly repeat of mistakes could befall a program if previous operational lessons, especially from operators of vehicles with similar GN&C systems, are not considered during the GN&C design phase of spacecraft. The information gained from operational consideration during the design can lead to improvements of the design, allow less ground support during operations, and prevent repetition of previous mistakes. However, this benefit can only occur if spacecraft operators adequately capture lessons learned that would improve future designs for operations and those who are designing spacecraft incorporate inputs from those that have previously operated similar GN&C systems.

Optimization techniques applied to passive measures for in-orbit spacecraft survivability

NASA Technical Reports Server (NTRS)

Mog, Robert A.; Price, D. Marvin

1987-01-01

Optimization techniques applied to passive measures for in-orbit spacecraft survivability, is a six-month study, designed to evaluate the effectiveness of the geometric programming (GP) optimization technique in determining the optimal design of a meteoroid and space debris protection system for the Space Station Core Module configuration. Geometric Programming was found to be superior to other methods in that it provided maximum protection from impact problems at the lowest weight and cost.

LANDSAT D instrument module study

NASA Technical Reports Server (NTRS)

1976-01-01

Spacecraft instrument module configurations which support an earth resource data gathering mission using a thematic mapper sensor were examined. The differences in size of these two experiments necessitated the development of two different spacecraft configurations. Following the selection of the best-suited configurations, a validation phase of design, analysis and modelling was conducted to verify feasibility. The chosen designs were then used to formulate definition for a systems weight, a cost range for fabrication and interface requirements for the thematic mapper (TM).

Reactor design and integration into a nuclear electric spacecraft

NASA Technical Reports Server (NTRS)

Phillips, W. M.; Koenig, D. R.

1978-01-01

One of the well-defined applications for nuclear power in space is nuclear electric propulsion (NEP). Mission studies have identified the optimum power level (400 kWe). A single Shuttle launch requirement and science-package integration have added additional constraints to the design. A reactor design which will meet these constraints has been studied. The reactor employs 90 fuel elements, each heat pipe cooled. Reactor control is obtained with BeO/B4C drums in a BeO reflector. The balance of the spacecraft is shielded from the reactor with LiH. Power conditioning and reactor control drum drives are located behind the LiH with the power conditioning. Launch safety, mechanical design and integration with the power conversion subsystem are discussed.

Theoretical Foundation of Copernicus: A Unified System for Trajectory Design and Optimization

NASA Technical Reports Server (NTRS)

Ocampo, Cesar; Senent, Juan S.; Williams, Jacob

2010-01-01

The fundamental methods are described for the general spacecraft trajectory design and optimization software system called Copernicus. The methods rely on a unified framework that is used to model, design, and optimize spacecraft trajectories that may operate in complex gravitational force fields, use multiple propulsion systems, and involve multiple spacecraft. The trajectory model, with its associated equations of motion and maneuver models, are discussed.

An unmanned probe to Pluto

NASA Technical Reports Server (NTRS)

1990-01-01

Now that Voyager II has completed its grand tour of the solar system, all the planets in the solar system, with the exception of Pluto, have been studied. Even now, missions to return to Mercury, Venus, Mars Jupiter, and Saturn are currently flying or are planned. However, a mission to explore Pluto is not, at the present time, being considered seriously. The design problem presented to the students was very general, i.e., design an unmanned mission to Pluto with a launch window constraint of the years 2000 to 2010. All other characteristics of the mission, such as mission type (flyby, orbiter, lander, penetrator), scientific objectives and payload, and the propulsion system were to be determined by the design teams. The design studies exposed several general problems to be solved. Due to the extreme distance to Pluto (and a corresponding travel time in the range of 10 to 25 years), the spacecraft had to be lighter and more robust than current spacecraft designs. In addition, advanced propulsion concepts had to be considered. These included the new generation of launch vehicles and upper stages and nuclear electric propulsion. The probe design offered an abundance of synthesis and analysis problems. These included sizing trade studies, selection of subsystem components, analysis of spacecraft dynamics, stability and control, structural design and material selection, trajectory design, and selection of scientific equipment. Since the characteristics of the mission, excluding the launch window, were to be determined by the design teams, the solutions varied widely.

Solar Probe thermal shield design and testing

NASA Technical Reports Server (NTRS)

Millard, Jerry M.; Miyake, Robert N.; Rainen, Richard A.

1992-01-01

This paper discusses the major thermal shield subsystem development activities in support of the Solar Probe study being conducted at JPL. The Solar Probe spacecraft will travel to within 4 solar radii of the sun's center to perform fundamental experiments in space physics. Exposure to 2900 earth suns at perihelion requires the spacecraft to be protected within the shadow envelope of a protective shield. In addition, the mass loss rate off of the shield at elevated temperature must comply with plasma instrument requirements and has become the driver of the shield design. This paper will focus on the analytical design work to size the shield and control the shield mass loss rate for the various spacecraft options under study, the application of carbon-carbon materials for shield components, development and preparation of carbon-carbon samples for materials testing, and a materials testing program for carbon-carbon and tungsten alloys to investigate thermal/optical properties, mass loss (carbon-carbon only), material integrity, and high velocity impact behavior.

First LDEF Post-Retrieval Symposium abstracts

NASA Technical Reports Server (NTRS)

Levine, Arlene S. (Compiler)

1991-01-01

The LDE facility was designed to better understand the environments of space and the effects of prolonged exposure in these environments on future spacecraft. The symposium abstracts presented here are organized according to the symposium agenda into five sessions. The first session provides an overview of the LDEF, the experiments, the mission, and the natural and induced environments the spacecraft and experiments encountered during the mission. The second session presents results to date from studies to better define the environments of near-Earth space. The third session addresses studies of the effects of the space environments on spacecraft materials. The fourth session addresses studies of the effects of the space environments on spacecraft systems. And the fifth session addresses other subjects such as results of the LDEF life science and crystal growth experiments.

SERT C project study

NASA Technical Reports Server (NTRS)

1974-01-01

The SERT C (Space Electric Rocket Test - C) project study defines a spacecraft mission that would demonstrate the technology readiness of ion thruster systems for primary propulsion and station keeping applications. As a low cost precursor, SERT C develops the components and systems required for subsequent Solar Electric Propulsion (SEP) applications. The SERT C mission requirements and preliminary spacecraft and subsystem design are described.

Effect of Voltage Level on Power System Design for Solar Electric Propulsion Missions

NASA Technical Reports Server (NTRS)

Kerslake, Thomas W.

2003-01-01

This paper presents study results quantifying the benefits of higher voltage, electric power system designs for a typical solar electric propulsion spacecraft Earth orbiting mission. A conceptual power system architecture was defined and design points were generated for system voltages of 28-V, 50-V, 120-V, and 300-V using state-of-the-art or advanced technologies. A 300-V 'direct-drive' architecture was also analyzed to assess the benefits of directly powering the electric thruster from the photovoltaic array without up-conversion. Fortran and spreadsheet computational models were exercised to predict the performance and size power system components to meet spacecraft mission requirements. Pertinent space environments, such as electron and proton radiation, were calculated along the spiral trajectory. In addition, a simplified electron current collection model was developed to estimate photovoltaic array losses for the orbital plasma environment and that created by the thruster plume. The secondary benefits of power system mass savings for spacecraft propulsion and attitude control systems were also quantified. Results indicate that considerable spacecraft wet mass savings were achieved by the 300-V and 300-V direct-drive architectures.

Spacecraft Antennas

NASA Technical Reports Server (NTRS)

Jamnejad, Vahraz; Manshadi, Farzin; Rahmat-Samii, Yahya; Cramer, Paul

1990-01-01

Some of the various categories of issues that must be considered in the selection and design of spacecraft antennas for a Personal Access Satellite System (PASS) are addressed, and parametric studies for some of the antenna concepts to help the system designer in making the most appropriate antenna choice with regards to weight, size, and complexity, etc. are provided. The question of appropriate polarization for the spacecraft as well as for the User Terminal Antenna required particular attention and was studied in some depth. Circular polarization seems to be the favored outcome of this study. Another problem that has generally been a complicating factor in designing the multiple beam reflector antennas, is the type of feeds (single vs. multiple element and overlapping vs. non-overlapping clusters) needed for generating the beams. This choice is dependent on certain system design factors, such as the required frequency reuse, acceptable interbeam isolation, antenna efficiency, number of beams scanned, and beam-forming network (BFN) complexity. This issue is partially addressed, but is not completely resolved. Indications are that it may be possible to use relatively simple non-overlapping clusters of only a few elements, unless a large frequency reuse and very stringent isolation levels are required.

Exploration Spacecraft and Space Suit Internal Atmosphere Pressure and Composition

NASA Technical Reports Server (NTRS)

Lange, Kevin; Duffield, Bruce; Jeng, Frank; Campbell, Paul

2005-01-01

The design of habitat atmospheres for future space missions is heavily driven by physiological and safety requirements. Lower EVA prebreathe time and reduced risk of decompression sickness must be balanced against the increased risk of fire and higher cost and mass of materials associated with higher oxygen concentrations. Any proposed increase in space suit pressure must consider impacts on space suit mass and mobility. Future spacecraft designs will likely incorporate more composite and polymeric materials both to reduce structural mass and to optimize crew radiation protection. Narrowed atmosphere design spaces have been identified that can be used as starting points for more detailed design studies and risk assessments.

Spacecraft Mission Design for the Mitigation of the 2017 PDC Hypothetical Asteroid Threat

NASA Technical Reports Server (NTRS)

Barbee, Brent W.; Sarli, Bruno V.; Lyzhoft, Joshua; Chodas, Paul W.; Englander, Jacob A.

2017-01-01

This paper presents a detailed mission design analysis results for the 2017 Planetary Defense Conference (PDC) Hypothetical Asteroid Impact Scenario, documented at https://cneos.jpl.nasa.gov/ pd/cs/pdc17/. The mission design includes campaigns for both reconnaissance (flyby or rendezvous) of the asteroid (to characterize it and the nature of the threat it poses to Earth) and mitigation of the asteroid, via kinetic impactor deflection, nuclear explosive device (NED) deflection, or NED disruption. Relevant scenario parameters are varied to assess the sensitivity of the design outcome, such as asteroid bulk density, asteroid diameter, momentum enhancement factor, spacecraft launch vehicle, and mitigation system type. Different trajectory types are evaluated in the mission design process from purely ballistic to those involving optimal midcourse maneuvers, planetary gravity assists, and/or lowthrust solar electric propulsion. The trajectory optimization is targeted around peak deflection points that were found through a novel linear numerical technique method. The optimization process includes constrain parameters, such as Earth departure date, launch declination, spacecraft/asteroid relative velocity and solar phase angle, spacecraft dry mass, minimum/maximum spacecraft distances from Sun and Earth, and Earth/spacecraft communications line of sight. Results show that one of the best options for the 2017 PDC deflection is solar electric propelled rendezvous mission with a single spacecraft using NED for the deflection

A Comparison of Photocatalytic Oxidation Reactor Performance for Spacecraft Cabin Trace Contaminant Control Applications

NASA Technical Reports Server (NTRS)

Perry, Jay L.; Frederick, Kenneth R.; Scott, Joseph P.; Reinermann, Dana N.

2011-01-01

Photocatalytic oxidation (PCO) is a maturing process technology that shows potential for spacecraft life support system application. Incorporating PCO into a spacecraft cabin atmosphere revitalization system requires an understanding of basic performance, particularly with regard to partial oxidation product production. Four PCO reactor design concepts have been evaluated for their effectiveness for mineralizing key trace volatile organic com-pounds (VOC) typically observed in crewed spacecraft cabin atmospheres. Mineralization efficiency and selectivity for partial oxidation products are compared for the reactor design concepts. The role of PCO in a spacecraft s life support system architecture is discussed.

Charging of the Van Allen Probes: Theory and Simulations

NASA Astrophysics Data System (ADS)

Delzanno, G. L.; Meierbachtol, C.; Svyatskiy, D.; Denton, M.

2017-12-01

The electrical charging of spacecraft has been a known problem since the beginning of the space age. Its consequences can vary from moderate (single event upsets) to catastrophic (total loss of the spacecraft) depending on a variety of causes, some of which could be related to the surrounding plasma environment, including emission processes from the spacecraft surface. Because of its complexity and cost, this problem is typically studied using numerical simulations. However, inherent unknowns in both plasma parameters and spacecraft material properties can lead to inaccurate predictions of overall spacecraft charging levels. The goal of this work is to identify and study the driving causes and necessary parameters for particular spacecraft charging events on the Van Allen Probes (VAP) spacecraft. This is achieved by making use of plasma theory, numerical simulations, and on-board data. First, we present a simple theoretical spacecraft charging model, which assumes a spherical spacecraft geometry and is based upon the classical orbital-motion-limited approximation. Some input parameters to the model (such as the warm plasma distribution function) are taken directly from on-board VAP data, while other parameters are either varied parametrically to assess their impact on the spacecraft potential, or constrained through spacecraft charging data and statistical techniques. Second, a fully self-consistent numerical simulation is performed by supplying these parameters to CPIC, a particle-in-cell code specifically designed for studying plasma-material interactions. CPIC simulations remove some of the assumptions of the theoretical model and also capture the influence of the full geometry of the spacecraft. The CPIC numerical simulation results will be presented and compared with on-board VAP data. This work will set the foundation for our eventual goal of importing the full plasma environment from the LANL-developed SHIELDS framework into CPIC, in order to more accurately predict spacecraft charging.

Trade-off study and computer simulation for assessing spacecraft pointing accuracy and stability capabilities

NASA Astrophysics Data System (ADS)

Algrain, Marcelo C.; Powers, Richard M.

1997-05-01

A case study, written in a tutorial manner, is presented where a comprehensive computer simulation is developed to determine the driving factors contributing to spacecraft pointing accuracy and stability. Models for major system components are described. Among them are spacecraft bus, attitude controller, reaction wheel assembly, star-tracker unit, inertial reference unit, and gyro drift estimators (Kalman filter). The predicted spacecraft performance is analyzed for a variety of input commands and system disturbances. The primary deterministic inputs are the desired attitude angles and rate set points. The stochastic inputs include random torque disturbances acting on the spacecraft, random gyro bias noise, gyro random walk, and star-tracker noise. These inputs are varied over a wide range to determine their effects on pointing accuracy and stability. The results are presented in the form of trade- off curves designed to facilitate the proper selection of subsystems so that overall spacecraft pointing accuracy and stability requirements are met.

Tracking and data relay satellite system configuration and tradeoff study. Volume 1: Study summary

NASA Technical Reports Server (NTRS)

Hill, T. E.

1973-01-01

A study was conducted to determine the configuration and tradeoffs of a tracking and data relay satellite. The study emphasized the design of a three axis stabilized satellite and a telecommunications system optimized for support of low and medium data rate user spacecraft. Telecommunications support to low and high, or low medium, and high data rate users, considering launches with the Delta 2914, the Atlas/Centaur, and the space shuttle was also considered. The following subjects are presented: (1) launch and deployment profile, (2) spacecraft mechanical and structural design, (3) attitude stabilization and control subsystem, and (4) reliability analysis.

NASA's Space Environments and Effects Program: Technology for the New Millennium

NASA Technical Reports Server (NTRS)

Hardage, Donna M.; Pearson, Steven D.

2000-01-01

Current trends in spacecraft development include the use of advanced technologies while maintaining the "faster, better, cheaper" philosophy. Spacecraft designers are continually designing with smaller and faster electronics as well as lighter and thinner materials providing better performance, lower weight, and ultimately lower costs. Given this technology trend, spacecraft will become increasingly susceptible to the harsh space environments, causing damaging or even disabling effects on space systems. NASA's Space Environments and Effects (SEE) Program defines the space environments and provides advanced technology development to support the design, development, and operation of spacecraft systems that will accommodate or mitigate effects due to the harsh space environments. This Program provides a comprehensive and focused approach to understanding the space environment, to define the best techniques for both flight and ground-based experimentation, to update the models which predict both the environments and the environmental effects on spacecraft, and finally to ensure that this multitudinous information is properly maintained and inserted into spacecraft design programs. A description of the SEE Program, its accomplishments, and future activities is provided.

Fire suppression in human-crew spacecraft

NASA Technical Reports Server (NTRS)

Friedman, Robert; Dietrich, Daniel L.

1991-01-01

Fire extinguishment agents range from water and foam in early-design spacecraft (Halon 1301 in the present Shuttle) to carbon dioxide proposed for the Space Station Freedom. The major challenge to spacecraft fire extinguishment design and operations is from the micro-gravity environment, which minimizes natural convection and profoundly influences combustion and extinguishing agent effectiveness, dispersal, and post-fire cleanup. Discussed here are extinguishment in microgravity, fire-suppression problems anticipated in future spacecraft, and research needs and opportunities.

Autonomous spacecraft maintenance study group

NASA Technical Reports Server (NTRS)

Marshall, M. H.; Low, G. D.

1981-01-01

A plan to incorporate autonomous spacecraft maintenance (ASM) capabilities into Air Force spacecraft by 1989 is outlined. It includes the successful operation of the spacecraft without ground operator intervention for extended periods of time. Mechanisms, along with a fault tolerant data processing system (including a nonvolatile backup memory) and an autonomous navigation capability, are needed to replace the routine servicing that is presently performed by the ground system. The state of the art fault handling capabilities of various spacecraft and computers are described, and a set conceptual design requirements needed to achieve ASM is established. Implementations for near term technology development needed for an ASM proof of concept demonstration by 1985, and a research agenda addressing long range academic research for an advanced ASM system for 1990s are established.

System Design under Uncertainty: Evolutionary Optimization of the Gravity Probe-B Spacecraft

NASA Technical Reports Server (NTRS)

Pullen, Samuel P.; Parkinson, Bradford W.

1994-01-01

This paper discusses the application of evolutionary random-search algorithms (Simulated Annealing and Genetic Algorithms) to the problem of spacecraft design under performance uncertainty. Traditionally, spacecraft performance uncertainty has been measured by reliability. Published algorithms for reliability optimization are seldom used in practice because they oversimplify reality. The algorithm developed here uses random-search optimization to allow us to model the problem more realistically. Monte Carlo simulations are used to evaluate the objective function for each trial design solution. These methods have been applied to the Gravity Probe-B (GP-B) spacecraft being developed at Stanford University for launch in 1999, Results of the algorithm developed here for GP-13 are shown, and their implications for design optimization by evolutionary algorithms are discussed.

A Comparison of Structurally Connected and Multiple Spacecraft Interferometers

NASA Technical Reports Server (NTRS)

Surka, Derek M.; Crawley, Edward F.

1996-01-01

Structurally connected and multiple spacecraft interferometers are compared in an attempt to establish the maximum baseline (referred to as the "cross-over baseline") for which it is preferable to operate a single-structure interferometer in space rather than an interferometer composed of numerous, smaller spacecraft. This comparison is made using the total launched mass of each configuration as the comparison metric. A framework of study within which structurally connected and multiple spacecraft interferometers can be compared is presented in block diagram form. This methodology is then applied to twenty-two different combinations of trade space parameters to investigate the effects of different orbits, orientations, truss materials, propellants, attitude control actuators, onboard disturbance sources, and performance requirements on the cross-over baseline. Rotating interferometers and the potential advantages of adding active structural control to the connected truss of the structurally connected interferometer are also examined. The minimum mass design of the structurally connected interferometer that meets all performance-requirements and satisfies all imposed constraints is determined as a function of baseline. This minimum mass design is then compared to the design of the multiple spacecraft interferometer. It is discovered that the design of the minimum mass structurally connected interferometer that meets all performance requirements and constraints in solar orbit is limited by the minimum allowable aspect ratio, areal density, and gage of the struts. In the formulation of the problem used in this study, there is no advantage to adding active structural control to the truss for interferometers in solar orbit. The cross-over baseline for missions of practical duration (ranging from one week to thirty years) in solar orbit is approximately 400 m for non-rotating interferometers and 650 m for rotating interferometers.

A Fully Integrated Micro-Magnometer/Microspacecraft for Multipoint Measurements: The Free-Flyer Magnometer

NASA Technical Reports Server (NTRS)

Goldstein, R.; Miller, L. N.; Fossum, E.; Pain, B.; Randolph, J. E.; Turner, P. R.; Cutting, E.

1995-01-01

In the decades since the advent of in situ plasma measurements on board spacecraft, the instrumentation has grown bigger, heavier, and more complex as our understanding of space plasmas improves and our appetite for more information increases...There has thus been a recent interest in the miniaturization of both spacecraft and the instrument payload... This paper describes the results and status of an ongoing design study to understand the problems and trade space of fully integrating an instrument into a micro-spacecraft.

A Control Concept for Large Flexible Spacecraft Using Order Reduction Techniques

NASA Technical Reports Server (NTRS)

Thieme, G.; Roth, H.

1985-01-01

Results found during the investigation of control problems of large flexible spacecraft are given. A triple plate configuration of such a spacecraft is defined and studied. The model is defined by modal data derived from infinite element modeling. The order reduction method applied is briefly described. An attitude control concept with low and high authority control has been developed to design an attitude controller for the reduced model. The stability and response of the original system together with the reduced controller is analyzed.

ANTARES: Spacecraft Simulation for Multiple User Communities and Facilities

NASA Technical Reports Server (NTRS)

Acevedo, Amanda; Berndt, Jon; Othon, William; Arnold, Jason; Gay, Robet

2007-01-01

The Advanced NASA Technology Architecture for Exploration Studies (ANTARES) simulation is the primary tool being used for requirements assessment of the NASA Orion spacecraft by the Guidance Navigation and Control (GN&C) teams at Johnson Space Center (JSC). ANTARES is a collection of packages and model libraries that are assembled and executed by the Trick simulation environment. Currently, ANTARES is being used for spacecraft design assessment, performance analysis, requirements validation, Hardware In the Loop (HWIL) and Human In the Loop (HIL) testing.

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

NASA Astrophysics Data System (ADS)

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.

KSC-2012-4353

NASA Image and Video Library

2012-08-10

CAPE CANAVERAL, Fla. – The Radiation Belt Storm Probes, or RBSP, spacecraft are moved inside their payload fairing on the payload transporter from the Astrotech payload processing facility in Titusville, Fla. to Space Launch Complex-41 at Cape Canaveral Air Force Station. The fairing, which holds the twin RBSP spacecraft, will be lifted to the top of a United Launch Alliance Atlas V rocket for launch later in August. The two spacecraft are designed to study the Van Allen radiation belts in unprecedented detail. Photo credit: NASA/Dmitri Gerondidakis

NPS alternate techsat satellite, design project for AE-4871

NASA Technical Reports Server (NTRS)

1993-01-01

This project was completed as part of AE-4871, Advanced Spacecraft Design. The intent of the course is to provide experience in the design of all the major components in a spacecraft system. Team members were given responsibility for the design of one of the six primary subsystems: power, structures, propulsion, attitude control, telemetry, tracking and control (TT&C), and thermal control. In addition, a single member worked on configuration control, launch vehicle integration, and a spacecraft test plan. Given an eleven week time constraint, a preliminary design of each subsystem was completed. Where possible, possible component selections were also made. Assistance for this project came principally from the Naval Research Laboratory's Spacecraft Technology Branch. Specific information on components was solicited from representatives in industry. The design project centers on a general purpose satellite bus that is currently being sought by the Strategic Defense Initiative.

JSC Design and Procedural Standards, JSC-STD-8080

NASA Technical Reports Server (NTRS)

Punch, Danny T.

2011-01-01

This document provides design and procedural requirements appropriate for inclusion in specifications for any human spaceflight program, project, spacecraft, system, or end item. The term "spacecraft" as used in the standards includes launch vehicles, orbital vehicles, non-terrestrial surface vehicles, and modules. The standards are developed and maintained as directed by Johnson Space Center (JSC) Policy Directive JPD 8080.2, JSC Design and Procedural Standards for Human Space Flight Equipment. The Design and Procedural Standards contained in this manual represent human spacecraft design and operational knowledge applicable to a wide range of spaceflight activities. These standards are imposed on JSC human spaceflight equipment through JPD 8080.2. Designers shall comply with all design standards applicable to their design effort.

Study to define an approach for developing a computer-based system capable of automatic, unattended assembly/disassembly of spacecraft, phase 1

NASA Technical Reports Server (NTRS)

Nevins, J. L.; Defazio, T. L.; Seltzer, D. S.; Whitney, D. E.

1981-01-01

The initial set of requirements for additional studies necessary to implement a space-borne, computer-based work system capable of achieving assembly, disassembly, repair, or maintenance in space were developed. The specific functions required of a work system to perform repair and maintenance were discussed. Tasks and relevant technologies were identified and delineated. The interaction of spacecraft design and technology options, including a consideration of the strategic issues of repair versus retrieval-replacement or destruction by removal were considered along with the design tradeoffs for accomplishing each of the options. A concept system design and its accompanying experiment or test plan were discussed.

Materials for Spacecraft. Chapter 6

NASA Technical Reports Server (NTRS)

Finckenor, Miria M.

2016-01-01

The general knowledge in this chapter is intended for a broad variety of spacecraft: manned or unmanned, low Earth to geosynchronous orbit, cis-lunar, lunar, planetary, or deep space exploration. Materials for launch vehicles are covered in chapter 7. Materials used in the fabrication of spacecraft hardware should be selected by considering the operational requirements for the particular application and the design engineering properties of the candidate materials. The information provided in this chapter is not intended to replace an in-depth materials study but rather to make the spacecraft designer aware of the challenges for various types of materials and some lessons learned from more than 50 years of spaceflight. This chapter discusses the damaging effects of the space environment on various materials and what has been successfully used in the past or what may be used for a more robust design. The material categories covered are structural, thermal control for on-orbit and re-entry, shielding against radiation and meteoroid/space debris impact, optics, solar arrays, lubricants, seals, and adhesives. Spacecraft components not directly exposed to space must still meet certain requirements, particularly for manned spacecraft where toxicity and flammability are concerns. Requirements such as fracture control and contamination control are examined, with additional suggestions for manufacturability. It is important to remember that the actual hardware must be tested to understand the real, "as-built" performance, as it could vary from the design intent. Early material trades can overestimate benefits and underestimate costs. An example of this was using graphite/epoxy composite in the International Space Station science racks to save weight. By the time the requirements for vibration isolation, Space Shuttle frequencies, and experiment operations were included, the weight savings had evaporated.

Design feasibility via ascent optimality for next-generation spacecraft

NASA Astrophysics Data System (ADS)

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).

Trajectory Design for the Phobos and Deimos & Mars Environment Spacecraft

NASA Technical Reports Server (NTRS)

Genova, Anthony L.; Korsmeyer, David J.; Loucks, Michel E.; Yang, Fan Yang; Lee, Pascal

2016-01-01

The presented trajectory design and analysis was performed for the Phobos and Deimos & Mars Environment (PADME) mission concept as part of a NASA proposal submission managed by NASA Ames Research Center in the 2014-2015 timeframe. The PADME spacecraft would be a derivative of the successfully flown Lunar Atmosphere & Dust Environment Explorer (LADEE) spacecraft. While LADEE was designed to enter low-lunar orbit, the PADME spacecraft would instead enter an elliptical Mars orbit of 2-week period. This Mars orbit would pass by Phobos near periapsis on successive orbits and then raise periapsis to yield close approaches of Deimos every orbit thereafter.

Thermal design of the IMP-I and H spacecraft

NASA Technical Reports Server (NTRS)

Hoffman, R. H.

1974-01-01

A description of the thermal subsystem of the IMP-I and H spacecraft is presented. These two spacecraft were of a larger and more advanced type in the Explorer series and were successfully launched in March 1971 and September 1972. The thermal requirements, analysis, and design of each spacecraft are described including several specific designs for individual experiments. Techniques for obtaining varying degrees of thermal isolation and contact are presented. The thermal control coatings including the spaceflight performance of silver-coated FEP Teflon are discussed. Predicted performance is compared to measured flight data. The good agreement between them verifies the validity of the thermal model and the selection of coatings.

The natural space environment: Effects on spacecraft

NASA Technical Reports Server (NTRS)

James, Bonnie F.; Norton, O. W. (Compiler); Alexander, Margaret B. (Editor)

1994-01-01

The effects of the natural space environments on spacecraft design, development, and operation are the topic of a series of NASA Reference Publications currently being developed by the Electromagnetics and Environments Branch, Systems Analysis and Integration Laboratory, Marshall Space Flight Center. This primer provides an overview of the natural space environments and their effect on spacecraft design, development, and operations, and also highlights some of the new developments in science and technology for each space environment. It is hoped that a better understanding of the space environment and its effect on spacecraft will enable program management to more effectively minimize program risks and costs, optimize design quality, and successfully achieve mission objectives.

Characterization of Multilayer Piezoelectric Actuators for Use in Active Isolation Mounts

NASA Technical Reports Server (NTRS)

Wise, Stephanie A.; Hooker, Matthew W.

1997-01-01

Active mounts are desirable for isolating spacecraft science instruments from on-board vibrational sources such as motors and release mechanisms. Such active isolation mounts typically employ multilayer piezoelectric actuators to cancel these vibrational disturbances. The actuators selected for spacecraft systems must consume minimal power while exhibiting displacements of 5 to 10 micron under load. This report describes a study that compares the power consumption, displacement, and load characteristics of four commercially available multilayer piezoelectric actuators. The results of this study indicate that commercially available actuators exist that meet or exceed the design requirements used in spacecraft isolation mounts.

KSC-98pc1353

NASA Image and Video Library

1998-10-16

KENNEDY SPACE CENTER, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), the Mars Polar Lander spacecraft is on display for the media, showing an almost fully installed set of components for its launch planned for Jan. 3, 1999. The solar-powered spacecraft is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

Systems cost/performance analysis (study 2.3). Volume 3: Programmer's manual and user's guide. [for unmanned spacecraft

NASA Technical Reports Server (NTRS)

Janz, R. F.

