Science.gov

Sample records for mission design options

  1. Mission design options for human Mars missions

    NASA Astrophysics Data System (ADS)

    Wooster, Paul D.; Braun, Robert D.; Ahn, Jaemyung; Putnam, Zachary R.

    Trajectory options for conjunction-class human Mars missions are examined, including crewed Earth-Mars trajectories with the option for abort to Earth, with the intent of serving as a resource for mission designers. An analysis of the impact of Earth and Mars entry velocities on aeroassist systems is included, and constraints are suggested for interplanetary trajectories based upon aeroassist system capabilities.

  2. Comparison of mission design options for manned Mars missions

    NASA Technical Reports Server (NTRS)

    Babb, Gus R.; Stump, William R.

    1986-01-01

    A number of manned Mars mission types, propulsion systems, and operational techniques are compared. Conjunction and opposition class missions for cryogenic, hybrid (cryo/storable), and NERVA propulsion concepts are addressed. In addition, both Earth and Mars orbit aerobraking, direct entry of landers, hyperbolic rendezvous, and electric propulsion cases are examined. A common payload to Mars was used for all cases. The basic figure of merit used was weight in low Earth orbit (LEO) at mission initiation. This is roughly proportional to launch costs.

  3. Direct transfer trajectory design options for interplanetary orbiter missions using an iterative patched conic method

    NASA Astrophysics Data System (ADS)

    Parvathi, S. P.; Ramanan, R. V.

    2017-04-01

    In a direct interplanetary transfer, the spacecraft moves from a parking orbit of the departure planet to a parking orbit of the arrival planet. The transfer trajectory must be designed such that the specified arrival parking orbit conditions are achieved. For a fixed departure epoch and flight duration, there are four distinct transfer trajectory design options in a direct transfer. The conventional patched conic method, the most widely used analytical trajectory design method, does not identify these design options. An iterative patched conic method that identifies these distinct design options is developed and presented in this paper. This method involves two iterative processes: (i) iteration on the hyperbolic orbit characteristics using an analytical tuning strategy to achieve the hyperbolic excess velocity vector at the patch point, (ii) iteration on the patch points at the sphere of influence. The performance of the proposed method is compared with the conventional and V-infinity tuned patched conic methods. A design analysis tool, based on the proposed method, is developed and tested in various orbiter mission scenarios.

  4. Superheavyweight missions SI versus DI: Ascent flight design options and recommendations

    NASA Technical Reports Server (NTRS)

    1990-01-01

    AFD has completed the trade study on Standard Insertion (SI) vs Direct Insertion (DI) for STS-50. RSOC Range Safety has developed acceptable DI targets from 130 n.mi. to 150 n.mi. and the corresponding performance assessment for these targets using STS-50 data has been completed. This mission has sufficient performance capability to perform this mission as a DI to 160 n.mi. A reduced OMS load corresponding to a DI mission is required for this option. The increase in altitude over the AFP baseline (SI to 145 n.mi.) is highly desirable for this mission. The orientation on orbit for the orbiter/USML-1 payload is such that orbital decay is maximized (maximum frontal cross-sectional area with vehicle normal to velocity vector). Increasing the operational altitude reduces the amount of vernier thruster firings necessary to maintain a constant gravity gradient. The results of this trade study can also be applied to other superheavyweight missions (EDO flights) and will allow for use of the DI technique for lower orbital altitudes, thereby eliminating the SI option for due east, low altitude missions. STSOC transmittal form no. 330-330-130, which documents the technical issues and assumptions used for this trade study effort in detail, should be referenced for further information. The main reason that a DI is desired for STS-50 and other superheavyweight flights (low altitude) is that ESMC range safety has expressed reservations about SI missions in general. The concern is that the current SI design underspeed exposes Africa and Madagascar to potential ET debris impact. In the past range safety has waived the requirement that these areas be protected in the event of an engine failure. With the advent of the pre-MECO OMS dump, the viability of DI and the high casualty expectations from the ACTA press to MECO hazard study, range safety has become more reluctant to approve SI flights. It is felt that to perform an SI mission there would have to be a large decrease in design

  5. Nuclear Thermal Rocket/Vehicle Design Options for Future NASA Missions to the Moon and Mars

    NASA Technical Reports Server (NTRS)

    Borowski, Stanley K.; Corban, Robert R.; Mcguire, Melissa L.; Beke, Erik G.

    1995-01-01

    The nuclear thermal rocket (NTR) provides a unique propulsion capability to planners/designers of future human exploration missions to the Moon and Mars. In addition to its high specific impulse (approximately 850-1000 s) and engine thrust-to-weight ratio (approximately 3-10), the NTR can also be configured as a 'dual mode' system capable of generating electrical power for spacecraft environmental systems, communications, and enhanced stage operations (e.g., refrigeration for long-term liquid hydrogen storage). At present the Nuclear Propulsion Office (NPO) is examining a variety of mission applications for the NTR ranging from an expendable, single-burn, trans-lunar injection (TLI) stage for NASA's First Lunar Outpost (FLO) mission to all propulsive, multiburn, NTR-powered spacecraft supporting a 'split cargo-piloted sprint' Mars mission architecture. Each application results in a particular set of requirements in areas such as the number of engines and their respective thrust levels, restart capability, fuel operating temperature and lifetime, cryofluid storage, and stage size. Two solid core NTR concepts are examined -- one based on NERVA (Nuclear Engine for Rocket Vehicle Application) derivative reactor (NDR) technology, and a second concept which utilizes a ternary carbide 'twisted ribbon' fuel form developed by the Commonwealth of Independent States (CIS). The NDR and CIS concepts have an established technology database involving significant nuclear testing at or near representative operating conditions. Integrated systems and mission studies indicate that clusters of two to four 15 to 25 klbf NDR or CIS engines are sufficient for most of the lunar and Mars mission scenarios currently under consideration. This paper provides descriptions and performance characteristics for the NDR and CIS concepts, summarizes NASA's First Lunar Outpost and Mars mission scenarios, and describes characteristics for representative cargo and piloted vehicles compatible with a

  6. Design concepts and options for the Thermal Infrared Imager (TIRI) as part of ESA's Asteroid Impact Mission.

    NASA Astrophysics Data System (ADS)

    Bowles, Neil; Calcutt, Simon; Licandro, Javier; Reyes, Marcos; Delbo, Marco; Donaldson Hanna, Kerri; Arnold, Jessica; Howe, Chris

    2016-04-01

    ESA's Asteroid Impact Mission (AIM) is being studied as part of the joint ESA/NASA AIDA mission for launch in 2020. AIDA's primary mission is to investigate the effect of a kinetic impactor on the secondary component of the binary asteroid 65803 Didymos in late 2022. AIM will characterise the Didymos system and monitor the response of the binary system to the impact. A multi-spectral, thermal-infrared imaging instrument (TIRI) will be an essential component of AIM's remote sensing payload, as it will provide key information on the nature of the surfaces (e.g. presence or absence of materials, degree of compaction, and rock abundance of the regolith) of both components in the Didymos system. The temperature maps provided by TIRI will be important for navigation and spacecraft health and safety for proximity/lander operations. By measuring the asteroids' diurnal thermal responses (thermal inertia) and their surface compositions via spectral signatures, TIRI will provide information on the origin and evolution of the binary system. In this presentation we will discuss possible instrument design for TIRI, exploring options that include imaging spectroscopy to broadband imaging. By using thermal models and compositional analogues of the Didymos system we will show how the performance of each design option compares to the wider scientific goals of the AIDA/AIM mission.

  7. Boulder Capture System Design Options for the Asteroid Robotic Redirect Mission Alternate Approach Trade Study

    NASA Technical Reports Server (NTRS)

    Belbin, Scott P.; Merrill, Raymond G.

    2014-01-01

    This paper presents a boulder acquisition and asteroid surface interaction electromechanical concept developed for the Asteroid Robotic Redirect Mission (ARRM) option to capture a free standing boulder on the surface of a 100 m or larger Near Earth Asteroid (NEA). It details the down select process and ranking of potential boulder capture methods, the evolution of a simple yet elegant articulating spaceframe, and ongoing risk reduction and concept refinement efforts. The capture system configuration leverages the spaceframe, heritage manipulators, and a new microspine technology to enable the ARRM boulder capture. While at the NEA it enables attenuation of terminal descent velocity, ascent to escape velocity, boulder collection and restraint. After departure from the NEA it enables, robotic inspection, sample caching, and crew Extra Vehicular Activities (EVA).

  8. RTGs Options for Pluto Fast Flyby Mission

    SciTech Connect

    Schock, Alfred

    1993-10-01

    A small spacecraft design for the Pluto Fast Flyby (PFF) Mission is under study by the Jet Propulsion Laboratory (JPL) for the National Aeronautics and Space Administration (NASA), for a possible launch as early as 1998. JPL's 1992 baseline design calls for a power source able to furnish an energy output of 3963 kWh and a power output of 69 watts(e) at the end of the 9.2-year mission. Satisfying those demands is made difficult because NASA management has set a goal of reducing the spacecraft mass from a baseline value of 166 kg to ~110 kg, which implies a mass goal of less than 10 kg for the power source. To support the ongoing NASA/JPL studies, the Department of Energy's Office of Special Applications (DOE/OSA) commissioned Fairchild Space to prepare and analyze conceptual designs of radioisotope power systems for the PFF mission. Thus far, a total of eight options employing essentially the same radioisotope heat source modules were designed and subjected to thermal, electrical, structural, and mass analyses by Fairchild. Five of these - employing thermoelectric converters - are described in the present paper, and three - employing free-piston Stirling converters - are described in the companion paper presented next. The system masses of the thermoelectric options ranged from 19.3 kg to 10.2 kg. In general, the options requiring least development are the heaviest, and the lighter options require more development with greater programmatic risk. There are four duplicate copies

  9. Vehicle and Mission Design Options for the Human Exploration of Mars/Phobos Using "Bimodal" NTR and LANTR Propulsion

    NASA Technical Reports Server (NTRS)

    Borowski, Stanley K.; Dudzinski, Leonard A.; McGuire, Melissa L.

    1998-01-01

    The nuclear thermal rocket (NTR) is one of the leading propulsion options for future human missions to Mars because of its high specific impulse (Isp-850-1000 s) capability and its attractive engine thrust-to-weight ratio (approximately equal 3-10). To stay within the available mass and payload volume limits of a "Magnum" heavy lift vehicle, a high performance propulsion system is required for trans-Mars injection (TMI). An expendable TMI stage, powered by three 15 thousand pounds force (klbf) NTR engines is currently under consideration by NASA for its Design Reference Mission (DRM). However, because of the miniscule burnup of enriched uranium-235 during the Earth departure phase (approximately 10 grams out of 33 kilograms in each NTR core), disposal of the TMI stage and its engines after a single use is a costly and inefficient use of this high performance stage. By reconfiguring the engines for both propulsive thrust and modest power generation (referred to as "bimodal" operation), a robust, multiple burn, "power-rich" stage with propulsive Mars capture and reuse capability is possible, A family of modular "bimodal" NTR (BNTR) vehicles are described which utilize a common "core" stage powered by three 15 klbf BNTRs that produce 50 kWe of total electrical power for crew life support, an active refrigeration / reliquification system for long term, "zero-boiloff" liquid hydrogen (LH2) storage, and high data rate communications. An innovative, spine-like "saddle truss" design connects the core stage and payload element and is open underneath to allow supplemental "in-line" propellant tanks and contingency crew consumables to be easily jettisoned to improve vehicle performance. A "modified" DRM using BNTR transfer vehicles requires fewer transportation system elements, reduces IMLEO and mission risk, and simplifies space operations. By taking the next logical step--use of the BNTR for propulsive capture of all payload elements into Mars orbit--the power available in

  10. Vehicle and Mission Design Options for the Human Exploration of Mars/Phobos Using "Bimodal" NTR and LANTR Propulsion

    NASA Astrophysics Data System (ADS)

    Borowski, Stanley K.; Dudzinski, Leonard A.; McGuire, Melissa L.

    2002-12-01

    The nuclear thermal rocket (NTR) is one of the leading propulsion options for future human missions to Mars because of its high specific impulse (1sp is approximately 850-1000 s) capability and its attractive engine thrust-to-weight ratio (approximately 3-10). To stay within the available mass and payload volume limits of a "Magnum" heavy lift vehicle, a high performance propulsion system is required for trans-Mars injection (TMI). An expendable TMI stage, powered by three 15 thousand pounds force (klbf) NTR engines is currently under consideration by NASA for its Design Reference Mission (DRM). However, because of the miniscule burnup of enriched uranium-235 during the Earth departure phase (approximately 10 grams out of 33 kilograms in each NTR core), disposal of the TMI stage and its engines after a single use is a costly and inefficient use of this high performance stage. By reconfiguring the engines for both propulsive thrust and modest power generation (referred to as "bimodal" operation), a robust, multiple burn, "power-rich" stage with propulsive Mars capture and reuse capability is possible. A family of modular bimodal NTR (BNTR) vehicles are described which utilize a common "core" stage powered by three 15 klbf BNTRs that produce 50 kWe of total electrical power for crew life support, an active refrigeration / reliquification system for long term, zero-boiloff liquid hydrogen (LH2) storage, and high data rate communications. An innovative, spine-like "saddle truss" design connects the core stage and payload element and is open underneath to allow supplemental "in-line" propellant tanks and contingency crew consumables to be easily jettisoned to improve vehicle performance. A "modified" DRM using BNTR transfer vehicles requires fewer transportation system elements, reduces IMLEO and mission risk, and simplifies space operations. By taking the next logical step--use of the BNTR for propulsive capture of all payload elements into Mars orbit--the power

  11. Implementation Options for the PROPEL Electrodynamic Tether Demonstration Mission

    NASA Technical Reports Server (NTRS)

    Bilen, Sven G.; Johnson, Les; Gilchrist, Brian E.; Hoyt, Robert P.; Elder, Craig H.; Fuhrhop, Keith P.; Scadera, Michael P.; Stone, Nobie H.

    2014-01-01

    The PROPEL flight mission concept will demonstrate the safe use of an electrodynamic tether for generating thrust. PROPEL is being designed to be a versatile electrodynamic-tether system for multiple end users and to be flexible with respect to platform. As such, several implementation options are being explored, including a comprehensive mission design for PROPEL with a mission duration of six months; a space demonstration mission concept design with configuration of a pair of tethered satellites, one of which is the Japanese H-II Transfer Vehicle; and an ESPA-based system. We report here on these possible implementation options for PROPEL. electrodynamic tether; PROPEL demonstration mission; propellantless propulsion

  12. Abort Options for Potential Mars Missions

    NASA Technical Reports Server (NTRS)

    Tartabini, P. V.; Striepe, S. A.; Powell, R. W.

    1994-01-01

    Mars trajectory design options were examined that would accommodate a premature termination of a nominal manned opposition class mission for opportunities between 2010 and 2025. A successful abort must provide a safe return to Earth in the shortest possible time consistent with mission constraints. In this study, aborts that provided a minimum increase in the initial vehicle mass in low Earth orbit (IMLEO) were identified by locating direct transfer nominal missions and nominal missions including an outbound or inbound Venus swing-by that minimized IMLEO. The ease with which these missions could be aborted while meeting propulsion and time constraints was investigated by examining free return (unpowered) and powered aborts. Further reductions in trip time were made to some aborts by the addition or removal of an inbound Venus swing-by. The results show that, although few free return aborts met the specified constraints, 85% of each nominal mission could be aborted as a powered abort without an increase in propellant. Also, in many cases, the addition or removal of a Venus swing-by increased the number of abort opportunities or decreased the total trip time during an abort.

  13. Improving Conceptual Design for Launch Vehicles. The Bimese Concept: A Study of Mission and Economic Options

    NASA Technical Reports Server (NTRS)

    Olds, John R.; Tooley, Jeffrey

    1999-01-01

    This report summarizes key activities conducted in the third and final year of the cooperative agreement NCC1-229 entitled "Improving Conceptual Design for Launch Vehicles." This project has been funded by the Vehicle Analysis Branch at NASA's Langley Research Center in Hampton, VA. Work has been performed by the Space Systems Design Lab (SSDL) at the Georgia Institute of Technology, Atlanta, GA. Accomplishments during the first and second years of this project have been previously reported in annual progress reports. This report will focus on the third and final year of the three year activity.

  14. A mission and system design option for the Orbiting Stellar Interferometer

    NASA Technical Reports Server (NTRS)

    Rayman, M. D.; Shao, M.

    1992-01-01

    The Orbiting Stellar Interferometer (OSI) is a proposed space-based observatory that will open exciting new vistas in astronomy and address fundamental scientific questions by making extremely accurate (3 to 30 microarcsecond) astrometric measurements of the positions of stars, quasars, and other astronomical objects as faint as magnitude 20. In addition, it will be able to image objects with a resolution of about 13 milliarcseconds. Using the lessons learned from the development of a point design for OSI, the Jet Propulsion Laboratory has developed a new conceptual design, with emphasis on reducing its cost and complexity while maximizing the return of valuable science.

  15. 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

  16. Asteroid Redirect Robotic Mission: Robotic Boulder Capture Option Overview

    NASA Technical Reports Server (NTRS)

    Mazanek, Daniel D.; Merrill, Raymond G.; Belbin, Scott P.; Reeves, David M.; Earle, Kevin D.; Naasz, Bo J.; Abell, Paul A.

    2014-01-01

    The National Aeronautics and Space Administration (NASA) is currently studying an option for the Asteroid Redirect Robotic Mission (ARRM) that would capture a multi-ton boulder (typically 2-4 meters in size) from the surface of a large (is approximately 100+ meter) Near-Earth Asteroid (NEA) and return it to cislunar space for subsequent human and robotic exploration. This alternative mission approach, designated the Robotic Boulder Capture Option (Option B), has been investigated to determine the mission feasibility and identify potential differences from the initial ARRM concept of capturing an entire small NEA (4-10 meters in size), which has been designated the Small Asteroid Capture Option (Option A). Compared to the initial ARRM concept, Option B allows for centimeter-level characterization over an entire large NEA, the certainty of target NEA composition type, the ability to select the boulder that is captured, numerous opportunities for mission enhancements to support science objectives, additional experience operating at a low-gravity planetary body including extended surface contact, and the ability to demonstrate future planetary defense strategies on a hazardous-size NEA. Option B can leverage precursor missions and existing Agency capabilities to help ensure mission success by targeting wellcharacterized asteroids and can accommodate uncertain programmatic schedules by tailoring the return mass.

  17. NASA's asteroid redirect mission: Robotic boulder capture option

    NASA Astrophysics Data System (ADS)

    Abell, P.; Nuth, J.; Mazanek, D.; Merrill, R.; Reeves, D.; Naasz, B.

    2014-07-01

    NASA is examining two options for the Asteroid Redirect Mission (ARM), which will return asteroid material to a Lunar Distant Retrograde Orbit (LDRO) using a robotic solar-electric-propulsion spacecraft, called the Asteroid Redirect Vehicle (ARV). Once the ARV places the asteroid material into the LDRO, a piloted mission will rendezvous and dock with the ARV. After docking, astronauts will conduct two extravehicular activities (EVAs) to inspect and sample the asteroid material before returning to Earth. One option involves capturing an entire small (˜4--10 m diameter) near-Earth asteroid (NEA) inside a large inflatable bag. However, NASA is also examining another option that entails retrieving a boulder (˜1--5 m) via robotic manipulators from the surface of a larger (˜100+ m) pre-characterized NEA. The Robotic Boulder Capture (RBC) option can leverage robotic mission data to help ensure success by targeting previously (or soon to be) well-characterized NEAs. For example, the data from the Japan Aerospace Exploration Agency's (JAXA) Hayabusa mission has been utilized to develop detailed mission designs that assess options and risks associated with proximity and surface operations. Hayabusa's target NEA, Itokawa, has been identified as a valid target and is known to possess hundreds of appropriately sized boulders on its surface. Further robotic characterization of additional NEAs (e.g., Bennu and 1999 JU_3) by NASA's OSIRIS REx and JAXA's Hayabusa 2 missions is planned to begin in 2018. This ARM option reduces mission risk and provides increased benefits for science, human exploration, resource utilization, and planetary defense.

  18. Radioisotope Stirling Generator Options for Pluto Fast Flyby Mission

    SciTech Connect

    Schock, Alfred

    1993-10-01

    The preceding paper described conceptual designs and analytical results for five Radioisotope Thermoelectric Generator (RTG) options for the Pluto Fast Flyby (PFF) mission, and the present paper describes three Radioisotope Stirling Generator (RSG) options for the same mission. The RSG options are based on essentially the same radioisotope heat source modules used in previously flown RTGs and on designs and analyses of a 75-watt free-piston Stirling engine produced by Mechanical Technology Incorporated (MTI) for NASA's Lewis Research Center. The integrated system design options presented were generated in a Fairchild Space study sponsored by the Department of Energy's Office of Special Applications, in support of ongoing PFF mission and spacecraft studies that the Jet Propulsion Laboratory (JPL) is conducting for the National Aeronautics and Space Administration (NASA). That study's NASA-directed goal is to reduce the spacecraft mass from its baseline value of 166 kg to ~110 kg, which implies a mass goal of less than 10 kg for a power source able to deliver 69 watts(e) at the end of the 9.2-year mission. In general, the Stirling options were found to be lighter than the thermoelectric options described in the preceding paper. But they are less mature, requiring more development, and entailing greater programmatic risk. The Stirling power system mass ranged from 7.3 kg (well below the 10-kg goal) for a non-redundant system to 11.3 kg for a redundant system able to maintain full power if one of its engines fails. In fact, the latter system could deliver as much as 115 watts(e) if desired by the mission planners. There are 5 copies in the file.

  19. Radioisotope Stirling Generator Options for Pluto Fast Flyby Mission

    SciTech Connect

    Schock, Alfred

    2012-01-19

    The preceding paper described conceptual designs and analytical results for five Radioisotope Thermoelectric Generator (RTG) options for the Pluto Fast Flyby (PFF) mission, and the present paper describes three Radioisotope Stirling Generator (RSG) options for the same mission. The RSG options are based on essentially the same radioisotope heat source modules used in previously flown RTGs and on designs and analyses of a 75-watt free-piston Stirling engine produced by Mechanical Technology Incorporated (MTI) for NASA's Lewis Research Center. The integrated system design options presented were generated in a Fairchild Space study sponsored by the Department of Energy's Office of Special Applications, in support of ongoing PFF mission and spacecraft studies that the Jet Propulsion Laboratory (JPL) is conducting for the National Aeronautics and Space Administration (NASA). That study's NASA-directed goal is to reduce the spacecraft mass from its baseline value of 166 kg to ~110 kg, which implies a mass goal of less than 10 kg for a power source able to deliver 69 watts(e) at the end of the 9.2-year mission. In general, the Stirling options were found to be lighter than the thermoelectric options described in the preceding paper. But they are less mature, requiring more development, and entailing greater programmatic risk. The Stirling power system mass ranged from 7.3 kg (well below the 10-kg goal) for a non-redundant system to 11.3 kg for a redundant system able to maintain full power if one of its engines fails. In fact, the latter system could deliver as much as 115 watts(e) if desired by the mission planners. There are 2 copies in the file.

  20. Implementation Options for the PROPEL Electrodynamic Tether Demonstration Mission

    NASA Technical Reports Server (NTRS)

    Bilen, Sven G.; Johnson, C. Les; Gilchrist, Brian E.; Hoyt, Robert P.; Elder, Craig H.; Fuhrhop, Keith P.; Scadera, Michael; Stone, Nobie

    2014-01-01

    . The ETPS builds on prior work on long-life, failure-resistant, conducting tethers and includes an instrument suite with demonstrated heritage capable of performing necessary diagnostics to measure performance against predictions for a given system size (to be determined) and boost rate. Mission designs in other configurations and launch vehicle options are being developed such that the system can be demonstration should a flight opportunity be identified. We will report on past and ongoing implementation options for PROPEL.

  1. NASA’s Asteroid Redirect Mission: The Boulder Capture Option

    NASA Astrophysics Data System (ADS)

    Abell, Paul; Nuth, Joseph A.; Mazanek, Dan D.; Merrill, Raymond G.; Reeves, David M.; Naasz, Bo J.

    2014-11-01

    NASA is examining two options for the Asteroid Redirect Mission (ARM), which will return asteroid material to a Lunar Distant Retrograde Orbit (LDRO) using a robotic solar-electric-propulsion spacecraft, called the Asteroid Redirect Vehicle (ARV). Once the ARV places the asteroid material into the LDRO, a piloted mission will rendezvous and dock with the ARV. After docking, astronauts will conduct two extravehicular activities (EVAs) to inspect and sample the asteroid material before returning to Earth. One option involves capturing an entire small (˜4-10 m diameter) near-Earth asteroid (NEA) inside a large inflatable bag. However, NASA is examining another option that entails retrieving a boulder (˜1-5 m) via robotic manipulators from the surface of a larger (˜100+ m) pre-characterized NEA. This option can leverage robotic mission data to help ensure success by targeting previously (or soon to be) well-characterized NEAs. For example, the data from the Hayabusa mission has been utilized to develop detailed mission designs that assess options and risks associated with proximity and surface operations. Hayabusa’s target NEA, Itokawa, has been identified as a valid target and is known to possess hundreds of appropriately sized boulders on its surface. Further robotic characterization of additional NEAs (e.g., Bennu and 1999 JU3) by NASA’s OSIRIS REx and JAXA’s Hayabusa 2 missions is planned to begin in 2018. The boulder option is an extremely large sample-return mission with the prospect of bringing back many tons of well-characterized asteroid material to the Earth-Moon system. The candidate boulder from the target NEA can be selected based on inputs from the world-wide science community, ensuring that the most scientifically interesting boulder be returned for subsequent sampling. This boulder option for NASA’s ARM can leverage knowledge of previously characterized NEAs from prior robotic missions, which provides more certainty of the target NEA

  2. NASA's Asteroid Redirect Mission: The Boulder Capture Option

    NASA Technical Reports Server (NTRS)

    Abell, Paul A.; Nuth, J.; Mazanek, D.; Merrill, R.; Reeves, D.; Naasz, B.

    2014-01-01

    NASA is examining two options for the Asteroid Redirect Mission (ARM), which will return asteroid material to a Lunar Distant Retrograde Orbit (LDRO) using a robotic solar-electric-propulsion spacecraft, called the Asteroid Redirect Vehicle (ARV). Once the ARV places the asteroid material into the LDRO, a piloted mission will rendezvous and dock with the ARV. After docking, astronauts will conduct two extravehicular activities (EVAs) to inspect and sample the asteroid material before returning to Earth. One option involves capturing an entire small (approximately 4-10 m diameter) near-Earth asteroid (NEA) inside a large inflatable bag. However, NASA is examining another option that entails retrieving a boulder (approximately 1-5 m) via robotic manipulators from the surface of a larger (approximately 100+ m) pre-characterized NEA. This option can leverage robotic mission data to help ensure success by targeting previously (or soon to be) well-characterized NEAs. For example, the data from the Hayabusa mission has been utilized to develop detailed mission designs that assess options and risks associated with proximity and surface operations. Hayabusa's target NEA, Itokawa, has been identified as a valid target and is known to possess hundreds of appropriately sized boulders on its surface. Further robotic characterization of additional NEAs (e.g., Bennu and 1999 JU3) by NASA's OSIRIS REx and JAXA's Hayabusa 2 missions is planned to begin in 2018. The boulder option is an extremely large sample-return mission with the prospect of bringing back many tons of well-characterized asteroid material to the Earth-Moon system. The candidate boulder from the target NEA can be selected based on inputs from the world-wide science community, ensuring that the most scientifically interesting boulder be returned for subsequent sampling. This boulder option for NASA's ARM can leverage knowledge of previously characterized NEAs from prior robotic missions, which provides more

  3. Ion Propulsion Module design and mission performance

    NASA Technical Reports Server (NTRS)

    Graf, J. E.; Boain, R. J.; Pawlik, E. V.; Pless, L. C.

    1978-01-01

    This paper describes the design options, processes and tradeoffs that occur during the establishment of viable Ion Drive vehicle and mission designs. The options identify those internal vehicle design alternatives which are being considered for future Ion Drive missions, such as sunlight concentrating arrays and direct drive thrust subsystems, and their effect on mission performance. Also, the highly interactive nature of the Ion Drive design process, which occurs between the spacecraft and mission designers, is described. The results of design tradeoffs, performed for three Ion Drive comet rendezvous missions, are presented. These results include the following: (1) the power profile is determined primarily by the trajectory while second order effects include the solar cell characteristics and array concentration factor and degradation; and (2) the dominant parameter in mission performance determination, Ion Propulsion Module (IPM) mass, and IPM design, is the total cell power evaluated without concentration, at the beginning of life and at 1 AU.

  4. Vehicle and Mission Design Options for the Human Exploration of Mars/Phobos Using "Bimodal" NTR and LANTR Propulsion. Revised Dec. 1998

    NASA Technical Reports Server (NTRS)

    Borowski, Stanley K.; Dudzinski, Leonard A.; McGuire, Melissa L.

    2002-01-01

    The nuclear thermal rocket (NTR) is one of the leading propulsion options for future human missions to Mars because of its high specific impulse (1sp is approximately 850-1000 s) capability and its attractive engine thrust-to-weight ratio (approximately 3-10). To stay within the available mass and payload volume limits of a "Magnum" heavy lift vehicle, a high performance propulsion system is required for trans-Mars injection (TMI). An expendable TMI stage, powered by three 15 thousand pounds force (klbf) NTR engines is currently under consideration by NASA for its Design Reference Mission (DRM). However, because of the miniscule burnup of enriched uranium-235 during the Earth departure phase (approximately 10 grams out of 33 kilograms in each NTR core), disposal of the TMI stage and its engines after a single use is a costly and inefficient use of this high performance stage. By reconfiguring the engines for both propulsive thrust and modest power generation (referred to as "bimodal" operation), a robust, multiple burn, "power-rich" stage with propulsive Mars capture and reuse capability is possible. A family of modular bimodal NTR (BNTR) vehicles are described which utilize a common "core" stage powered by three 15 klbf BNTRs that produce 50 kWe of total electrical power for crew life support, an active refrigeration / reliquification system for long term, zero-boiloff liquid hydrogen (LH2) storage, and high data rate communications. An innovative, spine-like "saddle truss" design connects the core stage and payload element and is open underneath to allow supplemental "in-line" propellant tanks and contingency crew consumables to be easily jettisoned to improve vehicle performance. A "modified" DRM using BNTR transfer vehicles requires fewer transportation system elements, reduces IMLEO and mission risk, and simplifies space operations. By taking the next logical step--use of the BNTR for propulsive capture of all payload elements into Mars orbit--the power

  5. Electric Propulsion Options for a Magnetospheric Mapping Mission

    NASA Technical Reports Server (NTRS)

    Oleson, Steven; Russell, Chris; Hack, Kurt; Riehl, John

    1998-01-01

    The Twin Electric Magnetospheric Probes Exploring on Spiral Trajectories mission concept was proposed as a Middle Explorer class mission. A pre-phase-A design was developed which utilizes the advantages of electric propulsion for Earth scientific spacecraft use. This paper presents propulsion system analyses performed for the proposal. The proposed mission required two spacecraft to explore near circular orbits 0.1 to 15 Earth radii in both high and low inclination orbits. Since the use of chemical propulsion would require launch vehicles outside the Middle Explorer class a reduction in launch mass was sought using ion, Hall, and arcjet electric propulsion system. Xenon ion technology proved to be the best propulsion option for the mission requirements requiring only two Pegasus XL launchers. The Hall thruster provided an alternative solution but required two larger, Taurus launch vehicles. Arcjet thrusters did not allow for significant launch vehicle reduction in the Middle Explorer class.

  6. Kepler Mission Design

    NASA Technical Reports Server (NTRS)

    Koch, David; Borucki, William; Lissauer, J.; Mayer, David; Voss, Janice; Basri, Gibor; Gould, Alan; Brown, Timothy; Cockran, William; Caldwell, Douglas

    2005-01-01

    The Kepler Mission is in the development phase with launch planned for 2007. The mission goal first off is to reliably detect a significant number of Earth-size planets in the habitable zone of solar-like stars. The mission design allows for exploring the diversity of planetary sizes, orbital periods, stellar spectral types, etc. In this paper we describe the technical approach taken for the mission design; describing the flight and ground system, the detection methodology, the photometer design and capabilities, and the way the data are taken and processed. (For Stellar Classification program. Finally the detection capability in terms of planet size and orbit are presented as a function of mission duration and stellar type.

  7. Propulsion Options for the LISA Mission

    NASA Technical Reports Server (NTRS)

    Cardiff, Eric H.; Marr, Gregory C.

    2004-01-01

    The LISA mission is a constellation of three spacecraft operating at 1 AU from the Sun in a position trailing the Earth. After launch, a propulsion module provides the AV necessary to reach this operational orbit, and separates from the spacecraft. A second propulsion system integrated with the spacecraft maintains the operational orbit and reduces nongravitational disturbances on the instruments. Both chemical and electrical propulsion systems were considered for the propulsion module, and this trade is presented to show the possible benefits of an EP system. Several options for the orbit maintenance and disturbance reduction system are also briefly discussed, along with several important requirements that suggest the use of a FEEP thruster system.

  8. Venus 2000 Mission Design

    NASA Astrophysics Data System (ADS)

    Folta, David; Marr, Greg; Vaughn, Frank; Houghton, Martin B.

    1997-05-01

    As part of the Discovery Program, National Aeronautics and Space Administration (NASA) has solicited proposals for inter-planetary research to conduct solar system exploration science investigations. A mission, called Venus 2000 (V2k), has been proposed for exploration of the Venus Atmosphere. This is NASAs first voyage to Venus to investigate key science objectives since Magellan and will be launched in summer 2002. In keeping with discovery program requirements to reduce total mission cost and utilize new technology, V2k mission design and control will focus on the use of innovative and proven trajectory analysis programs and control systems provided by the Goddard Space Flight Center (GSFC).

  9. Electric Power System Technology Options for Lunar Surface Missions

    NASA Technical Reports Server (NTRS)

    Kerslake, Thomas W.

    2005-01-01

    In 2004, the President announced a 'Vision for Space Exploration' that is bold and forward-thinking, yet practical and responsible. The vision explores answers to longstanding questions of importance to science and society and will develop revolutionary technologies and capabilities for the future, while maintaining good stewardship of taxpayer dollars. One crucial technology area enabling all space exploration is electric power systems. In this paper, the author evaluates surface power technology options in order to identify leading candidate technologies that will accomplish lunar design reference mission three (LDRM-3). LDRM-3 mission consists of multiple, 90-day missions to the lunar South Pole with 4-person crews starting in the year 2020. Top-level power requirements included a nominal 50 kW continuous habitat power over a 5-year lifetime with back-up or redundant emergency power provisions and a nominal 2-kW, 2-person unpressurized rover. To help direct NASA's technology investment strategy, this lunar surface power technology evaluation assessed many figures of merit including: current technology readiness levels (TRLs), potential to advance to TRL 6 by 2014, effectiveness of the technology to meet the mission requirements in the specified time, mass, stowed volume, deployed area, complexity, required special ground facilities, safety, reliability/redundancy, strength of industrial base, applicability to other LDRM-3 elements, extensibility to Mars missions, costs, and risks. For the 50-kW habitat module, dozens of nuclear, radioisotope and solar power technologies were down-selected to a nuclear fission heat source with Brayton, Stirling or thermoelectric power conversion options. Preferred energy storage technologies included lithium-ion battery and Proton Exchange Membrane (PEM) Regenerative Fuel Cells (RFC). Several AC and DC power management and distribution architectures and component technologies were defined consistent with the preferred habitat

  10. Mars Observer mission design

    NASA Technical Reports Server (NTRS)

    Beerer, Joseph G.; Horvat, Glen M.; Roncoli, Ralph B.

    1989-01-01

    The spacecraft for the Mars Observer mission is described, and an interplanetary trajectory design maximizing the spacecraft dry mass delivered into its mapping orbit is presented, along with an orbit insertion strategy minimizing spacecraft propulsive requirements. Emphasis is placed on the mapping orbit designed to meet the science requirements for a low-altitude near-circular near-polar orbit which is sun-synchronous with the dayside equatorial crossing at 2 PM local mean solar time. Additional requirements on the design are that the mapping orbit have a repeating groundtrack of no more than 7 sols and comply with the NASA Planetary Protection requirements. It is planned to operate the spacecraft and instruments in a repetitive fashion to minimize mission operation complexity and cost.

  11. Pioneer probe mission with orbiter option

    NASA Technical Reports Server (NTRS)

    1975-01-01

    A spacecraft is described which is based on Pioneer 10 and 11, and existing propulsion technology; it can transport and release a probe for entry into Jupiter's atmosphere, and subsequently maneuver to place the spacecraft in orbit about Jupiter. Orbital operations last 3 years and include maneuvers to provide multiple close satellite encounters which allow the orbit to be significantly changed to explore different parts of the magnetosphere. A mission summary, a guide to related documents, and background information about Jupiter are presented along with mission analysis over the complete mission profile. Other topics discussed include the launch, interplanetary flight, probe release and orbit deflection, probe entry, orbit selection, orbit insertion, periapsis raising, spacecraft description, and the effects of Jupiter's radiation belt on both orbiter and the probe.

  12. Lowest cost, nearest term options for a manned Mars mission

    NASA Technical Reports Server (NTRS)

    Sauls, Bob; Mortensen, Michael; Myers, Renee; Guacci, Giovanni; Montes, Fred

    1992-01-01

    This study is part of a NASA/USRA Advanced Design Program project executed for the purpose of examining the requirements of a first manned Mars mission. The mission, classified as a split/sprint mission, has been designed for a crew of six with a total manned trip time of one year.

  13. Current trajectory options for a comet nucleus sample return mission

    NASA Astrophysics Data System (ADS)

    Sauer, Carl G., Jr.

    1992-08-01

    A summary of the current trajectory options available for the ESA comet nucleus sample return mission, Rosetta, is presented. These options include direct trajectories, delta-V-EGA trajectories using a Titan IV/Centaur launch vehicle with upgraded solid rocket motors, a trajectory involving a gravity assist of the earth (VEGA) prior to comet rendezvous, and one involving an additional gravity assist of the earth (VEEGA). Other propulsion options proposed and discussed are solar electric propulsion/ballistic trajectory modes and nuclear electric propulsion trajectory modes. Tables of performance data for each of these trajectory options are given.

  14. AXTAR: Mission Design Concept

    NASA Technical Reports Server (NTRS)

    Ray, Paul S.; Chakrabarty, Deepto; Wilson-Hodge, Colleen A.; Philips, Bernard F.; Remillard, Ronald A.; Levine, Alan M.; Wood, Kent S.; Wolff, Michael T.; Gwon, Chul S.; Strohmayer, Tod E.; Briggs, Michael S.; Capizzo, Peter; Fabisinski, Leo; Hopkins, Randall C.; Hornsby, Linda S.; Johnson, Les; Maples, C. Dauphne; Miernik, Janie H.; Thomas, Dan; DeGeronimo, Gianluigi

    2010-01-01

    The Advanced X-ray Timing Array (AXTAR) is a mission concept for X-ray timing of compact objects that combines very large collecting area, broadband spectral coverage, high time resolution, highly flexible scheduling, and an ability to respond promptly to time-critical targets of opportunity. It is optimized for sub-millisecond timing of bright Galactic X-ray sources in order to study phenomena at the natural time scales of neutron star surfaces and black hole event horizons, thus probing the physics of ultra-dense matter, strongly curved spacetimes, and intense magnetic fields. AXTAR s main instrument, the Large Area Timing Array (LATA) is a collimated instrument with 2 50 keV coverage and over 3 square meters effective area. The LATA is made up of an array of super-modules that house 2-mm thick silicon pixel detectors. AXTAR will provide a significant improvement in effective area (a factor of 7 at 4 keV and a factor of 36 at 30 keV) over the RXTE PCA. AXTAR will also carry a sensitive Sky Monitor (SM) that acts as a trigger for pointed observations of X-ray transients in addition to providing high duty cycle monitoring of the X-ray sky. We review the science goals and technical concept for AXTAR and present results from a preliminary mission design study

  15. Advanced propulsion options for the Mars cargo mission

    NASA Technical Reports Server (NTRS)

    Frisbee, Robert H.; Blandino, John J.; Sercel, Joel C.; Sargent, Mark S.; Gowda, Nandini

    1990-01-01

    Several advanced propulsion options for a split-mission piloted Mars exploration scenario are presented. The primary study focus is on identifying concepts that can reduce total initial mass in low earth orbit (IMLEO) for the cargo delivery portion of the mission; in addition, concepts that can reduce the trip time of the piloted option are assessed. The propulsion options considered are nuclear thermal propulsion, solar sails, multimegawatt-class nuclear electric propulsion, solar electric propulsion, magnetic sails, mass drivers, rail guns, solar thermal rockets, beamed-energy propulsion systems, and tethers. For the cargo mission, solar sails are found to provide the greatest mass savings over the baseline chemical system, although they suffer from having very long trip times; a good performance compromise between a low IMLEO and a short trip time can be obtained using multimegawatt-class nuclear electric propulsion systems.

  16. Swarming UAVs mission design strategy

    NASA Astrophysics Data System (ADS)

    Lin, Kuo-Chi

    2007-04-01

    This paper uses a behavioral hierarchy approach to reduce the mission solution space and make the mission design easier. A UAV behavioral hierarchy is suggested, which is derived from three levels of behaviors: basic, individual and group. The individual UAV behavior is a combination of basic, lower level swarming behaviors with priorities. Mission design can be simplified by picking the right combination of individual swarming behaviors, which will emerge the needed group behaviors. Genetic Algorithm is used in both lower-level basic behavior design and mission design.

  17. 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.

  18. STEREO Mission Design Implementation

    NASA Technical Reports Server (NTRS)

    Guzman, Jose J.; Dunham, David W.; Sharer, Peter J.; Hunt, Jack W.; Ray, J. Courtney; Shapiro, Hongxing S.; Ossing, Daniel A.; Eichstedt, John E.

    2007-01-01

    STEREO (Solar-TErrestrial RElations Observatory) is the third mission in the Solar Terrestrial Probes program (STP) of the National Aeronautics and Space Administration (NASA) Science Mission Directorate Sun-Earth Connection theme. This paper describes the successful implementation (lunar swingby targeting) of the mission following the first phasing orbit to deployment into the heliocentric mission orbits following the two lunar swingbys. The STEREO Project had to make some interesting trajectory decisions in order to exploit opportunities to image a bright comet and an unusual lunar transit across the Sun.

  19. Design Evolution Study - Aging Options

    SciTech Connect

    P. McDaniel

    2002-04-05

    The purpose of this study is to identify options and issues for aging commercial spent nuclear fuel received for disposal at the Yucca Mountain Mined Geologic Repository. Some early shipments of commercial spent nuclear fuel to the repository may be received with high-heat-output (younger) fuel assemblies that will need to be managed to meet thermal goals for emplacement. The capability to age as much as 40,000 metric tons of heavy metal of commercial spent nuclear he1 would provide more flexibility in the design to manage this younger fuel and to decouple waste receipt and waste emplacement. The following potential aging location options are evaluated: (1) Surface aging at four locations near the North Portal; (2) Subsurface aging in the permanent emplacement drifts; and (3) Subsurface aging in a new subsurface area. The following aging container options are evaluated: (1) Complete Waste Package; (2) Stainless Steel inner liner of the waste package; (3) Dual Purpose Canisters; (4) Multi-Purpose Canisters; and (5) New disposable canister for uncanistered commercial spent nuclear fuel. Each option is compared to a ''Base Case,'' which is the expected normal waste packaging process without aging. A Value Engineering approach is used to score each option against nine technical criteria and rank the options. Open issues with each of the options and suggested future actions are also presented. Costs for aging containers and aging locations are evaluated separately. Capital costs are developed for direct costs and distributable field costs. To the extent practical, unit costs are presented. Indirect costs, operating costs, and total system life cycle costs will be evaluated outside of this study. Three recommendations for aging commercial spent nuclear fuel--subsurface, surface, and combined surface and subsurface are presented for further review in the overall design re-evaluation effort. Options that were evaluated but not recommended are: subsurface aging in a new

  20. A Titan exploration study: Science, technology and mission planning options, volume 1

    NASA Technical Reports Server (NTRS)

    Tindle, E. L.; Manning, L. A.; Sadin, S. R.; Edsinger, L. E.; Weissman, P. R.; Swenson, B. L.

    1976-01-01

    Mission concepts and technology advancements that can be used in the exploration of the outer planet satellites were examined. Titan, the seventh satellite of Saturn was selected as the target of interest. Science objectives for Titan exploration were identified, and recommended science payloads for four basic mission modes were developed (orbiter, atmospheric probe, surface penetrator and lander). Trial spacecraft and mission designs were produced for the various mission modes. Using these trial designs as a base, technology excursions were then made to find solutions to the problems resulting from these conventional approaches and to uncover new science, technology and mission planning options. Several mission modes were developed that take advantage of the unique conditions expected at Titan. They include a combined orbiter, atmosphere probe and lander vehicle, a combined probe and surface penetrator configuration and concepts for advanced remote sensing orbiters.

  1. Mission Options Scoping Tool for Mars Orbiters: Mass Cost Calculator (MC2)

    NASA Technical Reports Server (NTRS)

    Sturm, Eric J., II; Deutsch, Marie-Jose; Harmon, Corey; Nakagawa, Roy; Kinsey, Robert; Lopez, Nino; Kudrle, Paul; Evans, Alex

    2007-01-01

    Prior to developing the details of an advanced mission study, the mission architecture trade space is typically explored to assess the scope of feasible options. This paper describes the main features of an Excel-based tool, called the Mass-Cost-Calculator (MC2 ), which is used to perform rapid, high-level mass and cost options analyses of Mars orbiter missions. MC2 consists of a combination of databases, analytical solutions, and parametric relationships to enable quick evaluation of new mission concepts and comparison of multiple architecture options. The tool's outputs provide program management and planning teams with answers to "what if" queries, as well as an understanding of the driving mission elements, during the pre-project planning phase. These outputs have been validated against the outputs generated by the Advanced Projects Design Team (Team X) at NASA's Jet Propulsion Laboratory (JPL). The architecture of the tool allows for future expansion to other orbiters beyond Mars, and to non-orbiter missions, such as those involving fly-by spacecraft, probes, landers, rovers, or other mission elements.

  2. Power generation technology options for a Mars mission

    NASA Technical Reports Server (NTRS)

    Bozek, John M.; Cataldo, Robert L.

    1994-01-01

    The power requirements and resultant power system performances of an aggressive Mars mission are characterized. The power system technologies discussed will support both cargo and piloted space transport vehicles as well as a six-person crew on the Martian surface for 600 days. The mission uses materials transported by cargo vehicles and materials produced using in-situ planetary feed stock to establish a life-support cache and infrastructure for the follow-on piloted lander. Numerous power system technical options are sized to meet the mission power requirements using conventional and solar, nuclear, and wireless power transmission technologies for stationary, mobile surface, and space applications. Technology selections will depend on key criteria such as mass, volume, area, maturity, and application flexibility.

  3. Power Generation Technology Options for a Mars Mission

    NASA Astrophysics Data System (ADS)

    Bozek, John M.; Cataldo, Robert L.

    1994-07-01

    The power requirements and resultant power system performances of an aggressive Mars mission are characterized. The power system technologies discussed will support both cargo and piloted space transport vehicles as well as a six person crew on the Martian surface for 600 days. The mission utilizes materials transported by cargo vehicles and materials produced using in-situ planetary feed stock to establish a life-support cache and infrastructure for the follow-on piloted lander. Numerous power system technical options are sized to meet the mission power requirements using both conventional and novel solar, nuclear, and wireless power transmission technologies for both stationary and mobile surface applications as well as space applications. Technology selections will depend on key criteria such as mass, volume, area, maturity, and application flexibility.

  4. Evolution of Orion Mission Design for Exploration Mission 1 and 2

    NASA Technical Reports Server (NTRS)

    Gutkowski, Jeffrey P.; Dawn, Timothy F.; Jedrey, Richard M.

    2016-01-01

    The evolving mission design and concepts of NASA’s next steps have shaped Orion into the spacecraft that it is today. Since the initial inception of Orion, through the Constellation Program, and now in the Exploration Mission frame-work with the Space Launch System (SLS), each mission design concept and pro-gram goal have left Orion with a set of capabilities that can be utilized in many different mission types. Exploration Missions 1 and 2 (EM-1 and EM-2) have now been at the forefront of the mission design focus for the last several years. During that time, different Design Reference Missions (DRMs) were built, analyzed, and modified to solve or mitigate enterprise level design trades to ensure a viable mission from launch to landing. The resulting DRMs for EM-1 and EM-2 were then expanded into multi-year trajectory scans to characterize vehicle performance as affected by variations in Earth-Moon geometry. This provides Orion’s subsystems with stressing reference trajectories to help design their system. Now that Orion has progressed through the Preliminary and Critical Design Reviews (PDR and CDR), there is a general shift in the focus of mission design from aiding the vehicle design to providing mission specific products needed for pre-flight and real time operations. Some of the mission specific products needed include, large quantities of nominal trajectories for multiple monthly launch periods and abort options at any point in the mission for each valid trajectory in the launch window.

  5. Proximity Operations for the Robotic Boulder Capture Option for the Asteroid Redirect Mission

    NASA Technical Reports Server (NTRS)

    Reeves, David M.; Naasz, Bo J.; Wright, Cinnamon A.; Pini, Alex J.

    2014-01-01

    In September of 2013, the Asteroid Robotic Redirect Mission (ARRM) Option B team was formed to expand on NASA's previous work on the robotic boulder capture option. While the original Option A concept focuses on capturing an entire smaller Near-Earth Asteroid (NEA) using an inflatable bag capture mechanism, this design seeks to land on a larger NEA and retrieve a boulder off of its surface. The Option B team has developed a detailed and feasible mission concept that preserves many aspects of Option A's vehicle design while employing a fundamentally different technique for returning a significant quantity of asteroidal material to the Earth-Moon system. As part of this effort, a point of departure proximity operations concept was developed complete with a detailed timeline, as well as DeltaV and propellant allocations. Special attention was paid to the development of the approach strategy, terminal descent to the surface, controlled ascent with the captured boulder, and control during the Enhanced Gravity Tractor planetary defense demonstration. The concept of retrieving a boulder from the surface of an asteroid and demonstrating the Enhanced Gravity Tractor planetary defense technique is found to be feasible and within the proposed capabilities of the Asteroid Redirect Vehicle (ARV). While this point of departure concept initially focuses on a mission to Itokawa, the proximity operations design is also shown to be extensible to wide range of asteroids.

  6. US Decadal Survey Outer Solar System Missions: Trajectory Options

    NASA Astrophysics Data System (ADS)

    Spilker, T. R.; Atkinson, D. H.; Strange, N. J.; Landau, D.

    2012-04-01

    The report of the US Planetary Science Decadal Survey (PSDS), released in draft form March 7, 2011, identifies several mission concepts involving travel to high-priority outer solar system (OSS) destinations. These include missions to Europa and Jupiter, Saturn and two of its satellites, and Uranus. Because travel to the OSS involves much larger distances and larger excursions out of the sun's gravitational potential well than inner solar system (ISS) missions, transfer trajectories for OSS missions are stronger drivers of mission schedule and resource requirements than for ISS missions. Various characteristics of each planet system, such as obliquity, radiation belts, rings, deep gravity wells, etc., carry ramifications for approach trajectories or trajectories within the systems. The maturity of trajectory studies for each of these destinations varies significantly. Europa has been the focus of studies for well over a decade. Transfer trajectory options from Earth to Jupiter are well understood. Current studies focus on trajectories within the Jovian system that could reduce the total mission cost of a Europa orbiter mission. Three missions to the Saturn system received high priority ratings in the PSDS report: two flagship orbital missions, one to Titan and one to Enceladus, and a Saturn atmospheric entry probe mission for NASA's New Frontiers Program. The Titan Saturn System Mission (TSSM) studies of 2007-2009 advanced our understanding of trajectory options for transfers to Saturn, including solar electric propulsion (SEP) trajectories. But SEP trajectories depend more on details of spacecraft and propulsion system characteristics than chemical trajectories, and the maturity of SEP trajectory search tools has not yet caught up with chemical trajectory tools, so there is still more useful research to be done on Saturn transfers. The TSSM studies revealed much about Saturn-orbiting trajectories that yield efficient and timely delivery to Titan or Enceladus

  7. Preliminary design of an asteroid hopping mission

    NASA Astrophysics Data System (ADS)

    Scheppa, Michael D.

    In 2010, NASA announced that its new vision is to support private space launch operations. It is anticipated that this new direction will create the need for new and innovative ideas that push the current boundaries of space exploration and contain the promise of substantial gain, both in research and capital. The purpose of the study is to plan and estimate the feasibility of a mission to visit a number of near Earth asteroids (NEAs). The mission would take place before the end of the 21st century, and would only use commercially available technology. Throughout the mission design process, while holding astronaut safety paramount, it was the goal to maximize the return while keeping the cost to a minimum. A mission of the nature would appeal to the private space industry because it could be easily adapted and set into motion. The mission design was divided into three main parts; mission timeline, vehicle design and power sources, with emphasis on nuclear and solar electric power, were investigated. The timeline and associated trajectories were initially selected using a numerical estimation and then optimized using Satellite Tool Kit (STK) 9.s's Design Explorer Optimizer [1]. Next, the spacecraft was design using commercially available parts that would support the mission requirements. The Variable Specific Impulse Magnetoplasma Rocket (VASIMR) was and instrumental piece in maximizing the number of NEAs visited. Once the spacecraft was designed, acceptable power supply options were investigated. The VASIMR VX-200 requires 200 kilowatts of power to maintain thrust. This creates the need for a substantial power supply that consists of either a nuclear reactor of massive solar arrays. STK 9.1's Design Explorer Optimizer was able to create a mission time line that allowed for the exploration of seven NEAs in under two years, while keeping the total mission DeltaV under 71 kilometers per second. Based on these initial findings, it is determined that a mission of this

  8. Spitzer Space Telescope mission design

    NASA Technical Reports Server (NTRS)

    Kwok, Johnny H.; Garcia, Mark D.; Bonfiglio, Eugene; Long, Stacia M.

    2004-01-01

    This paper gives a description of the mission design, launch, orbit, and navigation results for the Spitzer space telescope mission. The Spitzer telescope was launched by the Delta II Heavy launch vehicle into a heliocentric Earth trailing orbit. This orbit is flown for the first time and will be used by several future astronomical missions such as Kepler, SIM, and LISA. This paper describes the launch strategy for a winter versus a summer launch and how it affects communications. It also describes how the solar orbit affects the design and operations of the Observatory. It describes the actual launch timeline, launch vehicle flight performance, and the long term behavior of the as flown orbit. It also provides the orbit knowledge from in-flight navigation data.

  9. Interplanetary mission design techniques for flagship-class missions

    NASA Astrophysics Data System (ADS)

    Kloster, Kevin W.

    Trajectory design, given the current level of propulsive technology, requires knowledge of orbital mechanics, computational resources, extensive use of tools such as gravity-assist and V infinity leveraging, as well as insight and finesse. Designing missions that deliver a capable science package to a celestial body of interest that are robust and affordable is a difficult task. Techniques are presented here that assist the mission designer in constructing trajectories for flagship-class missions in the outer Solar System. These techniques are applied in this work to spacecraft that are currently in flight or in the planning stages. By escaping the Saturnian system, the Cassini spacecraft can reach other destinations in the Solar System while satisfying planetary quarantine. The patched-conic method was used to search for trajectories that depart Saturn via gravity assist at Titan. Trajectories were found that fly by Jupiter to reach Uranus or Neptune, capture at Jupiter or Neptune, escape the Solar System, fly by Uranus during its 2049 equinox, or encounter Centaurs. A "grand tour," which visits Jupiter, Uranus, and Neptune, departs Saturn in 2014. New tools were built to search for encounters with Centaurs, small Solar System bodies between the orbits of Jupiter and Neptune, and to minimize the DeltaV to target these encounters. Cassini could reach Chiron, the first-discovered Centaur, in 10.5 years after a 2022 Saturn departure. For a Europa Orbiter mission, the strategy for designing Jovian System tours that include Io flybys differs significantly from schemes developed for previous versions of the mission. Assuming that the closest approach distance of the incoming hyperbola at Jupiter is below the orbit of Io, then an Io gravity assist gives the greatest energy pump-down for the least decrease in perijove radius. Using Io to help capture the spacecraft can increase the savings in Jupiter orbit insertion DeltaV over a Ganymede-aided capture. The tour design is

  10. Mission Architecture and Technology Options for a Flagship Class Venus In Situ Mission

    NASA Technical Reports Server (NTRS)

    Balint, Tibor S.; Kwok, Johnny H.; Kolawa, Elizabeth A.; Cutts, James A.; Senske, David A.