1974-01-01

The systems cost/performance model was implemented as a digital computer program to perform initial program planning, cost/performance tradeoffs, and sensitivity analyses. The computer is described along with the operating environment in which the program was written and checked, the program specifications such as discussions of logic and computational flow, the different subsystem models involved in the design of the spacecraft, and routines involved in the nondesign area such as costing and scheduling of the design. Preliminary results for the DSCS-II design are also included.

Communication analysis for the expendable explorer spacecraft

NASA Technical Reports Server (NTRS)

1990-01-01

This report provides the results of communication analysis for the baseline and enhanced performance spacecraft designs proposed for Expendable Explorer Spacecraft (EES) series of missions. Five classes of orbits (Geosynchronous, Circular-28 degree inclination, Polar-90 degree inclination, Sunsynchronous-97 degree inclination, Molniya orbit) and a set of candidate instrument payloads provided by the ESS Study Manager were used to formulate the basis for the ESS Communications Study. The study was performed to assess the feasibility of using Space Network or ground stations for supporting the communications, tracking and data handling of the candidate instruments that are proposed to be launched into the desired orbit.

Standardization and program effect analysis (Study 2.4). Volume 2: Equipment commonality analysis. [cost savings of using flight-proven components in designing spacecraft

NASA Technical Reports Server (NTRS)

Shiokari, T.

1975-01-01

The feasibility and cost savings of using flight-proven components in designing spacecraft were investigated. The components analyzed were (1) large space telescope, (2) stratospheric aerosol and gas equipment, (3) mapping mission, (4) solar maximum mission, and (5) Tiros-N. It is concluded that flight-proven hardware can be used with not-too-extensive modification, and significant savings can be realized. The cost savings for each component are presented.

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

NASA Astrophysics Data System (ADS)

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 of the asset on orbit.

Advanced very high resolution radiometer

NASA Technical Reports Server (NTRS)

1978-01-01

The program covered the design, construction, and test of a Breadboard Model, Engineering Model, Protoflight Model, Mechanical/Structural Model, and a Life Test Model. Special bench test and calibration equipment was also developed for use on the program. Initially, the instrument was to operate from a 906 n.mi. orbit and be thermally isolated from the spacecraft. The Breadboard Model and the Mechanical/Structural Model were designed and built to these requirements. The spacecraft altitude was changed to 450 n.mi., IFOVs and spectral characteristics were modified, and spacecraft interfaces were changed. The final spacecraft design provided a temperature-controlled Instrument Mounting Platform (IMP) to carry the AVHRR and other instruments. The design of the AVHRR was modified to these new requirements and the modifications were incorporated in the Engineering Model. The Protoflight Model and the Flight Models conform to this design.

Architecture Study on Telemetry Coverage for Immediate Post-Separation Phase

NASA Technical Reports Server (NTRS)

Cheung, Kar-Ming; Lee, Charles; Kellogg, Kent; Stocklin, Frank; Zillig, David; Fielhauer, Karl

2008-01-01

This document is the viewgraphs that accompanies a paper that presents the preliminary results of an architecture study that provides continuous telemetry coverage for NASA missions for immediate post-separation phase. After launch when the spacecraft separated from the upper stage, the spacecraft typically executes a number of mission-critical operations prior to the deployment of solar panels and the activation of the primary communication subsystem. JPL, GSFC, and APL have similar design principle statements that require continuous coverage of mission-critical telemetry during the immediate post-separation phase. To conform to these design principles, an architecture that consists of a separate spacecraft transmitter and a robust communication network capable of tracking the spacecraft signals is needed. The main results of this study are as follows: 1) At low altitude (< 10000 km) when most post-separation critical operations are executed, Earth-based network (e.g. Deep Space Network (DSN)) can only provide limited coverage, whereas space-based network (e.g. Space Network (SN)) can provide continuous coverage. 2) Commercial-off-the-shelf SN compatible transmitters are available for small satellite applications. In this paper we present the detailed coverage analysis of Earth-based and Space-based networks. We identify the key functional and performance requirements of the architecture, and describe the proposed selection criteria of the spacecraft transmitter. We conclude the paper with a proposed forward plan.

Semikinematic mount for spatially constrained high aspect ratio spacecraft fold mirrors

NASA Astrophysics Data System (ADS)

Sahu, Rupali; Arora, Hemant; Munjal, Bhawdeep Singh

2017-12-01

An attempt has been made to propose a passive flexure-based semikinematic optimized mounting design for mirror fixing devices (MFDs) to mount spacecraft mirrors made of brittle materials, especially for high aspect ratio mirrors with low available space for mounting in satellites. The traditionally used tangent cantilever spiders occupy a lot of space and are suitable only for small mirrors. Similarly, the efficiency of flexural bipods is lost if not placed 120 deg apart, which is not possible in high aspect ratio mirrors. Two mounting configurations, one with collinear MFDs and the other with staggered MFDs, have been studied. An optimization problem is set up with dimensions of the proposed design as design variables and constraints imposed on structural performance of the mirror assembly. Investigations indicate that both configurations have potential applications in spacecrafts as they have provided feasible results and have satisfactory optical performance as well.

The Magnetospheric Multiscale Constellation

NASA Technical Reports Server (NTRS)

Tooley, C. R.; Black, R. K.; Robertson, B. P.; Stone, J. M.; Pope, S. E.; Davis, G. T.

2015-01-01

The Magnetospheric Multiscale (MMS) mission is the fourth mission of the Solar Terrestrial Probe (STP) program of the National Aeronautics and Space Administration (NASA). The MMS mission was launched on March 12, 2015. The MMS mission consists of four identically instrumented spin-stabilized observatories which are flown in formation to perform the first definitive study of magnetic reconnection in space. The MMS mission was presented with numerous technical challenges, including the simultaneous construction and launch of four identical large spacecraft with 100 instruments total, stringent electromagnetic cleanliness requirements, closed-loop precision maneuvering and pointing of spinning flexible spacecraft, on-board GPS based orbit determination far above the GPS constellation, and a flight dynamics design that enables formation flying with separation distances as small as 10 km. This paper describes the overall mission design and presents an overview of the design, testing, and early on-orbit operation of the spacecraft systems and instrument suite.

Mission Design for the Innovative Interstellar Explorer Vision Mission

NASA Technical Reports Server (NTRS)

Fiehler, Douglas I.; McNutt, Ralph L.

2005-01-01

The Innovative Interstellar Explorer, studied under a NASA Vision Mission grant, examined sending a probe to a heliospheric distance of 200 Astronomical Units (AU) in a "reasonable" amount of time. Previous studies looked at the use of a near-Sun propulsive maneuver, solar sails, and fission reactor powered electric propulsion systems for propulsion. The Innovative Interstellar Explorer's mission design used a combination of a high-energy launch using current launch technology, a Jupiter gravity assist, and electric propulsion powered by advanced radioisotope power systems to reach 200 AU. Many direct and gravity assist trajectories at several power levels were considered in the development of the baseline trajectory, including single and double gravity assists utilizing the outer planets (Jupiter, Saturn, Uranus, and Neptune). A detailed spacecraft design study was completed followed by trajectory analyses to examine the performance of the spacecraft design options.

Some problems of the design of highly directional spacecraft antennas

NASA Technical Reports Server (NTRS)

Prigoda, B. A.

1974-01-01

Problems of optimization and selection of the most expedient forms of design of directional antenna systems encountered in spacecraft design are discussed. Selection of a given type of antenna depends on its characteristic size, weight, and potential.

Powersail High Power Propulsion System Design Study

NASA Astrophysics Data System (ADS)

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.

Bounding the Spacecraft Atmosphere Design Space for Future Exploration Missions

NASA Technical Reports Server (NTRS)

Lange, Kevin E.; Perka, Alan T.; Duffield, Bruce E.; Jeng, Frank F.

2005-01-01

The selection of spacecraft and space suit atmospheres for future human space exploration missions will play an important, if not critical, role in the ultimate safety, productivity, and cost of such missions. Internal atmosphere pressure and composition (particularly oxygen concentration) influence many aspects of spacecraft and space suit design, operation, and technology development. Optimal atmosphere solutions must be determined by iterative process involving research, design, development, testing, and systems analysis. A necessary first step in this process is the establishment of working bounds on the atmosphere design space.

Long life assurance study for manned spacecraft long life hardware. Volume 1: Summary of long life assurance guidelines

NASA Technical Reports Server (NTRS)

1972-01-01

A long life assurance program for the development of design, process, test, and application guidelines for achieving reliable spacecraft hardware was conducted. The study approach consisted of a review of technical data performed concurrently with a survey of the aerospace industry. The data reviewed included design and operating characteristics, failure histories and solutions, and similar documents. The topics covered by the guidelines are reported. It is concluded that long life hardware is achieved through meticulous attention to many details and no simple set of rules can suffice.

High Capacity Communications From Martian Distances. Part 1; Spacecraft Link Design Analysis

NASA Technical Reports Server (NTRS)

Vyas, Hemali N.; Schuchman, Leonard; Orr, Richard; Williams, Wallace Dan; Collins, Michael; Noreen, Gary

2006-01-01

High capacity space communications has been a desire for Human Exploration and Science missions. Current Mars missions operate at data rates of 120 kbps for telemetry downlink and it is desirable to study high rate communication links in the range of 100 Mbps to 1 Gbps data rates from Martian distances. This paper will present some assumed scenarios along with link design assumptions and link analysis for high capacity communications from Mars. The paper will focus on RF subsystems namely antenna and power for the downlink communication from a relay orbiter at Mars. The relay orbiter will communicate with the low orbit spacecrafts at Mars or any Martian surface elements such as robots, and relay the data back to the ground networks on Earth. The study will dive into the spacecraft downlink system design and communication link analysis between the relay orbiter and ground network on Earth for data rates ranging from 100 Mbps to 1 Gbps based on the assumed scenarios and link assumptions. With high rate links at larger distances, there will be a significant impact on the antenna and power requirements and the link design will make an attempt to minimize the mass of the RF subsystem on the spacecraft. The results of this study will be presented for three data rates 1 Gbps, 500 Mbps and 100 Mbps at maximum Mars to Earth distance of 2.67AU. The design will use a Ka-band downlink with 90% link availability, along with various ground network G/T assumptions and possible bandwidth efficient modulations. The paper will conclude with what types of high rate communication links are feasible from Martian distances and also identify a range of requirements for antenna and power technologies for these high capacity communications from Mars.

The Design of a Power System for the PETSAT Modular Small Spacecraft Bus

NASA Astrophysics Data System (ADS)

Clark, C. S.; Lopez Mazarias, A.; Kobayashi, C.; Nakasuka, S.

2008-08-01

There is considerable interest in the benefits of having a modular spacecraft where it is possible to construct a satellite using a number of modules with identical mechanical and electrical interfaces, but with each performing a specific function to achieve the required platform specification. In recent years, steps have been made towards modular spacecraft becoming a reality and the concept is due to be demonstrated in-orbit later this year with the first flight of the PETSAT spacecraft concept on the mission, SOHLA-2. This paper describes the approach to the design of the SOHLA-2 power system. The approach is significant; PETSAT is an excellent example of a modular approach to spacecraft design. The PETSAT concept consists of a number of 'Panel Modules', roughly the same size as a pizza box. The panels stack together in stowed configuration for launch, and unfold once in orbit. Apart from being a very novel approach to spacecraft design and construction, this concept offers advantages in power generation as, once unfolded, there is significant surface area on which to mount solar cells for power generation. The power system for PETSAT has been designed such that each Panel Module contains a power system that can either operate in isolation for the purpose of unit testing, or as part of a larger spacecraft power system once connected to other Panel Modules. When connected together, the power systems on each module share the energy from the solar arrays and the batteries. The approach to the design of the system has provided a simple solution to difficult problem.

NASA-STD-4005 and NASA-HDBK-4006, LEO Spacecraft Solar Array Charging Design Standard

NASA Technical Reports Server (NTRS)

Ferguson, Dale C.

2007-01-01

Two new NASA Standards are now official. They are the NASA LEO Spacecraft Charging Design Standard (NASA-STD-4005) and the NASA LEO Spacecraft Charging Design Handbook (NASA-HDBK-4006). They give the background and techniques for controlling solar array-induced charging and arcing in LEO. In this paper, a brief overview of the new standards is given, along with where they can be obtained and who should be using them.

Engineering design, stress and thermal analysis, and documentation for SATS program

NASA Technical Reports Server (NTRS)

1973-01-01

An in-depth analysis and mechanical design of the solar array stowage and deployment arrangements for use in Small Applications Technology Satellite spacecraft is presented. Alternate approaches for the major elements of work are developed and evaluated. Elements include array stowage and deployment arrangements, the spacecraft and array behavior in the spacecraft despin mode, and the design of the main hinge and segment hinge assemblies. Feasibility calculations are performed and the preferred approach is identified.

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

NASA Technical Reports Server (NTRS)

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.

Mechanical Design of Spacecraft

NASA Technical Reports Server (NTRS)

1962-01-01

In the spring of 1962, engineers from the Engineering Mechanics Division of the Jet Propulsion Laboratory gave a series of lectures on spacecraft design at the Engineering Design seminars conducted at the California Institute of Technology. Several of these lectures were subsequently given at Stanford University as part of the Space Technology seminar series sponsored by the Department of Aeronautics and Astronautics. Presented here are notes taken from these lectures. The lectures were conceived with the intent of providing the audience with a glimpse of the activities of a few mechanical engineers who are involved in designing, building, and testing spacecraft. Engineering courses generally consist of heavily idealized problems in order to allow the more efficient teaching of mathematical technique. Students, therefore, receive a somewhat limited exposure to actual engineering problems, which are typified by more unknowns than equations. For this reason it was considered valuable to demonstrate some of the problems faced by spacecraft designers, the processes used to arrive at solutions, and the interactions between the engineer and the remainder of the organization in which he is constrained to operate. These lecture notes are not so much a compilation of sophisticated techniques of analysis as they are a collection of examples of spacecraft hardware and associated problems. They will be of interest not so much to the experienced spacecraft designer as to those who wonder what part the mechanical engineer plays in an effort such as the exploration of space.

Weight and cost forecasting for advanced manned space vehicles

NASA Technical Reports Server (NTRS)

Williams, Raymond

1989-01-01

A mass and cost estimating computerized methology for predicting advanced manned space vehicle weights and costs was developed. The user friendly methology designated MERCER (Mass Estimating Relationship/Cost Estimating Relationship) organizes the predictive process according to major vehicle subsystem levels. Design, development, test, evaluation, and flight hardware cost forecasting is treated by the study. This methodology consists of a complete set of mass estimating relationships (MERs) which serve as the control components for the model and cost estimating relationships (CERs) which use MER output as input. To develop this model, numerous MER and CER studies were surveyed and modified where required. Additionally, relationships were regressed from raw data to accommodate the methology. The models and formulations which estimated the cost of historical vehicles to within 20 percent of the actual cost were selected. The result of the research, along with components of the MERCER Program, are reported. On the basis of the analysis, the following conclusions were established: (1) The cost of a spacecraft is best estimated by summing the cost of individual subsystems; (2) No one cost equation can be used for forecasting the cost of all spacecraft; (3) Spacecraft cost is highly correlated with its mass; (4) No study surveyed contained sufficient formulations to autonomously forecast the cost and weight of the entire advanced manned vehicle spacecraft program; (5) No user friendly program was found that linked MERs with CERs to produce spacecraft cost; and (6) The group accumulation weight estimation method (summing the estimated weights of the various subsystems) proved to be a useful method for finding total weight and cost of a spacecraft.

Unmanned Vehicle Material Flammability Test

NASA Technical Reports Server (NTRS)

Urban, David; Ruff, Gary A.; Fernandez-Pello, A. Carlos; T’ien, James S.; Torero, Jose L.; Cowlard, Adam; Rouvreau, Sebastian; Minster, Olivier; Toth, Balazs; Legros, Guillaume;

A Technique for the Assessment of Flight Operability Characteristics of Human Rated Spacecraft

NASA Technical Reports Server (NTRS)

Crocker, Alan

2010-01-01

In support of new human rated spacecraft development programs, the Mission Operations Directorate at NASA Johnson Space Center has implemented a formal method for the assessment of spacecraft operability. This "Spacecraft Flight Operability Assessment Scale" defines six key themes of flight operability, with guiding principles and goals stated for each factor. A standardized rating technique provides feedback that is useful to the operations, design and program management communities. Applicability of this concept across the program structure and life cycle is addressed. Examples of operationally desirable and undesirable spacecraft design characteristics are provided, as is a sample of the assessment scale product.

An investigation of a sterile access technique for the repair and adjustment of sterile spacecraft

NASA Technical Reports Server (NTRS)

Farmer, F. H.; Fuller, H. V.; Hueschen, R. M.

1973-01-01

A description is presented of a unique system for the sterilization and sterile repair of spacecraft and the results of a test program designed to assess the biological integrity and engineering reliability of the system. This trailer-mounted system, designated the model assembly sterilizer for testing (MAST), is capable of the dry-heat sterilization of spacecraft and/or components less than 2.3 meters in diameter at temperatures up to 433 K and the steam sterilization of components less than 0.724 meter in diameter. Sterile access to spacecraft is provided by two tunnel suits, called the bioisolator suit systems (BISS), which are contiguous with the walls of the sterilization chambers. The test program was designed primarily to verify the biological and engineering reliability of the MAST system by processing simulated space hardware. Each test cycle simulated the initial sterilization of a spacecraft, sterile repair of a failed component, removal of the spacecraft from the MAST for mating with the bus, and a sterile recycle repair.

Baseline spacecraft and mission design for the SP-100 flight experiment

NASA Technical Reports Server (NTRS)

Deininger, William D.; Vondra, Robert J.

1989-01-01

The design and performance of a spacecraft employing arcjet nuclear electric propulsion, suitable for use in the SP-100 Space Reactor Power System (SRPS) Flight Experiment, are outlined. The vehicle design is based on a 93 kWe ammonia arcjet system operating at an experimentally-measured specific impulse of 1030 s and an efficiency of 42 percent. The arcjet/gimbal assemblies, power conditioning subsystem, propellant feed system, propulsion system thermal control, spacecraft diagnostic instrumentation, and the telemetry requirements are described. A 100 kWe SRPS is assumed. The total spacecraft mass is baselined at 5675 kg excluding the propellant and propellant feed system. Four mission scenarios are described which are capable of demonstrating the full capability of the SRPS. The missions considered include spacecraft deployment to possible surveillance platform orbits, a spacecraft storage mission and an orbit raising round trip corresponding to possible orbit transfer vehicle missions. Launches from Kennedy Space Center using the Titan IV expendable launch vehicle are assumed.

Full-Scale Spacecraft Simulator Design for a 2D Zero Gravity Small Body Surface Sampling Validation

NASA Astrophysics Data System (ADS)

Mongelli, Marco

NASA is developing several Touch-And-Go (TAG) classes of missions. These types of missions like the OSIRIS-REx asteroid sample return [1] or a comet sample return mission (CSSR)[2], consist usually in three phases: propulsive approach to the target, sampling and propulsion to move the spacecraft away from the target. The development of TAG mission, from concept to realization, is usually divided in two phases: Phase I discusses the major trades that could affect the mission architecture; Phase II focuses in detail on the design. This project of a spacecraft emulator fits into phase II and specifically on the way the spacecraft could react in absence of gravity while the Sample Acquisition System (SAS) is collecting the sample. A full-scale spacecraft on a 2D zero-friction environment has been designed. Also a propulsion system has been implemented to re-create the full dynamics of a spacecraft in space.

Design of multi-mission chemical propulsion modules for planetary orbiters. Volume 1: Summary report

NASA Technical Reports Server (NTRS)

1975-01-01

Results are presented of a conceptual design and feasibility study of chemical propulsion stages that can serve as modular propulsion units, with little or no modification, on a variety of planetary orbit missions, including orbiters of Mercury, Saturn, and Uranus. Planetary spacecraft of existing design or currently under development, viz., spacecraft of the Pioneer and Mariner families, are assumed as payload vehicles. Thus, operating requirements of spin-stabilized and 3-axis stabilized spacecraft have to be met by the respective propulsion module designs. As launch vehicle for these missions the Shuttle orbiter and interplanetary injection stage, or Tug, plus solid-propellant kick motor was assumed. Accommodation constraints and interfaces involving the payloads and the launch vehicle are considered in the propulsion module design. The applicability and performance advantages were evaluated of the space-storable high-energy bipropellants. The incentive for using this advanced propulsion technology on planetary missions is the much greater performance potential when orbit insertion velocities in excess of 4 km/sec are required, as in the Mercury orbiter. Design analyses and performance tradeoffs regarding earth-storable versus space-storable propulsion systems are included. Cost and development schedules of multi-mission versus custom-designed propulsion modules are examined.

Parametric Study of Variable Emissivity Radiator Surfaces

NASA Technical Reports Server (NTRS)

Grob, Lisa M.; Swanson, Theodore D.

2000-01-01

The goal of spacecraft thermal design is to accommodate a high function satellite in a low weight and real estate package. The extreme environments that the satellite is exposed during its orbit are handled using passive and active control techniques. Heritage passive heat rejection designs are sized for the hot conditions and augmented for the cold end with heaters. The active heat rejection designs to date are heavy, expensive and/or complex. Incorporating an active radiator into the design that is lighter, cheaper and more simplistic will allow designers to meet the previously stated goal of thermal spacecraft design Varying the radiator's surface properties without changing the radiating area (as with VCHP), or changing the radiators' views (traditional louvers) is the objective of the variable emissivity (vary-e) radiator technologies. A parametric evaluation of the thermal performance of three such technologies is documented in this paper. Comparisons of the Micro-Electromechanical Systems (MEMS), Electrochromics, and Electrophoretics radiators to conventional radiators, both passive and active are quantified herein. With some noted limitations, the vary-e radiator surfaces provide significant advantages over traditional radiators and a promising alternative design technique for future spacecraft thermal systems.

Benefits of Spacecraft Level Vibration Testing

NASA Technical Reports Server (NTRS)

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.

Miniature star tracker for small remote sensing satellites

NASA Astrophysics Data System (ADS)

Cassidy, Lawrence W.; Schlom, Leslie

1995-01-01

Designers of future remote sensing spacecraft, including platforms for Mission to Planet Earth and small satellites, will be driven to provide spacecraft designs that maximize data return and minimize hardware and operating costs. The attitude determination subsystems of these spacecraft must likewise provide maximum capability and versatility at an affordable price. Hughes Danbury Optical Systems (HDOS) has developed the Model HD-1003 Miniature Star Tracker which combines high accuracy, high reliability and growth margin for `all-stellar' capability in a compact, radiation tolerant design that meets these future spacecraft needs and whose cost is competitive with horizon sensors and digital fine sum sensors. Begun in 1991, our HD-1003 development program has now entered the hardware qualification phase. This paper acquaints spacecraft designers with the design and performance capabilities of the HD- 1003 tracker. We highlight the tracker's unique features which include: (1) Very small size (165 cu. in.). (2) Low weight (7 lbs). (3) Multi-star tracking (6 stars simultaneously). (4) Eighteen arc-sec (3-sigma) accuracy. (5) Growth margin for `all-stellar' attitude reference.

Reengineering the JPL Spacecraft Design Process

NASA Technical Reports Server (NTRS)

Briggs, C.

1995-01-01

This presentation describes the factors that have emerged in the evolved process of reengineering the unmanned spacecraft design process at the Jet Propulsion Laboratory in Pasadena, California. Topics discussed include: New facilities, new design factors, new system-level tools, complex performance objectives, changing behaviors, design integration, leadership styles, and optimization.

Earth Orbit Raise Design for the Artemis Mission

NASA Technical Reports Server (NTRS)

Wiffen, Gregory J.; Sweetser, Theodore H.

2011-01-01

The Artemis mission is an extension of the Themis mission. The Themis mission1 consisted of five identical spacecraft in varying sized Earth orbits designed to make simultaneous measurements of the Earth's electric and magnetic environment. Themis was designed to observe geomagnetic storms resulting from solar wind's interaction with the Earth's magnetosphere. Themis was meant to answer the age old question of why the Earth's aurora can change rapidly on a global scale. The Themis spacecraft are spin stabilized with 20 meter long electric field booms as well as several shorter magnetometer booms. The goal of the Artemis2 mission extension is to deliver the field and particle measuring capabilities of two of the Themis spacecraft to the vicinity of the Moon. The Artemis mission required transferring two Earth orbiting Themis spacecraft on to two different low energy trans-lunar trajectories ultimately ending in lunar orbit. This paper describes the processes that resulted in successful orbit raise designs for both spacecraft.

Multiple spacecraft configuration designs for coordinated flight missions

NASA Astrophysics Data System (ADS)

Fumenti, Federico; Theil, Stephan

2018-06-01

Coordinated flight allows the replacement of a single monolithic spacecraft with multiple smaller ones, based on the principle of distributed systems. According to the mission objectives and to ensure a safe relative motion, constraints on the relative distances need to be satisfied. Initially, differential perturbations are limited by proper orbit design. Then, the induced differential drifts can be properly handled through corrective maneuvers. In this work, several designs are surveyed, defining the initial configuration of a group of spacecraft while counteracting the differential perturbations. For each of the investigated designs, focus is placed upon the number of deployable spacecraft and on the possibility to ensure safe relative motion through station keeping of the initial configuration, with particular attention to the required Δ V budget and the constraints violations.

Magnet-Based System for Docking of Miniature Spacecraft

NASA Technical Reports Server (NTRS)

Howard, Nathan; Nguyen, Hai D.

2007-01-01

A prototype system for docking a miniature spacecraft with a larger spacecraft has been developed by engineers at the Johnson Space Center. Engineers working on Mini AERCam, a free-flying robotic camera, needed to find a way to successfully dock and undock their miniature spacecraft to refuel the propulsion and recharge the batteries. The subsystems developed (see figure) include (1) a docking port, designed for the larger spacecraft, which contains an electromagnet, a ball lock mechanism, and a service probe; and (2) a docking cluster, designed for the smaller spacecraft, which contains either a permanent magnet or an electromagnet. A typical docking operation begins with the docking spacecraft maneuvering into position near the docking port on the parent vehicle. The electromagnet( s) are then turned on, and, if necessary, the docking spacecraft is then maneuvered within the capture envelope of the docking port. The capture envelope for this system is approximated by a 5-in. (12.7-cm) cube centered on the front of the docking-port electromagnet and within an angular misalignment of <30 . Thereafter, the magnetic forces draw the smaller spacecraft toward the larger one and this brings the spacecraft into approximate alignment prior to contact. Mechanical alignment guides provide the final rotational alignment into one of 12 positions. Once the docking vehicle has been captured magnetically in the docking port, the ball-lock mechanism is activated, which locks the two spacecraft together. At this point the electromagnet( s) are turned off, and the service probe extended if recharge and refueling are to be performed. Additionally, during undocking, the polarity of one electromagnet can be reversed to provide a gentle push to separate the two spacecraft. This system is currently being incorporated into the design of Mini AERCam vehicle.

HELIOS Third Joint Working Group Meeting

NASA Technical Reports Server (NTRS)

Ousley, Gilbert; Kutzer, Ants

1970-01-01

During the past six months since the Second Helios Joint Working Group Meeting held 27-30 April 1970 at Goddard Space Flight Center, the TDS Sub-Group supported the Helios Project Office and the other Sub-Groups in the timely disposition of action items and the dissemination of information pertinent to the development of interface documentation. Of particular importance during this time period was the Project's decision to incorporate a single-channel telemetry system design aboard the spacecraft. The TDS Sub-Group participated actively in the process that led to this decision. Still under active study with TDS participation is the pending Project Office decision regarding the incorporation of a ranging capability within the telecommunications design. The TDS Sub-Group assisted the Mission Analysis and Operations Sub-Group in establishment of a study effort concerning the Near-Earth Sequence of Events from launch to launch plus 8 hours. This study, which will provide valuable data for the spacecraft telecommunications design, will include participation by the Experiment, Launch Vehicle, Spacecraft, as well as the TDS and MA&O Sub-Groups. Also during the past 6-month period, the TDS, in conjunction with the Spacecraft Sub-Group, initiated activity to develop the Helios Spacecraft/TDS Compatibility Test Plans and Procedures. Activity concerning the foregoing interface discussions has been and will continue to be based upon the "TDS Estimated Capabilities Document for the Helios Missions" (613-1), and the "DSN/Flight Project Interface Design Handbook" (810-5). These will continue to be considered TDS controlling documents until specific Helios Project/TDS interface documentation is generated and signed off by the respective parties. In addition to the above, the DSN continued the Helios Trainee Program with seven GfW/DFVLR trainees in residence at JPL. Two trainees will complete their year's residency concurrent with the Third Helios Joint Working Group Meeting, while four new trainees are expected to arrive following the Third Helios Joint Working Group Meeting. These and other activities are reported in detail in the paragraphs that follow.

Distributed parameter modelling of flexible spacecraft: Where's the beef?

NASA Technical Reports Server (NTRS)

Hyland, D. C.

1994-01-01

This presentation discusses various misgivings concerning the directions and productivity of Distributed Parameter System (DPS) theory as applied to spacecraft vibration control. We try to show the need for greater cross-fertilization between DPS theorists and spacecraft control designers. We recommend a shift in research directions toward exploration of asymptotic frequency response characteristics of critical importance to control designers.

Ka-Band Autonomous Formation Flying Sensor

NASA Technical Reports Server (NTRS)

Tien, Jeffrey; Purcell, George, Jr.; Srinivasan, Jeffrey; Ciminera, Michael; Srinivasan, Meera; Meehan, Thomas; Young, Lawrence; Aung, MiMi; Amaro, Luis; Chong, Yong;

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

NASA Technical Reports Server (NTRS)

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 consideration and measurements of resistivity for many typical spacecraft charging scenarios.

A unified approach to computer analysis and modeling of spacecraft environmental interactions

NASA Technical Reports Server (NTRS)

Katz, I.; Mandell, M. J.; Cassidy, J. J.