    2008-01-01

    Venus, as part of the inner triad with Earth and Mars, represents an important exploration target if we want to learn more about solar system formation and evolution. Comparative planetology could also elucidate the differences between the past, present, and future of these three planets, and can help with the characterization of potential habitable zones in our solar system and, by extension, extrasolar systems. A long lived in situ Venus mission concept, called the Venus Mobile Explorer, was prominently featured in NASA's 2006 SSE Roadmap and supported in the community White Paper by the Venus Exploration Analysis Group (VEXAG). Long-lived in situ missions are expected to belong to the largest (Flagship) mission class, which would require both enabling and enhancing technologies beside mission architecture options. Furthermore, extreme environment mitigation technologies for Venus are considered long lead development items and are expected to require technology development through a dedicated program. To better understand programmatic and technology needs and the motivating science behind them, in this fiscal year (FY08) NASA is funding a Venus Flaghip class mission study, based on key science and technology drivers identified by a NASA appointed Venus Science and Technology Definition Team (STDT). These mission drivers are then assembled around a suitable mission architecture to further refine technology and cost elements. In this paper we will discuss the connection between the final mission architecture and the connected technology drivers from this NASA funded study, which - if funded - could enable a future Flagship class Venus mission and potentially drive a proposed Venus technology development program.

  11. MIOSAT Mission Scenario and Design

    NASA Astrophysics Data System (ADS)

    Agostara, C.; Dionisio, C.; Sgroi, G.; di Salvo, A.

    2008-08-01

    MIOSAT ("Mssione Ottica su microSATellite") is a low-cost technological / scientific microsatellite mission for Earth Observation, funded by Italian Space Agency (ASI) and managed by a Group Agreement between Rheinmetall Italia - B.U. Spazio - Contraves as leader and Carlo Gavazzi Space as satellite manufacturer. Several others Italians Companies, SME and Universities are involved in the development team with crucial roles. MIOSAT is a microsatellite weighting around 120 kg and placed in a 525 km altitude sun-synchronuos circular LEO orbit. The microsatellite embarks three innovative optical payloads: Sagnac multi spectral radiometer (IFAC-CNR), Mach Zehender spectrometer (IMM-CNR), high resolution pancromatic camera (Selex Galileo). In addition three technological experiments will be tested in-flight. The first one is an heat pipe based on Marangoni effect with high efficiency. The second is a high accuracy Sun Sensor using COTS components and the last is a GNSS SW receiver that utilizes a Leon2 processor. Finally a new generation of 28% efficiency solar cells will be adopted for the power generation. The platform is highly agile and can tilt along and cross flight direction. The pointing accuracy is in the order of 0,1° for each axe. The pointing determination during images acquisition is <0,02° for the axis normal to the boresight and 0,04° for the boresight. This paper deals with MIOSAT mission scenario and definition, highlighting trade-offs for mission implementation. MIOSAT mission design has been constrained from challenging requirements in terms of satellite mass, mission lifetime, instrument performance, that have implied the utilization of satellite agility capability to improve instruments performance in terms of S/N and resolution. The instruments provide complementary measurements that can be combined in effective ways to exploit new applications in the fields of atmosphere composition analysis, Earth emissions, antropic phenomena, etc. The Mission

  12. Evolution of Orion Mission Design for Exploration Mission 1 and 2

    NASA Technical Reports Server (NTRS)

    Gutkowski, Jeffrey P.; Dawn, Timothy F.; Jedrey, Richard M.

    2016-01-01

    The evolving mission design and concepts of NASA's next steps have shaped Orion into the spacecraft that it is today. Since the initial inception of Orion, through the Constellation Program, and now in the Exploration Mission frame-work with the Space Launch System (SLS), each mission design concept and program goal have left Orion with a set of capabilities that can be utilized in many different mission types. Exploration Missions 1 and 2 (EM-1 and EM-2) have now been at the forefront of the mission design focus for the last several years. During that time, different Design Reference Missions (DRMs) were built, analyzed, and modified to solve or mitigate enterprise level design trades to ensure a viable mission from launch to landing. The resulting DRMs for EM-1 and EM-2 were then expanded into multi-year trajectory scans to characterize vehicle performance and Earth-Moon geometry trends. This provides Orion's subsystems with stressing reference trajectories to help design their system. Now that Orion has progressed through the Preliminary and Critical Design Re-views (PDR and CDR) there is a general shift in the focus of mission design from aiding the vehicle design to providing mission specific products needed for pre-flight and real time operations. Some of the mission specific products need-ed include analysis of steering law performance, inputs into navigational accura-cy assessments, abort options at any point in the mission for each valid trajecto-ry in the launch window, recontact avoidance between the upper stage and Orion post nominal separation, etc.

  13. Pioneer Mars surface penetrator mission. Mission analysis and orbiter design

    NASA Technical Reports Server (NTRS)

    1974-01-01

    The Mars Surface Penetrator mission was designed to provide a capability for multiple and diverse subsurface science measurements at a low cost. Equipment required to adapt the Pioneer Venus spacecraft for the Mars mission is described showing minor modifications to hardware. Analysis and design topics which are similar and/or identical to the Pioneer Venus program are briefly discussed.

  14. Options for the Further Orbit Evolution of the Swarm Mission

    NASA Astrophysics Data System (ADS)

    Sieg, Detlef; Diekmann, Frank

    2016-08-01

    The three satellites of ESA's magnetic field mission Swarm were launched into a common low Earth circular orbit in November 2013 to measure precisely the magnetic signals from Earth's core, mantle, crust and oceans, as well as the ionosphere and magnetosphere. Since completion of the orbit acquisition phase in April 2014 one satellite (Swarm-B) is flying in a higher orbit with an inclination of 87.8deg and an altitude decaying from 520km. The other two satellites are Swarm-A (trailing) and Swarm-C (leading). They form the lower pair with an initial altitude of 473km, an inclination of 87.4 deg and an ascending node difference of 1.4 deg. The original mission analysis foresaw a decay of the lower pair down to 300km altitude within 4 years after launch. The target altitude of the launcher injection orbit was selected accordingly with some margin due to uncertainties in the solar activity prediction. However the final altitude selection had to be provided more than half a year before launch. Following several launch delays, the major part of the mission falls now beyond the maximum of the current solar cycle. Because of the lower radio flux and geomagnetic activity, the air drag forces are now much lower and the actual decay takes longer.As a first countermeasure the target for the inclination difference between Swarm-B and Swarm-A/C was reduced to 0.4deg shortly before the start of the orbit acquisition manoeuvre sequence early 2014 such that the LTAN drift between the orbit planes of B and A/C has been reduced to 1.5h per year to avoid a too large difference towards the end of the mission.First the paper describes the routine orbit determination approach by ESOC flight dynamics, which is used to determine absolute drag scale factors. Based on the in- flight calibrated values, long-term orbit predictions are calculated every half a year and can be compared against the actual observed decay. This gives good confidence for the prediction of the future altitude

  15. Early SP-100 flight mission designs

    NASA Astrophysics Data System (ADS)

    Josloff, Allan T.; Shepard, Neal F.; Kirpich, Aaron S.; Murata, Ronald; Smith, Michael A.; Stephen, James D.

    1993-01-01

    Early flight mission objectives can be met with a Space Reactor Power System (SRPS) using thermoelectric conversion in conjunction with fast spectrum, lithium-cooled reactors. This paper describes two system design options using thermoelectric technology to accommodate an early launch. In the first of these options, radiatively coupled Radioiosotope Thermoelectric Generator (RTG) unicouples are adapted for use with a SP-100-type reactor heat source. Unicouples have been widely used as the conversion technology in RTGs and have demonstrated the long-life characteristics necessary for a highly relible SRPS. The thermoelectric leg height is optimized in conjunction with the heat rejection temperature to provide a mass optimum 6-kWe system configured for launch on a Delta II launch vehicle. The flight-demonstrated status of this conversion technology provides a high confidence that such a system can be designed, assembled, tested, and launched by 1997. The use of a SP-100-type reactor assures compliance with safety requirements and expedites the flight safety approval process while, at the same time, providing flight performance verification for a heat source technology with the growth potential to meet future national needs for higher power levels. A 15-kW2, Atlas IIAS-launched system using the compact, conductively coupled multicouple converters being developed under the SP-100 program to support an early flight system launch also described. Both design concepts have been scaled to 20-kWe in order to support recent studies by DOE/NASA for higher power early launch missions.

  16. Trajectory design for a Mars Rover/Sample Return mission

    NASA Astrophysics Data System (ADS)

    Sweetser, Theodore H.

    This paper discusses two of the orbit design problems faced in the design of a Mars Rover/Sample Return mission, which is currently being studied at the Jet Propulsion Laboratory. The first is the problem of interplanetary transfer - what is the best trajectory for getting equipment to Mars and a sample back. Several kinds of trajectories are examined before the conclusion is made that straightforward direct transfers are best. The second orbit design problem is what kind of orbit around Mars is best for making high-resolution maps of sites where the rover could land and gather samples, and how can the same orbiter be used as a relay between a rover on Mars and ground stations on Earth. This question is examined in the context of alternate mission options being considered, and the answer depends on the requirements of the particular mission option.

  17. Trajectory Design Considerations for Exploration Mission 1

    NASA Technical Reports Server (NTRS)

    Dawn, Timothy F.; Gutkowski, Jeffrey P.; Batcha, Amelia L.

    2017-01-01

    Exploration Mission 1 (EM-1) will be the first mission to send an uncrewed Orion vehicle to cislunar space in 2018, targeted to a Distant Retrograde Orbit (DRO). Analysis of EM-1 DRO mission opportunities in 2018 help characterize mission parameters that are of interest to other subsystems (e.g., power, thermal, communications, flight operations, etc). Subsystems request mission design trades which include: landing lighting, addition of an Orion main engine checkout burn, and use of auxiliary thruster only cases. This paper examines the evolving trade studies that incorporate subsystem feedback and demonstrate the feasibility of these constrained mission trajectory designs and contingencies.

  18. L1C signal design options

    USGS Publications Warehouse

    Betz, J.W.; Cahn, C.R.; Dafesh, P.A.; Hegarty, C.J.; Hudnut, K.W.; Jones, A.J.; Keegan, R.; Kovach, K.; Lenahan, L.S.; Ma, H.H.; Rushanan, J.J.; Stansell, T.A.; Wang, C.C.; Yi, S.K.

    2006-01-01

    Design activities for a new civil signal centered at 1575.42 MHz, called L1C, began in 2003, and the Phase 1 effort was completed in 2004. The L1C signal design has evolved and matured during a Phase 2 design activity that began in 2005. Phase 2 has built on the initial design activity, guided by responses to international user surveys conducted during Phase 1. A common core of signal characteristics has been developed to provide advances in robustness and performance. The Phase 2 activity produced five design options, all drawing upon the core signal characteristics, while representing different blends of characteristics and capabilities. A second round of international user surveys was completed to solicit advice concerning these design options. This paper provides an update of the L1C design process, and describes the current L1C design options. Initial performance estimates are presented for each design option, displaying trades between signal tracking robustness, the speed and robustness of clock and ephemeris data, and the rate and robustness of other data message contents. Planned remaining activities are summarized, leading to optimization of the L1C design.

  19. Study of Power Options for Jupiter and Outer Planet Missions

    NASA Technical Reports Server (NTRS)

    Landis, Geoffrey A.; Fincannon, James

    2015-01-01

    Power for missions to Jupiter and beyond presents a challenging goal for photovoltaic power systems, but NASA missions including Juno and the upcoming Europa Clipper mission have shown that it is possible to operate solar arrays at Jupiter. This work analyzes photovoltaic technologies for use in Jupiter and outer planet missions, including both conventional arrays, as well as analyzing the advantages of advanced solar cells, concentrator arrays, and thin film technologies. Index Terms - space exploration, spacecraft solar arrays, solar electric propulsion, photovoltaic cells, concentrator, Fresnel lens, Jupiter missions, outer planets.

  20. Crew Transportation System Design Reference Missions

    NASA Technical Reports Server (NTRS)

    Mango, Edward J.

    2015-01-01

    Contains summaries of potential design reference mission goals for systems to transport humans to andfrom low Earth orbit (LEO) for the Commercial Crew Program. The purpose of this document is to describe Design Reference Missions (DRMs) representative of the end-to-end Crew Transportation System (CTS) framework envisioned to successfully execute commercial crew transportation to orbital destinations. The initial CTS architecture will likely be optimized to support NASA crew and NASA-sponsored crew rotation missions to the ISS, but consideration may be given in this design phase to allow for modifications in order to accomplish other commercial missions in the future. With the exception of NASA’s mission to the ISS, the remaining commercial DRMs are notional. Any decision to design or scar the CTS for these additional non-NASA missions is completely up to the Commercial Provider. As NASA’s mission needs evolve over time, this document will be periodically updated to reflect those needs.

  1. Mechanical design of the Mars Pathfinder mission

    NASA Technical Reports Server (NTRS)

    Eisen, Howard Jay; Buck, Carl W.; Gillis-Smith, Greg R.; Umland, Jeffrey W.

    1997-01-01

    The Mars Pathfinder mission and the Sojourner rover is reported on, with emphasis on the various mission steps and the performance of the technologies involved. The mechanical design of mission hardware was critical to the success of the entry sequence and the landing operations. The various mechanisms employed are considered.

  2. Space station needs, attributes, and architectural options: Mission requirements

    NASA Technical Reports Server (NTRS)

    Riel, F. D.

    1983-01-01

    Space station missions and their requirements are discussed. Analyses of the following four mission categories are summarized: (1) commercial, (2) technology, (3) operation, and (4) science and applications. The requirements determined by the study dictate a very strong need for a manned space station to satisfy the majority of the missions. The station is best located at a 28.5-deg inclination and initially (1992 era) requires a crew of four (three for mission payloads) and a mission power of 25 kW. A space platform in a polar orbit is needed to augment the station capability; it initially would be a 15-kW system, located in a sun-synchronous orbit.

  3. Lattice Design for ERL Options at SLAC

    SciTech Connect

    Nosochkov, Yuri; Cai, Yunhai; Huang, Xiaobiao; Wang, Min-Huey; /SLAC

    2011-06-02

    SLAC is investigating long-range options for building a high performance light source machine while reusing the existing linac and PEP-II tunnels. One previously studied option is the PEP-X low emittance storage ring. The alternative option is based on a superconducting Energy Recovery Linac (ERL) and the PEP-X design. The ERL advantages are the low beam emittance, short bunch length and small energy spread leading to better qualities of the X-ray beams. Two ERL configurations differed by the location of the linac have been studied. Details of the lattice design and the results of beam transport simulations with the coherent synchrotron radiation effects are presented.

  4. Trade Space Assessment for Human Exploration Mission Design

    NASA Technical Reports Server (NTRS)

    Joosten, B. Kent

    2006-01-01

    Many human space exploration mission architecture assessments have been performed over the years by diverse organizations and individuals. Direct comparison of metrics among these studies is extremely difficult due to widely varying assumptions involving projected technology readiness, mission goals, acceptable risk criteria, and socio-political environments. However, constant over the years have been the physical laws of celestial dynamics and rocket propulsion systems. A finite diverse yet finite architecture trade space should exist which captures methods of human exploration - particularly of the Moon and Mars - by delineating technical trades and cataloging the physically realizable options of each. A particular architectural approach should then have a traceable path through this "trade tree". It should be pointed out that not every permutation of paths will result in a physically realizable mission approach, but cataloging options that have been examined by past studies should help guide future analysis. This effort was undertaken in two phases by multi-center NASA working groups in the spring and summer of 2004 using more than thirty years of past studies to "flesh out" the Moon-Mars human exploration trade space. The results are presented, not as a "trade tree", which would be unwieldy, but as a "menu" of potential technical options as a function of mission phases. This is envisioned as a tool to aid future mission designers by offering guidance to relevant past analyses.

  5. Conceptual Design Methods and the Application of a Tradespace Modeling Tool for Deep Space Missions

    NASA Technical Reports Server (NTRS)

    Jones, Melissa A.; Chase, James P.

    2008-01-01

    Concept studies for deep space missions are typically time-consuming and costly, given the variety of missions and uniqueness of each design. Yet, in an increasingly cost-constrained environment, it is critical to identify the most scientifically valuable and cost-effective designs early in the design process. Modeling is an integral part in helping to identify the most desirable design option. While some spacecraft design models currently exist for Earth-orbiting spacecraft, there has been less success with deep space missions. Instead, these missions require a modified design and modeling approach to enable the same construction of a comprehensive, yet credible, mission tradespace. This paper presents an approach for efficiently constructing such a mission tradespace. In addition to a proposed design and modeling approach, three case study missions are presented including a solar orbiter, a Europa orbiter, and a near-Earth asteroid (NEA) sample return mission.

  6. Controlled Ecological Life Support Systems (CELSS) conceptual design option study

    NASA Technical Reports Server (NTRS)

    Oleson, Melvin; Olson, Richard L.

    1986-01-01

    Results are given of a study to explore options for the development of a Controlled Ecological Life Support System (CELSS) for a future Space Station. In addition, study results will benefit the design of other facilities such as the Life Sciences Research Facility, a ground-based CELSS demonstrator, and will be useful in planning longer range missions such as a lunar base or manned Mars mission. The objectives were to develop weight and cost estimates for one CELSS module selected from a set of preliminary plant growth unit (PGU) design options. Eleven Space Station CELSS module conceptual PGU designs were reviewed, components and subsystems identified and a sensitivity analysis performed. Areas where insufficient data is available were identified and divided into the categories of biological research, engineering research, and technology development. Topics which receive significant attention are lighting systems for the PGU, the use of automation within the CELSS system, and electric power requirements. Other areas examined include plant harvesting and processing, crop mix analysis, air circulation and atmosphere contaminant flow subsystems, thermal control considerations, utility routing including accessibility and maintenance, and nutrient subsystem design.

  7. Tracking system options for future altimeter satellite missions

    NASA Technical Reports Server (NTRS)

    Davis, G. W.; Rim, H. J.; Ries, J. C.; Tapley, B. D.

    1994-01-01

    Follow-on missions to provide continuity in the observation of the sea surface topography once the successful TOPEX/POSEIDON (T/P) oceanographic satellite mission has ended are discussed. Candidates include orbits which follow the ground tracks of T/P GEOSAT or ERS-1. The T/P precision ephemerides, estimated to be near 3 cm root-mean-square, demonstrate the radial orbit accuracy that can be achieved at 1300 km altitude. However, the radial orbit accuracy which can be achieved for a mission at the 800 km altitudes of GEOSAT and ERS-1 has not been established, and achieving an accuracy commensurate with T/P will pose a great challenge. This investigation focuses on the radial orbit accuracy that can be achieved for a mission in the GEOSAT orbit. Emphasis is given to characterizing the effects of force model errors on the estimated radial orbit accuracy, particularly those due to gravity and drag. The importance of global, continuous tracking of the satellite for reduction in these sources of orbit error is demonstrated with simulated GPS tracking data. A gravity tuning experiment is carried out to show how the effects of gravity error may be reduced. Assuming a GPS flight receiver with a full-sky tracking capability, the simulation results indicate that a 5 cm radial orbit accuracy for an altimeter satellite in GEOSAT orbit should be achievable during low-drag atmospheric conditions and after an acceptable tuning of the gravity model.

  8. Neptune aerocapture mission and spacecraft design overview

    NASA Technical Reports Server (NTRS)

    Bailey, Robert W.; Hall, Jeff L.; Spliker, Tom R.; O'Kongo, Nora

    2004-01-01

    A detailed Neptune aerocapture systems analysis and spacecraft design study was performed as part of NASA's In-Space Propulsion Program. The primary objectives were to assess the feasibility of a spacecraft point design for a Neptune/Triton science mission. That uses aerocapture as the Neptune orbit insertion mechanism. This paper provides an overview of the science, mission and spacecraft design resulting from that study.

  9. Network interface unit design options performance analysis

    NASA Technical Reports Server (NTRS)

    Miller, Frank W.

    1991-01-01

    An analysis is presented of three design options for the Space Station Freedom (SSF) onboard Data Management System (DMS) Network Interface Unit (NIU). The NIU provides the interface from the Fiber Distributed Data Interface (FDDI) local area network (LAN) to the DMS processing elements. The FDDI LAN provides the primary means for command and control and low and medium rate telemetry data transfers on board the SSF. The results of this analysis provide the basis for the implementation of the NIU.

  10. Transfemoral socket design and suspension options.

    PubMed

    Kapp, S L

    2000-08-01

    The transfemoral socket is the interface between the patient's residual limb and the remaining prosthesis. Whether the patient uses a technologically advanced knee mechanism and foot or a single axis, constant friction knee and solid ankle cushioned heel, no other prosthetic component is as crucial in assuring a comfortable and well-functioning prosthesis as the socket. This article discusses the rationale of quadrilateral and ischial containment socket designs and reviews suspension options.

  11. NASA'S RPS Design Reference Mission Set for Solar System Exploration

    NASA Astrophysics Data System (ADS)

    Balint, Tibor S.

    2007-01-01

    NASA's 2006 Solar System Exploration (SSE) Strategic Roadmap identified a set of proposed large Flagship, medium New Frontiers and small Discovery class missions, addressing key exploration objectives. These objectives respond to the recommendations by the National Research Council (NRC), reported in the SSE Decadal Survey. The SSE Roadmap is down-selected from an over-subscribed set of missions, called the SSE Design Reference Mission (DRM) set Missions in the Flagship and New Frontiers classes can consider Radioisotope Power Systems (RPSs), while small Discovery class missions are not permitted to use them, due to cost constraints. In line with the SSE DRM set and the SSE Roadmap missions, the RPS DRM set represents a set of missions, which can be enabled or enhanced by RPS technologies. At present, NASA has proposed the development of two new types of RPSs. These are the Multi-Mission Radioisotope Thermoelectric Generator (MMRTG), with static power conversion; and the Stirling Radioisotope Generator (SRG), with dynamic conversion. Advanced RPSs, under consideration for possible development, aim to increase specific power levels. In effect, this would either increase electric power generation for the same amount of fuel, or reduce fuel requirements for the same power output, compared to the proposed MMRTG or SRG. Operating environments could also influence the design, such that an RPS on the proposed Titan Explorer would use smaller fins to minimize heat rejection in the extreme cold environment; while the Venus Mobile Explorer long-lived in-situ mission would require the development of a new RPS, in order to tolerate the extreme hot environment, and to simultaneously provide active cooling to the payload and other electric components. This paper discusses NASA's SSE RPS DRM set, in line with the SSE DRM set. It gives a qualitative assessment regarding the impact of various RPS technology and configuration options on potential mission architectures, which could

  12. NASA's RPS Design Reference Mission Set for Solar System Exploration

    NASA Technical Reports Server (NTRS)

    Balint, Tibor S.

    2007-01-01

    NASA's 2006 Solar System Exploration (SSE) Strategic Roadmap identified a set of proposed large Flagship, medium New Frontiers and small Discovery class missions, addressing key exploration objectives. These objectives respond to the recommendations by the National Research Council (NRC), reported in the SSE Decadal Survey. The SSE Roadmap is down-selected from an over-subscribed set of missions, called the SSE Design Reference Mission (DRM) set. Missions in the Flagship and New Frontiers classes can consider Radioisotope Power Systems (RPSs), while small Discovery class missions are not permitted to use them, due to cost constraints. In line with the SSE DRM set and the SSE Roadmap missions, the RPS DRM set represents a set of missions, which can be enabled or enhanced by RPS technologies. At present, NASA has proposed the development of two new types of RPSs. These are the Multi-Mission Radioisotope Thermoelectric Generator (MMRTG), with static power conversion; and the Stirling Radioisotope Generator (SRG), with dynamic conversion. Advanced RPSs, under consideration for possible development, aim to increase specific power levels. In effect, this would either increase electric power generation for the same amount of fuel, or reduce fuel requirements for the same power output, compared to the proposed MMRTG or SRG. Operating environments could also influence the design, such that an RPS on the proposed Titan Explorer would use smaller fins to minimize heat rejection in the extreme cold environment; while the Venus Mobile Explorer long-lived in-situ mission would require the development of a new RPS, in order to tolerate the extreme hot environment, and to simultaneously provide active cooling to the payload and other electric components. This paper discusses NASA's SSE RPS DRM set, in line with the SSE DRM set. It gives a qualitative assessment regarding the impact of various RPS technology and configuration options on potential mission architectures, which could

  13. Xenia Mission: Spacecraft Design Concept

    NASA Technical Reports Server (NTRS)

    Hopkins, R. C.; Johnson, C. L.; Kouveliotou, C.; Jones, D.; Baysinger, M.; Bedsole, T.; Maples, C. C.; Benfield, P. J.; Turner, M.; Capizzo, P.; Fabinski, L.; Hornsby, L.; Thompson, K.; Miernik, J. H.; Percy, T.

    2009-01-01

    The proposed Xenia mission will, for the first time, chart the chemical and dynamical state of the majority of baryonic matter in the universe. using high-resolution spectroscopy, Xenia will collect essential information from major traces of the formation and evolution of structures from the early universe to the present time. The mission is based on innovative instrumental and observational approaches: observing with fast reaction gamma-ray bursts (GRBs) with a high spectral resolution. This enables the study of their (star-forming) environment from the dark to the local universe and the use of GRBs as backlight of large-scale cosmological structures, observing and surveying extended sources with high sensitivity using two wide field-of-view x-ray telescopes - one with a high angular resolution and the other with a high spectral resolution.

  14. Power System Options Evaluated for the Radiation and Technology Demonstration Mission

    NASA Technical Reports Server (NTRS)

    Kerslake, Thomas W.; Benson, Scott W.

    2000-01-01

    The Radiation and Technology Demonstration (RTD) Mission is under joint study by three NASA Centers: the NASA Johnson Space Center, the NASA Goddard Space Flight Center, and the NASA Glenn Research Center at Lewis Field. This Earth-orbiting mission, which may launch on a space shuttle in the first half of the next decade, has the primary objective of demonstrating high-power electric thruster technologies. Secondary objectives include better characterization of Earth's Van Allen trapped-radiation belts, measurement of the effectiveness of the radiation shielding for human protection, measurement of radiation effects on advanced solar cells, and demonstration of radiation-tolerant microelectronics. During the mission, which may continue up to 1 year, the 2000-kg RTD spacecraft will first spiral outward from the shuttle-deployed, medium-inclination, low Earth orbit. By the phased operation of a 10-kW Hall thruster and a 10-kW Variable Specific Impulse Magneto-Plasma Rocket, the RTD spacecraft will reach a low-inclination Earth orbit with a radius greater than five Earth radii. This will be followed by an inward spiraling orbit phase when the spacecraft deploys 8 to 12 microsatellites to map the Van Allen belts. The mission will conclude in low Earth orbit with the possible retrieval of the spacecraft by the space shuttle. A conceptual RTD spacecraft design showing two photovoltaic (PV) array wings, the Hall thruster with propellant tanks, and stowed microsatellites is presented. Early power system studies assessed five different PV array design options coupled with a 120-Vdc power management and distribution system (PMAD) and secondary lithium battery energy storage. Array options include (1) state-of-the-art 10-percent efficient three-junction amorphous SiGe thin-film cells on thin polymer panels deployed with an inflatable (or articulated) truss, (2) SCARLET array panels, (3) commercial state-of-the-art, planar PV array rigid panels with 25-percent efficient, three

  15. Shuttle mission simulator software conceptual design

    NASA Technical Reports Server (NTRS)

    Burke, J. F.

    1973-01-01

    Software conceptual designs (SCD) are presented for meeting the simulator requirements for the shuttle missions. The major areas of the SCD discussed include: malfunction insertion, flight software, applications software, systems software, and computer complex.

  16. Cloud Computing Techniques for Space Mission Design

    NASA Technical Reports Server (NTRS)

    Arrieta, Juan; Senent, Juan

    2014-01-01

    The overarching objective of space mission design is to tackle complex problems producing better results, and faster. In developing the methods and tools to fulfill this objective, the user interacts with the different layers of a computing system.

  17. Spacecraft design considerations for an Inner Magnetosphere Imager mission

    NASA Technical Reports Server (NTRS)

    Herrmann, Melody C.; Johnson, Charles L.