1986-01-01

A new, coordinated, unified approach to the development of spacecraft plasma interaction models is proposed. The objective is to eliminate the unnecessary duplicative work in order to allow researchers to concentrate on the scientific aspects. By streamlining the developmental process, the interchange between theories and experimentalists is enhanced, and the transfer of technology to the spacecraft engineering community is faster. This approach is called the UNIfied Spacecraft Interaction Model (UNISIM). UNISIM is a coordinated system of software, hardware, and specifications. It is a tool for modeling and analyzing spacecraft interactions. It will be used to design experiments, to interpret results of experiments, and to aid in future spacecraft design. It breaks a Spacecraft Ineraction analysis into several modules. Each module will perform an analysis for some physical process, using phenomenology and algorithms which are well documented and have been subject to review. This system and its characteristics are discussed.

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

NASA Technical Reports Server (NTRS)

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.

Time Frequency Analysis of Spacecraft Propellant Tank Spinning Slosh

NASA Technical Reports Server (NTRS)

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.

Probe interface design consideration. [for interplanetary spacecraft missions

NASA Technical Reports Server (NTRS)

Casani, E. K.

1974-01-01

Interface design between a probe and a spacecraft requires not only technical considerations but also management planning and mission analysis interactions. Two further aspects of importance are the flyby versus the probe trade-off, and the relay link design and data handling optimization.

A Web Based Collaborative Design Environment for Spacecraft

NASA Technical Reports Server (NTRS)

Dunphy, Julia

1998-01-01

In this era of shrinking federal budgets in the USA we need to dramatically improve our efficiency in the spacecraft engineering design process. We have come up with a method which captures much of the experts' expertise in a dataflow design graph: Seamlessly connectable set of local and remote design tools; Seamlessly connectable web based design tools; and Web browser interface to the developing spacecraft design. We have recently completed our first web browser interface and demonstrated its utility in the design of an aeroshell using design tools located at web sites at three NASA facilities. Multiple design engineers and managers are now able to interrogate the design engine simultaneously and find out what the design looks like at any point in the design cycle, what its parameters are, and how it reacts to adverse space environments.

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

NASA Technical Reports Server (NTRS)

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 conceptual design of a spacecraft. Engineering design is as much an art as it is a science. The critical thinking skills necessary to uncover lurking problems in an experimental design and creatively develop solutions are some of the same skills necessary to design a new spacecraft. Thus, I believe engineers unfamiliar with or removed from design and development need time to transition and develop the required skill set to be effective spacecraft designers. I believe the creative process necessary in design can be enhanced and even taught as early as grades K-12 and should continue to be nurtured and developed at the university level and beyond. I am going to present a strategy for developing learning teams to address complex multidisciplinary problems and to creatively develop solutions to those problems rapidly at minimal cost. I will frame a real problem, the development of on-orbit thermal protection system repair of the Space Shuttle, and step through the series of skills necessary to enhance the creative process. The case study I will illustrate is based on a real project, the R&D Reinforced Carbon-Carbon (RCC) Repair Team's development of on-orbit repair concepts for damaged Space Shuttle RCC nose cap and/or leading edges.

NASA's Space Environments and Effects (SEE) Program: Meteoroid and Orbital Debris Lesson Plan.

ERIC Educational Resources Information Center

National Aeronautics and Space Administration, Washington, DC.

The study of the natural space environment and its effects on spacecraft is one of the most important and least understood aspects of spacecraft design. The Space Environments and Effects (SEE) Program prepared the Meteoroids and Orbital Debris Lesson Plan, a SEE-focused high school curriculum to engage students in creative activities that will…

Tone Noise Predictions for a Spacecraft Cabin Ventilation Fan Ingesting Distorted Inflow and the Challenges of Validation

NASA Technical Reports Server (NTRS)

Koch, L. Danielle; Shook, Tony D.; Astler, Douglas T.; Bittinger, Samantha A.

2011-01-01

A fan tone noise prediction code has been developed at NASA Glenn Research Center that is capable of estimating duct mode sound power levels for a fan ingesting distorted inflow. This code was used to predict the circumferential and radial mode sound power levels in the inlet and exhaust duct of an axial spacecraft cabin ventilation fan. Noise predictions at fan design rotational speed were generated. Three fan inflow conditions were studied: an undistorted inflow, a circumferentially symmetric inflow distortion pattern (cylindrical rods inserted radially into the flowpath at 15deg, 135deg, and 255deg), and a circumferentially asymmetric inflow distortion pattern (rods located at 15deg, 52deg and 173deg). Noise predictions indicate that tones are produced for the distorted inflow cases that are not present when the fan operates with an undistorted inflow. Experimental data are needed to validate these acoustic predictions, as well as the aerodynamic performance predictions. Given the aerodynamic design of the spacecraft cabin ventilation fan, a mechanical and electrical conceptual design study was conducted. Design features of a fan suitable for obtaining detailed acoustic and aerodynamic measurements needed to validate predictions are discussed.

Tone Noise Predictions for a Spacecraft Cabin Ventilation Fan Ingesting Distorted Inflow and the Challenges of Validation

NASA Technical Reports Server (NTRS)

Koch, L. Danielle; Shook, Tony D.; Astler, Douglas T.; Bittinger, Samantha A.

2012-01-01

A fan tone noise prediction code has been developed at NASA Glenn Research Center that is capable of estimating duct mode sound power levels for a fan ingesting distorted inflow. This code was used to predict the circumferential and radial mode sound power levels in the inlet and exhaust duct of an axial spacecraft cabin ventilation fan. Noise predictions at fan design rotational speed were generated. Three fan inflow conditions were studied: an undistorted inflow, a circumferentially symmetric inflow distortion pattern (cylindrical rods inserted radially into the flowpath at 15deg, 135deg, and 255deg), and a circumferentially asymmetric inflow distortion pattern (rods located at 15deg, 52deg and 173deg). Noise predictions indicate that tones are produced for the distorted inflow cases that are not present when the fan operates with an undistorted inflow. Experimental data are needed to validate these acoustic predictions, as well as the aerodynamic performance predictions. Given the aerodynamic design of the spacecraft cabin ventilation fan, a mechanical and electrical conceptual design study was conducted. Design features of a fan suitable for obtaining detailed acoustic and aerodynamic measurements needed to validate predictions are discussed.

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

NASA Technical Reports Server (NTRS)

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.

Materials Characterization at Utah State University: Facilities and Knowledge-base of Electronic Properties of Materials Applicable to Spacecraft Charging

NASA Technical Reports Server (NTRS)

Dennison, J. R.; Thomson, C. D.; Kite, J.; Zavyalov, V.; Corbridge, Jodie

2004-01-01

In an effort to improve the reliability and versatility of spacecraft charging models designed to assist spacecraft designers in accommodating and mitigating the harmful effects of charging on spacecraft, the NASA Space Environments and Effects (SEE) Program has funded development of facilities at Utah State University for the measurement of the electronic properties of both conducting and insulating spacecraft materials. We present here an overview of our instrumentation and capabilities, which are particularly well suited to study electron emission as related to spacecraft charging. These measurements include electron-induced secondary and backscattered yields, spectra, and angular resolved measurements as a function of incident energy, species and angle, plus investigations of ion-induced electron yields, photoelectron yields, sample charging and dielectric breakdown. Extensive surface science characterization capabilities are also available to fully characterize the samples in situ. Our measurements for a wide array of conducting and insulating spacecraft materials have been incorporated into the SEE Charge Collector Knowledge-base as a Database of Electronic Properties of Materials Applicable to Spacecraft Charging. This Database provides an extensive compilation of electronic properties, together with parameterization of these properties in a format that can be easily used with existing spacecraft charging engineering tools and with next generation plasma, charging, and radiation models. Tabulated properties in the Database include: electron-induced secondary electron yield, backscattered yield and emitted electron spectra; He, Ar and Xe ion-induced electron yields and emitted electron spectra; photoyield and solar emittance spectra; and materials characterization including reflectivity, dielectric constant, resistivity, arcing, optical microscopy images, scanning electron micrographs, scanning tunneling microscopy images, and Auger electron spectra. Further details of the instrumentation used for insulator measurements and representative measurements of insulating spacecraft materials are provided in other Spacecraft Charging Conference presentations. The NASA Space Environments and Effects Program, the Air Force Office of Scientific Research, the Boeing Corporation, NASA Graduate Research Fellowships, and the NASA Rocky Mountain Space Grant Consortium have provided support.

The Outer Planetary Mission Design Project

NASA Astrophysics Data System (ADS)

Benfield, Michael; Turner, M. W.

2010-10-01

With the recent focus from the planetary science community on the outer planets of the solar system, The University of Alabama in Huntsville Integrated Product Team program is embarking on a new challenge to develop an outer planetary mission for the academic year 2010-2011. Currently four bodies are of interest for this mission: Titan, Europa, Triton, and Enceledus, with one body being chosen by the instructors by the beginning of the fall semester. This project will use the 2010 Discovery Announcement of Opportunity as its Request for Proposal (RFP). All of the teams competing in this project will use the AO to respond with a proposal to the instructors for their proposed mission and spacecraft concept. The project employs the two-semester design sequence of the IPT program to provide a framework for the development of this mission. This sequence is divided into four phases. Phase 1 - Requirements Development - focuses on the development of both the scientific and engineering requirements of the mission. During this phase the teams work very closely with the PI organization, represented by the College of Charleston. Phase 2 - Team Formation and Architecture Development - concentrates on the assessment of the overall mission architecture from the launch vehicle to the ground operations of the proposed spacecraft. Phase 3 - System Definition - provides for spacecraft subsystem trade studies and further refinement of the specific spacecraft to meet the scientific requirements and objectives developed in Phase 1. Phase 4 - Design - is the phase where the engineers provide the spacecraft design that is required for the mission of interest. At the conclusion of Phases 2 and 4, an external review board evaluates the proposed designs and chooses one winner of the competition.

An Earth-Moon Transfer Trajectory Design and Analysis Considering Spacecraft's Visibility from Daejeon Ground Station at TLI and LOI Maneuvers

NASA Astrophysics Data System (ADS)

Woo, Jin; Song, Young-Joo; Park, Sang-Young; Kim, Hae-Dong; Sim, Eun-Sup

2010-09-01

The optimal Earth-Moon transfer trajectory considering spacecraft's visibility from the Daejeon ground station visibility at both the trans lunar injection (TLI) and lunar orbit insertion (LOI) maneuvers is designed. Both the TLI and LOI maneuvers are assumed to be impulsive thrust. As the successful execution of the TLI and LOI maneuvers are crucial factors among the various lunar mission parameters, it is necessary to design an optimal lunar transfer trajectory which guarantees the visibility from a specified ground station while executing these maneuvers. The optimal Earth-Moon transfer trajectory is simulated by modifying the Korean Lunar Mission Design Software using Impulsive high Thrust Engine (KLMDS-ITE) which is developed in previous studies. Four different mission scenarios are established and simulated to analyze the effects of the spacecraft's visibility considerations at the TLI and LOI maneuvers. As a result, it is found that the optimal Earth-Moon transfer trajectory, guaranteeing the spacecraft's visibility from Daejeon ground station at both the TLI and LOI maneuvers, can be designed with slight changes in total amount of delta-Vs. About 1% difference is observed with the optimal trajectory when none of the visibility condition is guaranteed, and about 0.04% with the visibility condition is only guaranteed at the time of TLI maneuver. The spacecraft's mass which can delivered to the Moon, when both visibility conditions are secured is shown to be about 534 kg with assumptions of KSLV-2's on-orbit mass about 2.6 tons. To minimize total mission delta-Vs, it is strongly recommended that visibility conditions at both the TLI and LOI maneuvers should be simultaneously implemented to the trajectory optimization algorithm.

Low-Impact Mating System for Docking Spacecraft

NASA Technical Reports Server (NTRS)

Lewis, James L.; Robertson, Brandan; Carroll, Monty B.; Le, Thang; Morales, Ray

2008-01-01

A document describes a low-impact mating system suitable for both docking (mating of two free-flying spacecraft) and berthing (in which a robot arm in one spacecraft positions an object for mating with either spacecraft). The low-impact mating system is fully androgynous: it mates with a copy of itself, i.e., all spacecraft and other objects to be mated are to be equipped with identical copies of the system. This aspect of the design helps to minimize the number of unique parts and to standardize and facilitate mating operations. The system includes a closed-loop feedback control subsystem that actively accommodates misalignments between mating spacecraft, thereby attenuating spacecraft dynamics and mitigating the need for precise advance positioning of the spacecraft. The operational characteristics of the mating system can be easily configured in software, during operation, to enable mating of spacecraft having various masses, center-of-gravity offsets, and closing velocities. The system design provides multi-fault tolerance for critical operations: for example, to ensure unmating at a critical time, a redundant unlatching mechanism and two independent pyrotechnic release subsystems are included.

Space Station man-machine automation trade-off analysis

NASA Technical Reports Server (NTRS)

Zimmerman, W. F.; Bard, J.; Feinberg, A.

1985-01-01

The man machine automation tradeoff methodology presented is of four research tasks comprising the autonomous spacecraft system technology (ASST) project. ASST was established to identify and study system level design problems for autonomous spacecraft. Using the Space Station as an example spacecraft system requiring a certain level of autonomous control, a system level, man machine automation tradeoff methodology is presented that: (1) optimizes man machine mixes for different ground and on orbit crew functions subject to cost, safety, weight, power, and reliability constraints, and (2) plots the best incorporation plan for new, emerging technologies by weighing cost, relative availability, reliability, safety, importance to out year missions, and ease of retrofit. A fairly straightforward approach is taken by the methodology to valuing human productivity, it is still sensitive to the important subtleties associated with designing a well integrated, man machine system. These subtleties include considerations such as crew preference to retain certain spacecraft control functions; or valuing human integration/decision capabilities over equivalent hardware/software where appropriate.

Aerospace Engineering Space Mission Concept Feasibility Study: A Neptune Mission Design Example

NASA Technical Reports Server (NTRS)

Esper, Jaime

2007-01-01

This viewgraph document reviews the feasibility study of a mission to Neptune. Included are discussions of the science instruments, the design methodology, the trajectory, the spacecraft design, the alternative propulsion systems, (chemical, solar electric (SEP)), the communications systems, the power systems, the thermal system.

Common approach to solving SGEMP, DEMP, and ESD survivability

NASA Technical Reports Server (NTRS)

Ling, D.

1977-01-01

System Generated Electromagnetic Pulse (SGEMP) and Dispersed Electromagnetic Pulse DEMP) are nuclear generated spacecraft environments. Electrostatic discharge (ESD) is a natural spacecraft environment resulting from differential charging in magnetic substorms. All three phenomena, though differing in origin, result in the same problem to the spacecraft and that is Electromagnetic Interference (EMI). A common design approach utilizing a spacecraft structural Faraday Cage is presented which helps solve the EMI problem. Also, other system design techniques are discussed which minimize the magnitude of these environments through control of materials and electrical grounding configuration.

Genetic algorithm approaches for conceptual design of spacecraft systems including multi-objective optimization and design under uncertainty

NASA Astrophysics Data System (ADS)

Hassan, Rania A.

In the design of complex large-scale spacecraft systems that involve a large number of components and subsystems, many specialized state-of-the-art design tools are employed to optimize the performance of various subsystems. However, there is no structured system-level concept-architecting process. Currently, spacecraft design is heavily based on the heritage of the industry. Old spacecraft designs are modified to adapt to new mission requirements, and feasible solutions---rather than optimal ones---are often all that is achieved. During the conceptual phase of the design, the choices available to designers are predominantly discrete variables describing major subsystems' technology options and redundancy levels. The complexity of spacecraft configurations makes the number of the system design variables that need to be traded off in an optimization process prohibitive when manual techniques are used. Such a discrete problem is well suited for solution with a Genetic Algorithm, which is a global search technique that performs optimization-like tasks. This research presents a systems engineering framework that places design requirements at the core of the design activities and transforms the design paradigm for spacecraft systems to a top-down approach rather than the current bottom-up approach. To facilitate decision-making in the early phases of the design process, the population-based search nature of the Genetic Algorithm is exploited to provide computationally inexpensive---compared to the state-of-the-practice---tools for both multi-objective design optimization and design optimization under uncertainty. In terms of computational cost, those tools are nearly on the same order of magnitude as that of standard single-objective deterministic Genetic Algorithm. The use of a multi-objective design approach provides system designers with a clear tradeoff optimization surface that allows them to understand the effect of their decisions on all the design objectives under consideration simultaneously. Incorporating uncertainties avoids large safety margins and unnecessary high redundancy levels. The focus on low computational cost for the optimization tools stems from the objective that improving the design of complex systems should not be achieved at the expense of a costly design methodology.

On-orbit assembly of a team of flexible spacecraft using potential field based method

NASA Astrophysics Data System (ADS)

Chen, Ti; Wen, Hao; Hu, Haiyan; Jin, Dongping

2017-04-01

In this paper, a novel control strategy is developed based on artificial potential field for the on-orbit autonomous assembly of four flexible spacecraft without inter-member collision. Each flexible spacecraft is simplified as a hub-beam model with truncated beam modes in the floating frame of reference and the communication graph among the four spacecraft is assumed to be a ring topology. The four spacecraft are driven to a pre-assembly configuration first and then to the assembly configuration. In order to design the artificial potential field for the first step, each spacecraft is outlined by an ellipse and a virtual leader of circle is introduced. The potential field mainly depends on the attitude error between the flexible spacecraft and its neighbor, the radial Euclidian distance between the ellipse and the circle and the classical Euclidian distance between the centers of the ellipse and the circle. It can be demonstrated that there are no local minima for the potential function and the global minimum is zero. If the function is equal to zero, the solution is not a certain state, but a set. All the states in the set are corresponding to the desired configurations. The Lyapunov analysis guarantees that the four spacecraft asymptotically converge to the target configuration. Moreover, the other potential field is also included to avoid the inter-member collision. In the control design of the second step, only small modification is made for the controller in the first step. Finally, the successful application of the proposed control law to the assembly mission is verified by two case studies.

Self Assembling Mars Transfer Vehicles: The Preferred Concept of the Space Transfer Concepts and Analysis for Explorations Missions Study

NASA Technical Reports Server (NTRS)

Donahue, Benjamin

1994-01-01

Recently, one of the most comprehensive design studies of conceptual manned Mars vehicles, conducted since the Apollo era Mars mission studies of the 1960's, was completed. One of the tasks of the study involved the analysis of nuclear thermal propulsion spacecraft for Manned Mars exploration missions. This paper describes the specific effort aimed at vehicle configuration design. Over the course of the four year study, three configuration baselines were developed, each reflecting trade study cycle results of sequential phases of the study. Favorable attributes incorporated into the final concept, including a capability for on-orbit self-assembly and ease of launch vehicle packability, represent design solutions to configuration deficiencies plaguing nuclear propulsion Mars spacecraft design since the vehicle archetype originated in the 1950's. This paper contains a narrative summary of significant milestones in the effort, describes the evolution to the preferred configuration, and set forth the benefits derived from its utilization.

IDEAS: A multidisciplinary computer-aided conceptual design system for spacecraft

NASA Technical Reports Server (NTRS)

Ferebee, M. J., Jr.

1984-01-01

During the conceptual development of advanced aerospace vehicles, many compromises must be considered to balance economy and performance of the total system. Subsystem tradeoffs may need to be made in order to satisfy system-sensitive attributes. Due to the increasingly complex nature of aerospace systems, these trade studies have become more difficult and time-consuming to complete and involve interactions of ever-larger numbers of subsystems, components, and performance parameters. The current advances of computer-aided synthesis, modeling and analysis techniques have greatly helped in the evaluation of competing design concepts. Langley Research Center's Space Systems Division is currently engaged in trade studies for a variety of systems which include advanced ground-launched space transportation systems, space-based orbital transfer vehicles, large space antenna concepts and space stations. The need for engineering analysis tools to aid in the rapid synthesis and evaluation of spacecraft has led to the development of the Interactive Design and Evaluation of Advanced Spacecraft (IDEAS) computer-aided design system. The ADEAS system has been used to perform trade studies of competing technologies and requirements in order to pinpoint possible beneficial areas for research and development. IDEAS is presented as a multidisciplinary tool for the analysis of advanced space systems. Capabilities range from model generation and structural and thermal analysis to subsystem synthesis and performance analysis.

Space environmental interactions with spacecraft surfaces

NASA Technical Reports Server (NTRS)

Stevens, J. N.

1979-01-01

Environmental interactions are defined as the response of spacecraft surfaces to the charged-particle environment. These interactions are divided into two broad categories: spacecraft passive, in which the environment acts on the surfaces and spacecraft active, in which the spacecraft or a system on the spacecraft causes the interaction. The principal spacecraft passive interaction of concern is the spacecraft charging phenomenon. The spacecraft active category introduces the concept of interactions with the thermal plasma environment and Earth's magnetic fields, which are important at all altitudes and must be considered the designs of proposed large space structures and space power systems. The status of the spacecraft charging investigations is reviewed along with the spacecraft active interactions.

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

NASA Technical Reports Server (NTRS)

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.

Integrated Design System (IDS) Tools for the Spacecraft Aeroassist/Entry Vehicle Design Process

NASA Technical Reports Server (NTRS)

Olynick, David; Braun, Robert; Langhoff, Steven R. (Technical Monitor)

1997-01-01

The definition of the Integrated Design System technology focus area as presented in the NASA Information Technology center of excellence strategic plan is described. The need for IDS tools in the aeroassist/entry vehicle design process is illustrated. Initial and future plans for spacecraft IDS tool development are discussed.

Design, Development, Testing, and Evaluation: Human Factors Engineering

NASA Technical Reports Server (NTRS)

Adelstein, Bernard; Hobbs, Alan; OHara, John; Null, Cynthia

2006-01-01

While human-system interaction occurs in all phases of system development and operation, this chapter on Human Factors in the DDT&E for Reliable Spacecraft Systems is restricted to the elements that involve "direct contact" with spacecraft systems. Such interactions will encompass all phases of human activity during the design, fabrication, testing, operation, and maintenance phases of the spacecraft lifespan. This section will therefore consider practices that would accommodate and promote effective, safe, reliable, and robust human interaction with spacecraft systems. By restricting this chapter to what the team terms "direct contact" with the spacecraft, "remote" factors not directly involved in the development and operation of the vehicle, such as management and organizational issues, have been purposely excluded. However, the design of vehicle elements that enable and promote ground control activities such as monitoring, feedback, correction and reversal (override) of on-board human and automation process are considered as per NPR8705.2A, Section 3.3.

Thermoelectric Outer Planets Spacecraft (TOPS)

NASA Technical Reports Server (NTRS)

1973-01-01

The research and advanced development work is reported on a ballistic-mode, outer planet spacecraft using radioisotope thermoelectric generator (RTG) power. The Thermoelectric Outer Planet Spacecraft (TOPS) project was established to provide the advanced systems technology that would allow the realistic estimates of performance, cost, reliability, and scheduling that are required for an actual flight mission. A system design of the complete RTG-powered outer planet spacecraft was made; major technical innovations of certain hardware elements were designed, developed, and tested; and reliability and quality assurance concepts were developed for long-life requirements. At the conclusion of its active phase, the TOPS Project reached its principal objectives: a development and experience base was established for project definition, and for estimating cost, performance, and reliability; an understanding of system and subsystem capabilities for successful outer planets missions was achieved. The system design answered long-life requirements with massive redundancy, controlled by on-board analysis of spacecraft performance data.

Applications technology satellites advanced mission study

NASA Technical Reports Server (NTRS)

Gould, L. M.

1972-01-01

Three spacecraft configurations were designed for operation as a high powered synchronous communications satellite. Each spacecraft includes a 1 kw TWT and a 2 kw Klystron power amplifier feeding an antenna with multiple shaped beams. One of the spacecraft is designed to be boosted by a Thor-Delta launch vehicle and raised to synchronous orbit with electric propulsion. The other two are inserted into a elliptical transfer orbit with an Atlas Centaur and injected into final orbit with an apogee kick motor. Advanced technologies employed in the several configurations include tubes with multiple stage collectors radiating directly to space, multiple-contoured beam antennas, high voltage rollout solar cell arrays with integral power conditioning, electric propulsion for orbit raising and on-station attitude control and station-keeping, and liquid metal slip rings.

LEO Spacecraft Charging Guidelines

NASA Technical Reports Server (NTRS)

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

2002-01-01

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, has commissioned and funded a design guidelines document intended to capture the current state of understanding. We present here an overview of this document, which is now nearing completion.

Near Earth Asteroid Rendezvous (NEAR) Revised Eros Orbit Phase Trajectory Design

NASA Technical Reports Server (NTRS)

Helfrich, J; Miller, J. K.; Antreasian, P. G.; Carranza, E.; Williams, B. G.; Dunham, D. W.; Farquhar, R. W.; McAdams, J. V.

1999-01-01

Trajectory design of the orbit phase of the NEAR mission involves a new process that departs significantly from those procedures used in previous missions. In most cases, a precise spacecraft ephemeris is designed well in advance of arrival at the target body. For NEAR, the uncertainty in the dynamic environment around Eros does not allow the luxury of a precise spacecraft trajectory to be defined in advance. The principal cause of this uncertainty is the limited knowledge oi' the gravity field a,-id rotational state of Eros. As a result, the concept for the NEAR trajectory design is to define a number of rules for satisfying spacecraft, mission, and science constraints, and then apply these rules to various assumptions for the model of Eros. Nominal, high, and low Eros mass models are used for testing the trajectory design strategy and to bracket the ranges of parameter variations that are expected upon arrival at the asteroid. The final design is completed after arrival at Eros and determination of the actual gravity field and rotational state. As a result of the unplanned termination of the deep space rendezvous maneuver on December 20, 1998, the NEAR spacecraft passed within 3830 km of Eros on December 23, 1998. This flyby provided a brief glimpse of Eros, and allowed for a more accurate model of the rotational parameters and gravity field uncertainty. Furthermore, after the termination of the deep space rendezvous burn, contact with the spacecraft was lost and the NEAR spacecraft lost attitude control. During the subsequent gyrations of the spacecraft, hydrazine thruster firings were used to regain attitude control. This unplanned thruster activity used Much of the fuel margin allocated for the orbit phase. Consequently, minimizing fuel consumption is now even more important.

The community satellite 1. [social implications of ATS 6

NASA Technical Reports Server (NTRS)

1974-01-01

NASA's Applications Technology Satellite-6 is being used to test a variety of new space communications concepts requiring the use of a geosynchronous-orbit spacecraft. These include broadcast of health and education television programs to small, low-cost ground receiving units in remote regions. Other studies to be conducted are related to aeronautical and maritime communications, position-location, and traffic-control techniques. Questions concerning spacecraft tracking and data relay are also investigated. The 1,402 kg spacecraft consists essentially of an Earth Viewing Module connected to a deployable reflector antenna. Details regarding the planned experiments and the spacecraft design are discussed and a brief history of the ATS program is presented.

The systems impact of a concentrated solar array on a Jupiter orbiter

NASA Technical Reports Server (NTRS)

Rockey, D. E.; Bamford, R.; Hollars, M. G.; Klemetson, R. W.; Koerner, T. W.; Marsh, E. L.; Price, H.; Uphoff, C.

1981-01-01

Results of a study are presented suggesting that a Galileo Jupiter orbiting mission could be performed with a concentrated solar array power source. A baseline spacecraft design using concentrated arrays is given, and the overall spacecraft implications for attitude control, propulsion, power conditioning and the resultant spacecraft mass are examined. It is noted that while the concentrated array concept still requires extensive development effort, no insurmountable system level barriers preclude the use of a concentrated solar array on this difficult mission, with its stressing radiation environment, its lengthy periods of spacecraft shadowing as it passes behind Jupiter, and, finally, its large delta v burn required for orbital insertion.

Adaptive attitude control and momentum management for large-angle spacecraft maneuvers

NASA Technical Reports Server (NTRS)

Parlos, Alexander G.; Sunkel, John W.

1992-01-01

The fully coupled equations of motion are systematically linearized around an equilibrium point of a gravity gradient stabilized spacecraft, controlled by momentum exchange devices. These equations are then used for attitude control system design of an early Space Station Freedom flight configuration, demonstrating the errors caused by the improper approximation of the spacecraft dynamics. A full state feedback controller, incorporating gain-scheduled adaptation of the attitude gains, is developed for use during spacecraft on-orbit assembly or operations characterized by significant mass properties variations. The feasibility of the gain adaptation is demonstrated via a Space Station Freedom assembly sequence case study. The attitude controller stability robustness and transient performance during gain adaptation appear satisfactory.

Trajectory Design and Control for the Compton Gamma Ray Observatory Re-Entry

NASA Technical Reports Server (NTRS)

Hoge, Susan; Vaughn, Frank J., Jr.

2001-01-01

The Compton Gamma Ray Observatory (CGRO) controlled re-entry operation was successfully conducted in June of 2000. The surviving parts of the spacecraft landed in the Pacific Ocean within the nominal impact target zone. The design of the maneuvers to control the trajectory to accomplish this re-entry presented several challenges. These challenges included the timing and duration of the maneuvers, propellant management, post-maneuver state determination, collision avoidance with other spacecraft, accounting for the break-up of the spacecraft into several pieces with a wide range of ballistic coefficients, and ensuring that the impact footprint would remain within the desired impact target zone in the event of contingencies. This paper presents the initial re-entry trajectory design and traces the evolution of that design into the maneuver sequence used for the re-entry. The paper also discusses the spacecraft systems and operational constraints imposed on the trajectory design and the required modifications to the initial design based on those constraints. Data from the reentry operation are also presented.

Mars Polar Lander undergoes testing in SAEF-2

NASA Technical Reports Server (NTRS)

1998-01-01

In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), a KSC technician takes part in testing science instruments and basic spacecraft subsystems on the Mars Polar Lander. The solar- powered spacecraft, targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999, is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere.