    1992-01-01

    Imaging the Earth's magnetosphere from space will enable scientists to better understand the global shape of the inner magnetosphere, its components and processes. The proposed Inner Magnetosphere Imager (IMI) mission will obtain the first simultaneous images of the component regions of the inner magnetosphere and will enable scientists to relate these global images to internal and external influences as well as local observations. NASA's Marshall Space Flight Center (MSFC) is performing a concept definition study of the proposed mission. As currently envisioned, the baseline mission calls for an instrument complement of approximately seven imagers to be flown in an elliptical Earth orbit with an apogee of seven Earth Radii (RE). Several spacecraft concepts have been examined for the mission. The baseline concept utilizes a spinning spacecraft with a despun platform, the second uses a three-axis stabilized spacecraft with a spinning platform, while the third option splits the instruments onto two small satellites; a spinning spacecraft and a complementary three-axis stabilized spacecraft. This paper will address the mission objectives, the rationale for using proven spacecraft designs, and the preliminary concept definition study team results for all three options.

  18. Options for a Geostationary Science Demonstration Mission (GSDM)

    NASA Astrophysics Data System (ADS)

    Pougatchev, N. S.; Bingham, G. E.; Zollinger, L.; Hancock, J. J.

    2009-12-01

    Geostationary ultraspectral imager with spectral resolution comparable with the ones of the current advanced LEO sounders such as AIRS and IASI brings the potential for significant new products to improve our lives and protect property. These include: improved severe weather warnings and hurricane track prediction, troposphere wind profiles at 2 Km vertical resolution, and pollutant and water vapor flux profiles. The GSDM data combined with OCO and GOSAT data can provide local and regional CO, CO2 emissions. The potential value of a GSDM is so great that the resent NASA/NOAA Decadal Survey recommended they “Complete the GIFTS instrument, deliver it to orbit via a cost-effective launch and spacecraft opportunity, and evaluate its potential to be a prototype for the HES instrument…”. GOES-R mission costs led to the cancellation of the HES program. Development of an entirely new instrument and flying it as an operational payload is clearly outside of the NOAA budget profile. However a joint NASA/NOAA An out-of-the-box, Venture Class style, PI-led mission to satisfy the NASA/NOAA Decadal Survey recommendation can be funded and managed with today’s budgets. An ideal NASA/NOAA mission would combine NOAA’s spare “Q” Imager and the upgraded GIFTS EDU hardware on a free flyer, launched in 2014 to the GOES East position and using the developing GOES-R downlink and communications system. Because the Ultraspectral Imager/Sounder data pixels are independent, GSDM data can be easily segmented into subimages, processed by massively parallel Linux computers, and analyzed by NASA and NOAA Algorithm working groups and science teams. A well calibrated Ultraspectral Imager/Sounder in a Geo orbit would also become the ultimate calibration transfer standard to support the WMO Global Space-based Inter-Calibration System (GSICS) effort. This poster reviews the science payoff of a GSDM, the measured GIFTS EDU hardware performance, and suggests an affordable mission strategy.

  19. The OSIRIS-REx Asteroid Sample Return Mission Operations Design

    NASA Technical Reports Server (NTRS)

    Gal-Edd, Jonathan S.; Cheuvront, Allan

    2015-01-01

    OSIRIS-REx is an acronym that captures the scientific objectives: Origins, Spectral Interpretation, Resource Identification, and Security-Regolith Explorer. OSIRIS-REx will thoroughly characterize near-Earth asteroid Bennu (Previously known as 1019551999 RQ36). The OSIRIS-REx Asteroid Sample Return Mission delivers its science using five instruments and radio science along with the Touch-And-Go Sample Acquisition Mechanism (TAGSAM). All of the instruments and data analysis techniques have direct heritage from flown planetary missions. The OSIRIS-REx mission employs a methodical, phased approach to ensure success in meeting the mission's science requirements. OSIRIS-REx launches in September 2016, with a backup launch period occurring one year later. Sampling occurs in 2019. The departure burn from Bennu occurs in March 2021. On September 24, 2023, the Sample Return Capsule (SRC) lands at the Utah Test and Training Range (UTTR). Stardust heritage procedures are followed to transport the SRC to Johnson Space Center, where the samples are removed and delivered to the OSIRIS-REx curation facility. After a six-month preliminary examination period the mission will produce a catalog of the returned sample, allowing the worldwide community to request samples for detailed analysis. Traveling and returning a sample from an Asteroid that has not been explored before requires unique operations consideration. The Design Reference Mission (DRM) ties together spacecraft, instrument and operations scenarios. Asteroid Touch and Go (TAG) has various options varying from ground only to fully automated (natural feature tracking). Spacecraft constraints such as thermo and high gain antenna pointing impact the timeline. The mission is sensitive to navigation errors, so a late command update has been implemented. The project implemented lessons learned from other "small body" missions. The key lesson learned was 'expect the unexpected' and implement planning tools early in the lifecycle

  20. An Initial Comparison of Selected Earth Departure Options for Solar Electric Propulsion Missions

    NASA Technical Reports Server (NTRS)

    Merrill, Raymond Gabriel; Komar, D. R.; Qu, Min; Chrone, Jon; Strange, Nathan; Landau, Damon

    2012-01-01

    Earth departure options such as the location for deployment, aggregation, and crew rendezvous as well as the type of propulsion leveraged for each mission phase effect overall mission performance metrics such as number of critical maneuvers, mass of propellant to achieve departure, and initial mass required in low Earth orbit. This paper identifies and compares a subset of tactical options for deployment, crew rendezvous, and Earth departure that leverage electric propulsion and hybrid chemical electric propulsion with a goal of improving system efficiency. Departure maneuver specific limitations and penalties are then identified for missions to specific targets for human interplanetary missions providing a better understanding of the impact of decisions related to aggregation and rendezvous locations as well as Earth departure maneuvers on overall system performance.

  1. BRRISON Mission Design and Development

    NASA Astrophysics Data System (ADS)

    O'Malley, Terence; Kremic, T.; Adams, D.; Arnold, S.; Cheng, A.

    2013-10-01

    In September 2012, the comet C/2012 S1 “ISON” was discovered by Russian amateur astronomers. A team consisting of personnel from Glenn Research Center (GRC) Space Science Project Office, the Johns Hopkins University Applied Physics Lab (APL), and the Southwest Research Institute (SWRI) was established to identify the science return on a high altitude balloon mission to observe ISON, and develop a plan based on re-using most of the hardware from the Stratospheric Terahertz Observatory (STO). The team determined that measuring the comet’s H20/CO2 ratio with an infra-red Camera would be a high-value and unique scientific contribution of a balloon borne payload. The BRRISON scientific payload consists of a heritage 80-cm telescope, a near-ultraviolet visible optical bench and instruments, and an infrared optical bench and instruments. The telescope, which has flown on prior balloon missions, consists of a light-weighted f/1.5 hyperboloid 80 cm diameter primary and a secondary mirror to provide an f/17 beam. The near ultra-violet and visible cameras and associated instruments are being integrated to an optics bench by SwRI. These instruments consist of a fine steering mirror (FSM) and a CMOS high rate camera to provide sub-arcsec pointing, and a CCD camera for low noise science operation, and a dichroic for splitting the f/17 beam between the two cameras. The infrared optics bench and instruments consist of an optics bench, re-imaging optics and cold stop, filter wheel and filters, and an infrared camera that is sensitive over the required wavelengths of 2.5 - 5 microns. The IR optics bench and instruments will be enclosed in an aluminum housing, which will be cooled to reduce the thermal background contribution to the IR signal. The BRRISON gondola is composed of a metal frame that carries and protects the science payload and subsystems and is the structural interface with the balloon flight train. They are composed of a Command & Control system, a Pointing

  2. Guidelines and Capabilities for Designing Human Missions

    NASA Technical Reports Server (NTRS)

    Allen, Christopher S.; Burnett, Rebeka; Charles, John; Cucinotta, Frank; Fullerton, Richard; Goodman, Jerry R.; Griffith, Anthony D., Sr.; Kosmo, Joseph J.; Perchonok, Michele; Railsback, Jan; Rajulu, Sudhakar; Stilwell, Don; Thomas, Gretchen; Tri, Terry; Joshi, Jitendra; Wheeler, Ray; Rudisill, Marianne; Wilson, John; Mueller, Alyssa; Simmons, Anne

    2003-01-01

    These guidelines and capabilities identify the points of intersection between human spaceflight crews and mission considerations such as architecture, vehicle design, technologies, operations, and science requirements. In these chapters, we will provide clear, top-level guidelines for human-related exploration studies and technology research that will address common questions and requirements. As a result, we hope that ongoing mission trade studies will consider common, standard, and practical criteria for human interfaces.

  3. Cascade Distillation System Design for Safety and Mission Assurance

    NASA Technical Reports Server (NTRS)

    Sarguisingh, Miriam; Callahan, Michael R.; Okon, Shira

    2015-01-01

    Per the NASA Human Health, Life Support and Habitation System Technology Area 06 report "crewed missions venturing beyond Low-Earth Orbit (LEO) will require technologies with improved reliability, reduced mass, self-sufficiency, and minimal logistical needs as an emergency or quick-return option will not be feasible".1 To meet this need, the development team of the second generation Cascade Distillation System (CDS 2.0) chose a development approach that explicitly incorporate consideration of safety, mission assurance, and autonomy. The CDS 2.0 preliminary design focused on establishing a functional baseline that meets the CDS core capabilities and performance. The critical design phase is now focused on incorporating features through a deliberative process of establishing the systems failure modes and effects, identifying mitigation strategies, and evaluating the merit of the proposed actions through analysis and test. This paper details results of this effort on the CDS 2.0 design.

  4. Cascade Distillation System Design for Safety and Mission Assurance

    NASA Technical Reports Server (NTRS)

    Sargusingh, Miriam J.; Callahan, Michael R.

    2015-01-01

    Per the NASA Human Health, Life Support and Habitation System Technology Area 06 report "crewed missions venturing beyond Low-Earth Orbit (LEO) will require technologies with improved reliability, reduced mass, self-sufficiency, and minimal logistical needs as an emergency or quick-return option will not be feasible." To meet this need, the development team of the second generation Cascade Distillation System (CDS 2.0) opted a development approach that explicitely incorporate consideration of safety, mission assurance, and autonomy. The CDS 2.0 prelimnary design focused on establishing a functional baseline that meets the CDS core capabilities and performance. The critical design phase is now focused on incorporating features through a deliberative process of establishing the systems failure modes and effects, identifying mitigative strategies, and evaluating the merit of the proposed actions through analysis and test. This paper details results of this effort on the CDS 2.0 design.

  5. Experimental Design for the LATOR Mission

    NASA Technical Reports Server (NTRS)

    Turyshev, Slava G.; Shao, Michael; Nordtvedt, Kenneth, Jr.

    2004-01-01

    This paper discusses experimental design for the Laser Astrometric Test Of Relativity (LATOR) mission. LATOR is designed to reach unprecedented accuracy of 1 part in 10(exp 8) in measuring the curvature of the solar gravitational field as given by the value of the key Eddington post-Newtonian parameter gamma. This mission will demonstrate the accuracy needed to measure effects of the next post-Newtonian order (near infinity G2) of light deflection resulting from gravity s intrinsic non-linearity. LATOR will provide the first precise measurement of the solar quadrupole moment parameter, J(sub 2), and will improve determination of a variety of relativistic effects including Lense-Thirring precession. The mission will benefit from the recent progress in the optical communication technologies the immediate and natural step above the standard radio-metric techniques. The key element of LATOR is a geometric redundancy provided by the laser ranging and long-baseline optical interferometry. We discuss the mission and optical designs, as well as the expected performance of this proposed mission. LATOR will lead to very robust advances in the tests of Fundamental physics: this mission could discover a violation or extension of general relativity, or reveal the presence of an additional long range interaction in the physical law. There are no analogs to the LATOR experiment; it is unique and is a natural culmination of solar system gravity experiments.

  6. Shuttle mission simulator hardware conceptual design report

    NASA Technical Reports Server (NTRS)

    Burke, J. F.

    1973-01-01

    The detailed shuttle mission simulator hardware requirements are discussed. The conceptual design methods, or existing technology, whereby those requirements will be fulfilled are described. Information of a general nature on the total design problem plus specific details on how these requirements are to be satisfied are reported. The configuration of the simulator is described and the capabilities for various types of training are identified.

  7. Guidelines and Capabilities for Designing Human Missions

    NASA Technical Reports Server (NTRS)

    2002-01-01

    The human element is likely the most complex and difficult one of mission design; it significantly influences every aspect of mission planning, from the basic parameters like duration to the more complex tradeoffs between mass, volume, power, risk, and cost. For engineers who rely on precise specifications in data books and other such technical references, dealing with the uncertainty and the variability of designing for human beings can be frustrating. When designing for the human element, questions arise more often than definitive answers. Nonetheless, we do not doubt that the most captivating discoveries in future space missions will necessitate human explorers. These guidelines and capabilities are meant to identify the points of intersection between humans and mission considerations such as architecture, vehicle design, technologies, operations, and science requirements. We seek to provide clear, top-level guidelines for human-related exploration studies and technology research that address common questions and requirements. As a result, we hope that ongoing mission trade studies consider common, standard, and practical criteria for human interfaces.

  8. Guidelines and Capabilities for Designing Human Missions

    NASA Astrophysics Data System (ADS)

    2002-03-01

    The human element is likely the most complex and difficult one of mission design; it significantly influences every aspect of mission planning, from the basic parameters like duration to the more complex tradeoffs between mass, volume, power, risk, and cost. For engineers who rely on precise specifications in data books and other such technical references, dealing with the uncertainty and the variability of designing for human beings can be frustrating. When designing for the human element, questions arise more often than definitive answers. Nonetheless, we do not doubt that the most captivating discoveries in future space missions will necessitate human explorers. These guidelines and capabilities are meant to identify the points of intersection between humans and mission considerations such as architecture, vehicle design, technologies, operations, and science requirements. We seek to provide clear, top-level guidelines for human-related exploration studies and technology research that address common questions and requirements. As a result, we hope that ongoing mission trade studies consider common, standard, and practical criteria for human interfaces.

  9. 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.

  10. Active Debris Removal mission design in Low Earth Orbit

    NASA Astrophysics Data System (ADS)

    Martin, Th.; Pérot, E.; Desjean, M.-Ch.; Bitetti, L.

    2013-03-01

    Active Debris Removal (ADR) aims at removing large sized intact objects ― defunct satellites, rocket upper-stages ― from space crowded regions. Why? Because they constitute the main source of the long-term debris environment deterioration caused by possible future collisions with fragments and worse still with other intact but uncontrolled objects. In order to limit the growth of the orbital debris population in the future (referred to as the Kessler syndrome), it is now highly recommended to carry out such ADR missions, together with the mitigation measures already adopted by national agencies (such as postmission disposal). At the French Space Agency, CNES, and in the frame of advanced studies, the design of such an ADR mission in Low Earth Orbit (LEO) is under evaluation. A two-step preliminary approach has been envisaged. First, a reconnaissance mission based on a small demonstrator (˜500 kg) rendezvousing with several targets (observation and in-flight qualification testing). Secondly, an ADR mission based on a larger vehicle (inherited from the Orbital Transfer Vehicle (OTV) concept) being able to capture and deorbit several preselected targets by attaching a propulsive kit to these targets. This paper presents a flight dynamics level tradeoff analysis between different vehicle and mission concepts as well as target disposal options. The delta-velocity, times, and masses required to transfer, rendezvous with targets and deorbit are assessed for some propelled systems and propellant less options. Total mass budgets are then derived for two end-to-end study cases corresponding to the reconnaissance and ADR missions mentioned above.

  11. Mars rapid round trip mission design

    NASA Astrophysics Data System (ADS)

    Sarzi Amade', Nicola

    The present research is divided in two parts. The first part is a well defined mathematical problem, with exact rules and results, in which the basic constraints for interplanetary round trip travels are used to calculate an interplanetary train schedule (ITS) of missions to Mars, in the general case of orbits with non-zero eccentricity and non-zero inclination. Several possible options for round trip travels to Mars are considered. In particular, options at high energy, which allow rapid round trip missions, are discussed. These options have important applications for human travels to Mars. The second part of the research is about systems engineering aspects, which are intrinsically less exact, since they can change with time due, for example, to technology development or economic and political factors. For the case of a selected human rapid round trip mission to Mars, the development of a mission architecture, an assessment of the masses involved in the mission (such as the initial masses required in LEO), an estimate of the necessary number of launches, and a preliminary analysis of the radiation protection requirements, are performed. The main problem that justifies the existence of basic constraints for round trip missions is that by increasing the DeltaV of a mission, in general the total round trip time does not vary much, because a higher DeltaV can only reduce the transfer time and it simply increases the stay-time on the target planet. However, if the DeltaV is increased beyond a well-defined level, the total round trip time has a sudden drop in duration that makes fast round trips possible. This is due to the fact that the traveler can go back before the home planet makes one extra revolution around the Sun. For a sufficiently high DeltaV, a round trip to Mars can change in duration from 2.7 years to about 5 months. For Mars missions, the round trip times are calculated for different DeltaV's and for different transfer trajectories (T1, T2, etc.). An

  12. Mars ISRU for Production of Mission Critical Consumables - Options, Recent Studies, and Current State of the Art

    NASA Technical Reports Server (NTRS)

    Sanders, G. B.; Paz, A.; Oryshchyn, L.; Araghi, K.; Muscatello, A.; Linne, D.; Kleinhenz, J.; Peters, T.

    2015-01-01

    In 1978, a ground breaking paper titled, "Feasibility of Rocket Propellant Production on Mars" by Ash, Dowler, and Varsi discussed how ascent propellants could be manufactured on the Mars surface from carbon dioxide collected from the atmosphere to reduce launch mass. Since then, the concept of making mission critical consumables such as propellants, fuel cell reactants, and life support consumables from local resources, commonly known as In-Situ Resource Utilization (ISRU), for robotic and human missions to Mars has been studied many times. In the late 1990's, NASA initiated a series of Mars Human Design Reference Missions (DRMs), the first of which was released in 1997. These studies primarily focused on evaluating the impact of making propellants on Mars for crew ascent to Mars orbit, but creating large caches of life support consumables (water & oxygen) as a backup for regenerative life support systems for long-duration surface stays (>500 days) was also considered in Mars DRM 3.0. Until science data from the Mars Odyssey orbiter and subsequent robotic missions revealed that water may be widely accessable across the surface of Mars, prior Mars ISRU studies were limited to processing Mars atmospheric resources (carbon dioxide, nitrogen, argon, oxygen, and water vapor). In December 2007, NASA completed the Mars Human Design Reference Architecture (DRA) 5.0 study which considered water on Mars as a potential resource for the first time in a human mission architecture. While knowledge of both water resources on Mars and the hardware required to excavate and extract the water were very preliminary, the study concluded that a significant reduction in mass and significant enhancements to the mission architecture were possible if Mars water resources were utilized. Two subsequent Mars ISRU studies aimed at reexamining ISRU technologies, processing options, and advancements in the state-of-the-art since 2007 and to better understand the volume and packaging associated

  13. The Far Ultraviolet (FUV) auroral imager for the Inner Magnetospheric Imager (IMI) mission: Options

    NASA Technical Reports Server (NTRS)

    Wilson, Gordon R.

    1993-01-01

    The change from an intermediate mission (cost ceiling of $300 million) to a solar-terrestrial probe class mission (cost ceiling of $150 million) will require some major changes in the configuration of the IMI mission. One option being considered is to move to a small spin-stabilized spacecraft (with no despun platform) which could be launched with a smaller Taurus or Conestoga class booster. Such a change in spacecraft type would not present any fundamental problems (other than restrictions on mass and power) for the He plus 304 A plasmasphere imager, the high and low energy neutral atom imagers, and the geocoronal imager, but would present a challenge for the FUV auroral imager since the original plan called for this instrument to operate from a despun platform. Since the FUV instrument is part of the core payload it cannot be dropped from the instrument complement without jeopardizing the science goals of the mission. A way must be found to keep this instrument and to allow it to accomplish most, if not all, of its science objectives. One of the subjects discussed are options for building an FUV instrument for a spinning spacecraft. Since a number of spinning spacecraft have carried auroral imagers, a range of techniques exists. In addition, the option of flying the FUV imager on a separate micro-satellite launched with the main IMI spacecraft or with a separate pegasus launch, was considered and is discussed.

  14. 47 CFR 1.2103 - Competitive bidding design options.

    Code of Federal Regulations, 2012 CFR

    2012-10-01

    ... 47 Telecommunication 1 2012-10-01 2012-10-01 false Competitive bidding design options. 1.2103... design options. (a) The Commission will choose from one or more of the following types of auction designs...) The Commission may use real time bidding in all electronic auction designs....

  15. ETF Mission Statement document. ETF Design Center team

    SciTech Connect

    Not Available

    1980-04-01

    The Mission Statement document describes the results, activities, and processes used in preparing the Mission Statement, facility characteristics, and operating goals for the Engineering Test Facility (ETF). Approximately 100 engineers and scientists from throughout the US fusion program spent three days at the Knoxville Mission Workshop defining the requirements that should be met by the ETF during its operating life. Seven groups were selected to consider one major category each of design and operation concerns. Each group prepared the findings of the assigned area as described in the major sections of this document. The results of the operations discussed must provide the data, knowledge, experience, and confidence to continue to the next steps beyond the ETF in making fusion power a viable energy option. The results from the ETF mission (operations are assumed to start early in the 1990's) are to bridge the gap between the base of magnetic fusion knowledge at the start of operations and that required to design the EPR/DEMO devices.

  16. NASA's Asteroid Redirect Mission: A Robotic Boulder Capture Option for Science, Human Exploration, Resource Utilization, and Planetary Defense

    NASA Technical Reports Server (NTRS)

    Abell, P.; Nuth, J.; Mazanek, D.; Merrill, R.; Reeves, D.; Naasz, B.

    2014-01-01

    NASA is examining two options for the Asteroid Redirect Mission (ARM), which will return asteroid material to a Lunar Distant Retrograde Orbit (LDRO) using a robotic solar electric propulsion spacecraft, called the Asteroid Redirect Vehicle (ARV). Once the ARV places the asteroid material into the LDRO, a piloted mission will rendezvous and dock with the ARV. After docking, astronauts will conduct two extravehicular activities (EVAs) to inspect and sample the asteroid material before returning to Earth. One option involves capturing an entire small (4 - 10 m diameter) near-Earth asteroid (NEA) inside a large inflatable bag. However, NASA is also examining another option that entails retrieving a boulder (1 - 5 m) via robotic manipulators from the surface of a larger (100+ m) pre-characterized NEA. The Robotic Boulder Capture (RBC) option can leverage robotic mission data to help ensure success by targeting previously (or soon to be) well- characterized NEAs. For example, the data from the Japan Aerospace Exploration Agency's (JAXA) Hayabusa mission has been utilized to develop detailed mission designs that assess options and risks associated with proximity and surface operations. Hayabusa's target NEA, Itokawa, has been identified as a valid target and is known to possess hundreds of appropriately sized boulders on its surface. Further robotic characterization of additional NEAs (e.g., Bennu and 1999 JU3) by NASA's OSIRIS REx and JAXA's Hayabusa 2 missions is planned to begin in 2018. This ARM option reduces mission risk and provides increased benefits for science, human exploration, resource utilization, and planetary defense. Science: The RBC option is an extremely large sample-return mission with the prospect of bringing back many tons of well-characterized asteroid material to the Earth-Moon system. The candidate boulder from the target NEA can be selected based on inputs from the world-wide science community, ensuring that the most scientifically interesting

  17. Space station needs, attributes and architectural options. Volume 3, task 1: Mission requirements

    NASA Technical Reports Server (NTRS)

    1983-01-01

    The mission requirements of the space station program are investigated. Mission parameters are divided into user support from private industry, scientific experimentation, U.S. national security, and space operations away from the space station. These categories define the design and use of the space station. An analysis of cost estimates is included.

  18. Space station needs, attributes and architectural options study. Volume 2: Mission definition

    NASA Technical Reports Server (NTRS)

    1983-01-01

    The space applications and science programs appropriate to the era beyond 1990, those user missions which can utilize the Space Station to an advantage, and user mission concepts so that requirements, which will drive the Space Stations (SS) design are addressed.

  19. Lunar prospector mission design and trajectory support

    NASA Technical Reports Server (NTRS)

    Lozier, David; Galal, Ken; Folta, David; Beckman, Mark

    1998-01-01

    The Lunar Prospector mission is the first dedicated NASA lunar mapping mission since the Apollo Orbiter program which was flown over 25 years ago. Competitively selected under the NASA Discovery Program, Lunar Prospector was launched on January 7, 1998 on the new Lockheed Martin Athena 2 launch vehicle. The mission design of Lunar Prospector is characterized by a direct minimum energy transfer trajectory to the moon with three scheduled orbit correction maneuvers to remove launch and cislunar injection errors prior to lunar insertion. At lunar encounter, a series of three lunar orbit insertion maneuvers and a small circularization burn were executed to achieve a 100 km altitude polar mapping orbit. This paper will present the design of the Lunar Prospector transfer, lunar insertion and mapping orbits, including maneuver and orbit determination strategies in the context of mission goals and constraints. Contingency plans for handling transfer orbit injection and lunar orbit insertion anomalies are also summarized. Actual flight operations results are discussed and compared to pre-launch support analysis.

  20. Integrating Safety and Mission Assurance in Design

    NASA Technical Reports Server (NTRS)

    Cianciola, Chris; Crane, Kenneth

    2008-01-01

    This presentation describes how the Ares Projects are learning from the successes and failures of previous launch systems in order to maximize safety and reliability while maintaining fiscal responsibility. The Ares Projects are integrating Safety and Mission Assurance into design activities and embracing independent assessments by Quality experts in thorough reviews of designs and processes. Incorporating Lean thinking into the design process, Ares is also streamlining existing processes and future manufacturing flows which will yield savings during production. Understanding the value of early involvement of Quality experts, the Ares Projects are leading launch vehicle development into the 21st century.

  1. Logistics and Maintenance Options to Support the P-8A Poseidon’s Expeditionary Mission

    DTIC Science & Technology

    2010-06-01

    Comparison, P- 3 and 737-800 (From NAVAIR MER Facilities Document, 2009, p. 6) The P-8A is powered by two CFM International CFM56 -7B27A turbofan engines...REPORT DATE June 2010 3 . REPORT TYPE AND DATES COVERED MBA Professional Report 4. TITLE AND SUBTITLE Logistics and Maintenance Options to...being utilized for missions being performed by the P- 3 Orion, and to propose organizational and operational recommendations to better support the agile

  2. Another Option for the Asteroid Sample of the Asteroid Redirect Mission

    NASA Astrophysics Data System (ADS)

    Hou, Xiyun; Tang, Jingshi; Liu, Lin; Xin, Xiaosheng

    2016-07-01

    The asteroid redirect mission (ARM) consists of two phases: the asteroid redirect robotic mission (ARRM) and the asteroid redirect crewed mission (ARCM). The ARRM phase aims at capturing a boulder from the surface of an asteroid of hundred meters in diameter and returning it back to the Earth-Moon system. Currently, the option for the orbit of the returned sample is a large lunar distant retrograde orbit (LDRO) around the Moon. After the sample is returned to this LDRO, then the ARCM phase will send astronauts to the sample. The total energy cost consists of two parts: (1) from the orbit of an near-Earth asteroid to the LDRO, here as part I; (2) from the parking low Earth orbit (LEO) to the LDRO, here as part II. In the authors' work for stable motions in the real Earth-Moon system, we found that there are stable motions around the triangular libration points (TLP). Theoretically, these orbits can also be used as candidate orbits to hold the returned sample. Our previous preliminary works show that the energy of sending a manned probe from the LEO to these orbits is comparable to the option of sending it from the LEO to the LDRO. Besides, it's also possible for the sample to be returned from the orbit of a near-Earth asteroid to these stable orbits, with very small delta-V corrections. In this work, we'll study the energy cost of this option (i.e., using the stable orbits around the TLP as the orbits for the asteroid sample) in detail and compare this option with the LDRO option.

  3. Waste Management Options for Long-Duration Space Missions: When to Reject, Reuse, or Recycle

    NASA Technical Reports Server (NTRS)

    Linne, Diane L.; Palaszewski, Bryan A.; Gokoglu, Suleyman; Gallo, Christopher A.; Balasubramaniam, Ramaswamy; Hegde, Uday G.