KSC-98pc1371

NASA Image and Video Library

1998-10-22

KENNEDY SPACE CENTER, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), workers move the Mars Polar Lander to a work stand where it will undergo testing of the science instruments and basic spacecraft subsystems. The solar-powered spacecraft, targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999, is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

KSC-98pc1337

NASA Image and Video Library

1998-10-13

KENNEDY SPACE CENTER, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), technicians test the science instruments and the basic spacecraft subsystems on the Mars Polar Lander. The solar-powered spacecraft is targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999. It is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

KSC-98pc1338

NASA Image and Video Library

1998-10-13

KENNEDY SPACE CENTE, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), a technician tests the science instruments and the basic spacecraft subsystems on the Mars Polar Lander. The solar-powered spacecraft is targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999. It is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

KSC-98pc1607

NASA Image and Video Library

1998-10-29

KENNEDY SPACE CENTER, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), a KSC technician takes part in testing science instruments and basic spacecraft subsystems on the Mars Polar Lander. The solar-powered spacecraft, targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999, is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

Choosing an efficient option of the combined propulsion system and flight profile of the INTERHELIO-PROBE spacecraft

NASA Astrophysics Data System (ADS)

Platov, I. V.; Simonov, A. V.; Konstantinov, M. S.

2016-12-01

The paper is devoted to the design features of the prospective Russian INTERHELIO-PROBE spacecraft using, depending on the configuration version, an electric or chemical propulsion system as a sustainer. The scientific goal of the mission is the study of near-solar space from close distances (60-70 solar radii). The paper presents the description of several versions of the spacecraft options depending on the installed propulsion system, as well as the main characteristics of the flight profile depending on the engine type.

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

NASA Technical Reports Server (NTRS)

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.

Attitude control/momentum management and payload pointing in advanced space vehicles

NASA Technical Reports Server (NTRS)

Parlos, Alexander G.; Jayasuriya, Suhada

1990-01-01

The design and evaluation of an attitude control/momentum management system for highly asymmetric spacecraft configurations are presented. The preliminary development and application of a nonlinear control system design methodology for tracking control of uncertain systems, such as spacecraft payload pointing systems are also presented. Control issues relevant to both linear and nonlinear rigid-body spacecraft dynamics are addressed, whereas any structural flexibilities are not taken into consideration. Results from the first task indicate that certain commonly used simplifications in the equations of motions result in unstable attitude control systems, when used for highly asymmetric spacecraft configurations. An approach is suggested circumventing this problem. Additionally, even though preliminary results from the second task are encouraging, the proposed nonlinear control system design method requires further investigation prior to its application and use as an effective payload pointing system design technique.

MODEL CORRELATION STUDY OF A RETRACTABLE BOOM FOR A SOLAR SAIL SPACECRAFT

NASA Technical Reports Server (NTRS)

Adetona, O.; Keel, L. H.; Oakley, J. D.; Kappus, K.; Whorton, M. S.; Kim, Y. K.; Rakpczy, J. M.

2005-01-01

To realize design concepts, predict dynamic behavior and develop appropriate control strategies for high performance operation of a solar-sail spacecraft, we developed a simple analytical model that represents dynamic behavior of spacecraft with various sizes. Since motion of the vehicle is dominated by retractable booms that support the structure, our study concentrates on developing and validating a dynamic model of a long retractable boom. Extensive tests with various configurations were conducted for the 30 Meter, light-weight, retractable, lattice boom at NASA MSFC that is structurally and dynamically similar to those of a solar-sail spacecraft currently under construction. Experimental data were then compared with the corresponding response of the analytical model. Though mixed results were obtained, the analytical model emulates several key characteristics of the boom. The paper concludes with a detailed discussion of issues observed during the study.

Illumination from space with orbiting solar-reflector spacecraft

NASA Technical Reports Server (NTRS)

Canady, J. E., Jr.; Allen, J. L., Jr.

1982-01-01

The feasibility of using orbiting mirrors to reflect sunlight to Earth for several illumination applications is studied. A constellation of sixteen 1 km solar reflector spacecraft in geosynchronous orbit can illuminate a region 333 km in diameter to 8 lux, which is brighter than most existing expressway lighting systems. This constellation can serve one region all night long or can provide illumination during mornings and evenings to five regions across the United States. Preliminary cost estimates indicate such an endeavor is economically feasible. The studies also explain how two solar reflectors can illuminate the in-orbit nighttime operations of Space Shuttle. An unfurlable, 1 km diameter solar reflector spacecraft design concept was derived. This spacecraft can be packaged in the Space, Shuttle, transported to low Earth orbit, unfurled, and solar sailed to operational orbits up to geosynchronous. The necessary technical studies and improvements in technology are described, and potential environmental concerns are discussed.

Composite structures for magnetosphere imager spacecraft

NASA Technical Reports Server (NTRS)

Chu, Tsuchin

1994-01-01

Results of a trade study addressing the issues and benefits in using carbon fiber reinforced composites for the Magnetosphere Imager (MI) spacecraft are presented. The MI mission is now part of the Sun/Earth Connection Program. To qualify for this category, new technology and innovative methods to reduce the cost and size have to be considered. Topics addressed cover: (1) what is a composite, including advantages and disadvantages of composites and carbon/graphite fibers; and (2) structural design for MI, including composite design configuration, material selection, and analysis of composite structures.

Attitude Ground System (AGS) For The Magnetospheric Multi-Scale (MMS) Mission

NASA Technical Reports Server (NTRS)

Raymond, Juan C.; Sedlak, Joseph E.; Vint, Babak

2015-01-01

The Magnetospheric Multiscale (MMS) mission is a Solar-Terrestrial Probe mission consisting of four identically instrumented spin-stabilized spacecraft flying in an adjustable pyramid-like formation around the Earth. The formation of the MMS spacecraft allows for three-dimensional study of the phenomenon of magnetic reconnection, which is the primary objective of the mission. The MMS spacecraft were launched early on March 13, 2015 GMT. Due to the challenging and very constricted attitude and orbit requirements for performing the science, as well as the need to maintain the spacecraft formation, multiple ground functionalities were designed to support the mission. These functionalities were incorporated into a ground system known as the Attitude Ground System (AGS). Various AGS configurations have been used widely to support a variety of three-axis-stabilized and spin-stabilized spacecraft missions within the NASA Goddard Space Flight Center (GSFC). The original MMS operational concept required the AGS to perform highly accurate predictions of the effects of environmental disturbances on the spacecraft orientation and to plan the attitude maneuvers necessary to stay within the science attitude tolerance. The orbit adjustment requirements for formation control drove the need also to perform calibrations that have never been done before in support of NASA GSFC missions. The MMS mission required support analysts to provide fast and accurately calibrated values of the inertia tensor, center of mass, and accelerometer bias for each MMS spacecraft. During early design of the AGS functionalities, a Kalman filter for estimating the attitude, body rates, center of mass, and accelerometer bias, using only star tracker and accelerometer measurements, was heavily analyzed. A set of six distinct filters was evaluated and considered for estimating the spacecraft attitude and body rates using star tracker data only. Four of the six filters are closely related and were compared during support of the Time History of Events and Macroscale Interactions during Substorms (THEMIS) and Space Technology-5 (ST-5) missions. These analyses exposed high dependency and sensitivity on the knowledge of the spacecraft inertia tensor for both body rates and accelerometer bias estimation. The conclusion of the analysis led to the design of an inertia tensor calibration technique using only star tracker data. The second most important result of the analysis was the design of two separate Kalman filters to estimate the spacecraft attitude and body rates and the accelerometer bias instead of a single combined filter. In this paper, the calibration results of the mass properties, as well as the performance of the spacecraft attitude and body rates filters using flight data are presented and compared against the mission requirements.

Catalog of lunar mission data

NASA Technical Reports Server (NTRS)

Mantel, E. J. (Editor); Miller, E. R. (Editor)

1977-01-01

Several series of spacecraft were developed, designed, built and launched to determine different characteristics of the lunar surface and environment for a manned landing. Both unmanned and manned spacecrafts, spacecraft equipment and lunar missions are documented.

Preliminary design of the thermal protection system for solar probe

NASA Technical Reports Server (NTRS)

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.

Comparison of Analysis with Test for Static Loading of Two Hypersonic Inflatable Aerodynamic Decelerator Concepts

NASA Technical Reports Server (NTRS)

Lyle, Karen H.

2015-01-01

Acceptance of new spacecraft structural architectures and concepts requires validated design methods to minimize the expense involved with technology demonstration via flight-testing. Hypersonic Inflatable Aerodynamic Decelerator (HIAD) architectures are attractive for spacecraft deceleration because they are lightweight, store compactly, and utilize the atmosphere to decelerate a spacecraft during entry. However, designers are hesitant to include these inflatable approaches for large payloads or spacecraft because of the lack of flight validation. This publication summarizes results comparing analytical results with test data for two concepts subjected to representative entry, static loading. The level of agreement and ability to predict the load distribution is considered sufficient to enable analytical predictions to be used in the design process.

Spacecraft Design Thermal Control Subsystem

NASA Technical Reports Server (NTRS)

Miyake, Robert N.

2003-01-01

This slide presentation reviews the functions of the thermal control subsystem engineers in the design of spacecraft. The goal of the thermal control subsystem that will be used in a spacecraft is to maintain the temperature of all spacecraft components, subsystems, and all the flight systems within specified limits for all flight modes from launch to the end of the mission. For most thermal control subsystems the mass, power and control and sensing systems must be kept below 10% of the total flight system resources. This means that the thermal control engineer is involved in all other flight systems designs. The two concepts of thermal control, passive and active are reviewed and the use of thermal modeling tools are explained. The testing of the thermal control is also reviewed.

The Recovery of TOMS-EP

NASA Technical Reports Server (NTRS)

Robertson, Brent; Sabelhaus, Phil; Mendenhall, Todd; Fesq, Lorraine

1998-01-01

On December 13th 1998, the Total Ozone Mapping Spectrometer - Earth Probe (TOMS-EP) spacecraft experienced a Single Event Upset which caused the system to reconfigure and enter a Safe Mode. This incident occurred two and a half years after the launch of the spacecraft which was designed for a two year life. A combination of factors, including changes in component behavior due to age and extended use, very unfortunate initial conditions and the safe mode processing logic prevented the spacecraft from entering its nominal long term storage mode. The spacecraft remained in a high fuel consumption mode designed for temporary use. By the time the onboard fuel was exhausted, the spacecraft was Sun pointing in a high rate flat spin. Although the uncontrolled spacecraft was initially in a power and thermal safe orientation, it would not stay in this state indefinitely due to a slow precession of its momentum vector. A recovery team was immediately assembled to determine if there was time to develop a method of despinning the vehicle and return it to normal science data collection. A three stage plan was developed that used the onboard magnetic torque rods as actuators. The first stage was designed to reduce the high spin rate to within the linear range of the gyros. The second stage transitioned the spacecraft from sun pointing to orbit reference pointing. The final stage returned the spacecraft to normal science operation. The entire recovery scenario was simulated with a wide range of initial conditions to establish the expected behavior. The recovery sequence was started on December 28th 1998 and completed by December 31st. TOMS-EP was successfully returned to science operations by the beginning of 1999. This paper describes the TOMS-EP Safe Mode design and the factors which led to the spacecraft anomaly and loss of fuel. The recovery and simulation efforts are described. Flight data are presented which show the performance of the spacecraft during its return to science. Finally, lessons learned are presented.

Design and Flight Performance of NOAA-K Spacecraft Batteries

NASA Technical Reports Server (NTRS)

Rao, Gopalakrishna M.; Chetty, P. R. K.; Spitzer, Tom; Chilelli, P.

1999-01-01

The US National Oceanic and Atmospheric Administration (NOAA) operates the Polar Operational Environmental Satellite (POES) spacecraft (among others) to support weather forecasting, severe storm tracking, and meteorological research by the National Weather Service (NWS). The latest in the POES series of spacecraft, named as NOAA-KLMNN, is in orbit and four more are in various phases of development. The NOAA-K spacecraft was launched on May 13, 1998. Each of these spacecraft carry three Nickel-Cadmium batteries designed and manufactured by Lockheed Martin. The battery, which consists of seventeen 40 Ah cells manufactured by SAFT, provides the spacecraft power during the ascent phase, orbital eclipse and when the power demand is in excess of the solar array capability. The NOAA-K satellite is in a 98 degree inclination, 7:30AM ascending node orbit. In this orbit the satellite experiences earth occultation only 25% of the year. This paper provides a brief overview of the power subsystem, followed by the battery design and qualification, the cell life cycle test data, and the performance during launch and in orbit.

Design and Flight Performance of NOAA-K Spacecraft Batteries

NASA Technical Reports Server (NTRS)

Rao, Gopalakrishna M.; Chetty, P. R. K.; Spitzer, Tom; Chilelli, P.

1998-01-01

The US National Oceanic and Atmospheric Administration (NOAA) operates the Polar Operational Environmental Satellite (POES) spacecraft (among others) to support weather forecasting, severe storm tracking, and meteorological research by the National Weather Service (NWS). The latest in the POES series of spacecraft, named as NOAA-KLMNN', one is in orbit and four more are in various phases of development. The NOAA-K spacecraft was launched on May 13, 1998. Each of these spacecraft carry three Nickel-Cadmium batteries designed and manufactured by Lockheed Martin. The battery, which consists of seventeen 40 Ah cells manufactured by SAFT, provides the spacecraft power during the ascent phase, orbital eclipse and when the power demand is in excess of the solar array capability. The NOAA-K satellite is in a 98 degree inclination, 7:30AM ascending node orbit. In this orbit the satellite experiences earth occultation only 25% of the year. This paper provides a brief overview of the power subsystem, followed by the battery design and qualification, the cell life cycle test data, and the performance during launch and in orbit.

LDEF materials results for spacecraft applications: Executive summary

NASA Astrophysics Data System (ADS)

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.

LDEF materials results for spacecraft applications: Executive summary

NASA Technical Reports Server (NTRS)

Whitaker, A. F. (Compiler); Dooling, D. (Compiler)

1995-01-01

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.

Dynamic interaction of rotating momentum wheels with spacecraft elements

NASA Astrophysics Data System (ADS)

Shankar Narayan, S.; Nair, P. S.; Ghosal, Ashitava

2008-09-01

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

Spacecraft applications of advanced global positioning system technology

NASA Technical Reports Server (NTRS)

1988-01-01

This is the final report on the Texas Instruments Incorporated (TI) simulations study of Spacecraft Application of Advanced Global Positioning System (GPS) Technology. This work was conducted for the NASA Johnson Space Center (JSC) under contract NAS9-17781. GPS, in addition to its baselined capability as a highly accurate spacecraft navigation system, can provide traffic control, attitude control, structural control, and uniform time base. In Phase 1 of this program, another contractor investigated the potential of GPS in these four areas and compared GPS to other techniques. This contract was for the Phase 2 effort, to study the performance of GPS for these spacecraft applications through computer simulations. TI had previously developed simulation programs for GPS differential navigation and attitude measurement. These programs were adapted for these specific spacecraft applications. In addition, TI has extensive expertise in the design and production of advanced GPS receivers, including space-qualified GPS receivers. We have drawn on this background to augment the simulation results in the system level overview, which is Section 2 of this report.

Technicians prepare the AIM spacecraft for fairing installation

NASA Image and Video Library

2007-04-12

At Vandenberg Air Force Base in California, technicians prepare the AIM spacecraft for fairing installation. The fairing is a molded structure that fits around the spacecraft and forms an aerodynamically smooth nose cone, protecting the spacecraft during launch. Launch will be from a Pegasus XL rocket, carried and released by Orbital Sciences L-1011 jet aircraft. AIM, which stands for Aeronomy of Ice in the Mesosphere, is being prepared for integrated testing and a flight simulation. The AIM spacecraft will fly three instruments designed to study polar mesospheric clouds located at the edge of space, 50 miles above the Earth's surface in the coldest part of the planet's atmosphere. The mission's primary goal is to explain why these clouds form and what has caused them to become brighter and more numerous and appear at lower latitudes in recent years. AIM's results will provide the basis for the study of long-term variability in the mesospheric climate and its relationship to global climate change. Launch is scheduled for April 25.

CHIPSat spacecraft design: significant science on a low budget

NASA Astrophysics Data System (ADS)

Janicik, Jeffrey; Wolff, Jonathan

2003-12-01

The Cosmic Hot Interstellar Plasma Spectrometer satellite (CHIPSat) was launched on January 12, 2003 and is successfully accomplishing its mission. CHIPS is NASA"s first-ever University-Class Explorer (UNEX) project, and is performed through a grant to the University of California at Berkeley (UCB) Space Sciences Laboratory (SSL). As a small start-up aerospace company, SpaceDev was awarded responsibility for a low-cost spacecraft and mission design, build, integration and test, and mission operations. The company leveraged past small satellite mission experiences to help design a robust small spacecraft system architecture. In addition, they utilized common industry hardware and software standards to facilitate design implementation, integration, and test of the bus, including the use of TCP/IP protocols and the Internet for end-to-end satellite communications. The approach called for a single-string design except in critical areas, the use of COTS parts to incorporate the latest proven technologies in commercial electronics, and the establishment of a working system as quickly as possible in order to maximize test hours prior to launch. Furthermore, automated ground systems were combined with table-configured onboard software to allow for "hands-off" mission operations. During nominal operations, the CHIPSat spacecraft uses a 3-axis stabilized zero-momentum bias "Nominal" mode. The secondary mode is a "Safehold" mode where fixed "keep-alive" arrays maintain enough power to operate the essential spacecraft bus in any attitude and spin condition, and no a-priori attitude knowledge is required to recover. Due to the omnidirectional antenna design, communications are robust in "Safehold" mode, including the transmission of basic housekeeping data at a duty cycle that is adjusted based on available solar power. This design enables the entire mission to be spent in "Observation Mode" with timed pointing files mapping the sky as desired unless an anomalous event upsets the health of the bus such that the spacecraft system toggles back to "Safehold". In all conditions, spacecraft operations do not require any time-critical operator involvement. This paper will examine the results of the first six months of CHIPSat on-orbit operations and measure them against the expectations of the aforementioned design architecture. The end result will be a "lessons learned" account of a 3-axis sun-pointing small spacecraft design architecture that will be useful for future science missions.

Earth Observatory Satellite system definition study. Report 5: System design and specifications. Volume 4: Mission peculiar spacecraft segment and module specifications

NASA Technical Reports Server (NTRS)

1974-01-01

The specifications for the Earth Observatory Satellite (EOS) peculiar spacecraft segment and associated subsystems and modules are presented. The specifications considered include the following: (1) wideband communications subsystem module, (2) mission peculiar software, (3) hydrazine propulsion subsystem module, (4) solar array assembly, and (5) the scanning spectral radiometer.

Fundamentals of the orbit and response for TianQin

NASA Astrophysics Data System (ADS)

Hu, Xin-Chun; Li, Xiao-Hong; Wang, Yan; Feng, Wen-Fan; Zhou, Ming-Yue; Hu, Yi-Ming; Hu, Shou-Cun; Mei, Jian-Wei; Shao, Cheng-Gang

2018-05-01

TianQin is a space-based laser interferometric gravitational wave detector aimed at detecting gravitational waves at low frequencies (0.1 mHz–1 Hz). It is formed by three identical drag-free spacecrafts in an equilateral triangular constellation orbiting around the Earth. The distance between each pair of spacecrafts is approximately 1.7 × 105 ~km . The spacecrafts are interconnected by infrared laser beams forming up to three Michelson-type interferometers. The detailed mission design and the study of science objectives for the TianQin project depend crucially on the orbit and the response of the detector. In this paper, we provide the analytic expressions for the coordinates of the orbit for each spacecraft in the heliocentric-ecliptic coordinate system to the leading orders. This enables a sufficiently accurate study of science objectives and data analysis, and serves as a first step to further orbit design and optimization. We calculate the response of a single Michelson detector to plane gravitational waves in arbitrary waveform which is valid in the full range of the sensitive frequencies. It is then used to generate the more realistic sensitivity curve of TianQin. We apply this model on a reference white-dwarf binary as a proof of principle.

Long life assurance study for manned spacecraft long life hardware. Volume 2: Long life assurance studies of EEE parts and packaging

NASA Technical Reports Server (NTRS)

1972-01-01

Guidelines for the design, development, and fabrication of electronic components and circuits for use in spacecraft construction are presented. The subjects discussed involve quality control procedures and test methodology for the following subjects: (1) monolithic integrated circuits, (2) hybrid integrated circuits, (3) transistors, (4) diodes, (5) tantalum capacitors, (6) electromechanical relays, (7) switches and circuit breakers, and (8) electronic packaging.

The Galileo Attitude and Articulation Control System - A radiation-hard, high precision, state-of-the-art control system

NASA Technical Reports Server (NTRS)

Rhoads Stephenson, R.

1986-01-01

The Galileo Mission and Spacecraft design impose tight requirements on the Attitude and Articulation Control System (AACS). These requirements, coupled with the flexible spacecraft, the need for autonomy, and a severe radiation environment, pose a great challenge for the AACS designer. The resulting design and implementation are described, along with the discovery and solution of the Single-Event Upset problem. The status of the testing of the AACS in the Integration and Test Laboratory as well as at the spacecraft level is summarized.

Effects of arcing due to spacecraft charging on spacecraft survival

NASA Technical Reports Server (NTRS)

Rosen, A.; Sanders, N. L.; Ellen, J. M., Jr.; Inouye, G. T.

1978-01-01

A quantitative assessment of the hazard associated with spacecraft charging and arcing on spacecraft systems is presented. A literature survey on arc discharge thresholds and characteristics was done and gaps in the data and requirements for additional experiments were identified. Calculations of coupling of arc discharges into typical spacecraft systems were made and the susceptibility of typical spacecraft to disruption by arc discharges was investigated. Design guidelines and recommended practices to reduce or eliminate the threat of malfunction and failures due to spacecraft charging/arcing were summarized.

Assessment and control of spacecraft electromagnetic interference

NASA Technical Reports Server (NTRS)

1972-01-01

Design criteria are presented to provide guidance in assessing electromagnetic interference from onboard sources and establishing requisite control in spacecraft design, development, and testing. A comprehensive state-of-the-art review is given which covers flight experience, sources and transmission of electromagnetic interference, susceptible equipment, design procedure, control techniques, and test methods.

The STEIN Particle Detector

DTIC Science & Technology

2015-02-27

cubesats. The CINEMA (Cubesat for Ions Neutrals Electrons and Magnetic 1 Approved for public release; distribution is unlimited. fields) was the primary...intended host for STEIN. Additionally some calibration efforts were performed with the CINEMA spacecraft as an element of the readout. This resulted...designed to accept a clock from its host spacecraft (as was the design case for CINEMA ) of 8.38MHz (specifically 2^23 Hz). As well as a spacecraft

Design and research of sun sensor based on technology of optical fiber

NASA Astrophysics Data System (ADS)

Li, Ye; Zhou, Wang; Li, Dan

2010-08-01

A kind of sun sensor is designed based on the optical fiber. This project consists of three parts: optical head, photoelectric sensor and signal processing unit. The innovation of this design lies in the improvement of traditional sun sensor, where multi-fibers, used as a leader, are symmetrically distributed on the surface of a spacecraft. To determine the attitude of a spacecraft, the sun sensor should measure the direction of the sun. Because the fiber length can be adjusted according to the fact, photoelectric sensor can be placed deeply inside a spacecraft to protect the photoelectric sensor against the damage by the high-energy particles from outer space. The processing unit calculates the difference value of sun energy imported by each pair of opposite optical fiber so as to obtain the angle and the orientation between the spacecraft and the sun. This sun sensor can suit multi-field of view, both small and large. It improves the accuracy of small field of view and increases the precision of locating a spacecraft. This paper briefly introduces the design of processing unit. This sun sensor is applicable to detect the attitude of a spacecraft. In addition, it can also be used in solar tracking system of PV technology.

Trajectory options for the DART mission

NASA Astrophysics Data System (ADS)

Atchison, Justin A.; Ozimek, Martin T.; Kantsiper, Brian L.; Cheng, Andrew F.

2016-06-01

This study presents interplanetary trajectory options for the Double Asteroid Redirection Test (DART) spacecraft to reach the near Earth object, Didymos binary system, during its 2022 Earth conjunction. DART represents a component of a joint NASA-ESA mission to study near Earth object kinetic impact deflection. The DART trajectory must satisfy mission objectives for arrival timing, geometry, and lighting while minimizing launch vehicle and spacecraft propellant requirements. Chemical propulsion trajectories are feasible from two candidate launch windows in late 2020 and 2021. The 2020 trajectories are highly perturbed by Earth's orbit, requiring post-launch deep space maneuvers to retarget the Didymos system. Within these windows, opportunities exist for flybys of additional near Earth objects: Orpheus in 2021 or 2007 YJ in 2022. A second impact attempt, in the event that the first impact is unsuccessful, can be added at the expense of a shorter launch window and increased (∼3x) spacecraft ΔV . However, the second impact arrival geometry has poor lighting, high Earth ranges, and would require additional degrees of freedom for solar panel and/or antenna gimbals. A low-thrust spacecraft configuration increases the trajectory flexibility. A solar electric propulsion spacecraft could be affordably launched as a secondary spacecraft in an Earth orbit and spiral out to target the requisite interplanetary departure condition. A sample solar electric trajectory was constructed from an Earth geostationary transfer using a representative 1.5 kW thruster. The trajectory requires 9 months to depart Earth's sphere of influence, after which its interplanetary trajectory includes a flyby of Orpheus and a second Didymos impact attempt. The solar electric spacecraft implementation would impose additional bus design constraints, including large solar arrays that could pose challenges for terminal guidance. On the basis of this study, there are many feasible options for DART to meet its mission design objectives and enable this unique kinetic impact experiment.

Propellant-Less Spacecraft Formation-Flying and Maneuvering with Photonic Laser Thrusters

NASA Technical Reports Server (NTRS)

Bae, Young K.

2015-01-01

The present NIAC Phase II program explored an amplified photon thruster, Photonic Laser Thruster (PLT), as a means of enabling unprecedented maneuverability of small spacecraft, such as cubesats, and reducing space system SWaP for future NASA missions and other commercial and DoD space endeavors. In addition to its propellantless operation capability, PLT can provide orders of magnitude more precise controls in thrust magnitude and vector than conventional thrusters. Furthermore, PLT promises to enable innovative CONOPS (Concept of Operations) to change how some NASA missions are conceived and to represent a revolutionary departure from the "all-in-one" single-spacecraft approach, where a primary factor that dominates spacecraft design is a heavy and risk-intolerant mission-critical payload. Instead, the PLT CONOPS has evolved from a different path based on interbody dynamics via thrust and power beaming. As interbody atomic dynamics unfolds completely new classes of molecular structures that cannot be formed by solo acting atoms alone, the PLT interbody dynamics is predicted to unfold unprecedented multibody spacecraft structures. Therefore, the revolutionary path of the PLT CONOPS represents a technology push rather than a mission pull, and will enable an entirely new generation of planetary, heliospheric, and Earth-centric missions. The chief accomplishments of the present Phase II program are: 1) achievement of photon thrust up to 3.5 mN (100 times scaling up of Phase I PLT) and amplification factor up to 1,500 (15 times enhancement of Phase I PLT), 2) laboratory demonstration of propelling, slowing and stopping a 1U cubesat on an air track with PLT, 3) proof of feasibility on persistent out-of-plane formation flying with PLT in simulation studies, 4) preliminary SolidWorks designs of 1-mN class PLT, 5) establishment of SWaP for flight-ready PLT, 6) designs for proof-ofconcept missions of precision formation flying with cubesats, 7) definition of PLT-based NASA missions, such as Virtual Telescope. In sum, the present study conclusively demonstrated the potential of PLT to revolutionize future space endeavors by drastically enhancing maneuverability of spacecraft, reducing future space system SWaP by exploiting small spacecraft multi-system, and enabling innovative CONOPS.

Large-Scale Spacecraft Fire Safety Tests

NASA Technical Reports Server (NTRS)

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;

Galileo spacecraft power distribution and autonomous fault recovery

NASA Technical Reports Server (NTRS)

Detwiler, R. C.

1982-01-01

There is a trend in current spacecraft design to achieve greater fault tolerance through the implemenation of on-board software dedicated to detecting and isolating failures. A combination of hardware and software is utilized in the Galileo power system for autonomous fault recovery. Galileo is a dual-spun spacecraft designed to carry a number of scientific instruments into a series of orbits around the planet Jupiter. In addition to its self-contained scientific payload, it will also carry a probe system which will be separated from the spacecraft some 150 days prior to Jupiter encounter. The Galileo spacecraft is scheduled to be launched in 1985. Attention is given to the power system, the fault protection requirements, and the power fault recovery implementation.

Mars orbiter conceptual systems design study

NASA Technical Reports Server (NTRS)

Dixon, W.; Vogl, J.

1982-01-01

Spacecraft system and subsystem designs at the conceptual level to perform either of two Mars Orbiter missions, a Climatology Mission and an Aeronomy Mission were developed. The objectives of these missions are to obtain and return data.

Thermal design of the IUE hydrazine auxiliary propulsion system. [International Ultraviolet Explorer

NASA Technical Reports Server (NTRS)

Skladany, J. T.; Kelly, W. H.

1977-01-01

The International Ultraviolet Explorer is a large astronomical observatory scheduled to be placed in a three-axis stabilized synchronous orbit in the fourth quarter of 1977. The Hydrazine Auxiliary Propulsion System (HAPS) must perform a number of spacecraft maneuvers to achieve a successful mission. This paper describes the thermal design which accomplishes temperature control between 5 and 65 C for all orbital conditions by utilizing multilayer insulation and commandable component heaters. A primary design criteria was the minimization of spacecraft power by the selective use of the solar environment. The thermal design was carefully assessed and verified in both spacecraft thermal balance and subsystem solar simulation testing.

Design considerations for a space-borne ocean surface laser altimeter

NASA Technical Reports Server (NTRS)

Plotkin, H. H.

1972-01-01

Design procedures for using laser ranging systems in spacecraft to reflect ocean surface pulses vertically and measure spacecraft altitude with high precision are examined. Operating principles and performance experience of a prototype system are given.

Systems cost/performance analysis (study 2.3). Volume 2: Systems cost/performance model. [unmanned automated payload programs and program planning

NASA Technical Reports Server (NTRS)

Campbell, B. H.

1974-01-01

A methodology which was developed for balanced designing of spacecraft subsystems and interrelates cost, performance, safety, and schedule considerations was refined. The methodology consists of a two-step process: the first step is one of selecting all hardware designs which satisfy the given performance and safety requirements, the second step is one of estimating the cost and schedule required to design, build, and operate each spacecraft design. Using this methodology to develop a systems cost/performance model allows the user of such a model to establish specific designs and the related costs and schedule. The user is able to determine the sensitivity of design, costs, and schedules to changes in requirements. The resulting systems cost performance model is described and implemented as a digital computer program.

Enhanced Attitude Control Experiment for SSTI Lewis Spacecraft

NASA Technical Reports Server (NTRS)

Maghami, Peoman G.

1997-01-01

The enhanced attitude control system experiment is a technology demonstration experiment on the NASA's small spacecraft technology initiative program's Lewis spacecraft to evaluate advanced attitude control strategies. The purpose of the enhanced attitude control system experiment is to evaluate the feasibility of designing and implementing robust multi-input/multi-output attitude control strategies for enhanced pointing performance of spacecraft to improve the quality of the measurements of the science instruments. Different control design strategies based on modern and robust control theories are being considered for the enhanced attitude control system experiment. This paper describes the experiment as well as the design and synthesis of a mixed H(sub 2)/H(sub infinity) controller for attitude control. The control synthesis uses a nonlinear programming technique to tune the controller parameters and impose robustness and performance constraints. Simulations are carried out to demonstrate the feasibility of the proposed attitude control design strategy. Introduction

Boeing's High Voltage Solar Tile Test Results

NASA Astrophysics Data System (ADS)

Reed, Brian J.; Harden, David E.; Ferguson, Dale C.; Snyder, David B.

2002-10-01

Real concerns of spacecraft charging and experience with solar array augmented electrostatic discharge arcs on spacecraft have minimized the use of high voltages on large solar arrays despite numerous vehicle system mass and efficiency advantages. Boeing's solar tile (patent pending) allows high voltage to be generated at the array without the mass and efficiency losses of electronic conversion. Direct drive electric propulsion and higher power payloads (lower spacecraft weight) will benefit from this design. As future power demand grows, spacecraft designers must use higher voltage to minimize transmission loss and power cable mass for very large area arrays. This paper will describe the design and discuss the successful test of Boeing's 500-Volt Solar Tile in NASA Glenn's Tenney chamber in the Space Plasma Interaction Facility. The work was sponsored by NASA's Space Solar Power Exploratory Research and Technology (SERT) Program and will result in updated high voltage solar array design guidelines being published.

Boeing's High Voltage Solar Tile Test Results

NASA Technical Reports Server (NTRS)

Reed, Brian J.; Harden, David E.; Ferguson, Dale C.; Snyder, David B.

2002-01-01

Real concerns of spacecraft charging and experience with solar array augmented electrostatic discharge arcs on spacecraft have minimized the use of high voltages on large solar arrays despite numerous vehicle system mass and efficiency advantages. Boeing's solar tile (patent pending) allows high voltage to be generated at the array without the mass and efficiency losses of electronic conversion. Direct drive electric propulsion and higher power payloads (lower spacecraft weight) will benefit from this design. As future power demand grows, spacecraft designers must use higher voltage to minimize transmission loss and power cable mass for very large area arrays. This paper will describe the design and discuss the successful test of Boeing's 500-Volt Solar Tile in NASA Glenn's Tenney chamber in the Space Plasma Interaction Facility. The work was sponsored by NASA's Space Solar Power Exploratory Research and Technology (SERT) Program and will result in updated high voltage solar array design guidelines being published.

Integrated controls-structures design methodology development for a class of flexible spacecraft

NASA Technical Reports Server (NTRS)

Maghami, P. G.; Joshi, S. M.; Walz, J. E.; Armstrong, E. S.

1990-01-01

Future utilization of space will require large space structures in low-Earth and geostationary orbits. Example missions include: Earth observation systems, personal communication systems, space science missions, space processing facilities, etc., requiring large antennas, platforms, and solar arrays. The dimensions of such structures will range from a few meters to possibly hundreds of meters. For reducing the cost of construction, launching, and operating (e.g., energy required for reboosting and control), it will be necessary to make the structure as light as possible. However, reducing structural mass tends to increase the flexibility which would make it more difficult to control with the specified precision in attitude and shape. Therefore, there is a need to develop a methodology for designing space structures which are optimal with respect to both structural design and control design. In the current spacecraft design practice, it is customary to first perform the structural design and then the controller design. However, the structural design and the control design problems are substantially coupled and must be considered concurrently in order to obtain a truly optimal spacecraft design. For example, let C denote the set of the 'control' design variables (e.g., controller gains), and L the set of the 'structural' design variables (e.g., member sizes). If a structural member thickness is changed, the dynamics would change which would then change the control law and the actuator mass. That would, in turn, change the structural model. Thus, the sets C and L depend on each other. Future space structures can be roughly divided into four mission classes. Class 1 missions include flexible spacecraft with no articulated appendages which require fine attitude pointing and vibration suppression (e.g., large space antennas). Class 2 missions consist of flexible spacecraft with articulated multiple payloads, where the requirement is to fine-point the spacecraft and each individual payload while suppressing the elastic motion. Class 3 missions include rapid slewing of spacecraft without appendages, while Class 4 missions include general nonlinear motion of a flexible spacecraft with articulated appendages and robot arms. Class 1 and 2 missions represent linear mathematical modeling and control system design problems (except for actuator and sensor nonlinearities), while Class 3 and 4 missions represent nonlinear problems. The development of an integrated controls/structures design approach for Class 1 missions is addressed. The performance for these missions is usually specified in terms of (1) root mean square (RMS) pointing errors at different locations on the structure, and (2) the rate of decay of the transient response. Both of these performance measures include the contributions of rigid as well as elastic motion.

Leo Spacecraft Charging Design Guidelines: A Proposed NASA Standard

NASA Technical Reports Server (NTRS)

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.

Design and development of a brushless, direct drive solar array reorientation system

NASA Technical Reports Server (NTRS)

Jessee, R. D.

1972-01-01

This report covers the design and development of the laboratory model, and is essentially a compilation of reports covering the system and its various parts. To enhance completeness, the final report of Phase 1 covering circuit development of the controller is also included. A controller was developed for a brushless, direct-drive, single axis solar array reorientation system for earth-pointed, passively-stabilized spacecraft. A control systems was designed and breadboard circuits were built and tested for performance. The controller is designed to take over automatic control of the array on command after the spacecraft is stabilized in orbit. The controller will orient the solar array to the sun vector and automatically track to maintain proper orientation. So long as the orbit is circular, orientation toward the sun is maintained even though the spacecraft goes into the shadow of the earth. Particular attention was given in the design to limit reaction between the array and the spacecraft.

NASA Tropospheric Emission Spectrometer TES Instrument Onboard Aura

NASA Image and Video Library

2004-04-01

Technicians install NASA's Tropospheric Emission Spectrometer (TES) instrument on NASA's Aura spacecraft prior to launch. Launched in July 2004 and designed to fly for two years, the TES mission is currently in an extended operations phase. Mission managers at NASA's Jet Propulsion Laboratory, Pasadena, California, are evaluating an alternate way to collect and process science data from the Tropospheric Emission Spectrometer (TES) instrument on NASA's Aura spacecraft following the age-related failure of a critical instrument component. TES is an infrared sensor designed to study Earth's troposphere, the lowermost layer of Earth's atmosphere, which is where we live. The remainder of the TES instrument, and the Aura spacecraft itself, are operating as expected, and TES continues to collect science data. TES is one of four instruments on Aura, three of which are still operating. http://photojournal.jpl.nasa.gov/catalog/PIA15608

A conceptual design for an exoplanet imager

NASA Astrophysics Data System (ADS)

Hyland, David C.; Winkeller, Jon; Mosher, Robert; Momin, Anif; Iglesias, Gerardo; Donnellan, Quentin; Stanley, Jerry; Myers, Storm; Whittington, William G.; Asazuma, Taro; Slagle, Kami; Newton, Lindsay; Bourgeois, Scott; Tejeda, Donny; Young, Brian; Shaver, Nick; Cooper, Jacob; Underwood, Dennis; Perkins, James; Morea, Nathan; Goodnight, Ryan; Colunga, Aaron; Peltier, Scott; Singleton, Zane; Brashear, John; McPherson, Ronald; Guillory, Winston; Patel, Sunil; Stovall, Rachel; Meyer, Ryall; Eberle, Patrick; Morrison, Cole; Mong, Chun Yu

2007-09-01

This paper reports the results of a design study for an exoplanet imaging system. The design team consisted of the students in the "Electromagnetic Sensing for Space-Bourne Imaging" class taught by the principal author in the Spring, 2005 semester. The design challenge was to devise a space system capable of forming 10X10 pixel images of terrestrial-class planets out to 10 parsecs, observing in the 9.0 to 17.0 microns range. It was presumed that this system would operate after the Terrestrial Planet Finder had been deployed and had identified a number of planetary systems for more detailed imaging. The design team evaluated a large number of tradeoffs, starting with the use of a single monolithic telescope, versus a truss-mounted sparse aperture, versus a formation of free-flying telescopes. Having selected the free-flyer option, the team studied a variety of sensing technologies, including amplitude interferometry, intensity correlation imaging (ICI, based on the Brown-Twiss effect and phase retrieval), heterodyne interferometry and direct electric field reconstruction. Intensity correlation imaging was found to have several advantages. It does not require combiner spacecraft, nor nanometer-level control of the relative positions, nor diffraction-limited optics. Orbit design, telescope design, spacecraft structural design, thermal management and communications architecture trades were also addressed. A six spacecraft design involving non-repeating baselines was selected. By varying the overall scale of the baselines it was found possible to unambiguously characterize an entire multi-planet system, to image the parent star and, for the largest base scales, to determine 10X10 pixel images of individual planets.

A Study of Learning Curve Impact on Three Identical Small Spacecraft

NASA Technical Reports Server (NTRS)

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.

Planning Inmarsat's second generation of spacecraft

NASA Astrophysics Data System (ADS)

Williams, W. P.

1982-09-01

The next generation of studies of the Inmarsat service are outlined, such as traffic forecasting studies, communications capacity estimates, space segment design, cost estimates, and financial analysis. Traffic forecasting will require future demand estimates, and a computer model has been developed which estimates demand over the Atlantic, Pacific, and Indian ocean regions. Communications estimates are based on traffic estimates, as a model converts traffic demand into a required capacity figure for a given area. The Erlang formula is used, requiring additional data such as peak hour ratios and distribution estimates. Basic space segment technical requirements are outlined (communications payload, transponder arrangements, etc), and further design studies involve such areas as space segment configuration, launcher and spacecraft studies, transmission planning, and earth segment configurations. Cost estimates of proposed design parameters will be performed, but options must be reduced to make construction feasible. Finally, a financial analysis will be carried out in order to calculate financial returns.

A GNC Perspective of the Launch and Commissioning of NASA's SMAP (Soil Moisture Active Passive) Spacecraft

NASA Technical Reports Server (NTRS)

Brown, Todd S.

2016-01-01

The NASA Soil Moisture Active Passive (SMAP) spacecraft was designed to use radar and radiometer measurements to produce global soil moisture measurements every 2-3 days. The SMAP spacecraft is a complicated dual-spinning design with a large 6 meter deployable mesh reflector mounted on a platform that spins at 14.6 rpm while the Guidance Navigation and Control algorithms maintain precise nadir pointing for the de-spun portion of the spacecraft. After launching in early 2015, the Guidance Navigation and Control software and hardware aboard the SMAP spacecraft underwent an intensive spacecraft checkout and commissioning period. This paper describes the activities performed by the Guidance Navigation and Control team to confirm the health and phasing of subsystem hardware and the functionality of the guidance and control modes and algorithms. The operations tasks performed, as well as anomalies that were encountered during the commissioning, are explained and results are summarized.

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

NASA Technical Reports Server (NTRS)

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.

Position Extrema in Keplerian Relative Motion: A Gröbner Basis Approach

NASA Astrophysics Data System (ADS)

Allgeier, Shawn E.; Fitz-Coy, Norman G.; Erwin, R. Scott

2012-12-01

This paper analyzes the relative motion between two spacecraft in orbit. Specifically, the paper provides bounds for relative spacecraft position-based measures which impact spacecraft formation-flight mission design and analysis. Previous efforts have provided bounds for the separation distance between two spacecraft. This paper presents a methodology for bounding the local vertical, horizontal, and cross track components of the relative position vector in a spacecraft centered, rotating reference frame. Three metrics are derived and a methodology for bounding them is presented. The solution of the extremal equations for the metrics is formulated as an affine variety and obtained using a Gröbner basis reduction. No approximations are utilized and the only assumption is that the two spacecraft are in bound Keplerian orbits. Numerical examples are included to demonstrate the efficacy of the method. The metrics have utility to the mission designer of formation flight architectures, with relevance to Earth observation constellations.

Searching for organics on the dwarf planet Ceres

NASA Astrophysics Data System (ADS)

Nayak, Michael

The Herschel Space Observatory recently detected the presence of water vapor in observations of Ceres, bringing it into the crosshairs of the search for the building blocks of life in the solar system. I present a mission concept designed in collaboration with the NASA Ames Research Center for a two-probe mission to the dwarf planet Ceres, utilizing a pair of small low-cost spacecraft. The primary spacecraft will carry both a mass and an infrared spectrometer to characterize the detected vapor. Shortly after its arrival a second and largely similar spacecraft will impact Ceres to create an impact ejecta "plume" timed to enable a rendezvous and sampling by the primary spacecraft. This enables additional subsurface chemistry, volatile content and material characterization, and new science complementary to the Dawn spacecraft, the first to arrive at Ceres. Science requirements, candidate instruments, rendezvous trajectories, spacecraft design and comparison with Dawn science are detailed.

Nonlinear Dynamic Behavior in the Cassini Spacecraft Modal Survey

NASA Technical Reports Server (NTRS)

Carney, Kelly S.

1997-01-01

In October 1997, the 6-ton robotic spacecraft, Cassini, will lift off from Cape Canaveral atop a Titan IV B rocket, beginning a 7-year journey to Saturn. Upon completion of that voyage, Cassini will send the Huygens probe into the atmosphere of Saturn's largest moon, Titan. Cassini will then spend years studying Saturn's vast realm of rings, icy moons, and magnetic fields. The size and complexity of this endeavor mandates the involvement of many organizations. The Jet Propulsion Laboratory (JPL) manages the project for NASA and is responsible for the spacecraft design, development, and assembly. The NASA Lewis Research Center is the launch system integrator. As is typical for such a spacecraft, a test-verified finite element model is required for loads analysis. JPL had responsibility for the Cassini modal survey and the development of the spacecraft test-verified finite element model. Test verification is a complex and sometimes subjective process. Because of this, NASA Lewis independently verified and validated the Cassini spacecraft modal survey.

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

NASA Astrophysics Data System (ADS)

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.

KSC-98pc1234

NASA Image and Video Library

1998-10-03

KENNEDY SPACE CENTER, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), the top of the Mars Polar Lander is secured on a portable stand. The Lander will undergo testing, including a functional test of the science instruments and the basic spacecraft subsystems, before its launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999. The solar-powered spacecraft is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

KSC-98pc1230

NASA Image and Video Library

1998-10-03

KENNEDY SPACE CENTER, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), the top of the Mars Polar Lander is removed for testing, which includes a functional test of the science instruments and the basic spacecraft subsystems. The Mars Polar Lander is targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999. The solar-powered spacecraft is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

KSC-98pc1231

NASA Image and Video Library

1998-10-03

KENNEDY SPACE CENTER, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), the Mars Polar Lander is secured on a workstand for testing, which includes a functional test of the science instruments and the basic spacecraft subsystems. The Mars Polar Lander is targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999. The solar-powered spacecraft is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

KSC-98pc1235

NASA Image and Video Library

1998-10-03

KENNEDY SPACE CENTER, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), a technician begins testing on the Mars Polar Lander. The checkout includes a functional test of the science instruments and the basic spacecraft subsystems. The Mars Polar Lander is targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999. The solar-powered spacecraft is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

An optimization-based integrated controls-structures design methodology for flexible space structures

NASA Technical Reports Server (NTRS)

Maghami, Peiman G.; Joshi, Suresh M.; Armstrong, Ernest S.

1993-01-01

An approach for an optimization-based integrated controls-structures design is presented for a class of flexible spacecraft that require fine attitude pointing and vibration suppression. The integrated design problem is posed in the form of simultaneous optimization of both structural and control design variables. The approach is demonstrated by application to the integrated design of a generic space platform and to a model of a ground-based flexible structure. The numerical results obtained indicate that the integrated design approach can yield spacecraft designs that have substantially superior performance over a conventional design wherein the structural and control designs are performed sequentially. For example, a 40-percent reduction in the pointing error is observed along with a slight reduction in mass, or an almost twofold increase in the controlled performance is indicated with more than a 5-percent reduction in the overall mass of the spacecraft (a reduction of hundreds of kilograms).

Solar array study for solar electric propulsion spacecraft for the Encke rendezvous mission

NASA Technical Reports Server (NTRS)

Sequeira, E. A.; Patterson, R. E.

1974-01-01

The work is described which was performed on the design, analysis and performance of a 20 kW rollup solar array capable of meeting the design requirements of a solar electric spacecraft for the 1980 Encke rendezvous mission. To meet the high power requirements of the proposed electric propulsion mission, solar arrays on the order of 186.6 sq m were defined. Because of the large weights involved with arrays of this size, consideration of array configurations is limited to lightweight, large area concepts with maximum power-to-weight ratios. Items covered include solar array requirements and constraints, array concept selection and rationale, structural and electrical design considerations, and reliability considerations.

Spacecraft Fire Safety Research at NASA Glenn Research Center

NASA Technical Reports Server (NTRS)

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 results focusing on realistic fuel preparations and heating profiles with regards to early detection of smoke. SAFFIRE is the upcoming large-scale fire experiment which will be executed in a Cygnus vehicle after it undocks from the ISS.

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

NASA Technical Reports Server (NTRS)

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.;

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

NASA Technical Reports Server (NTRS)

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.

Operator Performance Evaluation of Fault Management Interfaces for Next-Generation Spacecraft

NASA Technical Reports Server (NTRS)

Hayashi, Miwa; Ravinder, Ujwala; Beutter, Brent; McCann, Robert S.; Spirkovska, Lilly; Renema, Fritz

2008-01-01

In the cockpit of the NASA's next generation of spacecraft, most of vehicle commanding will be carried out via electronic interfaces instead of hard cockpit switches. Checklists will be also displayed and completed on electronic procedure viewers rather than from paper. Transitioning to electronic cockpit interfaces opens up opportunities for more automated assistance, including automated root-cause diagnosis capability. The paper reports an empirical study evaluating two potential concepts for fault management interfaces incorporating two different levels of automation. The operator performance benefits produced by automation were assessed. Also, some design recommendations for spacecraft fault management interfaces are discussed.

Multi-kilowatt modularized spacecraft power processing system development

NASA Technical Reports Server (NTRS)

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

1975-01-01

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.

Study of a solid hydrogen cooler for spacecraft instruments and sensors

NASA Astrophysics Data System (ADS)

Sherman, A.

1980-08-01

The results of tests and studies to investigate the utilization of solid hydrogen for cooling of spacecraft instruments and sensors are presented. The results are presented in two sections; the first describing the tests in which an existing single stage solid cooler was filled and tested with solid hydrogen and the second which describes the analysis and design of a catalytic converter which will be tested in the vent line of the cooler.

Study of a solid hydrogen cooler for spacecraft instruments and sensors

NASA Technical Reports Server (NTRS)

Sherman, A.

1980-01-01

The results of tests and studies to investigate the utilization of solid hydrogen for cooling of spacecraft instruments and sensors are presented. The results are presented in two sections; the first describing the tests in which an existing single stage solid cooler was filled and tested with solid hydrogen and the second which describes the analysis and design of a catalytic converter which will be tested in the vent line of the cooler.

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

NASA Astrophysics Data System (ADS)

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

2017-08-01

ATHENA, Europe's next generation x-ray telescope, is currently under Assessment Phase study with parallel candidate industrial Prime contractors after selection for the 'L2' slot in ESA's Cosmic Vision Programme, with a mandate to address the 'Hot and Energetic Universe' Cosmic Vision science theme. This paper will consider the main technical requirements of the mission, and their mapping to resulting design choices at both mission and spacecraft level. The reference mission architecture and current reference spacecraft design will then be described, with particular emphasis given to description of the Science Instrument Module (SIM) design, currently under the responsibility of the ESA Study Team. The SIM is a very challenging item due primarily to the need to provide to the instruments (i) a soft ride during launch, and (ii) a very large ( 3 kW) heat dissipation capability at varying interface temperatures and locations.

Trajectory Optimization of an Interstellar Mission Using Solar Electric Propulsion

NASA Technical Reports Server (NTRS)

Kluever, Craig A.

1996-01-01

This paper presents several mission designs for heliospheric boundary exploration using spacecraft with low-thrust ion engines as the primary mode of propulsion The mission design goal is to transfer a 200-kg spacecraft to the heliospheric boundary in minimum time. The mission design is a combined trajectory and propulsion system optimization problem. Trajectory design variables include launch date, launch energy, burn and coast arc switch times, thrust steering direction, and planetary flyby conditions. Propulsion system design parameters include input power and specific impulse. Both SEP and NEP spacecraft arc considered and a wide range of launch vehicle options are investigated. Numerical results are presented and comparisons with the all chemical heliospheric missions from Ref 9 are made.

Electromagnetic braking for Mars spacecraft

NASA Technical Reports Server (NTRS)

Holt, A. C.

1986-01-01

Aerobraking concepts are being studied to improve performance and cost effectiveness of propulsion systems for Mars landers and Mars interplanetary spacecraft. Access to megawatt power levels (nuclear power coupled to high-storage inductive or capacitive devices) on a manned Mars interplanetary spacecraft may make feasible electromagnetic braking and lift modulation techniques which were previously impractical. Using pulsed microwave and magnetic field technology, potential plasmadynamic braking and hydromagnetic lift modulation techniques have been identified. Entry corridor modulation to reduce loads and heating, to reduce vertical descent rates, and to expand horizontal and lateral landing ranges are possible benefits. In-depth studies are needed to identify specific design concepts for feasibility assessments. Standing wave/plasma sheath interaction techniques appear to be promising. The techniques may require some tailoring of spacecraft external structures and materials. In addition, rapid response guidance and control systems may require the use of structurally embedded sensors coupled to expert systems or to artificial intelligence systems.

KSC-98pc1196

NASA Image and Video Library

1998-10-01

KENNEDY SPACE CENTER, FLA. -- At the Shuttle Landing Facility, the Mars Polar Lander is loaded onto a truck after its flight aboard an Air Force C-17 cargo plane that carried it from the Lockheed Martin Astronautics plant in Denver, CO. The lander is being transported to the Spacecraft Assembly and Encapsulation Facility-2(SAEF-2) in the KSC Industrial Area for testing, including a functional test of the science instruments and the basic spacecraft subsystems. The solar-powered spacecraft is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere. The Mars Polar Lander spacecraft is planned for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999

Direct conversion of infrared radiant energy for space power applications

NASA Technical Reports Server (NTRS)

Finke, R. C.

1982-01-01

A proposed technology to convert the earth radiant energy (infrared albedo) for spacecraft power is presented. The resultant system would eliminate energy storage requirements and simplify the spacecraft design. The design and performance of a infrared rectenna is discussed.

Optical metrology for Starlight Separated Spacecraft Stellar Interferometry Mission

NASA Technical Reports Server (NTRS)

Dubovitsky, S.; Lay, O. P.; Peters, R. D.; Abramovici, A.; Asbury, C. G.; Kuhnert, A. C.; Mulder, J. L.

2002-01-01

We describe a high-precision inter-spacecraft metrology system designed for NASA 's StarLight mission, a space-based separated-spacecraft stellar interferometer. It consists of dual-target linear metrology, based on a heterodyne interferometer with carrier phase modulation, and angular metrology designed to sense the pointing of the laser beam and provides bearing information. The dual-target operation enables one metrology beam to sense displacement of two targets independently. We present the current design, breadboard implementation of the Metrology Subsystem in a stellar interferometer testbed and the present state of development of flight qualifiable subsystem components.

Applications of Multi-Body Dynamical Environments: The ARTEMIS Transfer Trajectory Design

NASA Technical Reports Server (NTRS)

Folta, David C.; Woodard, Mark; Howell, Kathleen; Patterson, Chris; Schlei, Wayne

2010-01-01

The application of forces in multi-body dynamical environments to pennit the transfer of spacecraft from Earth orbit to Sun-Earth weak stability regions and then return to the Earth-Moon libration (L1 and L2) orbits has been successfully accomplished for the first time. This demonstrated transfer is a positive step in the realization of a design process that can be used to transfer spacecraft with minimal Delta-V expenditures. Initialized using gravity assists to overcome fuel constraints; the ARTEMIS trajectory design has successfully placed two spacecraft into EarthMoon libration orbits by means of these applications.

A novel approach to spacecraft re-entry and recovery

NASA Astrophysics Data System (ADS)

Patten, Richard; Hedgecock, Judson C.

1990-01-01

A deployable radiative heat shield design for spacecraft reentry is discussed. The design would allow the spacecraft to be cylindrical instead of the the traditional conical shape, providing a greater internal volume and thus enhancing mission capabilities. The heat shield uses a flexible thermal blanket material which is deployed in a manner similar to an umbrella. Based on the radiative properties of this blanket material, heating constraints have been established which allow a descent trajectory to be designed. The heat shield and capsule configuration are analyzed for resistance to heat flux and aerodynamic stability based on reentry trajectory. Experimental tests are proposed.

Environmental Verification Experiment for the Explorer Platform (EVEEP)

NASA Technical Reports Server (NTRS)

Norris, Bonnie; Lorentson, Chris

1992-01-01

Satellites and long-life spacecraft require effective contamination control measures to ensure data accuracy and maintain overall system performance margins. Satellite and spacecraft contamination can occur from either molecular or particulate matter. Some of the sources of the molecular species are as follows: mass loss from nonmetallic materials; venting of confined spacecraft or experiment volumes; exhaust effluents from attitude control systems; integration and test activities; and improper cleaning of surfaces. Some of the sources of particulates are as follows: leaks or purges which condense upon vacuum exposure; abrasion of movable surfaces; and micrometeoroid impacts. The Environmental Verification Experiment for the Explorer Platform (EVEEP) was designed to investigate the following aspects of spacecraft contamination control: materials selection; contamination modeling of existing designs; and thermal vacuum testing of a spacecraft with contamination monitors.

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

PubMed

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.

ProSEDS Telemetry System Utilization of GPS Position Data for Transmitter Cycling

NASA Technical Reports Server (NTRS)

Kennedy, Paul; Sims, Herb

2000-01-01

NASA Marshall Space Flight Center will launch the Propulsive Small Expendable Deployer System (ProSEDS) space experiment in late 2000. ProSEDS will demonstrate the use of an electrodynamic tether propulsion system and will utilize a conducting wire tether to generate limited spacecraft power. This paper will provide an overview of the ProSEDS mission and will discuss the design, development and test of the spacecraft telemetry system which utilizes a custom designed GPS subsystem to determine spacecraft position relative to ground station location and to control transmitter on/off cycling based on spacecraft state vector and ground station visibility.

Space environmental effects on spacecraft: LEO materials selection guide, part 1

NASA Astrophysics Data System (ADS)

Silverman, Edward M.

1995-08-01

This document provides performance properties on major spacecraft materials and subsystems that have been exposed to the low-Earth orbit (LEO) space environment. Spacecraft materials include metals, polymers, composites, white and black paints, thermal-control blankets, adhesives, and lubricants. Spacecraft subsystems include optical components, solar cells, and electronics. Information has been compiled from LEO short-term spaceflight experiments (e.g., space shuttle) and from retrieved satellites of longer mission durations (e.g., Long Duration Exposure Facility). Major space environment effects include atomic oxygen (AO), ultraviolet radiation, micrometeoroids and debris, contamination, and particle radiation. The main objective of this document is to provide a decision tool to designers for designing spacecraft and structures. This document identifies the space environments that will affect the performance of materials and components, e.g., thermal-optical property changes of paints due to UV exposures, AO-induced surface erosion of composites, dimensional changes due to thermal cycling, vacuum-induced moisture outgassing, and surface optical changes due to AO/UV exposures. Where appropriate, relationships between the space environment and the attendant material/system effects are identified. Part 1 covers spacecraft design considerations for the space environment; advanced composites; polymers; adhesives; metals; ceramics; protective coatings; and lubricants, greases, and seals.

Space environmental effects on spacecraft: LEO materials selection guide, part 1

NASA Technical Reports Server (NTRS)

Silverman, Edward M.

1995-01-01

This document provides performance properties on major spacecraft materials and subsystems that have been exposed to the low-Earth orbit (LEO) space environment. Spacecraft materials include metals, polymers, composites, white and black paints, thermal-control blankets, adhesives, and lubricants. Spacecraft subsystems include optical components, solar cells, and electronics. Information has been compiled from LEO short-term spaceflight experiments (e.g., space shuttle) and from retrieved satellites of longer mission durations (e.g., Long Duration Exposure Facility). Major space environment effects include atomic oxygen (AO), ultraviolet radiation, micrometeoroids and debris, contamination, and particle radiation. The main objective of this document is to provide a decision tool to designers for designing spacecraft and structures. This document identifies the space environments that will affect the performance of materials and components, e.g., thermal-optical property changes of paints due to UV exposures, AO-induced surface erosion of composites, dimensional changes due to thermal cycling, vacuum-induced moisture outgassing, and surface optical changes due to AO/UV exposures. Where appropriate, relationships between the space environment and the attendant material/system effects are identified. Part 1 covers spacecraft design considerations for the space environment; advanced composites; polymers; adhesives; metals; ceramics; protective coatings; and lubricants, greases, and seals.