    2014-01-01

    The amount of waste generated on long-duration space missions away from Earth orbit creates the daunting challenge of how to manage the waste through reuse, rejection, or recycle. The option to merely dispose of the solid waste through an airlock to space was studied for both Earth-moon libration point missions and crewed Mars missions. Although the unique dynamic characteristics of an orbit around L2 might allow some discarded waste to intersect the lunar surface before re-impacting the spacecraft, the large amount of waste needed to be managed and potential hazards associated with volatiles recondensing on the spacecraft surfaces make this option problematic. A second option evaluated is to process the waste into useful gases to be either vented to space or used in various propulsion systems. These propellants could then be used to provide the yearly station-keeping needs at an L2 orbit, or if processed into oxygen and methane propellants, could be used to augment science exploration by enabling lunar mini landers to the far side of the moon.

  4. A survey of propulsion options for cargo and piloted missions to Mars.

    PubMed

    Sankaran, K; Cassady, L; Kodys, A D; Choueiri, E Y

    2004-05-01

    In this paper, high-power electric propulsion options are surveyed in the context of cargo and piloted missions to Mars. A low-thrust trajectory optimization program (raptor) is utilized to analyze this mission. Candidate thrusters are chosen based upon demonstrated performance in the laboratory. Hall, self-field magnetoplasmadynamic (MPDT), self-field lithium Lorentz force accelerator (LiLFA), arcjet, and applied-field LiLFA systems are considered for this mission. In this first phase of the study, all thrusters are assumed to operate at a single power level (regardless of the efficiency-power curve), and the thruster specific mass and power plant specific mass are taken to be the same for all systems. Under these assumptions, for a 7.5 MW, 60 mT payload, piloted mission, the self-field LiLFA results in the shortest trip time (340 days) with a reasonable propellant mass fraction of 57% (129 mT). For a 150 kW, 9 mT payload, cargo mission, both the applied-field LiLFA and the Hall thruster seem reasonable choices with propellant mass fractions of 42 to 45%(7 to 8 mT). The Hall thrusters provide better trip times (530-570 days) compared to the applied-field LiLFA (710 days) for the relatively less demanding mission.

  5. Mars Mission Analysis Trades Based on Legacy and Future Nuclear Propulsion Options

    SciTech Connect

    Joyner, Russell; Lentati, Andrea; Cichon, Jaclyn

    2007-01-30

    The purpose of this paper is to discuss the results of mission-based system trades when using a nuclear thermal propulsion (NTP) system for Solar System exploration. The results are based on comparing reactor designs that use a ceramic-metallic (CERMET), graphite matrix, graphite composite matrix, or carbide matrix fuel element designs. The composite graphite matrix and CERMET designs have been examined for providing power as well as propulsion. Approaches to the design of the NTP to be discussed will include an examination of graphite, composite, carbide, and CERMET core designs and the attributes of each in regards to performance and power generation capability. The focus is on NTP approaches based on tested fuel materials within a prismatic fuel form per the Argonne National Laboratory testing and the ROVER/NERVA program. NTP concepts have been examined for several years at Pratt and Whitney Rocketdyne for use as the primary propulsion for human missions beyond earth. Recently, an approach was taken to examine the design trades between specific NTP concepts; NERVA-based (UC)C-Graphite, (UC,ZrC)C-Composite, (U,Zr)C-Solid Carbide and UO2-W CERMET. Using Pratt and Whitney Rocketdyne's multidisciplinary design analysis capability, a detailed mission and vehicle model has been used to examine how several of these NTP designs impact a human Mars mission. Trends for the propulsion system mass as a function of power level (i.e. thrust size) for the graphite-carbide and CERMET designs were established and correlated against data created over the past forty years. These were used for the mission trade study. The resulting mission trades presented in this paper used a comprehensive modeling approach that captures the mission, vehicle subsystems, and NTP sizing.

  6. Mars Mission Analysis Trades Based on Legacy and Future Nuclear Propulsion Options

    NASA Astrophysics Data System (ADS)

    Joyner, Russell; Lentati, Andrea; Cichon, Jaclyn

    2007-01-01

    The purpose of this paper is to discuss the results of mission-based system trades when using a nuclear thermal propulsion (NTP) system for Solar System exploration. The results are based on comparing reactor designs that use a ceramic-metallic (CERMET), graphite matrix, graphite composite matrix, or carbide matrix fuel element designs. The composite graphite matrix and CERMET designs have been examined for providing power as well as propulsion. Approaches to the design of the NTP to be discussed will include an examination of graphite, composite, carbide, and CERMET core designs and the attributes of each in regards to performance and power generation capability. The focus is on NTP approaches based on tested fuel materials within a prismatic fuel form per the Argonne National Laboratory testing and the ROVER/NERVA program. NTP concepts have been examined for several years at Pratt & Whitney Rocketdyne for use as the primary propulsion for human missions beyond earth. Recently, an approach was taken to examine the design trades between specific NTP concepts; NERVA-based (UC)C-Graphite, (UC,ZrC)C-Composite, (U,Zr)C-Solid Carbide and UO2-W CERMET. Using Pratt & Whitney Rocketdyne's multidisciplinary design analysis capability, a detailed mission and vehicle model has been used to examine how several of these NTP designs impact a human Mars mission. Trends for the propulsion system mass as a function of power level (i.e. thrust size) for the graphite-carbide and CERMET designs were established and correlated against data created over the past forty years. These were used for the mission trade study. The resulting mission trades presented in this paper used a comprehensive modeling approach that captures the mission, vehicle subsystems, and NTP sizing.

  7. Towards Risk Based Design for NASA's Missions

    NASA Technical Reports Server (NTRS)

    Tumer, Irem Y.; Barrientos, Francesca; Meshkat, Leila

    2004-01-01

    This paper describes the concept of Risk Based Design in the context of NASA s low volume, high cost missions. The concept of accounting for risk in the design lifecycle has been discussed and proposed under several research topics, including reliability, risk analysis, optimization, uncertainty, decision-based design, and robust design. This work aims to identify and develop methods to enable and automate a means to characterize and optimize risk, and use risk as a tradeable resource to make robust and reliable decisions, in the context of the uncertain and ambiguous stage of early conceptual design. This paper first presents a survey of the related topics explored in the design research community as they relate to risk based design. Then, a summary of the topics from the NASA-led Risk Colloquium is presented, followed by current efforts within NASA to account for risk in early design. Finally, a list of "risk elements", identified for early-phase conceptual design at NASA, is presented. The purpose is to lay the foundation and develop a roadmap for future work and collaborations for research to eliminate and mitigate these risk elements in early phase design.

  8. The HSCT mission analysis of waverider designs

    NASA Technical Reports Server (NTRS)

    1993-01-01

    The grant provided partial support for an investigation of wave rider design and analysis with application to High-Speed Civil Transport (HSCT) vehicles. Proposed was the development of the necessary computational fluid dynamics (CFD) tools for the direct simulation of the waverider vehicles, the development of two new wave rider design methods that would provide computational speeds and design flexibilities never before achieved in wave rider design studies, and finally the selection of a candidate waverider-based vehicle and the evaluation of the chosen vehicle for a canonical HSCT mission scenario. This, the final report, reiterates the proposed project objectives in moderate detail, and it outlines the state of completion of each portion of the study, providing references to current and forthcoming publications that resulted from this work.

  9. 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.

  10. Structural Design for a Neptune Aerocapture Mission

    NASA Technical Reports Server (NTRS)

    Dyke, R. Eric; Hrinda, Glenn A.

    2004-01-01

    A multi-center study was conducted in 2003 to assess the feasibility of and technology requirements for using aerocapture to insert a scientific platform into orbit around Neptune. The aerocapture technique offers a potential method of greatly reducing orbiter mass and thus total spacecraft launch mass by minimizing the required propulsion system mass. This study involved the collaborative efforts of personnel from Langley Research Center (LaRC), Johnson Space Flight Center (JSFC), Marshall Space Flight Center (MSFC), Ames Research Center (ARC), and the Jet Propulsion Laboratory (JPL). One aspect of this effort was the structural design of the full spacecraft configuration, including the ellipsled aerocapture orbiter and the in-space solar electric propulsion (SEP) module/cruise stage. This paper will discuss the functional and structural requirements for each of these components, some of the design trades leading to the final configuration, the loading environments, and the analysis methods used to ensure structural integrity. It will also highlight the design and structural challenges faced while trying to integrate all the mission requirements. Component sizes, materials, construction methods and analytical results, including masses and natural frequencies, will be presented, showing the feasibility of the resulting design for use in a Neptune aerocapture mission. Lastly, results of a post-study structural mass optimization effort on the ellipsled will be discussed, showing potential mass savings and their influence on structural strength and stiffness

  11. Design Options Study. Volume III. Qualitative Assessment.

    DTIC Science & Technology

    1980-09-01

    commercial commnonality by improving commercial economics and/or military cost effectiveness. This detailed examination of these options is described in...including those listed above, have the greatest potential effects on commercial economics , on military cost- effectiveness, and ultimately, on the...in a military environment as well as the commercial economics was carefully examined. The analysis produced aircraft configurations that are optimum

  12. Trajectory Design of the Lunar Impactor Mission Concept

    NASA Technical Reports Server (NTRS)

    Chung, Min-Kun J.; McElrath, Timothy P.; Roncoli, Ralph B.

    2006-01-01

    The National Aeronautics and Space Administration (NASA) solicited proposals in 2006 for an opportunity to include a small secondary payload with the launch of the Lunar Reconnaissance Orbiter (LRO) scheduled for October 2008. The cost cap of the proposal was between $50 and $80M, and the mass cap was 1,000 kilograms. JPL proposed a Lunar Impactor (LI) concept for this solicitation. The mission objective of LI was to impact the permanently shadowed region of a South polar crater ultimately to detect the presence of water. The detection of water ice would prove to be an important factor on future lunar exploration. NASA Ames Research Center also proposed a similar concept, the Lunar Crater observation and Sensing Satellite (LCROSS), which was selected by NASA for the mission. However, in this paper, the trajectory design of the LI proposed by JPL is considered. Since the LI spacecraft was to be launched on the LRO launch vehicle as a secondary payload, its initial trajectory must be diverted at some later time from the LRO trans-lunar trajectory for the subsequent impact. Several such trajectories have been considered, where each trajectory option fields some specific values for the mission parameters. The mission parameters include the availability of LRO instruments at the time of impact for the observation by LRO, the mission duration, the impact velocity, the impact angle, etc. It is possible for the LI to be deflected with a relatively low delta-V to impact a South polar crater at a reasonable impact velocity and impact angle directly with no delay. However, the instruments on-board LRO may not be ready for observation. Thus, several delayed trajectory options have been considered further. The lunar phase at the time of impact may also play an important factor for observation, especially from Earth. Several lunar flyby trajectory maneuvers have been identified to arrive at the Moon for impact at the desired lunar phase. By using a combination of these

  13. Design of a solar sail mission to Mars

    NASA Technical Reports Server (NTRS)

    Fleri, E. J., Jr.; Galliano, P. A.; Harrison, M. E.; Johnson, W. B.; Meyer, G. J.

    1989-01-01

    A new area of interest in space vehicles is the solar sail. Various applications for which it has been considered are attitude control of satellites, focusing light on the jungles of Vietnam, and a Halley's comet rendezvous. Although for various reasons these projects were never completed, new interest in solar sails has arisen. The solar sail is an alternative to the rocket-propelled space vehicle as an interplanetary cargo vehicle, and manufacture of solar sails on the space station is a possibility. Solar sails have several advantages over rockets, including an unlimited power supply and low maintenance. The purpose of this project is to design a solar sail mission to Mars. The spacecraft will efficiently journey to Mars powered only by a solar sail. The vehicle weighs 487.16 kg and will be launchable on an expendable launch vehicle. The project includes an investigation of options to minimize cost, weight, and flight duration. The design of the sail and its deployment system are a major part of the project, as is the actual mission planning. Various topics researched include solar power, material, space environment, thermal control, trajectories, and orbit transfer. Various configurations are considered in order to determine the optimal structure. Another design consideration is the control system of the vehicle. This system includes the attitude control and the communication system of the sail. This project will aid in determining the feasibility of a solar sail and will raise public interest in space research.

  14. Starshade Design for Occulter Based Exoplanet Missions

    NASA Technical Reports Server (NTRS)

    Thomson, Mark W.; Lisman, P. Douglas; Helms, Richard; Walkemeyer, Phil; Kissil, Andrew; Polanco, Otto; Lee, Siu-Chun

    2010-01-01

    We present a lightweight starshade design that delivers the requisite profile figure accuracy with a compact stowed volume that permits launching both the occulter system (starshade and spacecraft) and a 1 to 2m-class telescope system on a single existing launch vehicle. Optimal figure stability is achieved with a very stiff and mass-efficient deployable structure design that has a novel configuration. The reference design is matched to a 1.1m telescope and consists of a 15m diameter inner disc and 24 flower-like petals with 7.5m length. The total tip-to-tip diameter of 30m provides an inner working angle of 75 mas. The design is scalable to accommodate larger telescopes and several options have been assessed. A proof of concept petal is now in production at JPL for deployment demonstrations and as a testbed for developing additional elements of the design. Future plans include developing breadboard and prototype hardware of increasing fidelity for use in demonstrating critical performance capabilities such as deployed optical edge profile figure tolerances and stability thereof.

  15. Starshade design for occulter based exoplanet missions

    NASA Astrophysics Data System (ADS)

    Thomson, Mark W.; Lisman, P. Douglas; Helms, Richard; Walkemeyer, Phil; Kissil, Andrew; Polanco, Otto; Lee, Siu-Chun

    2010-07-01

    We present a lightweight starshade design that delivers the requisite profile figure accuracy with a compact stowed volume that permits launching both the occulter system (starshade and spacecraft) and a 1 to 2m-class telescope system on a single existing launch vehicle. Optimal figure stability is achieved with a very stiff and mass-efficient deployable structure design that has a novel configuration. The reference design is matched to a 1.1m telescope and consists of a 15m diameter inner disc and 24 flower-like petals with 7.5m length. The total tip-to-tip diameter of 30m provides an inner working angle of 75 mas. The design is scalable to accommodate larger telescopes and several options have been assessed. A proof of concept petal is now in production at JPL for deployment demonstrations and as a testbed for developing additional elements of the design. Future plans include developing breadboard and prototype hardware of increasing fidelity for use in demonstrating critical performance capabilities such as deployed optical edge profile figure tolerances and stability thereof.

  16. Electric Propulsion Options for 10 kW Class Earth-Space Missions

    NASA Technical Reports Server (NTRS)

    Patterson, M. J.; Curran, Francis M.

    1989-01-01

    Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment were evaluated. Analyses were conducted to determine first-order propulsion system performance and system component mass estimates. Overall mission performance of the electric propulsion systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA 2 6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 1.2x10 (exp 7) to 2.1x10 (exp 7) N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA 2 6920 vehicle range from approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 2.2x10 (exp 6) to 3.6x10 (exp 6) N-s, and approximately 662 to 1072 m/s, respectively.

  17. Electric propulsion options for 10 kW class earth space missions

    NASA Technical Reports Server (NTRS)

    Patterson, M. J.; Curran, Francis M.

    1989-01-01

    Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment have been evaluated. Analyses were conducted to determine first-order propulsion system performance and system component mass estimates. Overall mission performance of the electric propulsion systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA II 6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 1.2x10(7) to 2.1x10(7) N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA II 6920 vehicle range from approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 2.2x10(6) to 3.6x10(6) N-s, and approximately 662 to 1072 m/s, respectively.

  18. Waverly, Iowa Smart Planning Workshop: Policy Options and Project Designs

    EPA Pesticide Factsheets

    This report from the EPA-FEMA technical assistance project in Waverly, IA, presents policy options and project design ideas that could help new development be more resilient to disasters such as floods while bringing multiple benefits.

  19. Unique mission options available with a megawatt-class nuclear electric propulsion system

    SciTech Connect

    Coomes, E.P.; McCauley, L.A.; Christian, J.L.; Gomez, M.A.; Wong, W.A.

    1988-10-01

    The advantages of using electric propulsion systems are well-known in the aerospace community with the most common being its high specific impulse, lower propellant requirements, and lower system mass. But these advantages may not be as important as the overall unique mission options electric propulsion makes possible, especially if the system is powered by a megawatt-class nuclear electric power source. Although the lack of suitable electric power systems has been a major drawback to electric propulsion, recent efforts have shown megawatt-class nuclear electric power systems are feasible and could be available by the turn of the century. Coupling this with the resurgence in interest in free-space electromagnetic transmission of energy and technology developments in this area provide a whole new aspect to the view of electric propulsion. The propulsion system now has a second mission function that may be of more value than the well understood benefits of electric propulsion; that is providing large quantities of prime power in support of a broad spectrum of mission tasks. 30 refs., 9 figs.

  20. Risk analysis of earth return options for the Mars rover/sample return mission

    NASA Technical Reports Server (NTRS)

    1988-01-01

    Four options for return of a Mars surface sample to Earth were studied to estimate the risk of mission failure and the risk of a sample container breach that might result in the release of Martian life forms, should such exist, in the Earth's biosphere. The probabilities calculated refer only to the time period from the last midcourse correction burn to possession of the sample on Earth. Two extreme views characterize this subject. In one view, there is no life on Mars, therefore there is no significant risk and no serious effort is required to deal with back contamination. In the other view, public safety overrides any desire to return Martian samples, and any risk of damaging contamination greater than zero is unacceptable. Zero risk requires great expense to achieve and may prevent the mission as currently envisioned from taking place. The major conclusion is that risk of sample container breach can be reduced to a very low number within the framework of the mission as now envisioned, but significant expense and effort, above that currently planned is needed. There are benefits to the public that warrant some risk. Martian life, if it exists, will be a major discovery. If it does not, there is no risk.

  1. Designer Babies: Eugenics Repackaged or Consumer Options?

    ERIC Educational Resources Information Center

    Baird, Stephen L.

    2007-01-01

    "Designer babies" is a term used by journalists and commentators--not by scientists--to describe several different reproductive technologies. These technologies have one thing in common: they give parents more control over what their offspring will be like. Designer babies are made possible by progress in three fields: (1) Advanced…

  2. Simulation Packages Expand Aircraft Design Options

    NASA Technical Reports Server (NTRS)

    2013-01-01

    In 2001, NASA released a new approach to computational fluid dynamics that allows users to perform automated analysis on complex vehicle designs. In 2010, Palo Alto, California-based Desktop Aeronautics acquired a license from Ames Research Center to sell the technology. Today, the product assists organizations in the design of subsonic aircraft, space planes, spacecraft, and high speed commercial jets.

  3. Rural Schools Prototype Analysis. Volume I: Design, Determinants and Options.

    ERIC Educational Resources Information Center

    Construction Systems Management, Inc., Anchorage, AK.

    This resource guide presents Design Determinants and Options to be used by designers, school district personnel, and State officials in the programing and design of small rural secondary schools in the Alaska bush. The vast and unconventional educational and space planning challenge is compounded by: the need to provide most or all of the…

  4. Jovian plasma modeling for mission design

    NASA Technical Reports Server (NTRS)

    Garrett, Henry B.; Kim, Wousik; Belland, Brent; Evans, Robin

    2015-01-01

    The purpose of this report is to address uncertainties in the plasma models at Jupiter responsible for surface charging and to update the jovian plasma models using the most recent data available. The updated plasma environment models were then used to evaluate two proposed Europa mission designs for spacecraft charging effects using the Nascap-2k code. The original Divine/Garrett jovian plasma model (or "DG1", T. N. Divine and H. B. Garrett, "Charged particle distributions in Jupiter's magnetosphere," J. Geophys. Res., vol. 88, pp. 6889-6903,1983) has not been updated in 30 years, and there are known errors in the model. As an example, the cold ion plasma temperatures between approx.5 and 10 Jupiter radii (Rj) were found by the experimenters who originally published the data to have been underestimated by approx.2 shortly after publication of the original DG1 model. As knowledge of the plasma environment is critical to any evaluation of the surface charging at Jupiter, the original DG1 model needed to be updated to correct for this and other changes in our interpretation of the data so that charging levels could beproperly estimated using the Nascap-2k charging code. As an additional task, the Nascap-2k spacecraft charging tool has been adapted to incorporate the so-called Kappa plasma distribution function--an important component of the plasma model necessary to compute the particle fluxes between approx.5 keV and 100 keV (at the outset of this study,Nascap-2k did not directly incorporate this common representation of the plasma thus limiting the accuracy of our charging estimates). The updating of the DG1 model and its integration into the Nascap-2k design tool means that charging concerns can now be more efficiently evaluated and mitigated. (We note that, given the subsequent decision by the Europa project to utilize solar arrays for its baseline design, surface charging effects have becomeeven more of an issue for its mission design). The modifications and

  5. Jovian Plasma Modeling for Mission Design

    NASA Technical Reports Server (NTRS)

    Garrett, Henry B.; Kim, Wousik; Belland, Brent; Evans, Robin

    2015-01-01

    The purpose of this report is to address uncertainties in the plasma models at Jupiter responsible for surface charging and to update the jovian plasma models using the most recent data available. The updated plasma environment models were then used to evaluate two proposed Europa mission designs for spacecraft charging effects using the Nascap-2k code. The original Divine/Garrett jovian plasma model (or "DG1", T. N. Divine and H. B. Garrett, "Charged particle distributions in Jupiter's magnetosphere," J. Geophys. Res., vol. 88, pp. 6889-6903,1983) has not been updated in 30 years, and there are known errors in the model. As an example, the cold ion plasma temperatures between approx.5 and 10 Jupiter radii (Rj) were found by the experimenters who originally published the data to have been underestimated by approx.2 shortly after publication of the original DG1 model. As knowledge of the plasma environment is critical to any evaluation of the surface charging at Jupiter, the original DG1 model needed to be updated to correct for this and other changes in our interpretation of the data so that charging levels could beproperly estimated using the Nascap-2k charging code. As an additional task, the Nascap-2k spacecraft charging tool has been adapted to incorporate the so-called Kappa plasma distribution function--an important component of the plasma model necessary to compute the particle fluxes between approx.5 keV and 100 keV (at the outset of this study,Nascap-2k did not directly incorporate this common representation of the plasma thus limiting the accuracy of our charging estimates). The updating of the DG1 model and its integration into the Nascap-2k design tool means that charging concerns can now be more efficiently evaluated and mitigated. (We note that, given the subsequent decision by the Europa project to utilize solar arrays for its baseline design, surface charging effects have becomeeven more of an issue for its mission design). The modifications and

  6. Review of a design-build option

    SciTech Connect

    1999-06-01

    Development of the Nonproliferation and International Security Center (NISC) project was started originally in 1996. The NISC project is to provide 163,375 square feet of laboratory and office space for 465 technical and support staff, The total estimated cost for the facility is about sixty million dollars. DOE has a number of policies and procedures related to the authorization and construction of projects. These policies and procedures generally favor the Design, Bid and Build process. The process include the following key items: (1) Authorization to develop a Conceptual Design Report (CDR); (2) Development of the CDR; (3) Validation of the CDR; (4) Authorization and funding to design the facility; (5) Validation of the design; and (6) Authorization and funding to build the facility. This approach placed the project management team in the position of resolving conflicts between the A/E and the contractor. The project management often also has to resolve conflicts between the facility uses and the A/E and contractor. These activities often distracted from the project management`s main objectives of controlling costs, schedule, and quality assurance. The generally accepted definition of Design-Build is that it is a very condensed process ruled by a single contract between the owner and the design-build team (firm) from the beginning of the project. The main advantages to the design-build approach are single point of responsibility, accurate identification of project costs, faster delivery of facility, improved risk management, and better quality control. Each of these points will be discussed later as related to the NISC project at Los Alamos. Many of these points can be applied to other projects as well.

  7. Mission design for the low-cost Mariner Mark II missions

    NASA Technical Reports Server (NTRS)

    Wallace, R. A.; Blume, W. H.; Hulkower, N. D.; Yen, C. L.

    1982-01-01

    Mariner Mark II is a program of missions, now under study at JPL, which will maximize scientific return at substantially reduced cost. There will be 3 to 5 missions in the program investigating comets, asteroids, the outer planets and their satellites, and Mars in the 1990s. Mission opportunities for these targets in this time period are described in terms of launch vehicle, propulsion, and flight time requirements, as well as other mission constraints such as margin and launch period objectives. Example encounter designs as well as mission launch scenarios are also described.

  8. Design options for a bunsen reactor.

    SciTech Connect

    Moore, Robert Charles

    2013-10-01

    This work is being performed for Matt Channon Consulting as part of the Sandia National Laboratories New Mexico Small Business Assistance Program (NMSBA). Matt Channon Consulting has requested Sandia's assistance in the design of a chemical Bunsen reactor for the reaction of SO2, I2 and H2O to produce H2SO4 and HI with a SO2 feed rate to the reactor of 50 kg/hour. Based on this value, an assumed reactor efficiency of 33%, and kinetic data from the literature, a plug flow reactor approximately 1%E2%80%9D diameter and and 12 inches long would be needed to meet the specification of the project. Because the Bunsen reaction is exothermic, heat in the amount of approximately 128,000 kJ/hr would need to be removed using a cooling jacket placed around the tubular reactor. The available literature information on Bunsen reactor design and operation, certain support equipment needed for process operation and a design that meet the specification of Matt Channon Consulting are presented.

  9. Radioisotope Electric Propulsion Missions Utilizing a Common Spacecraft Design

    NASA Technical Reports Server (NTRS)

    Fiehler, Douglas; Oleson, Steven

    2004-01-01

    A study was conducted that shows how a single Radioisotope Electric Propulsion (REP) spacecraft design could be used for various missions throughout the solar system. This spacecraft design is based on a REP feasibility design from a study performed by NASA Glenn Research Center and the Johns Hopkins University Applied Physics Laboratory. The study also identifies technologies that need development to enable these missions. The mission baseline for the REP feasibility design study is a Trojan asteroid orbiter. This mission sends an REP spacecraft to Jupiter s leading Lagrange point where it would orbit and examine several Trojan asteroids. The spacecraft design from the REP feasibility study would also be applicable to missions to the Centaurs, and through some change of payload configuration, could accommodate a comet sample-return mission. Missions to small bodies throughout the outer solar system are also within reach of this spacecraft design. This set of missions, utilizing the common REP spacecraft design, is examined and required design modifications for specific missions are outlined.

  10. Attitude Design for the LADEE Mission

    NASA Technical Reports Server (NTRS)

    Galal, Ken; Nickel, Craig; Sherman, Ryan

    2015-01-01

    The Lunar Atmosphere and Dust Environment Explorer (LADEE) satellite successfully completed its 148-day science investigation in a low-altitude, near-equatorial lunar orbit on April 18, 2014. The LADEE spacecraft was built, managed and operated by NASA's Ames Research Center (ARC). The Mission Operations Center (MOC) was located at Ames and was responsible for activity planning, command sequencing, trajectory and attitude design, orbit determination, and spacecraft operations. The Science Operations Center (SOC) was located at Goddard Space Flight Center and was responsible for science planning, data archiving and distribution. This paper details attitude design and operations support for the LADEE mission. LADEE's attitude design was shaped by a wide range of instrument pointing requirements that necessitated regular excursions from the baseline one revolution per orbit "Ram" attitude. Such attitude excursions were constrained by a number of flight rules levied to protect instruments from the Sun, avoid geometries that would result in simultaneous occlusion of LADEE's two star tracker heads, and maintain the spacecraft within its thermal and power operating limits. To satisfy LADEE's many attitude requirements and constraints, a set of rules and conventions was adopted to manage the complexity of this design challenge and facilitate the automation of ground software that generated pointing commands spanning multiple days of operations at a time. The resulting LADEE Flight Dynamics System (FDS) that was developed used Visual Basic scripts that generated instructions to AGI's Satellite Tool Kit (STK) in order to derive quaternion commands at regular intervals that satisfied LADEE's pointing requirements. These scripts relied heavily on the powerful "align and constrain" capability of STK's attitude module to construct LADEE's attitude profiles and the slews to get there. A description of the scripts and the attitude modeling they embodied is provided. One particular

  11. Abort Options for Human Lunar Missions between Earth Orbit and Lunar Vicinity

    NASA Technical Reports Server (NTRS)

    Condon, Gerald L.; Senent, Juan S.; Llama, Eduardo Garcia

    2005-01-01

    Apollo mission design emphasized operational flexibility that supported premature return to Earth. However, that design was tailored to use expendable hardware for short expeditions to low-latitude sites and cannot be applied directly to an evolutionary program requiring long stay times at arbitrary sites. This work establishes abort performanc e requirements for representative onorbit phases of missions involvin g rendezvous in lunar-orbit, lunar-surface and at the Earth-Moon libr ation point. This study submits reference abort delta-V requirements and other Earth return data (e.g., entry speed, flight path angle) and also examines the effect of abort performance requirements on propul sive capability for selected vehicle configurations.