Total-System Approach To Design And Analysis Of Structures

NASA Technical Reports Server (NTRS)

Verderaime, V.

1995-01-01

Paper presents overview and study of, and comprehensive approach to, multidisciplinary engineering design and analysis of structures. Emphasizes issues related to design of semistatic structures in environments in which spacecraft launched, underlying concepts applicable to other structures within unique terrestrial, marine, or flight environments. Purpose of study to understand interactions among traditionally separate engineering design disciplines with view toward optimizing not only structure but also overall design process.

Habitability design for spacecraft

NASA Technical Reports Server (NTRS)

Franklin, G. C.

1978-01-01

Habitability is understood to mean those spacecraft design elements that involve a degree of comfort, quality or necessities to support man in space. These elements are environment, architecture, mobility, clothing, housekeeping, food and drink, personal hygiene, off-duty activities, each of which plays a substantial part in the success of a mission. Habitability design for past space flights is discussed relative to the Mercury, Gemini, Apollo, and Skylab spacecraft, with special emphasis on an examination of the Shuttle Orbiter cabin design from a habitability standpoint. Future projects must consider the duration and mission objectives to meet their habitability requirements. Larger ward rooms, improved sleeping quarters and more complete hygiene facilities must be provided for future prolonged space flights

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

NASA Astrophysics Data System (ADS)

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.

Overview and Software Architecture of the Copernicus Trajectory Design and Optimization System

NASA Technical Reports Server (NTRS)

Williams, Jacob; Senent, Juan S.; Ocampo, Cesar; Mathur, Ravi; Davis, Elizabeth C.

2010-01-01

The Copernicus Trajectory Design and Optimization System represents an innovative and comprehensive approach to on-orbit mission design, trajectory analysis and optimization. Copernicus integrates state of the art algorithms in optimization, interactive visualization, spacecraft state propagation, and data input-output interfaces, allowing the analyst to design spacecraft missions to all possible Solar System destinations. All of these features are incorporated within a single architecture that can be used interactively via a comprehensive GUI interface, or passively via external interfaces that execute batch processes. This paper describes the Copernicus software architecture together with the challenges associated with its implementation. Additionally, future development and planned new capabilities are discussed. Key words: Copernicus, Spacecraft Trajectory Optimization Software.

STARDUST and HAYABUSA: Sample Return Missions to Small Bodies in the Solar System

NASA Technical Reports Server (NTRS)

Sandford, S. A.

2005-01-01

There are currently two active spacecraft missions designed to return samples to Earth from small bodies in our Solar System. STARDUST will return samples from the comet Wild 2, and HAYABUSA will return samples from the asteroid Itokawa. On January 3,2004, the STARDUST spacecraft made the closest ever flyby (236 km) of the nucleus of a comet - Comet Wild 2. During the flyby the spacecraft collected samples of dust from the coma of the comet. These samples will be returned to Earth on January 15,2006. After a brief preliminary examination to establish the nature of the returned samples, they will be made available to the general scientific community for study. The HAYABUSA spacecraft arrived at the Near Earth Asteroid Itokawa in September 2005 and is currently involved in taking remote sensing data from the asteroid. Several practice landings have been made and a sample collection landing will be made soon. The collected sample will be returned to Earth in June 2007. During my talk I will discuss the scientific goals of the STARDUST and HAYABUSA missions and provide an overview of their designs and flights to date. I will also show some of the exciting data returned by these spacecraft during their encounters with their target objects.

Propulsion Trade Studies for Spacecraft Swarm Mission Design

NASA Technical Reports Server (NTRS)

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.

LEO cooperative multi-spacecraft refueling mission optimization considering J2 perturbation and target's surplus propellant constraint

NASA Astrophysics Data System (ADS)

Zhao, Zhao; Zhang, Jin; Li, Hai-yang; Zhou, Jian-yong

2017-01-01

The optimization of an LEO cooperative multi-spacecraft refueling mission considering the J2 perturbation and target's surplus propellant constraint is studied in the paper. First, a mission scenario is introduced. One service spacecraft and several target spacecraft run on an LEO near-circular orbit, the service spacecraft rendezvouses with some service positions one by one, and target spacecraft transfer to corresponding service positions respectively. Each target spacecraft returns to its original position after obtaining required propellant and the service spacecraft returns to its original position after refueling all target spacecraft. Next, an optimization model of this mission is built. The service sequence, orbital transfer time, and service position are used as deign variables, whereas the propellant cost is used as the design objective. The J2 perturbation, time constraint and the target spacecraft's surplus propellant capability constraint are taken into account. Then, a hybrid two-level optimization approach is presented to solve the formulated mixed integer nonlinear programming (MINLP) problem. A hybrid-encoding genetic algorithm is adopted to seek the near optimal solution in the up-level optimization, while a linear relative dynamic equation considering the J2 perturbation is used to obtain the impulses of orbital transfer in the low-level optimization. Finally, the effectiveness of the proposed model and method is validated by numerical examples.

Guidance and Navigation Software Architecture Design for the Autonomous Multi-Agent Physically Interacting Spacecraft (AMPHIS) Test Bed

DTIC Science & Technology

2006-12-01

NAVIGATION SOFTWARE ARCHITECTURE DESIGN FOR THE AUTONOMOUS MULTI-AGENT PHYSICALLY INTERACTING SPACECRAFT (AMPHIS) TEST BED by Blake D. Eikenberry...Engineer Degree 4. TITLE AND SUBTITLE Guidance and Navigation Software Architecture Design for the Autonomous Multi- Agent Physically Interacting...iii Approved for public release; distribution is unlimited GUIDANCE AND NAVIGATION SOFTWARE ARCHITECTURE DESIGN FOR THE AUTONOMOUS MULTI

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

NASA Astrophysics Data System (ADS)

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.

Handbook for Designing MMOD Protection

NASA Technical Reports Server (NTRS)

Arnold, Jim; Christiansen, Eric L.; Davis, Alan; Hyde, James; Lear, Dana; Liou, J.C.; Lyons, Frankel; Prior, Thomas; Studor, George; Ratliff, Martin;

Spacecraft Charging: Hazard Causes, Hazard Effects, Hazard Controls

NASA Technical Reports Server (NTRS)

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.

Design Study for a Mars Geyser Hopper

NASA Technical Reports Server (NTRS)

Landis, Geoffrey A.; Oleson, Steven J.; McGuire, Melissa

2012-01-01

The Mars Geyser Hopper is a design reference missions (DRMs) for a Discovery-class spacecraft using Advanced Stirling Radioisotope Generator (ASRG) power source. The Geyser Hopper is a mission concept that will investigate the springtime carbon-dioxide geysers found in regions around the south pole of Mars. The Geyser Hopper design uses Phoenix heritage systems and approach, but uses a single ASRG as the power source, rather than twin solar arrays, and is designed to last over a one-year stay on the South Pole. The spacecraft will land at a target landing area near the south pole of Mars, and have the ability to "hop" after a summertime landing to reposition itself close to a geyser site, and wait through the winter until the first sunlight of spring to witness first-hand the geyser phenomenon.

A membrane-based subsystem for very high recoveries of spacecraft waste waters

NASA Technical Reports Server (NTRS)

Ray, Roderick J.; Retzlaff, Sandra E.; Radke-Mitchell, Lyn; Newbold, David D.; Price, Donald F.

1986-01-01

This paper describes the continued development of a membrane-based subsystem designed to recover up to 99.5 percent of the water from various spacecraft waste waters. Specifically discussed are: (1) the design and fabrication of an energy-efficient reverse-osmosis (RO) breadboard subsystem; (2) data showing the performance of this subsystem when operated on a synthetic wash-water solution - including the results of a 92-day test; and (3) the results of pasteurization studies, including the design and operation of an in-line pasteurizer. Also included in this paper is a discussion of the design and performance of a second RO stage. This second stage results in higher-purity product water at a minimal energy requirement and provides a substantial redundancy factor to this subsystem.

Conceptual Launch Vehicle and Spacecraft Design for Risk Assessment

NASA Technical Reports Server (NTRS)

Motiwala, Samira A.; Mathias, Donovan L.; Mattenberger, Christopher J.

2014-01-01

One of the most challenging aspects of developing human space launch and exploration systems is minimizing and mitigating the many potential risk factors to ensure the safest possible design while also meeting the required cost, weight, and performance criteria. In order to accomplish this, effective risk analyses and trade studies are needed to identify key risk drivers, dependencies, and sensitivities as the design evolves. The Engineering Risk Assessment (ERA) team at NASA Ames Research Center (ARC) develops advanced risk analysis approaches, models, and tools to provide such meaningful risk and reliability data throughout vehicle development. The goal of the project presented in this memorandum is to design a generic launch 7 vehicle and spacecraft architecture that can be used to develop and demonstrate these new risk analysis techniques without relying on other proprietary or sensitive vehicle designs. To accomplish this, initial spacecraft and launch vehicle (LV) designs were established using historical sizing relationships for a mission delivering four crewmembers and equipment to the International Space Station (ISS). Mass-estimating relationships (MERs) were used to size the crew capsule and launch vehicle, and a combination of optimization techniques and iterative design processes were employed to determine a possible two-stage-to-orbit (TSTO) launch trajectory into a 350-kilometer orbit. Primary subsystems were also designed for the crewed capsule architecture, based on a 24-hour on-orbit mission with a 7-day contingency. Safety analysis was also performed to identify major risks to crew survivability and assess the system's overall reliability. These procedures and analyses validate that the architecture's basic design and performance are reasonable to be used for risk trade studies. While the vehicle designs presented are not intended to represent a viable architecture, they will provide a valuable initial platform for developing and demonstrating innovative risk assessment capabilities.

Thermionic reactor power conditioner design for nuclear electric propulsion.

NASA Technical Reports Server (NTRS)

Jacobsen, A. S.; Tasca, D. M.

1971-01-01

Consideration of the effects of various thermionic reactor parameters and requirements upon spacecraft power conditioning design. A basic spacecraft is defined using nuclear electric propulsion, requiring approximately 120 kWe. The interrelationships of reactor operating characteristics and power conditioning requirements are discussed and evaluated, and the effects on power conditioner design and performance are presented.

Attitude Determination and Control Subsystem (ADCS) Preparations for the EPOXI Flyby of Comet Haley 2

NASA Technical Reports Server (NTRS)

Luna, Michael E.; Collins, Stephen M.

2011-01-01

On November 4, 2010 the already "in-flight" Deep Impact spacecraft flew within 700km of comet 103P/Hartley 2 as part of its extended mission EPOXI, the 5th time to date any spacecraft visited a comet. In 2005, the spacecraft had previously imaged a probe impact comet Tempel 1. The EPOXI flyby marked the first time in history that two comets were explored with the same instruments on a re-used spacecraft-with hardware and software originally designed and optimized for a different mission. This made the function of the attitude determination and control subsystem (ADCS) critical to the successful execution of the EPOXI flyby. As part of the spacecraft team preparations, the ADCS team had to perform thorough sequence reviews, key spacecraft activities and onboard calibrations. These activities included: review of background sequences for the initial conditions vector, sun sensor coefficients, and reaction wheel assembly (RWA) de-saturations; design and execution of 10 trajectory correction maneuvers; science calibration of the two telescope instruments; a flight demonstration of the fastest turns conducted by the spacecraft between Earth and comet point; and assessment of RWA health (given RWA problems on other spacecraft).

An experiment to study energetic particle fluxes in and beyond the earth's outer magnetosphere

NASA Technical Reports Server (NTRS)

Anderson, K. A.; Lin, R. P.; Paoli, R. J.; Parks, G. K.; Lin, C. S.; Reme, H.; Bosqued, J. M.; Martel, F.; Cotin, F.; Cros, A.

1978-01-01

This experiment is designed to take advantage of the ISEE Mother/Daughter dual spacecraft system to study energetic particle phenomena in the earth's outer magnetosphere and beyond. Large geometric factor fixed voltage electrostatic analyzers and passively cooled semiconductor detector telescopes provide high time resolution coverage of the energy range from 1.5 to 300 keV for both ions and electrons. Essentially identical instrumentation is placed on the two spacecraft to separate temporal from spatial effects in the observed particle phenomena.

Meteoroid Protection Methods for Spacecraft Radiators Using Heat Pipes

NASA Technical Reports Server (NTRS)

Ernst, D. M.

1979-01-01

Various aspects of achieving a low mass heat pipe radiator for the nuclear electric propulsion spacecraft were studied. Specific emphasis was placed on a concept applicable to a closed Brayton cycle power sub-system. Three aspects of inter-related problems were examined: (1) the armor for meteoroid protection, (2) emissivity of the radiator surface, and (3) the heat pipe itself. The study revealed several alternatives for the achievement of the stated goal, but a final recommendation for the best design requires further investigation.

A study of an orbital radar mapping mission to Venus. Volume 1: Summary

NASA Technical Reports Server (NTRS)

1973-01-01

A preliminary design of a Venus radar mapping orbiter mission and spacecraft was developed. The important technological problems were identified and evaluated. The study was primarily concerned with trading off alternate ways of implementing the mission and examining the most attractive concepts in order to assess technology requirements. Compatible groupings of mission and spacecraft parameters were analyzed by examining the interaction of their functioning elements and assessing their overall cost effectiveness in performing the mission.

Three Generations of Tracking and Data Relay Satellite (TDRS) Spacecraft

NASA Technical Reports Server (NTRS)

Zaleski, Ron

2016-01-01

The current Tracking and Data Relay Satellite configuration consists of nine in-orbit satellites (four first generation, three second generation and two third generation satellites) globally distributed in geosynchronous orbit to provide near continuous data relay service to missions like Hubble Space Telescope and the International Space Station. The 1st generation spacecraft were designed by TRW/Northrop Grumman with their launches of the five spacecraft ranging from 1983 through 1995. The 2nd and 3rd generation spacecraft were designed by Boeing with their launches ranging 2000 - 2002 and 2013 - 2017 respectively. TDRS-3 is now 27 years on orbit, continues to be a capable asset for the TDRS constellation. Lack of need for inclination control combined with large fuel reserves and redundancy on critical elements provides spacecraft that operate well past design life, all of which contributes to expanded TDRS constellation support capabilities. All spacecraft generations have issues. Significant issues will be summarized with the focus on the Boeing related problems. Degradations and failures are continually assessed and provide the foundation for yearly updates to spacecraft reliability models, constellation service projections and deorbit plans (in order to meet NASAs mandate of limiting orbital debris). Even when accounting for degradations and failures, the life expectancy for the Boeing delivered 2nd generation TDRS-8, 9 10 TDRS are anticipated to be 25+ years.

Design and development of the Cassini main engine assembly Gimbal mechanism

NASA Technical Reports Server (NTRS)

Rudolph, Dale

1996-01-01

Cassini is an international cooperative effort between NASA, which is producing the orbiter spacecraft, the European Space Agency, which is providing the Huygens Probe, and the Italian Space Agency, which is responsible for the spacecraft radio antenna and portions of three scientific experiments. In the U.S., the mission is managed by NASA's Jet Propulsion Laboratory (JPL) in Pasadena, California. Lockheed-Martin successfully bid on the contract to build the PMS (Propulsion Module Subsystem) for this project. The Cassini spacecraft will be launched on an expedition to Saturn in October, 1997. Its mission is to enter orbit around Saturn in July, 2004, and to explore its moons, rings, and magnetic environment for four years. Cassini will carry the Huygens probe, an instrument package equipped with a parachute, which is designed to study the atmosphere and surface of Saturn's largest moon, Titan.

Heat pipe radiator. [for spacecraft waste heat rejection

NASA Technical Reports Server (NTRS)

Swerdling, B.; Alario, J.

1973-01-01

A 15,000 watt spacecraft waste heat rejection system utilizing heat pipe radiator panels was investigated. Of the several concepts initially identified, a series system was selected for more in-depth analysis. As a demonstration of system feasibility, a nominal 500 watt radiator panel was designed, built and tested. The panel, which is a module of the 15,000 watt system, consists of a variable conductance heat pipe (VCHP) header, and six isothermalizer heat pipes attached to a radiating fin. The thermal load to the VCHP is supplied by a Freon-21 liquid loop via an integral heat exchanger. Descriptions of the results of the system studies and details of the radiator design are included along with the test results for both the heat pipe components and the assembled radiator panel. These results support the feasibility of using heat pipes in a spacecraft waste heat rejection system.

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

NASA Astrophysics Data System (ADS)

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.

Intelligent systems and advanced user interfaces for design, operation, and maintenance of command management systems

NASA Technical Reports Server (NTRS)

Potter, William J.; Mitchell, Christine M.

1993-01-01

Historically, command management systems (CMS) have been large and 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 to develop a more generic CMS system. New technologies, in addition to a core CMS common to a range of spacecraft, may facilitate the training and enhance the efficiency of CMS operations. Current mission operations center (MOC) hardware and software include Unix workstations, the C/C++ programming languages, and an X window interface. This configuration provides the power and flexibility to support sophisticated and intelligent user interfaces that exploit state-of-the-art technologies in human-machine interaction, 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 these issues in CMS design and operation suggests that application of technologies such as intelligent planning, case-based reasoning, human-machine systems design and analysis tools (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 the CMS for a specific spacecraft as well as continuity for CMS design and development across spacecraft. The first six months of this research saw a broad investigation by Georgia Tech researchers into the function, design, and operation of current and planned command management systems at Goddard Space Flight Center. As the first step, the researchers attempted to understand the current and anticipated horizons of command management systems at Goddard. Preliminary results are given on CMS commonalities and causes of low re-use, and methods are proposed to facilitate increased re-use.

Formation flying for electric sails in displaced orbits. Part I: Geometrical analysis

NASA Astrophysics Data System (ADS)

Wang, Wei; Mengali, Giovanni; Quarta, Alessandro A.; Yuan, Jianping

2017-09-01

We present a geometrical methodology for analyzing the formation flying of electric solar wind sail based spacecraft that operate in heliocentric, elliptic, displaced orbits. The spacecraft orbit is maintained by adjusting its propulsive acceleration modulus, whose value is estimated using a thrust model that takes into account a variation of the propulsive performance with the sail attitude. The properties of the relative motion of the spacecraft are studied in detail and a geometrical solution is obtained in terms of relative displaced orbital elements, assumed to be small quantities. In particular, for the small eccentricity case (i.e. for a near-circular displaced orbit), the bounds characterized by the extreme values of relative distances are analytically calculated, thus providing an useful mathematical tool for preliminary design of the spacecraft formation structure.

Comet flyby sample return

NASA Technical Reports Server (NTRS)

Tsou, P.; Albee, A.

1985-01-01

The results of a joint JPL/CSFC feasability study of a low-cost comet sample return flyby mission are presented. It is shown that the mission could be undertaken using current earth orbiter spacecraft technology in conjunction with pathfinder or beacon spacrcraft. Detailed scenarios of missions to the comets Honda-Mrkos-Pajdusakova (HMP), comet Kopff, and comet Giacobini-Zinner (GZ) are given, and some crossectional diagrams of the spacecraft designs are provided.

Spacecraft design sensitivity for a disaster warning satellite system

NASA Technical Reports Server (NTRS)

Maloy, J. E.; Provencher, C. E.; Leroy, B. E.; Braley, R. C.; Shumaker, H. A.

1977-01-01

A disaster warning satellite (DWS) is described for warning the general public of impending natural catastrophes. The concept is responsive to NOAA requirements and maximizes the use of ATS-6 technology. Upon completion of concept development, the study was extended to establishing the sensitivity of the DWSS spacecraft power, weight, and cost to variations in both warning and conventional communications functions. The results of this sensitivity analysis are presented.

Bayesian Approach for Reliability Assessment of Sunshield Deployment on JWST

NASA Technical Reports Server (NTRS)

Kaminskiy, Mark P.; Evans, John W.; Gallo, Luis D.

2013-01-01

Deployable subsystems are essential to mission success of most spacecraft. These subsystems enable critical functions including power, communications and thermal control. The loss of any of these functions will generally result in loss of the mission. These subsystems and their components often consist of unique designs and applications, for which various standardized data sources are not applicable for estimating reliability and for assessing risks. In this study, a Bayesian approach for reliability estimation of spacecraft deployment was developed for this purpose. This approach was then applied to the James Webb Space Telescope (JWST) Sunshield subsystem, a unique design intended for thermal control of the observatory's telescope and science instruments. In order to collect the prior information on deployable systems, detailed studies of "heritage information", were conducted extending over 45 years of spacecraft launches. The NASA Goddard Space Flight Center (GSFC) Spacecraft Operational Anomaly and Reporting System (SOARS) data were then used to estimate the parameters of the conjugative beta prior distribution for anomaly and failure occurrence, as the most consistent set of available data and that could be matched to launch histories. This allows for an emperical Bayesian prediction for the risk of an anomaly occurrence of the complex Sunshield deployment, with credibility limits, using prior deployment data and test information.

A Bayesian Framework for Reliability Analysis of Spacecraft Deployments

NASA Technical Reports Server (NTRS)

Evans, John W.; Gallo, Luis; Kaminsky, Mark

2012-01-01

Deployable subsystems are essential to mission success of most spacecraft. These subsystems enable critical functions including power, communications and thermal control. The loss of any of these functions will generally result in loss of the mission. These subsystems and their components often consist of unique designs and applications for which various standardized data sources are not applicable for estimating reliability and for assessing risks. In this study, a two stage sequential Bayesian framework for reliability estimation of spacecraft deployment was developed for this purpose. This process was then applied to the James Webb Space Telescope (JWST) Sunshield subsystem, a unique design intended for thermal control of the Optical Telescope Element. Initially, detailed studies of NASA deployment history, "heritage information", were conducted, extending over 45 years of spacecraft launches. This information was then coupled to a non-informative prior and a binomial likelihood function to create a posterior distribution for deployments of various subsystems uSing Monte Carlo Markov Chain sampling. Select distributions were then coupled to a subsequent analysis, using test data and anomaly occurrences on successive ground test deployments of scale model test articles of JWST hardware, to update the NASA heritage data. This allowed for a realistic prediction for the reliability of the complex Sunshield deployment, with credibility limits, within this two stage Bayesian framework.

The Magnetospheric Multiscale Mission...Resolving Fundamental Processes in Space Plasmas

NASA Technical Reports Server (NTRS)

Curtis, S.

1999-01-01

The Magnetospheric Multiscale (MMS) mission is a multiple-spacecraft Solar-Terrestrial Probe designed to study the microphysics of magnetic reconnection, charged particle acceleration, and turbulence in key boundary regions of Earth's magnetosphere. These three processes, which control the flow of energy, mass, and momentum within and across plasma boundaries, occur throughout the universe and are fundamental to our understanding of astrophysical and solar system plasmas. Only in Earth's magnetosphere, however, are they readily accessible for sustained study through in-situ measurement. MMS will employ five co-orbiting spacecraft identically instrumented to measure electric and magnetic fields, plasmas, and energetic particles. The initial parameters of the individual spacecraft orbits will be designed so that the spacecraft formation will evolve into a three-dimensional configuration near apogee, allowing MMS to differentiate between spatial and temporal effects and to determine the three dimensional geometry of plasma, field, and current structures. In order to sample all of the magnetospheric boundary regions, MMS will employ a unique four-phase orbital strategy involving carefully sequenced changes in the local time and radial distance of apogee and, in the third phase, a change in orbit inclination from 10 degrees to 90 degrees. The nominal mission operational lifetime is two years. Launch is currently scheduled for 2006.

A Low Cost Spacecraft Architecture for Robotic Lunar Exploration Projects

NASA Technical Reports Server (NTRS)

Lemke, Lawrence G.; Gonzales, Andrew A.

2006-01-01

A program of frequent, capable, but affordable lunar robotic missions prior to return of humans to the moon can contribute to the Vision for Space Exploration (VSE) NASA is tasked to execute. The Lunar Reconnaissance Orbiter (LRO) and its secondary payload are scheduled to orbit the moon, and impact it, respectively, in 2008. It is expected that the sequence of missions occurring for approximately the decade after 2008 will place an increasing emphasis on soft landed payloads. These missions are requited to explore intrinsic characteristics of the moon, such as hydrogen distribution in the regolith, and levitated dust, to demonstrate the ability to access and process in-situ resources, and to demonstrate functions critical to supporting human presence, such as automated precision navigation and landing. Additional factors governing the design of spacecraft to accomplish this diverse set of objectives are: operating within a relatively modest funding profile, the need tb visit multiple sites (both polar and equatorial) repeatedly, and to use the current generation of launch vehicles. In the US, this implies use of the Evolved Expendable Launch Vehicles, or EELVs, although this design philosophy may be extended to launch vehicles of other nations, as well. Many of these factors are seemingly inconsistent with each other. For example, the cost of a spacecraft usually increases with mass; therefore the desire to fly frequent, modestly priced spacecraft seems to imply small spacecraft (< 1 Mt, injected mass). On the other hand, the smallest of the EELVs will inject approx. 3 Mt. on a Trans Lunar Injection (TLI) trajectory md would therefore be wasteful or launching a single, small spacecraft. Increasing the technical capability of a spacecraft (such as autonomous navigation and soft landing) also usually increases cost. A strategy for spacecraft design that meets these conflicting requirements is presented. Taken together, spacecraft structure and propulsion subsystems constitute the majority of spacecraft mass; saving development and integration cost on these elements is critical to controlling cost. Therefore, a low cost, modular design for spacecraft structure and propulsion subsystems is presented which may be easily scaled up or down for either insertion into lunar orbit or braking for landing on the lunar surface. In order to effectively use the approx.3 Mt mass-to-TLI of the EELV, two low cost spacecraft will be manifested on the same launch. One spacecraft will be located on top of the other for launch and the two will have to be released in sequence in order to achieve all mission objectives. The two spacecraft could both be landers, both orbiters, or one lander and one orbiter. In order to achieve mass efficiency, the body of the spacecraft will serve the dual purposes of carrying launch loads and providing attachment points for all the spacecraft subsystems. In order to avoid unaffordable technology development costs, small liquid propulsion components and autonomous, scene-matching navigation cameras may be adapted from military missile programs in order to execute precision soft landings.

Pluto Fast Flyby: An Overview of the Mission and Spacecraft Design, Advanced Technology Insertion Efforts, and Student Involvement Opportunities

NASA Technical Reports Server (NTRS)

Abraham, D. S.; Staehle, R.; Brewster, S.; Caldwell, D.; Carraway, J.; Henry, P.; Herman, M.; Kissel, G.; Peak, S.; Randolph, V.;

Applying Formal Methods to NASA Projects: Transition from Research to Practice

NASA Technical Reports Server (NTRS)

Othon, Bill

2009-01-01

NASA project managers attempt to manage risk by relying on mature, well-understood process and technology when designing spacecraft. In the case of crewed systems, the margin for error is even tighter and leads to risk aversion. But as we look to future missions to the Moon and Mars, the complexity of the systems will increase as the spacecraft and crew work together with less reliance on Earth-based support. NASA will be forced to look for new ways to do business. Formal methods technologies can help NASA develop complex but cost effective spacecraft in many domains, including requirements and design, software development and inspection, and verification and validation of vehicle subsystems. To realize these gains, the technologies must be matured and field-tested so that they are proven when needed. During this discussion, current activities used to evaluate FM technologies for Orion spacecraft design will be reviewed. Also, suggestions will be made to demonstrate value to current designers, and mature the technology for eventual use in safety-critical NASA missions.

Trajectory Optimization for Missions to Small Bodies with a Focus on Scientific Merit.

PubMed

Englander, Jacob A; Vavrina, Matthew A; Lim, Lucy F; McFadden, Lucy A; Rhoden, Alyssa R; Noll, Keith S

2017-01-01

Trajectory design for missions to small bodies is tightly coupled both with the selection of targets for a mission and with the choice of spacecraft power, propulsion, and other hardware. Traditional methods of trajectory optimization have focused on finding the optimal trajectory for an a priori selection of destinations and spacecraft parameters. Recent research has expanded the field of trajectory optimization to multidisciplinary systems optimization that includes spacecraft parameters. The logical next step is to extend the optimization process to include target selection based not only on engineering figures of merit but also scientific value. This paper presents a new technique to solve the multidisciplinary mission optimization problem for small-bodies missions, including classical trajectory design, the choice of spacecraft power and propulsion systems, and also the scientific value of the targets. This technique, when combined with modern parallel computers, enables a holistic view of the small body mission design process that previously required iteration among several different design processes.

Microspacecraft and Earth observation: Electrical field (ELF) measurement project

NASA Technical Reports Server (NTRS)

Olsen, Tanya; Elkington, Scot; Parker, Scott; Smith, Grover; Shumway, Andrew; Christensen, Craig; Parsa, Mehrdad; Larsen, Layne; Martinez, Ranae; Powell, George

1990-01-01

The Utah State University space system design project for 1989 to 1990 focuses on the design of a global electrical field sensing system to be deployed in a constellation of microspacecraft. The design includes the selection of the sensor and the design of the spacecraft, the sensor support subsystems, the launch vehicle interface structure, on board data storage and communications subsystems, and associated ground receiving stations. Optimization of satellite orbits and spacecraft attitude are critical to the overall mapping of the electrical field and, thus, are also included in the project. The spacecraft design incorporates a deployable sensor array (5 m booms) into a spinning oblate platform. Data is taken every 0.1 seconds by the electrical field sensors and stored on-board. An omni-directional antenna communicates with a ground station twice per day to down link the stored data. Wrap-around solar cells cover the exterior of the spacecraft to generate power. Nine Pegasus launches may be used to deploy fifty such satellites to orbits with inclinations greater than 45 deg. Piggyback deployment from other launch vehicles such as the DELTA 2 is also examined.

Avionics for a Small Robotic Inspection Spacecraft

NASA Technical Reports Server (NTRS)

Abbott, Larry; Shuler, Robert L., Jr.

2005-01-01

A report describes the tentative design of the avionics of the Mini-AERCam -- a proposed 7.5-in. (approximately 19-cm)-diameter spacecraft that would contain three digital video cameras to be used in visual inspection of the exterior of a larger spacecraft (a space shuttle or the International Space Station). The Mini-AERCam would maneuver by use of its own miniature thrusters under radio control by astronauts inside the larger spacecraft. The design of the Mini-AERCam avionics is subject to a number of constraints, most of which can be summarized as severely competing requirements to maximize radiation hardness and maneuvering, image-acquisition, and data-communication capabilities while minimizing cost, size, and power consumption. The report discusses the design constraints, the engineering approach to satisfying the constraints, and the resulting iterations of the design. The report places special emphasis on the design of a flight computer that would (1) acquire position and orientation data from a Global Positioning System receiver and a microelectromechanical gyroscope, respectively; (2) perform all flight-control (including thruster-control) computations in real time; and (3) control video, tracking, power, and illumination systems.

High Capacity Battery Cell Flight Qualified

NASA Technical Reports Server (NTRS)

McKissock, Barbara I.

1997-01-01

The High Capacity Battery Cell project is an effort equally funded by the NASA Lewis Research Center and Hughes Space and Communications Company (a unit of Hughes Aircraft Company) to develop and flight qualify a higher capacity nickel hydrogen battery for continuing use on commercial spacecraft. The larger diameter, individual pressure vessel cell will provide approximately twice the power, while occupying the same volume, as the current state-of-the-art nickel hydrogen cell. These cells are also anticipated to reduce battery cost by 20 percent. The battery is currently booked for use on 26 spacecraft, with the first flight scheduled in 1997. A strong requirement for batteries with higher power levels (6 to 12 kW), long life, and reduced cost was identified in studies of the needs of commercial communications spacecraft. With the design developed in this effort, the higher power level was accommodated without having to modify the rest of the existing spacecraft bus. This design scaled-up the existing state-of-the-art nickel hydrogen battery cell from a 3.5-in., 50-Ahr cell to a 5.5-in., 350-Ahr cell. An improvement in cycle life was also achieved by the use of the 26-percent KOH electrolyte design developed by NASA Lewis. The cell design was completed, and flight batteries were built and flight qualified by Hughes Space and Communications Company with input from NASA Lewis. Two batteries were shipped in September 1996 to undergo life cycle testing under the purview of NASA Lewis.

Trajectory Design from GTO to Near-Equatorial Lunar Orbit for the Dark Ages Radio Explorer (DARE) Spacecraft

NASA Technical Reports Server (NTRS)

Genova, Anthony L.; Yang Yang, Fan; Perez, Andres Dono; Galal, Ken F.; Faber, Nicolas T.; Mitchell, Scott; Landin, Brett; Burns, Jack O.

2015-01-01

The trajectory design for the Dark Ages Radio Explorer (DARE) mission concept involves launching the DARE spacecraft into a geosynchronous transfer orbit (GTO) as a secondary payload. From GTO, the spacecraft then transfers to a lunar orbit that is stable (i.e., no station-keeping maneuvers are required with minimum perilune altitude always above 40 km) and allows for more than 1,000 cumulative hours for science measurements in the radio-quiet region located on the lunar farside.

Spacecraft load, design and test philosophies

NASA Technical Reports Server (NTRS)

Wada, B. K.

1986-01-01

The development of spacecraft loads, design and test philosophies at the Jet Propulsion Laboratory (JPL) during the past 25 years is presented. Examples from the JPL's Viking, Voyager and Galileo spacecraft are used to explain the changes in philosophy necessary to meet the program requirements with a reduction in cost and schedule. Approaches to validate mathematical models of large structures which can't be ground tested as an overall system because of size and/or adverse effects of terrestrial conditions such as gravity are presented.

An Application of Linear Covariance Analysis to the Design of Responsive Near-Rendezvous Missions

DTIC Science & Technology

2007-06-01

accurately before making large ma- neuvers. A fifth type of error is maneuver knowledge error (MKER). This error accounts for how well a spacecraft is able...utilized due in a large part to the cost of designing and launching spacecraft , in a market where currently there are not many options for launching...is then ordered to fire its thrusters to increase its orbital altitude to 800 km. Before the maneuver the spacecraft is moving with some velocity, V

Environmental parameters of shuttle support for life sciences experiments

NASA Technical Reports Server (NTRS)

Waligora, J. M.

1976-01-01

The environments provided by the Orbiter vehicle and by the Spacelab will differ substantially from the environment provided by prior spacecraft. The specific design limits for each environmental parameter and expected operating characteristics are presented for both the Orbiter and the Spacelab. The environments are compared with those of earlier spacecraft and with the normal earth laboratory. Differences between the spacecraft environments and the normal laboratory environment and the impact of these differences on experiments and equipment design are discussed.

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

NASA Technical Reports Server (NTRS)

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.

System concepts and design examples for optical communication with planetary spacecraft

NASA Astrophysics Data System (ADS)

Lesh, James R.

Systems concepts for optical communication with future deep-space (planetary) spacecraft are described. These include not only the optical transceiver package aboard the distant spacecraft, but the earth-vicinity optical-communications receiving station as well. Both ground-based, and earth-orbiting receivers are considered. Design examples for a number of proposed or potential deep-space missions are then presented. These include an orbital mission to Saturn, a Lander and Rover mission to Mars, and an astronomical mission to a distance of 1000 astronomical units.

Computer memory power control for the Galileo spacecraft

NASA Technical Reports Server (NTRS)

Detwiler, R. C.

1983-01-01

The developmental history, major design drives, and final topology of the computer memory power system on the Galileo spacecraft are described. A unique method of generating memory backup power directly from the fault current drawn during a spacecraft power overload or fault condition allows this system to provide continuous memory power. This concept provides a unique solution to the problem of volatile memory loss without the use of a battery of other large energy storage elements usually associated with uninterrupted power supply designs.

Spacecraft System Integration and Test: SSTI Lewis critical design audit

NASA Technical Reports Server (NTRS)

Brooks, R. P.; Cha, K. K.

1995-01-01

The Critical Design Audit package is the final detailed design package which provides a comprehensive description of the SSTI mission. This package includes the program overview, the system requirements, the science and applications activities, the ground segment development, the assembly, integration and test description, the payload and technology demonstrations, and the spacecraft bus subsystems. Publication and presentation of this document marks the final requirements and design freeze for SSTI.

Interplanetary charged particle models (1974). [and the effects of cosmic exposure upon spacecraft and spacecraft components

NASA Technical Reports Server (NTRS)

Divine, N.

1975-01-01

The design of space vehicles for operation in interplanetary space is given, based on descriptions of solar wind, solar particle events, and galactic cosmic rays. A state-of-the-art review is presented and design criteria are developed from experiment findings aboard interplanetary and high-altitude earth-orbiting spacecraft. Solar cells were found to be particularly sensitive. Solar protons may also impact the reliability of electric propulsion systems and spacecraft surfaces, as well as causing interference, detector saturation, and spurious signals. Galactic cosmic-ray impact can lead to similar electronic failure and interference and may register in photographic films and other emulsions. It was concluded that solar wind electron measurements might result from differential charging when shadowed portions of the spacecraft acquired a negative charge from electron impact.

Applying Contamination Modelling to Spacecraft Propulsion Systems Designs and Operations

NASA Technical Reports Server (NTRS)

Chen, Philip T.; Thomson, Shaun; Woronowicz, Michael S.

2000-01-01

Molecular and particulate contaminants generated from the operations of a propulsion system may impinge on spacecraft critical surfaces. Plume depositions or clouds may hinder the spacecraft and instruments from performing normal operations. Firing thrusters will generate both molecular and particulate contaminants. How to minimize the contamination impact from the plume becomes very critical for a successful mission. The resulting effect from either molecular or particulate contamination of the thruster firing is very distinct. This paper will discuss the interconnection between the functions of spacecraft contamination modeling and propulsion system implementation. The paper will address an innovative contamination engineering approach implemented from the spacecraft concept design, manufacturing, integration and test (I&T), launch, to on- orbit operations. This paper will also summarize the implementation on several successful missions. Despite other contamination sources, only molecular contamination will be considered here.

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

NASA Technical Reports Server (NTRS)

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.

Investigation of crew motion disturbances on Skylab-Experiment T-013. [for future manned spacecraft design

NASA Technical Reports Server (NTRS)

Conway, B. A.

1974-01-01

Astronaut crew motions can produce some of the largest disturbances acting on a manned spacecraft which can affect vehicle attitude and pointing. Skylab Experiment T-013 was developed to investigate the magnitude and effects of some of these disturbances on the Skylab spacecraft. The methods and techniques used to carry out this experiment are discussed, and preliminary results of data analysis presented. Initial findings indicate that forces on the order of 300 N were exerted during vigorous soaring activities, and that certain experiment activities produced spacecraft angular rate excursions 0.03 to 0.07 deg/sec. Results of Experiment T-013 will be incorporated into mathematical models of crew-motion disturbances, and are expected to be of significant aid in the sizing, design, and analysis of stabilization and control systems for future manned spacecraft.

Description and User Instructions for the Quaternion_to_Orbit_v3 Software

NASA Technical Reports Server (NTRS)

Strekalov, Dmitry V.; Kruizinga, Gerhard L.; Paik, Meegyeong; Yuan, Dah-Ning; Asmar, Sami W.

2012-01-01

For a given inertial frame of reference, the software combines the spacecraft orbits with the spacecraft attitude quaternions, and rotates the body-fixed reference frame of a particular spacecraft to the inertial reference frame. The conversion assumes that the two spacecraft are aligned with respect to the mutual line of sight, with a parameterized time tag. The software is implemented in Python and is completely open source. It is very versatile, and may be applied under various circumstances and for other related purposes. Based on the solid linear algebra analysis, it has an extra option for compensating the linear pitch. This software has been designed for simulation of the calibration maneuvers performed by the two spacecraft comprising the GRAIL mission to the Moon, but has potential use for other applications. In simulations of formation flights, one needs to coordinate the spacecraft orbits represented in an appropriate inertial reference frame and the spacecraft attitudes. The latter are usually given as the time series of quaternions rotating the body-fixed reference frame of a particular spacecraft to the inertial reference frame. It is often desirable to simulate the same maneuver for different segments of the orbit. It is also useful to study various maneuvers that could be performed at the same orbit segment. These two lines of study are more timeand labor-efficient if the attitude and orbit data are generated independently, so that the part of the data that has not been changed can be recycled in the course of multiple simulations.

KSC-98pc1229

NASA Image and Video Library

1998-10-03

KENNEDY SPACE CENTER, FLA. -- In the Spacecraft Assembly and Encapsulation Facility-2 (SAEF-2), the top of the Mars Polar Lander is removed to prepare the Lander for testing, including a functional test of the science instruments and the basic spacecraft subsystems. The Mars Polar Lander is targeted for launch from Cape Canaveral Air Station aboard a Delta II rocket on Jan. 3, 1999. The solar-powered spacecraft is designed to touch down on the Martian surface near the northern-most boundary of the south pole in order to study the water cycle there. The lander also will help scientists learn more about climate change and current resources on Mars, studying such things as frost, dust, water vapor and condensates in the Martian atmosphere

Dynamic analysis of Apollo-Salyut/Soyuz docking

NASA Technical Reports Server (NTRS)

Schliesing, J. A.

1972-01-01

The use of a docking-system computer program in analyzing the dynamic environment produced by two impacting spacecraft and the attitude control systems is discussed. Performance studies were conducted to determine the mechanism load and capture sensitivity to parametric changes in the initial impact conditions. As indicated by the studies, capture latching is most sensitive to vehicle angular-alinement errors and is least sensitive to lateral-miss error. As proved by load-sensitivity studies, peak loads acting on the Apollo spacecraft are considerably lower than the Apollo design-limit loads.

ARC-2009-ACD09-0141-006

NASA Image and Video Library

2009-07-16

A plaque presented to Harvey Allen in recognition of his outstanding solution of the reentry heating problem which has been indispensable to the design of the Mercury, Gemini, and Apollo spacecraft (Manned Spacecraft Center, November 14, 1968) Plaque contains samples of tested materials and models of spacecraft.

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

NASA Technical Reports Server (NTRS)

Boyer, Jeffrey S.

1994-01-01

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.

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

NASA Astrophysics Data System (ADS)

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.

Model of spacecraft atomic oxygen and solar exposure microenvironments

NASA Technical Reports Server (NTRS)

Bourassa, R. J.; Pippin, H. G.

1993-01-01

Computer models of environmental conditions in Earth orbit are needed for the following reasons: (1) derivation of material performance parameters from orbital test data, (2) evaluation of spacecraft hardware designs, (3) prediction of material service life, and (4) scheduling spacecraft maintenance. To meet these needs, Boeing has developed programs for modeling atomic oxygen (AO) and solar radiation exposures. The model allows determination of AO and solar ultraviolet (UV) radiation exposures for spacecraft surfaces (1) in arbitrary orientations with respect to the direction of spacecraft motion, (2) overall ranges of solar conditions, and (3) for any mission duration. The models have been successfully applied to prediction of experiment environments on the Long Duration Exposure Facility (LDEF) and for analysis of selected hardware designs for deployment on other spacecraft. The work on these models has been reported at previous LDEF conferences. Since publication of these reports, a revision has been made to the AO calculation for LDEF, and further work has been done on the microenvironments model for solar exposure.

New NASA SEE LEO Spacecraft Charging Design Guidelines: How to Survive in LEO Rather Than GEO

NASA Technical Reports Server (NTRS)

Ferguson, Dale C.; Hillard, G. Barry

2003-01-01

It has been almost two solar cycles since the 1984 GEO Guidelines of Purvis, Garrett, Whittlesey, and Stevens were published. In that time, interest in high voltage LEO systems has increased. Correct and conventional wisdom has been that LEO conditions are sufficiently different from GEO that the GEO Guidelines (and other GEO and POLAR documents produced since then) should not be used for LEO spacecraft. Because of significant recent GEO spacecraft failures that have been shown in ground testing to be likely to also occur on LEO spacecraft, the SEE program commissioned the production of the new LEO Spacecraft Charging Design Guidelines. Now available in CD-ROM form, the LEO Guidelines highlight mitigation techniques to prevent spacecraft arcing on LEO solar arrays and other systems. We compare and contrast the mitigation techniques for LEO and GEO in this paper. We also discuss the extensive bibliography included in the LEO Guidelines, so results can be found in their primary sources.

New NASA SEE LEO Spacecraft Charging Design Guidelines: How to Survive in LEO Rather than GEO

NASA Technical Reports Server (NTRS)

Ferguson, Dale C.; Hillard, G. Barry

2004-01-01

It has been almost two solar cycles since the GEO Guidelines of Purvis et al (1984) were published. In that time, interest in high voltage LEO systems has increased. The correct and conventional wisdom has been that LEO conditions are sufficiently different from GEO that the GEO Guidelines (and other GEO and POLAR documents produced since then) should not be used for LEO spacecraft. Because of significant recent GEO spacecraft failures that have been shown in ground testing to be likely to also occur on LEO spacecraft, the SEE program commissioned the production of the new LEO Spacecraft Charging Design Guidelines (hereafter referred to as the LEO Guidelines). Now available in CD-ROM form, the LEO Guidelines highlight mitigation techniques to prevent spacecraft arcing on LEO solar arrays and other systems. We compare and contrast the mitigation techniques for LEO and GEO in this paper. We also discuss the extensive bibliography included in the LEO Guidelines, so results can be found in their primary sources.

Design method of combined protective against space environmental effects on spacecraft

NASA Astrophysics Data System (ADS)

Shen, Zicai; Gong, Zizheng; Ding, Yigang; Liu, Yuming; Liu, Yenan

2016-01-01

During its projected extended stay in LEO, spacecraft will encounter many environmental factors including energetic particles, ultraviolet radiation, atomic oxygen, and space debris and meteoroids, together with some induced environments such as contamination and discharging. These space environments and their effects have threat to the reliability and lifetime of spacecraft. So, it is important to give a combined design against the threat from space environments and their effects. The space environments and effects are reviewed in this paper firstly. Secondly, the design process and method against space environments are discussed. At last, some advices about protective structure and materials are proposed.

Synchronized Lunar Pole Impact Plume Sample Return Trajectory Design

NASA Technical Reports Server (NTRS)

Genova, Anthony L.; Foster, Cyrus; Colaprete, Tony

2016-01-01

The presented trajectory design enables two maneuverable spacecraft launched onto the same trans-lunar injection trajectory to coordinate a steep impact of a lunar pole and subsequent sample return of the ejecta plume to Earth. To demonstrate this concept, the impactor is assumed to use the LCROSS missions trajectory and spacecraft architecture, thus the permanently-shadowed Cabeus crater on the lunar south pole is assumed as the impact site. The sample-return spacecraft is assumed to be a CubeSat that requires a complimentary trajectory design that avoids lunar impact after passing through the ejecta plume to enable sample-return to Earth via atmospheric entry.

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

NASA Technical Reports Server (NTRS)

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

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.

Trajectory design for a cislunar CubeSat leveraging dynamical systems techniques: The Lunar IceCube mission

NASA Astrophysics Data System (ADS)

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

2018-03-01

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, is expected to be deployed from the upcoming Exploration Mission-1 vehicle. 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.

Design and Implementation of the ARTEMIS Lunar Transfer Using Multi-Body Dynamics

NASA Technical Reports Server (NTRS)

Folta, David; Woodard, Mark; Sweetser, Theodore; Broschart, Stephen B.; Cosgrove, Daniel

2011-01-01

The use of multi-body dynamics to design the transfer of spacecraft from Earth elliptical orbits to the Earth-Moon libration (L(sub 1) and L(sub 2)) orbits has been successfully demonstrated by the Acceleration Reconnection and Turbulence and Electrodynamics of the Moon's Interaction with the Sun (ARTEMIS) mission. Operational support of the two ARTEMIS spacecraft is a final step in the realization of a design process that can be used to transfer spacecraft with restrictive operational constraints and fuel limitations. The focus of this paper is to describe in detail the processes and implementation of this successful approach.

Studying Spacecraft Charging via Numerical Simulations

NASA Astrophysics Data System (ADS)

Delzanno, G. L.; Moulton, D.; Meierbachtol, C.; Svyatskiy, D.; Vernon, L.

2015-12-01

The electrical charging of spacecraft due to bombarding charged particles can affect their performance and operation. We study this charging using CPIC; a particle-in-cell code specifically designed for studying plasma-material interactions [1]. CPIC is based on multi-block curvilinear meshes, resulting in near-optimal computational performance while maintaining geometric accuracy. Relevant plasma parameters are imported from the SHIELDS framework (currently under development at LANL), which simulates geomagnetic storms and substorms in the Earth's magnetosphere. Simulated spacecraft charging results of representative Van Allen Probe geometries using these plasma parameters will be presented, along with an overview of the code. [1] G.L. Delzanno, E. Camporeale, J.D. Moulton, J.E. Borovsky, E.A. MacDonald, and M.F. Thomsen, "CPIC: A Curvilinear Particle-In-Cell Code for Plasma-Material Interaction Studies," IEEE Trans. Plas. Sci., 41 (12), 3577 (2013).

MarcoPolo-R: Mission and Spacecraft Design

NASA Astrophysics Data System (ADS)

Peacocke, L.; Kemble, S.; Chapuy, M.; Scheer, H.

2013-09-01

The MarcoPolo-R mission is a candidate for the European Space Agency's medium-class Cosmic Vision programme, with the aim to obtain a 100 g sample of asteroid surface material and return it safely to the Earth. Astrium is one of two industrial contractors currently studying the mission to Phase A level, and the team has been working on the mission and spacecraft design since January 2012. Asteroids are some of the most primitive bodies in our solar system and are key to understanding the formation of the Earth, Sun and other planetary bodies. A returned sample would allow extensive analyses in the large laboratory-sized instruments here on Earth that are not possible with in-situ instruments. This analysis would also increase our understanding of the composition and structure of asteroids, and aid in plans for asteroid deflection techniques. In addition, the mission would be a valuable precursor for missions such as Mars Sample Return, demonstrating a high speed Earth re-entry and hard landing of an entry capsule. Following extensive mission analysis of both the baseline asteroid target 1996 FG3 and alternatives, a particularly favourable trajectory was found to the asteroid 2008 EV5 resulting in a mission duration of 4.5 to 6 years. In October 2012, the MarcoPolo-R baseline target was changed to 2008 EV5 due to its extremely primitive nature, which may pre-date the Sun. This change has a number of advantages: reduced DeltaV requirements, an orbit with a more benign thermal environment, reduced communications distances, and a reduced complexity propulsion system - all of which simplify the spacecraft design significantly. The single spacecraft would launch between 2022 and 2024 on a Soyuz-Fregat launch vehicle from Kourou. Solar electric propulsion is necessary for the outward and return transfers due to the DeltaV requirements, to minimise propellant mass. Once rendezvous with the asteroid is achieved, an observation campaign will begin to characterise the asteroid properties and map the surface in detail. Five potential sampling sites will be selected and closely observed in a local characterisation phase, leading to a single preferred sampling site being identified. The baseline instruments are a Narrow Angle Camera, a Mid-Infrared Spectrometer, a Visible Near-Infrared Spectrometer, a Radio Science Experiment, and a Close-up Camera. For the sampling phase, the spacecraft will perform a touch-and-go manoeuvre. A boom with a sampling mechanism at the end will be deployed, and the spacecraft will descend using visual navigation to touch the asteroid for some seconds. The rotary brush sampling mechanism will be activated on touchdown to obtain a good quality sample comprising regolith dust and pebbles. Low touchdown velocities and collision avoidance are critical at this point to prevent damage to the spacecraft and solar arrays. The spacecraft will then move away, returning to a safe orbit, and the sample will be transferred to an Earth Re-entry Capsule. After a final post-sampling characterisation campaign, the spacecraft will perform the return transfer to Earth. The Earth Re-entry Capsule will be released to directly enter the Earth's atmosphere, and is designed to survive a hard landing with no parachute deceleration. Once recovered, the asteroid sample would be extracted in a sample curation facility in preparation for the full analysis campaign. This presentation will describe Astrium's MarcoPolo-R mission and spacecraft design, with a focus on the innovative aspects of the design.

Spacecraft dielectric material properties and spacecraft charging

NASA Technical Reports Server (NTRS)

Frederickson, A. R.; Wall, J. A.; Cotts, D. B.; Bouquet, F. L.

1986-01-01

The physics of spacecraft charging is reviewed, and criteria for selecting and testing semiinsulating polymers (SIPs) to avoid charging are discussed and illustrated. Chapters are devoted to the required properties of dielectric materials, the charging process, discharge-pulse phenomena, design for minimum pulse size, design to prevent pulses, conduction in polymers, evaluation of SIPs that might prevent spacecraft charging, and the general response of dielectrics to space radiation. SIPs characterized include polyimides, fluorocarbons, thermoplastic polyesters, poly(alkanes), vinyl polymers and acrylates, polymers containing phthalocyanine, polyacene quinones, coordination polymers containing metal ions, conjugated-backbone polymers, and 'metallic' conducting polymers. Tables summarizing the results of SIP radiation tests (such as those performed for the NASA Galileo Project) are included.

Developing Sustainable Spacecraft Water Management Systems

NASA Technical Reports Server (NTRS)

Thomas, Evan A.; Klaus, David M.

2009-01-01

It is well recognized that water handling systems used in a spacecraft are prone to failure caused by biofouling and mineral scaling, which can clog mechanical systems and degrade the performance of capillary-based technologies. Long duration spaceflight applications, such as extended stays at a Lunar Outpost or during a Mars transit mission, will increasingly benefit from hardware that is generally more robust and operationally sustainable overtime. This paper presents potential design and testing considerations for improving the reliability of water handling technologies for exploration spacecraft. Our application of interest is to devise a spacecraft wastewater management system wherein fouling can be accommodated by design attributes of the management hardware, rather than implementing some means of preventing its occurrence.

Derivation of Design Loads and Random Vibration specifications for Spacecraft Instruments and Sub-Units

NASA Astrophysics Data System (ADS)

Fransen, S.; Yamawaki, T.; Akagi, H.; Eggens, M.; van Baren, C.

2014-06-01

After a first estimation based on statistics, the design loads for instruments are generally estimated by coupled spacecraft/instrument sine analysis once an FE-model of the spacecraft is available. When the design loads for the instrument have been derived, the next step in the process is to estimate the random vibration environment at the instrument base and to compute the RMS load at the centre of gravity of the instrument by means of vibro-acoustic analysis. Finally the design loads of the light-weight sub-units of the instrument can be estimated through random vibration analysis at instrument level, taking into account the notches required to protect the instrument interfaces in the hard- mounted random vibration test. This paper presents the aforementioned steps of instrument and sub-units loads derivation in the preliminary design phase of the spacecraft and identifies the problems that may be encountered in terms of design load consistency between low-frequency and high-frequency environments. The SpicA FAR-infrared Instrument (SAFARI) which is currently developed for the Space Infrared Telescope for Cosmology and Astrophysics (SPICA) will be used as a guiding example.

CisLunar Habitat Internal Architecture Design Criteria

NASA Technical Reports Server (NTRS)

Jones, R.; Kennedy, K.; Howard, R.; Whitmore, M.; Martin, C.; Garate, J.

2017-01-01

BACKGROUND: In preparation for human exploration to Mars, there is a need to define the development and test program that will validate deep space operations and systems. In that context, a Proving Grounds CisLunar habitat spacecraft is being defined as the next step towards this goal. This spacecraft will operate differently from the ISS or other spacecraft in human history. The performance envelope of this spacecraft (mass, volume, power, specifications, etc.) is being defined by the Future Capabilities Study Team. This team has recognized the need for a human-centered approach for the internal architecture of this spacecraft and has commissioned a CisLunar Phase-1 Habitat Internal Architecture Study Team to develop a NASA reference configuration, providing the Agency with a "smart buyer" approach for future acquisition. THE CISLUNAR HABITAT INTERNAL ARCHITECTURE STUDY: Overall, the CisLunar Habitat Internal Architecture study will address the most significant questions and risks in the current CisLunar architecture, habitation, and operations concept development. This effort is achieved through definition of design criteria, evaluation criteria and process, design of the CisLunar Habitat Phase-1 internal architecture, and the development and fabrication of internal architecture concepts combined with rigorous and methodical Human-in-the-Loop (HITL) evaluations and testing of the conceptual innovations in a controlled test environment. The vision of the CisLunar Habitat Internal Architecture Study is to design, build, and test a CisLunar Phase-1 Habitat Internal Architecture that will be used for habitation (e.g. habitability and human factors) evaluations. The evaluations will mature CisLunar habitat evaluation tools, guidelines, and standards, and will interface with other projects such as the Advanced Exploration Systems (AES) Program integrated Power, Avionics, Software (iPAS), and Logistics for integrated human-in-the-loop testing. The mission of the CisLunar Habitat Internal Architecture Study is to become a forcing function to establish a common understanding of CisLunar Phase-1 Habitation Internal Architecture design criteria, processes, and tools. The scope of the CisLunar Habitat Internal Architecture study is to design, develop, demonstrate, and evaluate a Phase-1 CisLunar Habitat common module internal architecture based on design criteria agreed to by NASA, the International Partners, and Commercial Exploration teams. This task is to define the CisLunar Phase-1 Internal Architecture Government Reference Design, assist NASA in becoming a "smart buyer" for Phase-1 Habitat Concepts, and ultimately to derive standards and requirements from the Internal Architecture Design Process. The first step was to define a Habitat Internal Architecture Design Criteria and create a structured philosophy to be used by design teams as a filter by which critical aspects of consideration would be identified for the purpose of organizing and utilizing interior spaces. With design criteria in place, the team will develop a series of iterative internal architecture concept designs which will be assessed by means of an evaluation criteria and process. These assessments will successively drive and refine the design, leading to the combination and down-selection of design concepts. A single refined reference design configuration will be developed into in a medium-to-high fidelity mockup. A multi-day human-in-the-loop mission test will fully evaluate the reference design and validate its configuration. Lessons learned from the design and evaluation will enable the team to identify appropriate standards for Phase-1 CisLunar Habitat Internal Architecture and will enable NASA to develop derived requirements in support of maturing CisLunar Habitation capabilities. This paper will describe the criteria definition process, workshop event, and resulting CisLunar Phase-1 Habitat Internal Architecture Design Criteria.

High precision active nutation control for a flexible momentum biased spacecraft

NASA Technical Reports Server (NTRS)

Laskin, R. A.; Kopf, E. H.

1984-01-01

The controller design for the Solar Dynamics Observatory (SDO) is presented. SDO is a momentum biased spacecraft with three flexible appendages. Its primary scientific instrument, the solar oscillations imager (SOI), is rigidly attached to the spacecraft bus and has arc-second pointing requirements. Meeting these requirements necessitates the use of an active nutation controller (ANC) which is here mechanized with a small reaction wheel oriented along a bus transverse axis. The ANC does its job by orchestrating the transfer of angular momentum out of the bus transverse axes and into the momentum wheel. A simulation study verifies that the controller provides quick, stable, and accurate response.

Salt materials testing for a spacecraft adiabatic demagnetization refrigerator

NASA Technical Reports Server (NTRS)

Savage, M. L.; Kittel, P.; Roellig, T.

1990-01-01

As part of a technology development effort to qualify adiabatic demagnetization refrigerators for use in a NASA spacecraft, such as the Space Infrared Telescope Facility, a study of low temperature characteristics, heat capacity and resistance to dehydration was conducted for different salt materials. This report includes results of testing with cerrous metaphosphate, several synthetic rubies, and chromic potassium alum (CPA). Preliminary results show that CPA may be suitable for long-term spacecraft use, provided that the salt is property encapsulated. Methods of salt pill construction and testing for all materials are discussed, as well as reliability tests. Also, the temperature regulation scheme and the test cryostat design are briefly discussed.