  12. Space Station needs, attributes and architectural options. Volume 2, book 1, part 1: Mission requirements

    NASA Technical Reports Server (NTRS)

    1983-01-01

    The baseline mission model used to develop the space station mission-related requirements is described as well as the 90 civil missions that were evaluated, (including the 62 missions that formed the baseline model). Mission-related requirements for the space station baseline are defined and related to space station architectural development. Mission-related sensitivity analyses are discussed.

  13. Mission Design for the Lunar Reconnaissance Orbiter

    NASA Technical Reports Server (NTRS)

    Beckman, Mark

    2007-01-01

    The Lunar Reconnaissance Orbiter (LRO) will be the first mission under NASA's Vision for Space Exploration. LRO will fly in a low 50 km mean altitude lunar polar orbit. LRO will utilize a direct minimum energy lunar transfer and have a launch window of three days every two weeks. The launch window is defined by lunar orbit beta angle at times of extreme lighting conditions. This paper will define the LRO launch window and the science and engineering constraints that drive it. After lunar orbit insertion, LRO will be placed into a commissioning orbit for up to 60 days. This commissioning orbit will be a low altitude quasi-frozen orbit that minimizes stationkeeping costs during commissioning phase. LRO will use a repeating stationkeeping cycle with a pair of maneuvers every lunar sidereal period. The stationkeeping algorithm will bound LRO altitude, maintain ground station contact during maneuvers, and equally distribute periselene between northern and southern hemispheres. Orbit determination for LRO will be at the 50 m level with updated lunar gravity models. This paper will address the quasi-frozen orbit design, stationkeeping algorithms and low lunar orbit determination.

  14. 2016 Mars Insight Mission Design and Navigation

    NASA Technical Reports Server (NTRS)

    Abilleira, Fernando; Frauenholz, Ray; Fujii, Ken; Wallace, Mark; You, Tung-Han

    2014-01-01

    Scheduled for a launch in the 2016 Earth to Mars opportunity, the Interior Exploration using Seismic Investigations, Geodesy, and Heat Transport (InSight) Mission will arrive to Mars in late September 2016 with the primary objective of placing a science lander on the surface of the Red Planet followed by the deployment of two science instruments to investigate the fundamental processes of terrestrial planet formation and evolution. In order to achieve a successful landing, the InSight Project has selected a launch/arrival strategy that satisfies the following key and driving requirements: (1) Deliver a total launch mass of 727 kg, (2) target a nominal landing site with a cumulative Delta V99 less than 30 m/s, and (3) approach EDL with a V-infinity upper limit of 3.941 km/s and (4) an entry flight-path angle (EFPA) of -12.5 +/- 0.26 deg, 3-sigma; the InSight trajectories have been designed such that they (5) provide UHF-band communications via Direct-To-Earth and MRO from Entry through landing plus 60 s, (6) with injection aimpoints biased away from Mars such that the probability of the launch vehicle upper stage impacting Mars is less than 1.0 X 10(exp 4) for fifty years after launch, and (7) non-nominal impact probabilities due to failure during the Cruise phase less than 1.0 X 10(exp 2).

  15. Cloud Computing for Mission Design and Operations

    NASA Technical Reports Server (NTRS)

    Arrieta, Juan; Attiyah, Amy; Beswick, Robert; Gerasimantos, Dimitrios

    2012-01-01

    The space mission design and operations community already recognizes the value of cloud computing and virtualization. However, natural and valid concerns, like security, privacy, up-time, and vendor lock-in, have prevented a more widespread and expedited adoption into official workflows. In the interest of alleviating these concerns, we propose a series of guidelines for internally deploying a resource-oriented hub of data and algorithms. These guidelines provide a roadmap for implementing an architecture inspired in the cloud computing model: associative, elastic, semantical, interconnected, and adaptive. The architecture can be summarized as exposing data and algorithms as resource-oriented Web services, coordinated via messaging, and running on virtual machines; it is simple, and based on widely adopted standards, protocols, and tools. The architecture may help reduce common sources of complexity intrinsic to data-driven, collaborative interactions and, most importantly, it may provide the means for teams and agencies to evaluate the cloud computing model in their specific context, with minimal infrastructure changes, and before committing to a specific cloud services provider.

  16. Mission design for Human Outer Planet Exploration (HOPE) using a magnetoplasma spacecraft

    NASA Astrophysics Data System (ADS)

    Park, Sang-Young; Seywald, Hans; Krizan, Shawn A.; Stillwagen, Frederic H.

    2006-08-01

    To send humans beyond Mars, a Human Outer Planet Exploration (HOPE) mission has been studied for new spacecraft concepts and technologies. In this paper, an interplanetary trajectory and a preliminary spacecraft design are presented for the HOPE visit to Callisto, one of Jupiter's moons. To design a round-trip trajectory for the mission, the characteristics of the spacecraft and its trajectories are analyzed. A detailed optimization approach is formulated to utilize a Variable Specific Impulse Magnetoplasma Rocket (VASIMR) engine with capabilities of variable specific impulse, variable engine efficiency, and engine on-off control. It is mainly illustrated that a 30 MW powered spacecraft can make the mission possible in a 5-year round trip constraint around the year 2045. Trajectories with different power and reactor options are also discussed. The results obtained in this study can be used for formulating an overall concept for the mission.

  17. ExoMars Mission Analysis and Design - Launch, Cruise and Arrival Analyses

    NASA Technical Reports Server (NTRS)

    Cano, Juan L.; Cacciatore, Francesco

    2007-01-01

    ExoMars is ESA s next mission to planet Mars. The probe is aimed for launch either in 2013 or in 2016. The project is currently undergoing Phase B1 studies under ESA management and Thales Alenia Space Italia project leadership. In that context, DEIMOS Space is responsible for the Mission Analysis and Design for the interplanetary and the entry, descent and landing (EDL) activities. The present mission baseline is based on an Ariane 5 or Proton M launch in 2013 of a spacecraft Composite bearing a Carrier Module (CM) and a Descent Module (DM). A back-up option is proposed in 2016. This paper presents the current status of the interplanetary mission design from launch up to the start of the EDL phase.

  18. Modeling and Simulation for Mission Operations Work System Design

    NASA Technical Reports Server (NTRS)

    Sierhuis, Maarten; Clancey, William J.; Seah, Chin; Trimble, Jay P.; Sims, Michael H.

    2003-01-01

    Work System analysis and design is complex and non-deterministic. In this paper we describe Brahms, a multiagent modeling and simulation environment for designing complex interactions in human-machine systems. Brahms was originally conceived as a business process design tool that simulates work practices, including social systems of work. We describe our modeling and simulation method for mission operations work systems design, based on a research case study in which we used Brahms to design mission operations for a proposed discovery mission to the Moon. We then describe the results of an actual method application project-the Brahms Mars Exploration Rover. Space mission operations are similar to operations of traditional organizations; we show that the application of Brahms for space mission operations design is relevant and transferable to other types of business processes in organizations.

  19. Designing Medical Support for a Near-Earth Asteroid Mission

    NASA Technical Reports Server (NTRS)

    Watkins, S. D.; Charles, J. B.; Kundrot, C. E.; Barr, Y. R.; Barsten, K. N.; Chin, D. A.; Kerstman, E. L.; Otto, C.

    2011-01-01

    This panel will discuss the design of medical support for a mission to a near-Earth asteroid (NEA) from a variety of perspectives. The panelists will discuss the proposed parameters for a NEA mission, the NEA medical condition list, recommendations from the NASA telemedicine workshop, an overview of the Exploration Medical System Demonstration planned for the International Space Station, use of predictive models for mission planning, and mission-related concerns for behavioral health and performance. This panel is intended to make the audience aware of the multitude of factors influencing medical support during a NEA mission.

  20. Design Reference Missions for Deep-Space Optical Communication

    NASA Astrophysics Data System (ADS)

    Breidenthal, J.; Abraham, D.

    2016-05-01

    We examined the potential, but uncertain, NASA mission portfolio out to a time horizon of 20 years, to identify mission concepts that potentially could benefit from optical communication, considering their communications needs, the environments in which they would operate, and their notional size, weight, and power constraints. A set of 12 design reference missions was selected to represent the full range of potential missions. These design reference missions span the space of potential customer requirements, and encompass the wide range of applications that an optical ground segment might eventually be called upon to serve. The design reference missions encompass a range of orbit types, terminal sizes, and positions in the solar system that reveal the chief system performance variables of an optical ground segment, and may be used to enable assessments of the ability of alternative systems to meet various types of customer needs.

  1. The Space Infrared Interferometric Telescope (SPIRIT): The Mission Design Solution Space and the Art of the Possible

    NASA Technical Reports Server (NTRS)

    Leisawitz, David; Hyde, T. Tupper; Rinehart, Stephen A.; Weiss, Michael

    2008-01-01

    Although the Space Infrared Interferometric Telescope (SPIRIT) was studied as a candidate NASA Origins Probe mission, the real world presents a broader set of options, pressures, and constraints. Fundamentally, SPIRIT is a far-IR observatory for high-resolution imaging and spectroscopy designed to address a variety of compelling scientific questions. How do planetary systems form from protostellar disks, dousing some planets in water while leaving others dry? Where do planets form, and why are some ice giants while others are rocky? How did high-redshift galaxies form and merge to form the present-day population of galaxies? This paper takes a pragmatic look at the mission design solution space for SPIRIT, presents Probe-class and facility-class mission scenarios, and describes optional design changes. The costs and benefits of various mission design alternatives are roughly evaluated, giving a basis for further study and to serve as guidance to policy makers.

  2. Mission design considerations for nuclear risk mitigation

    NASA Technical Reports Server (NTRS)

    Stancati, Mike; Collins, John

    1993-01-01

    Strategies for the mitigation of the nuclear risk associated with two specific mission operations are discussed. These operations are the safe return of nuclear thermal propulsion reactors to earth orbit and the disposal of lunar/Mars spacecraft reactors.

  3. Mixed mode missions - Designing payloads to match flight opportunities

    NASA Technical Reports Server (NTRS)

    Noblitt, B. G.; Kissin, K.; Moke, R.; Mcannally, R.

    1980-01-01

    Shuttle missions have usually been viewed as being devoted to a single mission mode. But full utilization of the STS capabilities as a platform for space research dictates a flexible strategy toward payload integration and mission design. Simple interfaces, single disciplines, and large design and resource margins point the way toward larger payload integration and operation costs. Payloads for early missions (i.e., OSTA-1, OSS-1, and MEA) are used as illustrations to support this argument. The implications of this approach for future NASA plans as well as for new STS operational concepts are considered.

  4. Mission design applications of QUICK. [software for interactive trajectory calculation

    NASA Technical Reports Server (NTRS)

    Skinner, David L.; Bass, Laura E.; Byrnes, Dennis V.; Cheng, Jeannie T.; Fordyce, Jess E.; Knocke, Philip C.; Lyons, Daniel T.; Pojman, Joan L.; Stetson, Douglas S.; Wolf, Aron A.

    1990-01-01

    An overview of an interactive software environment for space mission design termed QUICK is presented. This stand-alone program provides a programmable FORTRAN-like calculator interface to a wide range of both built-in and user defined functions. QUICK has evolved into a general-purpose software environment that can be intrinsically and dynamically customized for a wide range of mission design applications. Specific applications are described for some space programs, e.g., the earth-Venus-Mars mission, the Cassini mission to Saturn, the Mars Observer, the Galileo Project, and the Magellan Spacecraft.

  5. JPL Mission Design Software: Current Efforts to Support Low-Cost Missions

    NASA Technical Reports Server (NTRS)

    Fordyce, J.

    1994-01-01

    Over the last several decades, engineers at the Jet Propulsion Laboratory have developed a collection of analytical tools to design missions to Earth orbit, the moon, sun, planets and various other bodies in our solar system, and beyond.

  6. Collaborative Mission Design at NASA Langley Research Center

    NASA Technical Reports Server (NTRS)

    Gough, Kerry M.; Allen, B. Danette; Amundsen, Ruth M.

    2005-01-01

    NASA Langley Research Center (LaRC) has developed and tested two facilities dedicated to increasing efficiency in key mission design processes, including payload design, mission planning, and implementation plan development, among others. The Integrated Design Center (IDC) is a state-of-the-art concurrent design facility which allows scientists and spaceflight engineers to produce project designs and mission plans in a real-time collaborative environment, using industry-standard physics-based development tools and the latest communication technology. The Mission Simulation Lab (MiSL), a virtual reality (VR) facility focused on payload and project design, permits engineers to quickly translate their design and modeling output into enhanced three-dimensional models and then examine them in a realistic full-scale virtual environment. The authors were responsible for envisioning both facilities and turning those visions into fully operational mission design resources at LaRC with multiple advanced capabilities and applications. In addition, the authors have created a synergistic interface between these two facilities. This combined functionality is the Interactive Design and Simulation Center (IDSC), a meta-facility which offers project teams a powerful array of highly advanced tools, permitting them to rapidly produce project designs while maintaining the integrity of the input from every discipline expert on the project. The concept-to-flight mission support provided by IDSC has shown improved inter- and intra-team communication and a reduction in the resources required for proposal development, requirements definition, and design effort.

  7. 47 CFR 1.2202 - Competitive bidding design options.

    Code of Federal Regulations, 2014 CFR

    2014-10-01

    ... 47 Telecommunication 1 2014-10-01 2014-10-01 false Competitive bidding design options. 1.2202 Section 1.2202 Telecommunication FEDERAL COMMUNICATIONS COMMISSION GENERAL PRACTICE AND PROCEDURE Grants by Random Selection Competitive Bidding Proceedings Broadcast Television Spectrum Reverse Auction §...

  8. Trajectory optimization software for planetary mission design

    NASA Technical Reports Server (NTRS)

    D'Amario, Louis A.

    1989-01-01

    The development history and characteristics of the interactive trajectory-optimization programs MOSES (D'Amario et al., 1981) and PLATO (D'Amario et al., 1982) are briefly reviewed, with an emphasis on their application to the Galileo mission. The requirements imposed by a mission involving flybys of several planetary satellites or planets are discussed; the formulation of the parameter-optimization problem is outlined; and particular attention is given to the use of multiconic methods to model the gravitational attraction of Jupiter in MOSES. Diagrams and tables of numerical data are included.

  9. A Neptune Orbiter Mission

    NASA Technical Reports Server (NTRS)

    Wallace, R. A.; Spilker, T. R.

    1998-01-01

    This paper describes the results of new analyses and mission/system designs for a low cost Neptune Orbiter mission. Science and measurement objectives, instrumentation, and mission/system design options are described and reflect an aggressive approach to the application of new advanced technologies expected to be available and developed over the next five to ten years.

  10. Interplanetary Trajectory Design for the Asteroid Robotic Redirect Mission Alternate Approach Trade Study

    NASA Technical Reports Server (NTRS)

    Merrill, Raymond Gabriel; Qu, Min; Vavrina, Matthew A.; Englander, Jacob A.; Jones, Christopher A.

    2014-01-01

    This paper presents mission performance analysis methods and results for the Asteroid Robotic Redirect Mission (ARRM) option to capture a free standing boulder on the surface of a 100 m or larger NEA. It details the optimization and design of heliocentric low-thrust trajectories to asteroid targets for the ARRM solar electric propulsion spacecraft. Extensive searches were conducted to determine asteroid targets with large pick-up mass potential and potential observation opportunities. Interplanetary trajectory approximations were developed in method based tools for Itokawa, Bennu, 1999 JU3, and 2008 EV5 and were validated by end-to-end integrated trajectories.

  11. 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.

  12. Solar Power System Design for the Solar Probe+ Mission

    NASA Technical Reports Server (NTRS)

    Landis, Geoffrey A.; Schmitz, Paul C.; Kinnison, James; Fraeman, Martin; Roufberg, Lew; Vernon, Steve; Wirzburger, Melissa

    2008-01-01

    Solar Probe+ is an ambitious mission proposed to the solar corona, designed to make a perihelion approach of 9 solar radii from the surface of the sun. The high temperature, high solar flux environment makes this mission a significant challenge for power system design. This paper summarizes the power system conceptual design for the solar probe mission. Power supplies considered included nuclear, solar thermoelectric generation, solar dynamic generation using Stirling engines, and solar photovoltaic generation. The solar probe mission ranges from a starting distance from the sun of 1 AU, to a minimum distance of about 9.5 solar radii, or 0.044 AU, from the center of the sun. During the mission, the solar intensity ranges from one to about 510 times AM0. This requires power systems that can operate over nearly three orders of magnitude of incident intensity.

  13. Designing Electrostatic Accelerometers for Next Gravity Missions

    NASA Astrophysics Data System (ADS)

    Huynh, Phuong-Anh; Foulon, Bernard; Christophe, Bruno; Liorzou, Françoise; Boulanger, Damien; Lebat, Vincent

    2016-04-01

    Square cuboid electrostatic accelerometers sensor core have been used in various combinations in recent and still flying missions (CHAMP, GRACE, GOCE). ONERA is now in the process of delivering such accelerometers for the GRACE Follow-On mission. The goal is to demonstrate the performance benefits of an interferometry laser ranging method for future low-low satellite to satellite missions. The electrostatic accelerometer becoming thus the system main performance limiter, we propose for future missions a new symmetry which will allow for three ultrasensitive axes instead of two. This implies no performance ground testing, as the now cubic proof-mass will be too heavy, but only free fall tests in catapult mode, taking advantage of the additional microgravity testing time offered by the updated ZARM tower. The updated mission will be in better adequacy with the requirements of a next generation of smaller and drag compensated micro-satellites. In addition to the measurement of the surface forces exerted on the spacecraft by the atmospheric drag and by radiation pressures, the accelerometer will become a major part of the attitude and orbit control system by acting as drag free sensor and by accurately measuring the angular accelerations. ONERA also works on a hybridization of the electrostatic accelerometer with an atomic interferometer to take advantage of the absolute nature of the atomic interferometer acceleration measurement and its great accuracy in the [5-100] mHz bandwidth. After a description of the improvement of the GRACE-FO accelerometer with respect to the still in-orbit previous models and a status of its development, the presentation will describe the new cubic configuration and how its operations and performances can be verified in the Bremen drop tower.

  14. PFERD Mission: Pluto Flyby Exploration/Research Design

    NASA Technical Reports Server (NTRS)

    Lemke, Gary; Zayed, Husni; Herring, Jason; Fuehne, Doug; Sutton, Kevin; Sharkey, Mike

    1990-01-01

    The Pluto Flyby Exploration/Research Design (PFERD) mission will consist of a flyby spacecraft to Pluto and its satellite, Charon. The mission lifetime is expected to be 18 years. The Titan 4 with a Centaur upper stage will be utilized to launch the craft into the transfer orbit. The proposal was divided into six main subsystems: (1) scientific instrumentation; (2) command, communications, and control: (3) altitude and articulation control; (4) power and propulsion; (5) structures and thermal control; and (6) mission management and costing. Tradeoff studies were performed to optimize all factors of design, including survivability, performance, cost, and weight. Problems encountered in the design are also presented.

  15. 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.

  16. Design and application of electromechanical actuators for deep space missions

    NASA Technical Reports Server (NTRS)

    Haskew, Tim A.; Wander, John

    1993-01-01

    The annual report Design and Application of Electromechanical Actuators for Deep Space Missions is presented. The reporting period is 16 Aug. 1992 to 15 Aug. 1993. However, the primary focus will be work performed since submission of our semi-annual progress report in Feb. 1993. Substantial progress was made. We currently feel confident in providing guidelines for motor and control strategy selection in electromechanical actuators to be used in thrust vector control (TVC) applications. A small portion was presented in the semi-annual report. At this point, we have implemented highly detailed simulations of various motor/drive systems. The primary motor candidates were the brushless dc machine, permanent magnet synchronous machine, and the induction machine. The primary control implementations were pulse width modulation and hysteresis current control. Each of the two control strategies were applied to each of the three motor choices. With either pulse width modulation or hysteresis current control, the induction machine was always vector controlled. A standard test position command sequence for system performance evaluation is defined. Currently, we are gathering all of the necessary data for formal presentation of the results. Briefly stated for TVC application, we feel that the brushless dc machine operating under PWM current control is the best option. Substantial details on the topic, with supporting simulation results, will be provided later, in the form of a technical paper prepared for submission and also in the next progress report with more detail than allowed for paper publication.

  17. Design and application of electromechanical actuators for deep space missions

    NASA Astrophysics Data System (ADS)

    Haskew, Tim A.; Wander, John

    1993-09-01

    The annual report Design and Application of Electromechanical Actuators for Deep Space Missions is presented. The reporting period is 16 Aug. 1992 to 15 Aug. 1993. However, the primary focus will be work performed since submission of our semi-annual progress report in Feb. 1993. Substantial progress was made. We currently feel confident in providing guidelines for motor and control strategy selection in electromechanical actuators to be used in thrust vector control (TVC) applications. A small portion was presented in the semi-annual report. At this point, we have implemented highly detailed simulations of various motor/drive systems. The primary motor candidates were the brushless dc machine, permanent magnet synchronous machine, and the induction machine. The primary control implementations were pulse width modulation and hysteresis current control. Each of the two control strategies were applied to each of the three motor choices. With either pulse width modulation or hysteresis current control, the induction machine was always vector controlled. A standard test position command sequence for system performance evaluation is defined. Currently, we are gathering all of the necessary data for formal presentation of the results. Briefly stated for TVC application, we feel that the brushless dc machine operating under PWM current control is the best option. Substantial details on the topic, with supporting simulation results, will be provided later, in the form of a technical paper prepared for submission and also in the next progress report with more detail than allowed for paper publication.

  18. Orbit Options for an Orion-Class Spacecraft Mission to a Near-Earth Object

    NASA Astrophysics Data System (ADS)

    Shupe, Nathan C.

    Based on the recommendations of the Augustine Commission, President Obama has proposed a vision for U.S. human spaceflight in the post-Shuttle era which includes a manned mission to a Near-Earth Object (NEO). A 2006-2007 study commissioned by the Constellation Program Advanced Projects Office investigated the feasibility of sending a crewed Orion spacecraft to a NEO using different combinations of elements from the latest launch system architecture at that time. The study found a number of suitable mission targets in the database of known NEOs, and predicted that the number of candidate NEOs will continue to increase as more advanced observatories come online and execute more detailed surveys of the NEO population. The objective of this thesis is to pick up where the previous Constellation study left off by considering what orbit options are available for an Orion-class spacecraft upon arrival at a NEO. A model including multiple perturbations (solar radiation pressure, solar gravity, non-spherical mass distribution of the central body) to two-body dynamics is constructed to numerically integrate the motion of a satellite in close proximity to a small body in an elliptical orbit about the Sun. Analytical limits derived elsewhere in the literature for the thresholds on the size of the satellite orbit required to maintain stability in the presence of these perturbing forces are verified by the numerical model. Simulations about NEOs possessing various physical parameters (size, shape, rotation period) are then used to empirically develop general guidelines for establishing orbits of an Orion-class spacecraft about a NEO. It is found that an Orion-class spacecraft can orbit NEOs at any distance greater than the NEO surface height and less than the maximum semi-major axis allowed by the solar radiation pressure perturbation, provided that the ellipticity perturbation is sufficiently weak (this condition is met if the NEO is relatively round and/or has a long rotation

  19. Solar Probe Plus: Mission design challenges and trades

    NASA Astrophysics Data System (ADS)

    Guo, Yanping

    2010-11-01

    NASA plans to launch the first mission to the Sun, named Solar Probe Plus, as early as 2015, after a comprehensive feasibility study that significantly changed the original Solar Probe mission concept. The original Solar Probe mission concept, based on a Jupiter gravity assist trajectory, was no longer feasible under the new guidelines given to the mission. A complete redesign of the mission was required, which called for developing alternative trajectories that excluded a flyby of Jupiter. Without the very powerful gravity assist from Jupiter it was extremely difficult to get to the Sun, so designing a trajectory to reach the Sun that is technically feasible under the new mission guidelines became a key enabler to this highly challenging mission. Mission design requirements and challenges unique to this mission are reviewed and discussed, including various mission scenarios and six different trajectory designs utilizing various planetary gravity assists that were considered. The V 5GA trajectory design using five Venus gravity assists achieves a perihelion of 11.8 solar radii ( RS) in 3.3 years without any deep space maneuver (DSM). The V 7GA trajectory design reaches a perihelion of 9.5 RS using seven Venus gravity assists in 6.39 years without any DSM. With nine Venus gravity assists, the V 9GA trajectory design shows a solar orbit at inclination as high as 37.9° from the ecliptic plane can be achieved with the time of flight of 5.8 years. Using combined Earth and Venus gravity assists, as close as 9 RS from the Sun can be achieved in less than 10 years of flight time at moderate launch C3. Ultimately the V 7GA trajectory was chosen as the new baseline mission trajectory. Its design allowing for science investigation right after launch and continuing for nearly 7 years is unprecedented for interplanetary missions. The redesigned Solar Probe Plus mission is not only feasible under the new guidelines but also significantly outperforms the original mission concept

  20. Revised Point of Departure Design Options for Nuclear Thermal Propulsion

    NASA Technical Reports Server (NTRS)

    Fittje, James E.; Borowski, Stanley K.; Schnitzler, Bruce

    2015-01-01

    In an effort to further refine potential point of departure nuclear thermal rocket engine designs, four proposed engine designs representing two thrust classes and utilizing two different fuel matrix types are designed and analyzed from both a neutronics and thermodynamic cycle perspective. Two of these nuclear rocket engine designs employ a tungsten and uranium dioxide cermet (ceramic-metal) fuel with a prismatic geometry based on the ANL-200 and the GE-710, while the other two designs utilize uranium-zirconium-carbide in a graphite composite fuel and a prismatic fuel element geometry developed during the Rover/NERVA Programs. Two engines are analyzed for each fuel type, a small criticality limited design and a 111 kN (25 klbf) thrust class engine design, which has been the focus of numerous manned mission studies, including NASA's Design Reference Architecture 5.0. slightly higher T/W ratios, but they required substantially more 235U.

  1. Multi-mission nuclear electric propulsion stage design.

    NASA Technical Reports Server (NTRS)

    Prickett, W. Z.; Stearns, J. W.

    1973-01-01

    Results of a mission engineering analysis of nuclear-thermionic electric propulsion spacecraft for unmanned interplanetary and geocentric missions. Critical technologies assessed are associated with the development of nuclear electric propulsion (NEP), and the impact of its availability on future space programs. Specific areas of investigation include outer planet and comet rendezvous mission analysis, NEP stage design for geocentric and interplanetary missions, and technology requirements for NEP stage development. A multimission NEP stage can be developed to perform both multiple geocentric and interplanetary missions for a 1983 launch. Identified pacing NEP technology requirements are the development of 20,000 full power hour ion thrustors and thermionic reactor and the development of related power conditioning. The resulting NEP stage design provides both inherent reliability and high payload mass capability.