Xenia Spacecraft Study

NASA Technical Reports Server (NTRS)

Hopkins, Randy

2009-01-01

This slide presentation reviews the proposed design for the Xenia mission spacecraft. The goal of this study is to perform a mission concept study for the mission. Included in this study are: the overall ground rules and assumptions (GR&A), a mission analysis, the configuration, the mass properties, the guidance, Navigation and control, the proposed avionics, the power system, the thermal protection system, the propulsion system, and the proposed structures. Conclusions from the study indicate that the observatory fits within the Falcon 9 mass and volume envelope for launching from Omelek, the pointing, slow slewing, and fast slewing requirements and the thermal requirements are met.

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

NASA Technical Reports Server (NTRS)

Dakermanji, George; Burns, Michael; Lee, Leonine; Lyons, John; Kim, David; Spitzer, Thomas; Kercheval, Bradford

2016-01-01

The Global Precipitation Measurement (GPM) spacecraft was jointly developed by National Aeronautics and Space Administration (NASA) and Japan Aerospace Exploration Agency (JAXA). It is a Low Earth Orbit (LEO) spacecraft launched on February 27, 2014. The spacecraft is in a circular 400 Km altitude, 65 degrees inclination nadir pointing orbit with a three year basic mission life. The solar array consists of two sun tracking wings with cable wraps. The panels are populated with triple junction cells of nominal 29.5% efficiency. One axis is canted by 52 degrees to provide power to the spacecraft at high beta angles. 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 x 84p batteries operated in parallel as a single battery. 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.

FASTSAT a Mini-Satellite Mission...A Way Ahead

NASA Technical Reports Server (NTRS)

Boudreaux, Mark; Pearson, Steve; Casas, Joseph

2012-01-01

The Fast Affordable Science and Technology Spacecraft (FASTSAT) is a mini-satellite weighing less than 150 kg. FASTSAT was developed as government-industry collaborative research and development flight project targeting rapid access to space to provide an alternative, low cost platform for a variety of scientific, research, and technology payloads. The initial spacecraft was designed to carry six instruments and launch as a secondary rideshare payload. This design approach greatly reduced overall mission costs while maximizing the on-board payload accommodations. FASTSAT was designed from the ground up to meet a challenging short schedule using modular components with a flexible, configurable layout to enable a broad range of payloads at a lower cost and shorter timeline than scaling down a more complex spacecraft. The integrated spacecraft along with its payloads were readied for launch 15 months from authority to proceed. As an ESPA-class spacecraft, FASTSAT is compatible with many different launch vehicles, including Minotaur I, Minotaur IV, Delta IV, Atlas V, Pegasus, Falcon 1/1e, and Falcon 9. These vehicles offer an array of options for launch sites and provide for a variety of rideshare possibilities.

Line-of-sight pointing accuracy/stability analysis and computer simulation for small spacecraft

NASA Astrophysics Data System (ADS)

Algrain, Marcelo C.; Powers, Richard M.

1996-06-01

This paper presents a case study where a comprehensive computer simulation is developed to determine the driving factors contributing to spacecraft pointing accuracy and stability. The simulation is implemented using XMATH/SystemBuild software from Integrated Systems, Inc. The paper is written in a tutorial manner and models for major system components are described. Among them are spacecraft bus, attitude controller, reaction wheel assembly, star-tracker unit, inertial reference unit, and gyro drift estimators (Kalman filter). THe predicted spacecraft performance is analyzed for a variety of input commands and system disturbances. The primary deterministic inputs are desired attitude angles and rate setpoints. The stochastic inputs include random torque disturbances acting on the spacecraft, random gyro bias noise, gyro random walk, and star-tracker noise. These inputs are varied over a wide range to determine their effects on pointing accuracy and stability. The results are presented in the form of trade-off curves designed to facilitate the proper selection of subsystems so that overall spacecraft pointing accuracy and stability requirements are met.

Exterior spacecraft subsystem protective shielding analysis and design

NASA Technical Reports Server (NTRS)

Schonberg, William P.; Taylor, Roy A.

1990-01-01

All spacecraft are susceptible to impacts by meteoroids and pieces of orbiting space debris. An effective mechanism is developed to protect external spacecraft subsystems against damage by ricochet particles formed during such impacts. Equations and design procedures for protective shield panels are developed based on observed ricochet phenomena and calculated ricochet particle sizes and speeds. It is found that the diameter of the most damaging ricochet debris particle can be as large as 40 percent of the original project tile diameter, and can travel at speeds between 24 and 36 percent of the original projectile impact velocity. Panel dimensions are shown to be strongly dependent on their inclination to the impact velocity vector and on their distribution around a spacecraft module. It is concluded that obliquity effects of high-speed impacts must be considered in the design of any structure exposed to the meteoroid and space debris environment.

Multipurpose hardened spacecraft insulation

NASA Technical Reports Server (NTRS)

Steimer, Carlos H.

1990-01-01

A Multipurpose Hardened Spacecraft Multilayer Insulation (MLI) system was developed and implemented to meet diverse survivability and performance requirements. Within the definition and confines of a MLI assembly (blanket), the design: (1) provides environmental protection from natural and induced nuclear, thermal, and electromagnetic radiation; (2) provides adequate electrostatic discharge protection for a geosynchronous satellite; (3) provides adequate shielding to meet radiated emission needs; and (4) will survive ascent differential pressure loads between enclosed volume and space. The MLI design is described which meets these requirements and design evolution and verification is discussed. The application is for MLI blankets which closeout the area between the laser crosslink subsystem (LCS) equipment and the DSP spacecraft cabin. Ancillary needs were implemented to ease installation at launch facility and to survive ascent acoustic and vibration loads. Directional venting accommodations were also incorporated to avoid contamination of LCS telescope, spacecraft sensors, and second surface mirrors (SSMs).

Control of flexible structures

NASA Technical Reports Server (NTRS)

Russell, R. A.

1985-01-01

The requirements for future space missions indicate that many of these spacecraft will be large, flexible, and in some applications, require precision geometries. A technology program that addresses the issues associated with the structure/control interactions for these classes of spacecraft is discussed. The goal of the NASA control of flexible structures technology program is to generate a technology data base that will provide the designer with options and approaches to achieve spacecraft performance such as maintaining geometry and/or suppressing undesired spacecraft dynamics. This technology program will define the appropriate combination of analysis, ground testing, and flight testing required to validate the structural/controls analysis and design tools. This work was motivated by a recognition that large minimum weight space structures will be required for many future missions. The tools necessary to support such design included: (1) improved structural analysis; (2) modern control theory; (3) advanced modeling techniques; (4) system identification; and (5) the integration of structures and controls.

Space Environments and Effects (SEE) Program: Spacecraft Charging Technology Development Activities

NASA Technical Reports Server (NTRS)

Kauffman, Billy; Hardage, Donna; Minor, Jody

2003-01-01

Reducing size and weight of spacecraft, along with demanding increased performance capabilities, introduces many uncertainties in the engineering design community on how materials and spacecraft systems will perform in space. The engineering design community is forever behind on obtaining and developing new tools and guidelines to mitigate the harmful effects of the space environment. Adding to this complexity is the continued push to use Commercial-off-the-shelf (COTS) microelectronics, potential usage of unproven technologies such as large solar sail structures and nuclear electric propulsion. In order to drive down these uncertainties, various programs are working together to avoid duplication, save what resources are available in this technical area and possess a focused agenda to insert these new developments into future mission designs. This paper will introduce the SEE Program, briefly discuss past and currently sponsored spacecraft charging activities and possible future endeavors.

Space Environments and Effects (SEE) Program: Spacecraft Charging Technology Development Activities

NASA Technical Reports Server (NTRS)

Kauffman, B.; Hardage, D.; Minor, J.

2004-01-01

Reducing size and weight of spacecraft, along with demanding increased performance capabilities, introduces many uncertainties in the engineering design community on how materials and spacecraft systems will perform in space. The engineering design community is forever behind on obtaining and developing new tools and guidelines to mitigate the harmful effects of the space environment. Adding to this complexity is the continued push to use Commercial-off-the-Shelf (COTS) microelectronics, potential usage of unproven technologies such as large solar sail structures and nuclear electric propulsion. In order to drive down these uncertainties, various programs are working together to avoid duplication, save what resources are available in this technical area and possess a focused agenda to insert these new developments into future mission designs. This paper will introduce the SEE Program, briefly discuss past and currently sponsored spacecraft charging activities and possible future endeavors.

Design consideration for a nuclear electric propulsion system

NASA Technical Reports Server (NTRS)

Phillips, W. M.; Pawlik, E. V.

1978-01-01

A study is currently underway to design a nuclear electric propulsion vehicle capable of performing detailed exploration of the outer-planets. Primary emphasis is on the power subsystem. Secondary emphasis includes integration into a spacecraft, and integration with the thrust subsystem and science package or payload. The results of several design iterations indicate an all-heat-pipe system offers greater reliability, elimination of many technology development areas and a specific weight of under 20 kg/kWe at the 400 kWe power level. The system is compatible with a single Shuttle launch and provides greater safety than could be obtained with designs using pumped liquid metal cooling. Two configurations, one with the reactor and power conversion forward on the spacecraft with the ion engines aft and the other with reactor, power conversion and ion engines aft were selected as dual baseline designs based on minimum weight, minimum required technology development and maximum growth potential and flexibility.

Evolutionary Optimization of a Quadrifilar Helical Antenna

NASA Technical Reports Server (NTRS)

Lohn, Jason D.; Kraus, William F.; Linden, Derek S.; Clancy, Daniel (Technical Monitor)

2002-01-01

Automated antenna synthesis via evolutionary design has recently garnered much attention in the research literature. Evolutionary algorithms show promise because, among search algorithms, they are able to effectively search large, unknown design spaces. NASA's Mars Odyssey spacecraft is due to reach final Martian orbit insertion in January, 2002. Onboard the spacecraft is a quadrifilar helical antenna that provides telecommunications in the UHF band with landed assets, such as robotic rovers. Each helix is driven by the same signal which is phase-delayed in 90 deg increments. A small ground plane is provided at the base. It is designed to operate in the frequency band of 400-438 MHz. Based on encouraging previous results in automated antenna design using evolutionary search, we wanted to see whether such techniques could improve upon Mars Odyssey antenna design. Specifically, a co-evolutionary genetic algorithm is applied to optimize the gain and size of the quadrifilar helical antenna. The optimization was performed in-situ in the presence of a neighboring spacecraft structure. On the spacecraft, a large aluminum fuel tank is adjacent to the antenna. Since this fuel tank can dramatically affect the antenna's performance, we leave it to the evolutionary process to see if it can exploit the fuel tank's properties advantageously. Optimizing in the presence of surrounding structures would be quite difficult for human antenna designers, and thus the actual antenna was designed for free space (with a small ground plane). In fact, when flying on the spacecraft, surrounding structures that are moveable (e.g., solar panels) may be moved during the mission in order to improve the antenna's performance.

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

NASA Technical Reports Server (NTRS)

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.

Relating MBSE to Spacecraft Development: A NASA Pathfinder

NASA Technical Reports Server (NTRS)

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.

Estimating the Reliability of Electronic Parts in High Radiation Fields

NASA Technical Reports Server (NTRS)

Everline, Chester; Clark, Karla; Man, Guy; Rasmussen, Robert; Johnston, Allan; Kohlhase, Charles; Paulos, Todd

2008-01-01

Radiation effects on materials and electronic parts constrain the lifetime of flight systems visiting Europa. Understanding mission lifetime limits is critical to the design and planning of such a mission. Therefore, the operational aspects of radiation dose are a mission success issue. To predict and manage mission lifetime in a high radiation environment, system engineers need capable tools to trade radiation design choices against system design and reliability, and science achievements. Conventional tools and approaches provided past missions with conservative designs without the ability to predict their lifetime beyond the baseline mission.This paper describes a more systematic approach to understanding spacecraft design margin, allowing better prediction of spacecraft lifetime. This is possible because of newly available electronic parts radiation effects statistics and an enhanced spacecraft system reliability methodology. This new approach can be used in conjunction with traditional approaches for mission design. This paper describes the fundamentals of the new methodology.

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

NASA Technical Reports Server (NTRS)

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.

Global precipitation measurement (GPM) mission core spacecraft systems engineering challenges

NASA Astrophysics Data System (ADS)

Bundas, David J.; O'Neill, Deborah; Rhee, Michael; Feild, Thomas; Meadows, Gary; Patterson, Peter

2006-09-01

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.

Simplified spacecraft vulnerability assessments at component level in early design phase at the European Space Agency's Concurrent Design Facility

NASA Astrophysics Data System (ADS)

Kempf, Scott; Schäfer, Frank K.; Cardone, Tiziana; Ferreira, Ivo; Gerené, Sam; Destefanis, Roberto; Grassi, Lilith

2016-12-01

During recent years, the state-of-the-art risk assessment of the threat posed to spacecraft by micrometeoroids and space debris has been expanded to the analysis of failure modes of internal spacecraft components. This method can now be used to perform risk analyses for satellites to assess various failure levels - from failure of specific sub-systems to catastrophic break-up. This new assessment methodology is based on triple-wall ballistic limit equations (BLEs), specifically the Schäfer-Ryan-Lambert (SRL) BLE, which is applicable for describing failure threshold levels for satellite components following a hypervelocity impact. The methodology is implemented in the form of the software tool Particle Impact Risk and vulnerability Analysis Tool (PIRAT). During a recent European Space Agency (ESA) funded study, the PIRAT functionality was expanded in order to provide an interface to ESA's Concurrent Design Facility (CDF). The additions include a geometry importer and an OCDT (Open Concurrent Design Tool) interface. The new interface provides both the expanded geometrical flexibility, which is provided by external computer aided design (CAD) modelling, and an ease of import of existing data without the need for extensive preparation of the model. The reduced effort required to perform vulnerability analyses makes it feasible for application during early design phase, at which point modifications to satellite design can be undertaken with relatively little extra effort. The integration of PIRAT in the CDF represents the first time that vulnerability analyses can be performed in-session in ESA's CDF and the first time that comprehensive vulnerability studies can be applied cost-effectively in early design phase in general.

Trajectory Design Considerations for Small Body Touch-and-Go

NASA Technical Reports Server (NTRS)

Wallace, Mark; Broschart, Stephen; Bonfiglio, Eugene; Bhaskharan, Shyam; Cangahuala, Alberto

2011-01-01

Outline: (1) Trajectory Description (2) Design Drivers: (2a) Dynamics (2b) Environment (2c) Spacecraft and Ground and System Capabilities (2d) Mission Objectives (3) Design Choices (4) Historical Precedents (5) Case Studies. What is Touch-and-Go (TAG)? (1) Descent to the surface (2) Brief contact (3) Ascends to a safe distance

Characterization of spacecraft and environmental disturbances on a SmallSat

NASA Technical Reports Server (NTRS)

Johnson, Thomas A.; Nguyen, Dung Phu Chi; Cuda, Vince; Freesland, Doug

1994-01-01

The objective of this study is to model the on-orbit vibration environment encountered by a SmallSat. Vibration control issues are common to the Earth observing, imaging, and microgravity communities. A spacecraft may contain dozens of support systems and instruments each a potential source of vibration. The quality of payload data depends on constraining vibration so that parasitic disturbances do not affect the payload's pointing or microgravity requirement. In practice, payloads are designed incorporating existing flight hardware in many cases with nonspecific vibration performance. Thus, for the development of a payload, designers require a thorough knowledge of existing mechanical devices and their associated disturbance levels. This study evaluates a SmallSat mission and seeks to answer basic questions concerning on-orbit vibration. Payloads were considered from the Earth observing, microgravity, and imaging communities. Candidate payload requirements were matched to spacecraft bus resources of present day SmallSats. From the set of candidate payloads, the representative payload GLAS (Geoscience Laser Altimeter System) was selected. The requirements of GLAS were considered very stringent for the 150 - 500 kg class of payloads. Once the payload was selected, a generic SmallSat was designed in order to accommodate the payload requirements (weight, size, power, etc.). This study seeks to characterize the on-orbit vibration environment of a SmallSat designed for this type of mission and to determine whether a SmallSat can provide the precision pointing and jitter control required for earth observing payloads.

Best Practices for Reliable and Robust Spacecraft Structures

NASA Technical Reports Server (NTRS)

Raju, Ivatury S.; Murthy, P. L. N.; Patel, Naresh R.; Bonacuse, Peter J.; Elliott, Kenny B.; Gordon, S. A.; Gyekenyesi, J. P.; Daso, E. O.; Aggarwal, P.; Tillman, R. F.

2007-01-01

A study was undertaken to capture the best practices for the development of reliable and robust spacecraft structures for NASA s next generation cargo and crewed launch vehicles. In this study, the NASA heritage programs such as Mercury, Gemini, Apollo, and the Space Shuttle program were examined. A series of lessons learned during the NASA and DoD heritage programs are captured. The processes that "make the right structural system" are examined along with the processes to "make the structural system right". The impact of technology advancements in materials and analysis and testing methods on reliability and robustness of spacecraft structures is studied. The best practices and lessons learned are extracted from these studies. Since the first human space flight, the best practices for reliable and robust spacecraft structures appear to be well established, understood, and articulated by each generation of designers and engineers. However, these best practices apparently have not always been followed. When the best practices are ignored or short cuts are taken, risks accumulate, and reliability suffers. Thus program managers need to be vigilant of circumstances and situations that tend to violate best practices. Adherence to the best practices may help develop spacecraft systems with high reliability and robustness against certain anomalies and unforeseen events.

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

NASA Technical Reports Server (NTRS)

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 manned travel to Jupiter's moon, Callisto, and assessed that round trip journey. This engine, although unoptimized, produced 10 pounds force of thrust at a delivered specific impulse of 527,000 seconds for the entire 15-year mission while providing enormous amounts of electrical power to the spacecraft. A payload of 60 metric tons, included in the 300 metric ton vehicle, was carried to Callisto and back; the propellant tanks holding the 4 metric tons of fuel were not jettisoned in the process. The study concluded that the engine and spacecraft are within today's technology, could be built, tested, launched on several SLS (Space Launch System) (or similar) launchers, integrated, checked out, moved to an in-space base such as at a Lagrange point and operated for decades.

Multi-Mission Power Analysis Tool (MMPAT) Version 3

NASA Technical Reports Server (NTRS)

Wood, Eric G.; Chang, George W.; Chen, Fannie C.

2012-01-01

The Multi-Mission Power Analysis Tool (MMPAT) simulates a spacecraft power subsystem including the power source (solar array and/or radioisotope thermoelectric generator), bus-voltage control, secondary battery (lithium-ion or nickel-hydrogen), thermostatic heaters, and power-consuming equipment. It handles multiple mission types including heliocentric orbiters, planetary orbiters, and surface operations. Being parametrically driven along with its user-programmable features can reduce or even eliminate any need for software modifications when configuring it for a particular spacecraft. It provides multiple levels of fidelity, thereby fulfilling the vast majority of a project s power simulation needs throughout the lifecycle. It can operate in a stand-alone mode with a graphical user interface, in batch mode, or as a library linked with other tools. This software can simulate all major aspects of a spacecraft power subsystem. It is parametrically driven to reduce or eliminate the need for a programmer. Added flexibility is provided through user-designed state models and table-driven parameters. MMPAT is designed to be used by a variety of users, such as power subsystem engineers for sizing power subsystem components; mission planners for adjusting mission scenarios using power profiles generated by the model; system engineers for performing system- level trade studies using the results of the model during the early design phases of a spacecraft; and operations personnel for high-fidelity modeling of the essential power aspect of the planning picture.

Optimization of Closed Loop Eigenvalues: Maneuvering, Vibration Control, and Structure/Control Design Iteration for Flexible Spacecraft.

DTIC Science & Technology

1986-05-31

Nonlinear Feedback Control 8-16 for Spacecraft Attitude Maneuvers" 2. " Spacecraft Attitude Control Using 17-35... nonlinear state feedback control laws are developed for space- craft attitude control using the Euler parameters and conjugate angular momenta. Time... Nonlinear Feedback Control for Spacecraft Attitude Maneuvers," to appear in AIAA J. of Guidance, Control, and Dynamics, (AIAA Paper No. 83-2230-CP,

Active spacecraft potential control: An ion emitter experiment. [Cluster mission

NASA Technical Reports Server (NTRS)

Riedler, W.; Goldstein, R.; Hamelin, M.; Maehlum, B. N.; Troim, J.; Olsen, R. C.; Pedersen, A.; Grard, R. J. L.; Schmidt, R.; Rudenauer, F.

1988-01-01

The cluster spacecraft are instrumented with ion emitters for charge neutralization. The emitters produce indium ions at 6 keV. The ion current is adjusted in a feedback loop with instruments measuring the spacecraft potential. The system is based on the evaporation of indium in the apex field of a needle. The design of the active spacecraft potential control instruments, and the ion emitters is presented.

Pattern Recognition Control Design

NASA Technical Reports Server (NTRS)

Gambone, Elisabeth

2016-01-01

Spacecraft control algorithms must know the expected spacecraft response to any command to the available control effectors, such as reaction thrusters or torque devices. Spacecraft control system design approaches have traditionally relied on the estimated vehicle mass properties to determine the desired force and moment, as well as knowledge of the effector performance to efficiently control the spacecraft. A pattern recognition approach can be used to investigate the relationship between the control effector commands and the spacecraft responses. Instead of supplying the approximated vehicle properties and the effector performance characteristics, a database of information relating the effector commands and the desired vehicle response can be used for closed-loop control. A Monte Carlo simulation data set of the spacecraft dynamic response to effector commands can be analyzed to establish the influence a command has on the behavior of the spacecraft. A tool developed at NASA Johnson Space Center (Ref. 1) to analyze flight dynamics Monte Carlo data sets through pattern recognition methods can be used to perform this analysis. Once a comprehensive data set relating spacecraft responses with commands is established, it can be used in place of traditional control laws and gains set. This pattern recognition approach can be compared with traditional control algorithms to determine the potential benefits and uses.

A small spacecraft for multipoint measurement of ionospheric plasma.

PubMed

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.

A small spacecraft for multipoint measurement of ionospheric plasma

NASA Astrophysics Data System (ADS)

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.

A spacecraft attitude and articulation control system design for the Comet Halley intercept mission

NASA Technical Reports Server (NTRS)

Key, R. W.

1981-01-01

An attitude and articulation control system design for the Comet Halley 1986 intercept mission is presented. A spacecraft dynamics model consisting of five hinge-connected rigid bodies is used to analyze the spacecraft attitude and articulation control system performance. Inertial and optical information are combined to generate scan platform pointing commands. The comprehensive spacecraft model has been developed into a digital computer simulation program, which provides performance characteristics and insight pertaining to the control and dynamics of a Halley Intercept spacecraft. It is shown that scan platform pointing error has a maximum value of 1.8 milliradians during the four minute closest approach interval. It is also shown that the jitter or scan platform pointing rate error would have a maximum value of 2.5 milliradians/second for the nominal 1000 km closest approach distance trajectory and associated environment model.

Design and architecture of the Mars relay network planning and analysis framework

NASA Technical Reports Server (NTRS)

Cheung, K. M.; Lee, C. H.

2002-01-01

In this paper we describe the design and architecture of the Mars Network planning and analysis framework that supports generation and validation of efficient planning and scheduling strategy. The goals are to minimize the transmitting time, minimize the delaying time, and/or maximize the network throughputs. The proposed framework would require (1) a client-server architecture to support interactive, batch, WEB, and distributed analysis and planning applications for the relay network analysis scheme, (2) a high-fidelity modeling and simulation environment that expresses link capabilities between spacecraft to spacecraft and spacecraft to Earth stations as time-varying resources, and spacecraft activities, link priority, Solar System dynamic events, the laws of orbital mechanics, and other limiting factors as spacecraft power and thermal constraints, (3) an optimization methodology that casts the resource and constraint models into a standard linear and nonlinear constrained optimization problem that lends itself to commercial off-the-shelf (COTS)planning and scheduling algorithms.

ARC-2009-ACD09-0141-001

NASA Image and Video Library

2009-07-16

John W. 'Jack Boyd holds a plaque presented to Harvey Allen in recognition of his outstanding solution of the reentry heating problem which has been indispensable to the design of the Mercury, Gemini, and Apollo spacecraft (Manned Spacecraft Center, November 14, 1968) Plaque contains samples of tested materials and models of spacecraft.

Cost-Effective Icy Bodies Exploration using Small Satellite Missions

NASA Technical Reports Server (NTRS)

Jonsson, Jonas; Mauro, David; Stupl, Jan; Nayak, Michael; Aziz, Jonathan; Cohen, Aaron; Colaprete, Anthony; Dono-Perez, Andres; Frost, Chad; Klamm, Benjamin;

High Earth orbit design for lunar assisted small Explorer class missions

NASA Technical Reports Server (NTRS)

Mathews, M.; Hametz, M.; Cooley, J.; Skillman, D.

1994-01-01

Small Expendable launch vehicles are capable of injecting modest payloads into high Earth orbits having apogee near the lunar distance. However, lunar and solar perturbations can quickly lower perigee and cause premature reentry. Costly perigee raising maneuvers by the spacecraft are required to maintain the orbit. In addition, the range of inclinations achievable is limited to those of launch sites unless costly spacecraft maneuvers are performed. This study investigates the use of a lunar swingby in a near-Hohmann transfer trajectory to raise perigee into the 8 to 25 solar radius range and reach a wide variety of inclinations without spacecraft maneuvers. It is found that extremely stable orbits can be obtained if the postencounter spacecraft orbital period is one-half of a lunar sidereal revolution and the Earth-vehicle-Moon geometry is within a specified range. Criteria for achieving stable orbits with various perigee heights and ecliptic inclinations are developed, and the sensitivity of the resulting mission orbits to transfer trajectory injection (TTI) errors is examined. It is shown that carefully designed orbits yield lifetimes of several years, with excellent ground station coverage characteristics and minimal eclipses. A phasing loop error correction strategy is considered with the spacecraft propulsion system delta V demand for TTI error correction and a postlunar encounter apogee trim maneuver typically in the 30 to 120 meters per second range.

KSC-05pd2273

NASA Image and Video Library

2005-09-29

KENNEDY SPACE CENTER, FLA. - On the Shuttle Landing Facility at NASA Kennedy Space Center, the Atlas V fairing halves for the New Horizons spacecraft have been offloaded from the Russian cargo plane (background). The fairing halves will be transported to Astrotech Space Operations in Titusville. The fairing later will be placed around the New Horizons spacecraft in the Payload Hazardous Service Facility. A fairing protects a spacecraft during launch and flight through the atmosphere. Once in space, it is jettisoned. The Lockheed Martin Atlas V is the launch vehicle for the New Horizons spacecraft, which is designed to make the first reconnaissance of Pluto and Charon - a "double planet" and the last planet in our solar system to be visited by spacecraft. The mission will then visit one or more objects in the Kuiper Belt region beyond Neptune. New Horizons is scheduled to launch in January 2006, swing past Jupiter for a gravity boost and scientific studies in February or March 2007, and reach Pluto and its moon, Charon, in July 2015.

KSC-05pd2274

NASA Image and Video Library

2005-09-29

KENNEDY SPACE CENTER, FLA. - On the Shuttle Landing Facility at NASA Kennedy Space Center, one of the Atlas V fairing halves for the New Horizons spacecraft is offloaded from the Russian cargo plane. The fairing halves will be transported to Astrotech Space Operations in Titusville. The fairing later will be placed around the New Horizons spacecraft in the Payload Hazardous Service Facility. A fairing protects a spacecraft during launch and flight through the atmosphere. Once in space, it is jettisoned. The Lockheed Martin Atlas V is the launch vehicle for the New Horizons spacecraft, which is designed to make the first reconnaissance of Pluto and Charon - a "double planet" and the last planet in our solar system to be visited by spacecraft. The mission will then visit one or more objects in the Kuiper Belt region beyond Neptune. New Horizons is scheduled to launch in January 2006, swing past Jupiter for a gravity boost and scientific studies in February or March 2007, and reach Pluto and its moon, Charon, in July 2015.

Dynamic performance of an aero-assist spacecraft - AFE

NASA Technical Reports Server (NTRS)

Chang, Ho-Pen; French, Raymond A.

1992-01-01

Dynamic performance of the Aero-assist Flight Experiment (AFE) spacecraft was investigated using a high-fidelity 6-DOF simulation model. Baseline guidance logic, control logic, and a strapdown navigation system to be used on the AFE spacecraft are also modeled in the 6-DOF simulation. During the AFE mission, uncertainties in the environment and the spacecraft are described by an error space which includes both correlated and uncorrelated error sources. The principal error sources modeled in this study include navigation errors, initial state vector errors, atmospheric variations, aerodynamic uncertainties, center-of-gravity off-sets, and weight uncertainties. The impact of the perturbations on the spacecraft performance is investigated using Monte Carlo repetitive statistical techniques. During the Solid Rocket Motor (SRM) deorbit phase, a target flight path angle of -4.76 deg at entry interface (EI) offers very high probability of avoiding SRM casing skip-out from the atmosphere. Generally speaking, the baseline designs of the guidance, navigation, and control systems satisfy most of the science and mission requirements.

KSC-05pd2271

NASA Image and Video Library

2005-09-29

KENNEDY SPACE CENTER, FLA. - A Russian cargo plane sits on the Shuttle Landing Facility at NASA Kennedy Space Center with the Atlas V fairing for the New Horizons spacecraft inside. The two fairing halves will be removed, loaded onto trucks and transported to Astrotech Space Operations in Titusville. The fairing later will be placed around the New Horizons spacecraft in the Payload Hazardous Service Facility. A fairing protects a spacecraft during launch and flight through the atmosphere. Once in space, it is jettisoned. The Lockheed Martin Atlas V is the launch vehicle for the New Horizons spacecraft, which is designed to make the first reconnaissance of Pluto and Charon - a "double planet" and the last planet in our solar system to be visited by spacecraft. The mission will then visit one or more objects in the Kuiper Belt region beyond Neptune. New Horizons is scheduled to launch in January 2006, swing past Jupiter for a gravity boost and scientific studies in February or March 2007, and reach Pluto and its moon, Charon, in July 2015.