  2. Design of human missions to Mars and robotic missions to Jupiter

    NASA Astrophysics Data System (ADS)

    Okutsu, Masataka

    We consider human missions to Mars and robotic missions to Jupiter for launch dates in the near- and far-future. For the near-future, we design trajectories for currently proposed space missions that have well-defined spacecraft and mission requirements. For example, for early human missions to Mars we assume that the constraints used in NASA's design reference missions are indicative of current and near-future technologies, which of course limit our capabilities to explore Mars--and these limits make the problem challenging. Similarly, in the case of robotic exploration of Jupiter, we consider that the technology levels assumed for the proposed Europa Orbiter mission represent reasonable limits. For the far-future (two to three decades from now), we take the best estimates from current literature about the capabilities that may be available in nuclear-powered electric propulsion. We consider hardware capabilities (in terms of specific mass, specific impulse, thrust, power, etc.) for low-thrust trajectories, which range froth near-term to far-future technologies. In designing such missions, several techniques are found useful. For example, the Tisserand Graph, which tracks the changes in orbital shapes and energies, provides insight in designing Jovian tours for the Europa Orbiter mission. The graph is also useful in analyzing abort trajectories for human missions to Mars. Furthermore, a patched-conic propagator, which can generate thousands of potential trajectories, plays a vital role in three of four chapters of this thesis. For launches in the next three decades, we discovered a class of Earth- Mars-Venus-Earth free returns (which appear only four times in the 100-year period), Jovian tours involving ten to twenty flybys of the Galilean satellites, and low-thrust trajectories to Jupiter via gravity assists from Venus, Earth, and Mars. In addition, our continuation method, in which a solution for a conic trajectory is gradually converted into that for a low

  3. Planning Coverage Campaigns for Mission Design and Analysis: Clasp for the Proposed DESDynI Mission

    NASA Technical Reports Server (NTRS)

    Knight, Russell; McLaren, David; Hu, Steven

    2012-01-01

    Mission design and analysis present challenges in that almost all variables are in constant flux, yet the goal is to achieve an acceptable level of performance against a concept of operations, which might also be in flux. To increase responsiveness, our approach is to use automated planning tools that allow for the continual modification of spacecraft, ground system, staffing, and concept of operations while returning metrics that are important to mission evaluation, such as area covered, peak memory usage, and peak data throughput. We have applied this approach to DESDynI (Deformation, Ecosystem Structure, and Dynamics of Ice) mission design concept using the CLASP (Compressed Large-scale Activity Scheduler/Planner) planning system [7], but since this adaptation many techniques have changed under the hood for CLASP and the DESDynI mission concept has undergone drastic changes, including that it has been renamed the Earth Radar Mission. Over the past two years, we have run more than fifty simulations with the CLASP-DESDynI adaptation, simulating different mission scenarios with changing parameters including targets, swaths, instrument modes, and data and downlink rates. We describe the evolution of simulations through the DESDynI MCR (Mission Concept Review) and afterwards.

  4. Overview of Mission Design for NASA Asteroid Redirect Robotic Mission Concept

    NASA Technical Reports Server (NTRS)

    Strange, Nathan; Landau, Damon; McElrath, Timothy; Lantoine, Gregory; Lam, Try; McGuire, Melissa; Burke, Laura; Martini, Michael; Dankanich, John

    2013-01-01

    Part of NASA's new asteroid initiative would be a robotic mission to capture a roughly four to ten meter asteroid and redirect its orbit to place it in translunar space. Once in a stable storage orbit at the Moon, astronauts would then visit the asteroid for science investigations, to test in space resource extraction, and to develop experience with human deep space missions. This paper discusses the mission design techniques that would enable the redirection of a 100-1000 metric ton asteroid into lunar orbit with a 40-50 kW Solar Electric Propulsion (SEP) system.

  5. Orion Entry, Descent, and Landing Performance and Mission Design

    NASA Technical Reports Server (NTRS)

    Broome, Joel M.; Johnson, Wyatt

    2007-01-01

    The Orion Vehicle is the next spacecraft to take humans into space and will include missions to ISS as well as missions to the Moon. As part of that challenge, the vehicle will have to accommodate multiple mission design concepts, since return from Low Earth Orbit and return from the Moon can be quite different. Commonality between the different missions as it relates to vehicle systems, guidance capability, and operations concepts is the goal. Several unique mission design concepts include the specification of multiple land-based landing sites for a vehicle with closed-loop direct and skip entry guidance, followed by a parachute descent and landing attenuation system. This includes the ability of the vehicle to accurately target and land at a designated landing site, including site location aspects, landing site size, and landing opportunities assessments. Analyses associated with these mission design and flight performance challenges and constraints will be discussed as well as potential operational concepts to provide feasibility and/or mission commonality.

  6. Design of the ARES Mars Airplane and Mission Architecture

    NASA Technical Reports Server (NTRS)

    Braun, Robert D.; Wright, Henry S.; Croom, Mark A.; Levine, Joel S.; Spencer, David A.

    2006-01-01

    Significant technology advances have enabled planetary aircraft to be considered as viable science platforms. Such systems fill a unique planetary science measurement gap, that of regional-scale, near-surface observation, while providing a fresh perspective for potential discovery. Recent efforts have produced mature mission and flight system concepts, ready for flight project implementation. This paper summarizes the development of a Mars airplane mission architecture that balances science, implementation risk and cost. Airplane mission performance, flight system design and technology maturation are described. The design, analysis and testing completed demonstrates the readiness of this science platform for use in a Mars flight project.

  7. Analysis and design of aeroassisted interplanetary missions

    NASA Astrophysics Data System (ADS)

    Johnson, Wyatt R.

    An aeroassisted mission uses atmospheric forces to effect a spacecraft delta-V, which could allow for substantial propellant savings. This research focuses on aero-maneuvers useful for interplanetary flight. The aerogravity assist maneuver uses aerodynamic lift to achieve a greater delta-V than with gravity alone. Aerobraking and aerocapture both use aerodynamic drag to supplement or replace propulsive planetary captures. Optimal aerogravity assist trajectories are found, using a combination of analytic and graphical techniques. Simple control schemes are developed to manage angular momentum during aerobraking and to guide a spacecraft to a desired final orbit during aerocapture.

  8. The Ninevah Mission: A design summary for an unmanned mission to Venus, volume 1

    NASA Technical Reports Server (NTRS)

    1988-01-01

    The design summary for an unmanned mission to the planet Venus, with code name Ninevah, is presented. The design includes a Hohmann transfer trajectory analysis, propulsion trade study, an overview of the communication and instrumentation systems, power requirements, probe and lander analysis, and a weight and cost analysis.

  9. Design of Superconducting Gravity Gradiometer Cryogenic System for Mars Mission

    NASA Technical Reports Server (NTRS)

    Li, X.; Lemoine, F. G.; Paik, H. J.; Zagarola, M.; Shirron, P. J.; Griggs, C. E.; Moody, M. V.; Han, S.-C.

    2016-01-01

    Measurement of a planet's gravity field provides fundamental information about the planet's mass properties. The static gravity field reveals information about the internal structure of the planet, including crustal density variations that provide information on the planet's geological history and evolution. The time variations of gravity result from the movement of mass inside the planet, on the surface, and in the atmosphere. NASA is interested in a Superconducting Gravity Gradiometer (SGG) with which to measure the gravity field of a planet from orbit. An SGG instrument is under development with the NASA PICASSO program, which will be able to resolve the Mars static gravity field to degree 200 in spherical harmonics, and the time-varying field on a monthly basis to degree 20 from a 255 x 320 km orbit. The SGG has a precision two orders of magnitude better than the electrostatic gravity gradiometer that was used on the ESA's GOCE mission. The SGG operates at the superconducting temperature lower than 6 K. This study developed a cryogenic thermal system to maintain the SGG at the design temperature in Mars orbit. The system includes fixed radiation shields, a low thermal conductivity support structure and a two-stage cryocooler. The fixed radiation shields use double aluminized polyimide to emit heat from the warm spacecraft into the deep space. The support structure uses carbon fiber reinforced plastic, which has low thermal conductivity at cryogenic temperature and very high stress. The low vibration cryocooler has two stages, of which the high temperature stage operates at 65 K and the low temperature stage works at 6 K, and the heat rejection radiator works at 300 K. The study also designed a second option with a 4-K adiabatic demagnetization refrigerator (ADR) and two-stage 10-K turbo-Brayton cooler.

  10. Design of Superconducting Gravity Gradiometer Cryogenic System for Mars Mission

    NASA Technical Reports Server (NTRS)

    Li, X.; Lemoine, F. G.; Shirron, P. J.; Paik, H. J.; Griggs, C. E.; Moody, M. V.; Han, S. C.; Zagarola, M.

    2016-01-01

    Measurement of a planets gravity field provides fundamental information about the planets mass properties. The static gravity field reveals information about the internal structure of the planet, including crustal density variations that provide information on the planets geological history and evolution. The time variations of gravity result from the movement of mass inside the planet, on the surface, and in the atmosphere. NASA is interested in a Superconducting Gravity Gradiometer (SGG) with which to measure the gravity field of a planet from orbit. An SGG instrument is under development with the NASA PICASSO program, which will be able to resolve the Mars static gravity field to degree 200 in spherical harmonics, and the time-varying field on a monthly basis to degree 20 from a 255 x 320 km orbit. The SGG has a precision two orders of magnitude better than the electrostatic gravity gradiometer that was used on the ESAs GOCE mission. The SGG operates at the superconducting temperature lower than 6 K. This study developed a cryogenic thermal system to maintain the SGG at the design temperature in Mars orbit. The system includes fixed radiation shields, a low thermal conductivity support structure and a two-stage cryocooler. The fixed radiation shields use double aluminized polyimide to emit heat from the warm spacecraft into the deep space. The support structure uses carbon fiber reinforced plastic, which has low thermal conductivity at cryogenic temperature and very high stress. The low vibration cryocooler has two stages, of which the high temperature stage operates at 65 K and the low temperature stage works at 6 K, and the heat rejection radiator works at 300 K. The study also designed a second option with a 4-K adiabatic demagnetization refrigerator (ADR) and two-stage 10-K turbo-Brayton cooler.

  11. Nuclear Cryogenic Propulsion Stage Conceptual Design and Mission Analysis

    NASA Technical Reports Server (NTRS)

    Kos, Larry D.; Russell, Tiffany E.

    2014-01-01

    The Nuclear Cryogenic Propulsion Stage (NCPS) is an in-space transportation vehicle, comprised of three main elements, designed to support a long-stay human Mars mission architecture beginning in 2035. The stage conceptual design and the mission analysis discussed here support the current nuclear thermal propulsion going on within partnership activity of NASA and the Department of Energy (DOE). The transportation system consists of three elements: 1) the Core Stage, 2) the In-line Tank, and 3) the Drop Tank. The driving mission case is the piloted flight to Mars in 2037 and will be the main point design shown and discussed. The corresponding Space Launch System (SLS) launch vehicle (LV) is also presented due to it being a very critical aspect of the NCPS Human Mars Mission architecture due to the strong relationship between LV lift capability and LV volume capacity.

  12. Libration Orbit Mission Design: Applications of Numerical & Dynamical Methods

    NASA Technical Reports Server (NTRS)

    Bauer, Frank (Technical Monitor); Folta, David; Beckman, Mark

    2002-01-01

    Sun-Earth libration point orbits serve as excellent locations for scientific investigations. These orbits are often selected to minimize environmental disturbances and maximize observing efficiency. Trajectory design in support of libration orbits is ever more challenging as more complex missions are envisioned in the next decade. Trajectory design software must be further enabled to incorporate better understanding of the libration orbit solution space and thus improve the efficiency and expand the capabilities of current approaches. The Goddard Space Flight Center (GSFC) is currently supporting multiple libration missions. This end-to-end support consists of mission operations, trajectory design, and control. It also includes algorithm and software development. The recently launched Microwave Anisotropy Probe (MAP) and upcoming James Webb Space Telescope (JWST) and Constellation-X missions are examples of the use of improved numerical methods for attaining constrained orbital parameters and controlling their dynamical evolution at the collinear libration points. This paper presents a history of libration point missions, a brief description of the numerical and dynamical design techniques including software used, and a sample of future GSFC mission designs.

  13. Integrated Human-Robotic Missions to the Moon and Mars: Mission Operations Design Implications

    NASA Technical Reports Server (NTRS)

    Mishkin, Andrew; Lee, Young; Korth, David; LeBlanc, Troy

    2007-01-01

    For most of the history of space exploration, human and robotic programs have been independent, and have responded to distinct requirements. The NASA Vision for Space Exploration calls for the return of humans to the Moon, and the eventual human exploration of Mars; the complexity of this range of missions will require an unprecedented use of automation and robotics in support of human crews. The challenges of human Mars missions, including roundtrip communications time delays of 6 to 40 minutes, interplanetary transit times of many months, and the need to manage lifecycle costs, will require the evolution of a new mission operations paradigm far less dependent on real-time monitoring and response by an Earthbound operations team. Robotic systems and automation will augment human capability, increase human safety by providing means to perform many tasks without requiring immediate human presence, and enable the transfer of traditional mission control tasks from the ground to crews. Developing and validating the new paradigm and its associated infrastructure may place requirements on operations design for nearer-term lunar missions. The authors, representing both the human and robotic mission operations communities, assess human lunar and Mars mission challenges, and consider how human-robot operations may be integrated to enable efficient joint operations, with the eventual emergence of a unified exploration operations culture.

  14. Integrated Human-Robotic Missions to the Moon and Mars: Mission Operations Design Implications

    NASA Technical Reports Server (NTRS)

    Korth, David; LeBlanc, Troy; Mishkin, Andrew; Lee, Young

    2006-01-01

    For most of the history of space exploration, human and robotic programs have been independent, and have responded to distinct requirements. The NASA Vision for Space Exploration calls for the return of humans to the Moon, and the eventual human exploration of Mars; the complexity of this range of missions will require an unprecedented use of automation and robotics in support of human crews. The challenges of human Mars missions, including roundtrip communications time delays of 6 to 40 minutes, interplanetary transit times of many months, and the need to manage lifecycle costs, will require the evolution of a new mission operations paradigm far less dependent on real-time monitoring and response by an Earthbound operations team. Robotic systems and automation will augment human capability, increase human safety by providing means to perform many tasks without requiring immediate human presence, and enable the transfer of traditional mission control tasks from the ground to crews. Developing and validating the new paradigm and its associated infrastructure may place requirements on operations design for nearer-term lunar missions. The authors, representing both the human and robotic mission operations communities, assess human lunar and Mars mission challenges, and consider how human-robot operations may be integrated to enable efficient joint operations, with the eventual emergence of a unified exploration operations culture.

  15. A GNM mission and system design proposal

    NASA Technical Reports Server (NTRS)

    Bailey, Stephen

    1990-01-01

    Here, the author takes an advocacy position for the proposed Mars Global Network Mission (GNM); it is not intended to be an objective review, although both pros and cons are presented in summary. The mission consists of launches from earth in the '96, '98, and '01 opportunities on Delta-class launch vehicles (approx. 1000 kg injected to Mars in 8 to 10 ft diameter shroud). The trans Mars boost stage injects a stack of small independent, aeroshelled spacecraft. The stack separates from the boost stage and each rigid (as opposed to deployable) aeroshell flies to Mars on its own, performing midcourse maneuvers as necessary. Each spacecraft flies a unique trajectory which is targeted to achieve approach atmospheric interface at the desired latitude and lighting conditions; arrival times may vary by a month or more. A direct entry is performed, there is no propulsive orbit capture. The aeroshelled rough-landers are targeted to achieve a desired attitude and entry flight path angle, and then follow a passive ballistic trajectory until terminal descent. Based on sensed acceleration (integrated to deduce altitude), the aft aeroshell skirt is jettisoned; a short later a supersonic parachute is deployed. The ballistic coefficient of the parachute is sized to achieve terminal velocity at about 8 km. However the parachute is not deployed until a few Km above the surface to minimize wind-induced drift. The nose cap descent imaging begins, a laser altimeter also measures true altitude. Based on range and range rate to the surface, the parachute is jettisoned and the lander uses descent engines to achieve touchdown velocity. A contact sensor shuts down the motors to avoid cratering, and the lander rough-lands at less than 5 m/sec. The remaining aeroshell and a deployable bladder attenuate landing loads and minimize the possibility of tip over. Science instruments are deployed and activated, and the network is established.

  16. Conceptual design of a communications system for Mars exploration missions

    NASA Technical Reports Server (NTRS)

    Badi, Deborah M.; Farmer, Jeffery T.; Martin, Gary L.; Garn, Paul A.

    1989-01-01

    Future missions to Mars will require a communications system to link activities on the Martian surface with each other and with mission controllers on earth. This paper documents the preliminary design of an areosynchronous communications satellite to provide such links. As designed, the satellite will provide almost continuous communication between Mars and earth, as well as continuous Mars surface-to-surface links. The capacity will exist for voice, data/command, and video transmission. The mission scenario assumed for this design is described, followed by the configuration design. The communications systems conceptual design and its impact on the overall spacecraft requirements are then presented. Supporting subsystems including electrical power; thermal control; guidance, navigation, and control; propulsion; structures; and configuration are reviewed. A description of the operating orbit and the assumed method of orbit acquisition are discussed.

  17. Game Changing: NASA's Space Launch System and Science Mission Design

    NASA Technical Reports Server (NTRS)

    Creech, Stephen D.

    2013-01-01

    NASA s Marshall Space Flight Center (MSFC) is directing efforts to build the Space Launch System (SLS), a heavy-lift rocket that will carry the Orion Multi-Purpose Crew Vehicle (MPCV) and other important payloads far beyond Earth orbit (BEO). Its evolvable architecture will allow NASA to begin with Moon fly-bys and then go on to transport humans or robots to distant places such as asteroids and Mars. Designed to simplify spacecraft complexity, the SLS rocket will provide improved mass margins and radiation mitigation, and reduced mission durations. These capabilities offer attractive advantages for ambitious missions such as a Mars sample return, by reducing infrastructure requirements, cost, and schedule. For example, if an evolved expendable launch vehicle (EELV) were used for a proposed mission to investigate the Saturn system, a complicated trajectory would be required - with several gravity-assist planetary fly-bys - to achieve the necessary outbound velocity. The SLS rocket, using significantly higher C3 energies, can more quickly and effectively take the mission directly to its destination, reducing trip time and cost. As this paper will report, the SLS rocket will launch payloads of unprecedented mass and volume, such as "monolithic" telescopes and in-space infrastructure. Thanks to its ability to co-manifest large payloads, it also can accomplish complex missions in fewer launches. Future analyses will include reviews of alternate mission concepts and detailed evaluations of SLS figures of merit, helping the new rocket revolutionize science mission planning and design for years to come.

  18. Schooling by Design: Mission, Action, and Achievement

    ERIC Educational Resources Information Center

    Wiggins, Grant; McTighe, Jay

    2007-01-01

    An essential part of moving forward with the Understanding by Design[R] framework is to make sure its principles and strategies are reflected in all aspects of your school improvement efforts, including curriculum planning, leadership, teacher professional development, and action research. Here's a book designed to help you. From creating your…

  19. 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.

  20. Preliminary Report on Mission Design and Operations for Critical Events

    NASA Technical Reports Server (NTRS)

    Hayden, Sandra C.; Tumer, Irem

    2005-01-01

    Mission-critical events are defined in the Jet Propulsion Laboratory s Flight Project Practices as those sequences of events which must succeed in order to attain mission goals. These are dependent on the particular operational concept and design reference mission, and are especially important when committing to irreversible events. Critical events include main engine cutoff (MECO) after launch; engine cutoff or parachute deployment on entry, descent, and landing (EDL); orbital insertion; separation of payload from vehicle or separation of booster segments; maintenance of pointing accuracy for power and communication; and deployment of solar arrays and communication antennas. The purpose of this paper is to report on the current practices in handling mission-critical events in design and operations at major NASA spaceflight centers. The scope of this report includes NASA Johnson Space Center (JSC), NASA Goddard Space Flight Center (GSFC), and NASA Jet Propulsion Laboratory (JPL), with staff at each center consulted on their current practices, processes, and procedures.

  1. Satellite Servicing in Mission Design Studies at the NASA GSFC

    NASA Technical Reports Server (NTRS)

    Leete, Stephen J.

    2003-01-01

    Several NASA missions in various stages of development have undergone one-week studies in the National Aeronautics and Space Administration (NASA) Goddard Space Flight Center (GSFC) Integrated Mission Design Center (IMDC), mostly in preparation for proposals. The possible role of satellite servicing has been investigated for several of these missions, applying the lessons learned from Hubble Space Telescope (HST) servicing, taking into account the current state of the art, projecting into the future, and implementing NASA long-range plans, and is presented here. The general benefits and costs of injecting satellite servicing are detailed, including components such as mission timeline, mass, fuel, spacecraft design, risk abatement, life extension, and improved performance. The approach taken in addressing satellite servicing during IMDC studies is presented.

  2. Flight Path Control Design for the Cassini Solstice Mission

    NASA Technical Reports Server (NTRS)

    Ballard, Christopher G.; Ionasescu, Rodica

    2011-01-01

    The Cassini spacecraft has been in orbit around Saturn for just over 7 years, with a planned 7-year extension, called the Solstice Mission, which started on September 27, 2010. The Solstice Mission includes 205 maneuvers and 70 flybys which consist of the moons Titan, Enceladus, Dione, and Rhea. This mission is designed to use all available propellant with a statistical margin averaging 0.6 m/s per encounter, and the work done to prove and ensure the viability of this margin is highlighted in this paper.

  3. Mission options for rendezvous with the most accessible Near-Earth Asteroid - 1989 ML

    NASA Astrophysics Data System (ADS)

    McAdams, Jim V.

    1992-08-01

    The recent discovery of the Amor-class 1989 ML, the most accessible known asteroid for minimum-energy rendezvous missions, has expedited the search for frequent, low-cost Near-Earth Asteroid rendezvous and round-trip missions. This paper identifies trajectory characteristics and assesses mass performance for low Delta V ballistic rendezvous opportunities to 1989 ML during the period 1996-2010. This asteroid also offers occasional unique extended mission opportunities, such as the lowest known Delta V requirement for any asteroid sample return mission as well as pre-rendezvous asteroid flyby and post-rendezvous comet flyby opportunities requiring less than 5.25 km/sec total Delta V. This paper also briefly comments concerning mission opportunities for asteroid 1991 JW, which recently replaced other known asteroids as the most accessible Near-Earth Asteroid for fast rendezvous and round-trip missions.

  4. Mission options for rendezvous with the most accessible Near-Earth Asteroid - 1989 ML

    NASA Technical Reports Server (NTRS)

    Mcadams, Jim V.

    1992-01-01

    The recent discovery of the Amor-class 1989 ML, the most accessible known asteroid for minimum-energy rendezvous missions, has expedited the search for frequent, low-cost Near-Earth Asteroid rendezvous and round-trip missions. This paper identifies trajectory characteristics and assesses mass performance for low Delta V ballistic rendezvous opportunities to 1989 ML during the period 1996-2010. This asteroid also offers occasional unique extended mission opportunities, such as the lowest known Delta V requirement for any asteroid sample return mission as well as pre-rendezvous asteroid flyby and post-rendezvous comet flyby opportunities requiring less than 5.25 km/sec total Delta V. This paper also briefly comments concerning mission opportunities for asteroid 1991 JW, which recently replaced other known asteroids as the most accessible Near-Earth Asteroid for fast rendezvous and round-trip missions.

  5. Designing planetary protection into the Mars Observer mission.

    PubMed

    Sweetser, T H; Halsell, C A; Cesarone, R J

    1995-03-01

    Planetary protection has been an important consideration during the process of designing the Mars Observer mission. It affected trajectory design of both the interplanetary transfer and the orbits at Mars; these in turn affected the observation strategies developed for the mission. The Project relied mainly on the strategy of collision avoidance to prevent contamination of Mars. Conservative estimates of spacecraft reliability and Martian atmosphere density were used to evaluate decisions concerning the interplanetary trajectory, the orbit insertion phase at Mars, and operations in orbit at Mars and afterwards. Changes in the trajectory design, especially in the orbit insertion phase, required a refinement of those estimates.

  6. 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.

  7. Mask Design for the Space Interferometry Mission Internal Metrology

    NASA Technical Reports Server (NTRS)

    Marx, David; Zhao, Feng; Korechoff, Robert

    2005-01-01

    This slide presentation reviews the mask design used for the internal metrology of the Space Interferometry Mission (SIM). Included is information about the project, the method of measurements with SIM, the internal metrology, numerical model of internal metrology, wavefront examples, performance metrics, and mask design

  8. Preliminary design trade-offs for a multi-mission stored cryogen cooler

    NASA Technical Reports Server (NTRS)

    Sherman, A.

    1978-01-01

    Preliminary design studies were performed for a multi-mission solid cryogen cooler having a wide range of application for both the shuttle sortie and free flyer missions. This multi-mission cooler (MMC) is designed to be utilized with various solid cryogens to meet a wide range of instrument cooling from 10 K (with solid hydrogen) to 90 K. The baseline cooler utilizes two stages of solid cryogen and incorporates an optional, higher temperature third stage which is cooled by either a passive radiator or a thermoelectric cooler. The MMC has an interface which can accommodate a wide variety of instrument configurations. A shrink fit adapter is incorporated which allows a drop-in instrument integration. The baseline design provides cooling of approximately 1 watt over a 60 to 100 K temperature range and about 0.5 watts from 15 to 60 K for a one year lifetime. For low cooling loads and with use of the optional radiator shield, cooling lifetimes as great as 8 years are predicted.

  9. Spacecraft and mission design for the SpaceNuclear PowerSystem Reference Mission

    NASA Technical Reports Server (NTRS)

    Deininger, William D.; Vondra, Robert J.

    1987-01-01

    The design and performance of a spacecraft employing arcjet nuclear electric propulsion, suitable for use in the Space Nuclear Power System Reference Mission, are outlined. The vehicle design is based on a 92 kW ammonia arcjet system operating at an I(sp) of 1050 s and an efficiency of 45 percent. The arcjet/gimbal system, power processing unit, and propellant feed-system are described. A 100 kW(e) space nuclear power system is assumed and the spacecraft mass is baselined at 5250 kg excluding the propellant, propellent feed system, and integrated chemical boost engine. A radiation/arcjet efflux diagnostics package is included in the performance analysis. Three mission scenarios are described and are capable of demonstrating the full capability of the space nuclear power source. The missions considered include power system deployment to possible SDI platform orbits and a spacecraft storage mission to an orbit of three times geosynchronous (GEO) with return to GEO corresponding to Delta V's between 7400 m/s, and 7900 m/s. This spacecraft meets the Reference Mission constraint of low developmental risk and is scaleable to power levels projected for future space platforms.

  10. Asteroid Redirect Crewed Mission Nominal Design and Performance

    NASA Technical Reports Server (NTRS)

    Condon, Gerald; williams, Jacob

    2014-01-01

    Mission (ARCM) nominal design and performance costs associated with an Orion based crewed rendezvous mission to a captured asteroid in an Earth-Moon DRO. The ARM study includes two fundamental mission phases: 1) The Asteroid Redirect Robotic Mission (ARRM) and 2) the ARCM. The ARRM includes a solar electric propulsion based robotic asteroid return vehicle (ARV) sent to rendezvous with a selected near Earth asteroid, capture it, and return it to a DRO in the Earth-Moon vicinity. The DRO is selected over other possible asteroid parking orbits due to its achievability (by both the robotic and crewed vehicles) and by its stability (e.g., no orbit maintenance is required). After the return of the asteroid to the Earth-Moon vicinity, the ARCM is executed and carries a crew of two astronauts to a DRO to rendezvous with the awaiting ARV with the asteroid. The outbound and inbound transfers employ lunar gravity assist (LGA) flybys to reduce the Orion propellant requirement for the overall nominal mission, which provides a nominal mission with some reserve propellant for possible abort situations. The nominal mission described in this report provides a better understanding of the mission considerations as well as the feasibility of such a crewed mission, particularly with regard to spacecraft currently undergoing development, such as the Orion vehicle and the Space Launch System (SLS).

  11. Concept designs for NASA's Solar Electric Propulsion Technology Demonstration Mission

    NASA Technical Reports Server (NTRS)

    Mcguire, Melissa L.; Hack, Kurt J.; Manzella, David H.; Herman, Daniel A.

    2014-01-01

    Multiple Solar Electric Propulsion Technology Demonstration Mission were developed to assess vehicle performance and estimated mission cost. Concepts ranged from a 10,000 kilogram spacecraft capable of delivering 4000 kilogram of payload to one of the Earth Moon Lagrange points in support of future human-crewed outposts to a 180 kilogram spacecraft capable of performing an asteroid rendezvous mission after launched to a geostationary transfer orbit as a secondary payload. Low-cost and maximum Delta-V capability variants of a spacecraft concept based on utilizing a secondary payload adapter as the primary bus structure were developed as were concepts designed to be co-manifested with another spacecraft on a single launch vehicle. Each of the Solar Electric Propulsion Technology Demonstration Mission concepts developed included an estimated spacecraft cost. These data suggest estimated spacecraft costs of $200 million - $300 million if 30 kilowatt-class solar arrays and the corresponding electric propulsion system currently under development are used as the basis for sizing the mission concept regardless of launch vehicle costs. The most affordable mission concept developed based on subscale variants of the advanced solar arrays and electric propulsion technology currently under development by the NASA Space Technology Mission Directorate has an estimated cost of $50M and could provide a Delta-V capability comparable to much larger spacecraft concepts.

  12. Space station needs, attributes and architectural options study. Volume 3: Mission requirements

    NASA Technical Reports Server (NTRS)

    1983-01-01

    User missions that are enabled or enhanced by a manned space station are identified. The mission capability requirements imposed on the space station by these users are delineated. The accommodation facilities, equipment, and functional requirements necessary to achieve these capabilities are identified, and the economic, performance, and social benefits which accrue from the space station are defined.

  13. Space station needs, attributes and architectural options study. Volume 2: Mission analysis

    NASA Technical Reports Server (NTRS)

    1983-01-01

    Space environment studies, astrophysics, Earth environment, life sciences, and material sciences are discussed. Commercial communication, materials processing, and Earth observation missions are addressed. Technology development, space operations, scenarios of operational capability, mission requirements, and benefits analysis results for space-produced gallium arsenide crystals, direct broadcasting satellite systems, and a high inclination space station are covered.

  14. Space station needs, attributes, and architectural options study. Volume 1: Missions and requirements

    NASA Technical Reports Server (NTRS)

    1983-01-01

    Science and applications, NOAA environmental observation, commercial resource observations, commercial space processing, commercial communications, national security, technology development, and GEO servicing are addressed. Approach to time phasing of mission requirements, system sizing summary, time-phased user mission payload support, space station facility requirements, and integrated time-phased system requirements are also addressed.

  15. The HYDROS Mission: Requirements and Baseline System Design

    NASA Technical Reports Server (NTRS)

    Njoku, Eni; Spencer, Michael; McDonald, Kyle; Smith, Joel; Houser, Paul; Doiron, Terence; O'Neill, Peggy; Girard, Ralph; Entekhabi, Dara

    2003-01-01

    The HYDROS mission is under development by NASA as part of its Earth System Science Pathfinder (ESSP) program. HYDROS is designed to provide global maps of the Earth's soil moisture and freezel/thaw state every 2-3 days, for weather and climate prediction, water and carbon cycle studies, natural hazards monitoring, and national security applications. HYDROS uses a unique active and passive L-band microwave system that optimizes measurement accuracy, spatial resolution, and coverage. It provides measurements in nearly all weather conditions, regardless of solar illumination. The designs of the radar and radiometer electronics, antenna feedhorn and reflector, and science data system, are driven by specific mission and science objectives. These objectives impose requirements on the frequencies, polarizations, sampling, spatial resolution, and accuracy of the system. In this paper we describe the HYDROS mission requirements, baseline design, and measurement capabilities.

  16. Design of A Lander For The Bepicolombo Mission

    NASA Astrophysics Data System (ADS)

    Pansart, O.; Anselmi, A.

    The BepiColombo ESA mission will deploy two orbiters (Mercury Planetary Orbiter and Mercury Magnetospheric Orbiter, the latter provided by ISAS, Japan) and the Mercury Surface Element (MSE), the first lander on the surface of Mercury. The main purpose of MSE is to investigate the physical and chemical properties of a spot on the Mercury surface. Alcatel Space Industries has been designing the MSE and its associated propulsion module in the frame of a Definition Study led by Alenia Spazio, under ESA contract. The MSE mission starts at separation from the orbital composite spacecraft. During the descent and landing phase, a high-thrust bipropellant engine decelerates the MSE, and an airbag system is deployed to ensure a safe landing after a short free-fall phase. The MSE then starts its scientific mission on the Mercury surface. The paper addresses the following topics: Mission overview - Descent and Landing phase - MSE surface mission phase - System preliminary design and performances. The presented MSE baseline was studied as part of the industrial Definition Study. The mission itself is currently going through a redefinition phase which affects the whole ESA Scientific Program, and which may lead to a revised version by the May-June 2002 time frame.

  17. The OSIRIS-Rex Asteroid Sample Return: Mission Operations Design

    NASA Technical Reports Server (NTRS)

    Gal-Edd, Jonathan; Cheuvront, Allan

    2014-01-01

    The OSIRIS-REx mission employs a methodical, phased approach to ensure success in meeting the missions science requirements. OSIRIS-REx launches in September 2016, with a backup launch period occurring one year later. Sampling occurs in 2019. The departure burn from Bennu occurs in March 2021. On September 24, 2023, the SRC lands at the Utah Test and Training Range (UTTR). Stardust heritage procedures are followed to transport the SRC to Johnson Space Center, where the samples are removed and delivered to the OSIRIS-REx curation facility. After a six-month preliminary examination period the mission will produce a catalog of the returned sample, allowing the worldwide community to request samples for detailed analysis.Traveling and returning a sample from an Asteroid that has not been explored before requires unique operations consideration. The Design Reference Mission (DRM) ties together space craft, instrument and operations scenarios. The project implemented lessons learned from other small body missions: APLNEAR, JPLDAWN and ESARosetta. The key lesson learned was expected the unexpected and implement planning tools early in the lifecycle. In preparation to PDR, the project changed the asteroid arrival date, to arrive one year earlier and provided additional time margin. STK is used for Mission Design and STKScheduler for instrument coverage analysis.

  18. Jovian tour design for orbiter and lander missions to Europa

    NASA Astrophysics Data System (ADS)

    Campagnola, Stefano; Buffington, Brent B.; Petropoulos, Anastassios E.

    2014-07-01

    Europa is one of the most interesting targets for solar system exploration, as its ocean of liquid water could harbor life. Following the recommendation of the Planetary Decadal Survey, NASA commissioned a study for a multiple flyby only mission, an orbiter mission, and a lander mission. This paper presents the moon tours for the lander and orbiter concepts. The total Δv and radiation dose would be reduced when compared to previous designs by exploiting multi-body dynamics and avoiding multi-revolution transfers in the Ganymede-to-Europa transfer. Tours 11-O3, 12-L1 and 12-L4 and their performances compared to other tours from previous Europa mission studies are presented in detail.

  19. Lunar Orbit Stability for Small Satellite Mission Design

    NASA Technical Reports Server (NTRS)

    Dono, Andres

    2015-01-01

    The irregular nature of the lunar gravity field will severely affect the orbit lifetime and behavior of future lunar small satellite missions. These spacecraft need stable orbits that do not require large deltaV budgets for station-keeping maneuvers. The initial classical elements of any lunar orbit are critical to address its stability and to comply with mission requirements. This publication identifies stable regions according to different initial conditions at the time of lunar orbit insertion (LOI). High fidelity numerical simulations with two different gravity models were performed. We focus in low altitude orbits where the dominant force in orbit propagation is the existence of unevenly distributed lunar mass concentrations. These orbits follow a periodic oscillation in some of the classical elements that is particularly useful for mission design. A set of orbital maintenance strategies for various mission concepts is presented.

  20. Mission design for the infrared astronomical satellite /IRAS/

    NASA Technical Reports Server (NTRS)

    Lundy, S. A.; Mclaughlin, W. I.; Pouw, A.

    1979-01-01

    IRAS, a joint United States, Netherlands, United Kingdom astronomical satellite, is scheduled to be launched early in 1981 with the purpose of completing an all-sky survey in the infrared wavelengths from 8 to 120 microns and to observe objects of special interest. The mission design is driven by thermal constraints primarily determined by the Sun and Earth; the orbit and survey strategy must be chosen so as to satisfy the mission requirements before the cryogenic system is depleted of its liquid helium. Computer graphics help the designer choose valid survey strategies and evaluate resulting sky coverage.

  1. 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.

  2. Science operations planning expertise: A neglected component of mission design

    NASA Astrophysics Data System (ADS)

    Chaizy, P. A.; Dimbylow, T. G.; Allan, P. M.; Hapgood, M. A.

    2011-09-01

    In this paper, Science Operations Planning Expertise (SOPE) is defined as the expertise that is held by people who have the two following qualities. First they have both theoretical and practical experience in operations planning, in general, and in space science operations planning in particular. Second, they can be used, on request and at least, to provide with advice the teams that design and implement science operations systems in order to optimise the performance and productivity of the mission. However, the relevance and use of such SOPE early on during the Mission Design Phase (MDP) is not sufficiently recognised. As a result, science operations planning is often neglected or poorly assessed during the mission definition phases. This can result in mission architectures that are not optimum in terms of cost and scientific returns, particularly for missions that require a significant amount of science operations planning. Consequently, science operations planning difficulties and cost underestimations are often realised only when it is too late to design and implement the most appropriate solutions. In addition, higher costs can potentially reduce both the number of new missions and the chances of existing ones to be extended. Moreover, the quality, and subsequently efficiency, of SOPE can vary greatly. This is why we also believe that the best possible type of SOPE requires a structure similar to the ones of existing bodies of expertise dedicated to the data processing such as the International Planetary Data Alliance (IPDA), the Space Physics Archive Search and Extract (SPASE) or the Planetary Data System (PDS). Indeed, this is the only way of efficiently identifying science operations planning issues and their solutions as well as of keeping track of them in order to apply them to new missions. Therefore, this paper advocates for the need to allocate resources in order to both optimise the use of SOPE early on during the MDP and to perform, at least, a

  3. Space station data system analysis/architecture study. Task 2: Options development, DR-5. Volume 2: Design options

    NASA Technical Reports Server (NTRS)

    1985-01-01

    The primary objective of Task 2 is the development of an information base that will support the conduct of trade studies and provide sufficient data to make key design/programmatic decisions. This includes: (1) the establishment of option categories that are most likely to influence Space Station Data System (SSDS) definition; (2) the identification of preferred options in each category; and (3) the characterization of these options with respect to performance attributes, constraints, cost and risk. This volume contains the options development for the design category. This category comprises alternative structures, configurations and techniques that can be used to develop designs that are responsive to the SSDS requirements. The specific areas discussed are software, including data base management and distributed operating systems; system architecture, including fault tolerance and system growth/automation/autonomy and system interfaces; time management; and system security/privacy. Also discussed are space communications and local area networking.

  4. Lean Mission Operations Systems Design - Using Agile and Lean Development Principles for Mission Operations Design and Development

    NASA Technical Reports Server (NTRS)

    Trimble, Jay Phillip

    2014-01-01

    The Resource Prospector Mission seeks to rove the lunar surface with an in-situ resource utilization payload in search of volatiles at a polar region. The mission operations system (MOS) will need to perform the short-duration mission while taking advantage of the near real time control that the short one-way light time to the Moon provides. To maximize our use of limited resources for the design and development of the MOS we are utilizing agile and lean methods derived from our previous experience with applying these methods to software. By using methods such as "say it then sim it" we will spend less time in meetings and more time focused on the one outcome that counts - the effective utilization of our assets on the Moon to meet mission objectives.

  5. Critical early mission design considerations for lunar data systems architecture

    NASA Technical Reports Server (NTRS)

    Hei, Donald J., Jr.; Stephens, Elaine

    1992-01-01

    This paper outlines recent early mission design activites for a lunar data systems architecture. Each major functional element is shown to be strikingly similar when viewed in a common reference system. While this similarity probably deviates with lower levels of decomposition, the sub-functions can always be arranged into similar and dissimilar categories. Similar functions can be implemented as objects - implemented once and reused several times like today's advanced integrated circuits. This approach to mission data systems, applied to other NASA programs, may result in substantial agency implementation and maintenance savings. In today's zero-sum-game budgetary environment, this approach could help to enable a lunar exploration program in the next decade. Several early mission studies leading to such an object-oriented data systems design are recommended.

  6. Space station needs, attributes and architectural options study. Volume 7-2: Data book. Commercial missions

    NASA Technical Reports Server (NTRS)

    1983-01-01

    The history of NASA's materials processing in space activities is reviewed. Market projections, support requirements, orbital operations issues, cost estimates and candidate systems (orbiter sortie flight, orbiter serviced free flyer, space station, space station serviced free flyer) for the space production of semiconductor crystals are examined. Mission requirements are identified for materials processing, communications missions, bioprocessing, and for transferring aviation maintenance training technology to spacecraft.

  7. Design and Analysis of RTGs for CRAF and Cassini Missions

    SciTech Connect

    Schock, Alfred; Noravian, Heros; Sankarankandath

    1990-11-30

    This report consists of two parts. Part 1 describes the development of novel analytical methods needed to predict the BOM performance and the subsequent performance degradation of the mutually obstructed RTGs for the CRAF and Cassini missions. Part II applies those methods to the two missions, presents the resultant predictions, and discusses their programmatic implications. The results indicate that JPL's original power demand goals could have been met with two standard GPHS RTGs for each mission. But subsequently JPL significantly increased both the power level and the mission duration for both missions, so that they can no longer by met by two standard RTGs. The resultant power gap must be closed either by reducing JPL's power demand (e.g., by decreasing contingency reserves) and/or by increasing the power system's output. One way under active consideration which more than meets the system power goal would be the addition of a third RTG for each mission. However, the author concluded that it may be possible to meet or closely approach the CRAF power demand goals with just two RTGs by relatively modest modification of their design and/or operating conditions. To explore that possibility, the effect of various modifications - either singly or in combination - was analyzed by Fairchild. The results indicate that modest modifications can meet or come very close to meeting the CRAF power goals with just two RTGs. Elimination of the third RTG would yield substantial cost and schedule savings. There are three copies in the file.

  8. HSCT mission analysis of waverider designs

    NASA Technical Reports Server (NTRS)

    1992-01-01

    In the second quarter the development of the two waverider design tools was continued, and the groundwork necessary for the incorporation of waverider technology into the realm of the High Speed Civil Transports (HSCT's) was laid out. Advances in each of these areas is summarized. Work on the WIPAR code included the addition of an upper surface geometry generator and characteristic flow solver and the inclusion of viscous analysis in the performance computations. Details of these changes are given. In the course of the second project quarter, much of the analysis performed during the first quarter was incorporated into a working computer code. To date, utilities were developed for the definition of arbitrary 3-D shock surfaces, the computation of post-shock flow conditions, and the marching of the solution in a roughly cross-stream direction away from the shock surface. These utilities are briefly described. During the second quarter groundwork for the analysis of complete configurations was initiated. This involved the development of computational utilities for the integration of powerplants with the waverider forebodies, and the acquisition of a number of configuration analysis software packages. Work in these areas is discussed.

  9. NASA'S Space Launch System: Opening Opportunities for Mission Design

    NASA Technical Reports Server (NTRS)

    Robinson, Kimberly F.; Hefner, Keith; Hitt, David

    2015-01-01

    Designed to meet the stringent requirements of human exploration missions into deep space and to Mars, NASA's Space Launch System (SLS) vehicle represents a unique new launch capability opening new opportunities for mission design. While SLS's super-heavy launch vehicle predecessor, the Saturn V, was used for only two types of missions - launching Apollo spacecraft to the moon and lofting the Skylab space station into Earth orbit - NASA is working to identify new ways to use SLS to enable new missions or mission profiles. In its initial Block 1 configuration, capable of launching 70 metric tons (t) to low Earth orbit (LEO), SLS is capable of not only propelling the Orion crew vehicle into cislunar space, but also delivering small satellites to deep space destinations. With a 5-meter (m) fairing consistent with contemporary Evolved Expendable Launch Vehicles (EELVs), the Block 1 configuration can also deliver science payloads to high-characteristic-energy (C3) trajectories to the outer solar system. With the addition of an upper stage, the Block 1B configuration of SLS will be able to deliver 105 t to LEO and enable more ambitious human missions into the proving ground of space. This configuration offers opportunities for launching co-manifested payloads with the Orion crew vehicle, and a new class of secondary payloads, larger than today's cubesats. The evolved configurations of SLS, including both Block 1B and the 130 t Block 2, also offer the capability to carry 8.4- or 10-m payload fairings, larger than any contemporary launch vehicle. With unmatched mass-lift capability, payload volume, and C3, SLS not only enables spacecraft or mission designs currently impossible with contemporary EELVs, it also offers enhancing benefits, such as reduced risk and operational costs associated with shorter transit time to destination and reduced risk and complexity associated with launching large systems either monolithically or in fewer components. As this paper will

  10. Asteroid Deflection Mission Design Considering On-Ground Risks

    NASA Astrophysics Data System (ADS)

    Rumpf, Clemens; Lewis, Hugh G.; Atkinson, Peter

    The deflection of an Earth-threatening asteroid requires high transparency of the mission design process. The goal of such a mission is to move the projected point of impact over the face of Earth until the asteroid is on a miss trajectory. During the course of deflection operations, the projected point of impact will match regions that were less affected before alteration of the asteroid’s trajectory. These regions are at risk of sustaining considerable damage if the deflecting spacecraft becomes non-operational. The projected impact point would remain where the deflection mission put it at the time of mission failure. Hence, all regions that are potentially affected by the deflection campaign need to be informed about this risk and should be involved in the mission design process. A mission design compromise will have to be found that is acceptable to all affected parties (Schweickart, 2004). A software tool that assesses the on-ground risk due to deflection missions is under development. It will allow to study the accumulated on-ground risk along the path of the projected impact point. The tool will help determine a deflection mission design that minimizes the on-ground casualty and damage risk due to deflection operations. Currently, the tool is capable of simulating asteroid trajectories through the solar system and considers gravitational forces between solar system bodies. A virtual asteroid may be placed at an arbitrary point in the simulation for analysis and manipulation. Furthermore, the tool determines the asteroid’s point of impact and provides an estimate of the population at risk. Validation has been conducted against the solar system ephemeris catalogue HORIZONS by NASA’s Jet Propulsion Laboratory (JPL). Asteroids that are propagated over a period of 15 years show typical position discrepancies of 0.05 Earth radii relative to HORIZONS’ output. Ultimately, results from this research will aid in the identification of requirements for

  11. Designing an Alternate Mission Operations Control Room

    NASA Technical Reports Server (NTRS)

    Montgomery, Patty; Reeves, A. Scott

    2014-01-01

    The Huntsville Operations Support Center (HOSC) is a multi-project facility that is responsible for 24x7 real-time International Space Station (ISS) payload operations management, integration, and control and has the capability to support small satellite projects and will provide real-time support for SLS launches. The HOSC is a service-oriented/ highly available operations center for ISS payloads-directly supporting science teams across the world responsible for the payloads. The HOSC is required to endure an annual 2-day power outage event for facility preventive maintenance and safety inspection of the core electro-mechanical systems. While complete system shut-downs are against the grain of a highly available sub-system, the entire facility must be powered down for a weekend for environmental and safety purposes. The consequence of this ground system outage is far reaching: any science performed on ISS during this outage weekend is lost. Engineering efforts were focused to maximize the ISS investment by engineering a suitable solution capable of continuing HOSC services while supporting safety requirements. The HOSC Power Outage Contingency (HPOC) System is a physically diversified compliment of systems capable of providing identified real-time services for the duration of a planned power outage condition from an alternate control room. HPOC was designed to maintain ISS payload operations for approximately three continuous days during planned HOSC power outages and support a local Payload Operations Team, International Partners, as well as remote users from the alternate control room located in another building.

  12. System design from mission definition to flight validation

    NASA Technical Reports Server (NTRS)

    Batill, S. M.

    1992-01-01

    Considerations related to the engineering systems design process and an approach taken to introduce undergraduate students to that process are presented. The paper includes details on a particular capstone design course. This course is a team oriented aircraft design project which requires the students to participate in many phases of the system design process, from mission definition to validation of their design through flight testing. To accomplish this in a single course requires special types of flight vehicles. Relatively small-scale, remotely piloted vehicles have provided the class of aircraft considered in this course.

  13. The HYDROS mission: requirements and baseline system design

    NASA Technical Reports Server (NTRS)

    Njoku, Eni; Spencer, Michael; McDonald, Kyle; Smith, Joel; Houser, Paul; Doiron, Terence; ONeill, Peggy; Girard, Ralph; Entekhabi, Dara

    2004-01-01

    The HYDROS mission is under development by NASA as part of its Earth System Science Pathfinder program. HYDROS is designed to provide global maps of the Earth's soil moisture and freeze/thaw state every 2-3 days, for weather and climate prediction, water and carbon cycle studies, natural hazards monitoring, and national security applications.

  14. Mission Control Technologies: A New Way of Designing and Evolving Mission Systems

    NASA Technical Reports Server (NTRS)

    Trimble, Jay; Walton, Joan; Saddler, Harry

    2006-01-01

    Current mission operations systems are built as a collection of monolithic software applications. Each application serves the needs of a specific user base associated with a discipline or functional role. Built to accomplish specific tasks, each application embodies specialized functional knowledge and has its own data storage, data models, programmatic interfaces, user interfaces, and customized business logic. In effect, each application creates its own walled-off environment. While individual applications are sometimes reused across multiple missions, it is expensive and time consuming to maintain these systems, and both costly and risky to upgrade them in the light of new requirements or modify them for new purposes. It is even more expensive to achieve new integrated activities across a set of monolithic applications. These problems impact the lifecycle cost (especially design, development, testing, training, maintenance, and integration) of each new mission operations system. They also inhibit system innovation and evolution. This in turn hinders NASA's ability to adopt new operations paradigms, including increasingly automated space systems, such as autonomous rovers, autonomous onboard crew systems, and integrated control of human and robotic missions. Hence, in order to achieve NASA's vision affordably and reliably, we need to consider and mature new ways to build mission control systems that overcome the problems inherent in systems of monolithic applications. The keys to the solution are modularity and interoperability. Modularity will increase extensibility (evolution), reusability, and maintainability. Interoperability will enable composition of larger systems out of smaller parts, and enable the construction of new integrated activities that tie together, at a deep level, the capabilities of many of the components. Modularity and interoperability together contribute to flexibility. The Mission Control Technologies (MCT) Project, a collaboration of

  15. 77 FR 34121 - Designation and Determination Pursuant to the Foreign Missions Act Concerning the Designation of...

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-06-08

    ... From the Federal Register Online via the Government Publishing Office DEPARTMENT OF STATE Designation and Determination Pursuant to the Foreign Missions Act Concerning the Designation of Entities in the United States That Are Substantially Owned or Effectively Controlled by the Government of Azerbaijan as Foreign Missions and the...

  16. Next-Generation Space Telescope design reference mission

    NASA Astrophysics Data System (ADS)

    Smith, Eric P.; Mather, John C.; Stockman, Hervey S.; Bely, Pierre Y.; Stiavelli, Massimo; Burg, Richard

    1998-08-01

    The Next Generation Space Telescope (NGST) Design Reference Mission (DRM) represents a suite of potential astronomical programs and targets along with their expected physical properties, and desired observation modes. This broad science program is being used to drive the observatory design in a way as fundamental as traditional engineering parameters. Astronomers use the DRM to communicate their desires in a quantitative fashion to the engineers who will eventually construct the observatory. The DRM is also the primary tool used to measure the relative value of NGST mission architectures and technological readiness of the program. Specifically, the fraction of the DRM completed by a given observatory configuration in a given time is, to first order, a measure of the value of the design. Those designs which complete a higher fraction of the observations listed below are more capable than those complete lesser fractions.

  17. The Messenger Spacecraft Power System Design and Early Mission Performance

    NASA Astrophysics Data System (ADS)

    Dakermanji, G.; Person, C.; Jenkins, J.; Kennedy, L.; Temkin, D.

    2005-05-01

    The MESSENGER (MErcury Surface, Space ENvironment, GEochemistry, and Ranging) spacecraft was launched on August 3, 2004. The spacecraft will be inserted into Mercury orbit in March 2011 for one year of orbital operation. During the mission, the spacecraft distance to the Sun will vary between approximately 1 and 0.3 Astronomical Units (AU), imposing severe requirements on the spacecraft thermal and power systems design. The spacecraft is maintained behind a sunshade. The two single-axis, gimbaled solar array panels are designed to withstand the expected high temperatures. A peak power tracking system has been selected to allow operation over the widely varying solar array I-V curves. In order to reduce cost and risk while increasing the likelihood of mission success, the approach taken in the power system design, including the solar arrays, was to use conventional design, materials, and fabrication techniques.

  18. TARDIS: An Automation Framework for JPL Mission Design and Navigation

    NASA Technical Reports Server (NTRS)

    Roundhill, Ian M.; Kelly, Richard M.

    2014-01-01

    Mission Design and Navigation at the Jet Propulsion Laboratory has implemented an automation framework tool to assist in orbit determination and maneuver design analysis. This paper describes the lessons learned from previous automation tools and how they have been implemented in this tool. In addition this tool has revealed challenges in software implementation, testing, and user education. This paper describes some of these challenges and invites others to share their experiences.

  19. Moon Age and Regolith Explorer (MARE) Mission Design and Performance

    NASA Technical Reports Server (NTRS)

    Condon, Gerald L.; Lee, David E.; Carson, John M., III

    2017-01-01

    On December 11, 1972, Apollo 17 marked the last controlled U.S. lunar landing and was followed by an absence of methodical in-situ investigation of the lunar surface. The Moon Age and Regolith Explorer (MARE) proposal provides scientific measurement of the age and composition of a relatively young portion of the lunar surface near Aristarchus Plateau and the first post-Apollo U.S. soft lunar landing. It includes the first demonstration of a crew survivability-enhancing autonomous hazard detection and avoidance system. This report focuses on the mission design and performance associated with the MARE robotic lunar landing subject to mission and trajectory constraints.

  20. Mission Design and Simulation Considerations for ADReS-A

    NASA Astrophysics Data System (ADS)

    Peters, S.; Förstner, R.; Fiedler, H.

    2016-09-01

    Space debris in general has become a major problem for modern space activities. Guidelines to mitigate the threat have been recommended, better prediction models are developed and an advanced observation of objects orbiting Earth is in progress. And still - without the implementation of active debris removal (ADR), the number of debris in space will exponentially increase. To support the ongoing research on ADR-missions, this paper presents the updated mission design of ADReS-A (Autonomous Debris Removal Satellite - #A) - one possible concept for the multiple active removal of large debris in Low Earth orbit, in this case especially of rocket bodies of the SL-8-type. ADReS-A as chaser satellite is supported by at least 5 de-orbit kits, allowing for the same number of targets to be removed. While ADReS-A is conceived for handling of the target, the kit's task is the controlled re-entry of the designated rocket body. The presented mission design forms the basis for the simulation environment in progress. The simulation shall serve as testbed to test multiple scenarios in terms of approach and abort optimization or different tumbling modes of the target. The ultimate goal is the test of autonomous behaviors of the spacecraft in case of unforeseen failures during the approach phase. Considerations to create a simulation for the described mission are presented and discussed. A first visualization of pre-calculated aboard trajectories can be found at the end of this paper.

  1. Lunar missions using chemical propulsion: System design issues

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan

    1991-01-01

    To transport lunar base elements to the Moon, large high-energy propulsion systems will be required. Advanced propulsion systems for lunar missions can significantly reduce launch mass and increase the delivered payload, resulting in significant launch cost savings. In this report, the masses in low Earth orbit (LEO) are compared for several propulsion systems: nitrogen tetroxide/monomethyl hydrazine (NTO/MMH), oxygen/methane (O2/CH4), oxygen/hydrogen (O2/H2), and metallized O2/H2/Al propellants. Also addressed are payload mass increases enabled with these systems; system design issues involving the engine thrust levels, engine commonality between the transfer vehicle and the excursion vehicle; the number of launches to place the lunar mission vehicles into LEO; and analyses of small lunar missions launched from a single Space Transportation System-Cargo (STS-C) flight.

  2. Designing astrophysics missions for NASA's Space Launch System

    NASA Astrophysics Data System (ADS)

    Stahl, H. Philip; Hopkins, Randall C.; Schnell, Andrew; Smith, David Alan; Jackman, Angela; Warfield, Keith R.

    2016-10-01

    Large space telescope missions have always been limited by their launch vehicle's mass and volume capacities. The Hubble Space Telescope was specifically designed to fit inside the Space Shuttle and the James Webb Space Telescope was specifically designed to fit inside an Ariane 5. Astrophysicists desire even larger space telescopes. NASA's "Enduring Quests Daring Visions" report calls for an 8- to 16-m Large UV-Optical-IR (LUVOIR) Surveyor mission to enable ultrahigh-contrast spectroscopy and coronagraphy. Association of Universities for Research in Astronomy's "From Cosmic Birth to Living Earth" report calls for a 12-m class High-Definition Space Telescope to pursue transformational scientific discoveries. NASA's "Planning for the 2020 Decadal Survey" calls for a Habitable Exoplanet Imaging (HabEx) and an LUVOIR as well as Far-IR and an X-ray Surveyor missions. Packaging larger space telescopes into existing launch vehicles is a significant engineering complexity challenge that drives cost and risk. NASA's planned Space Launch System (SLS), with its 8- or 10-m diameter fairings and ability to deliver 35 to 45 mt of payload to Sun-Earth-Lagrange-2, mitigates this challenge by fundamentally changing the design paradigm for large space telescopes. This paper introduces the mass and volume capacities of the planned SLS, provides a simple mass allocation recipe for designing large space telescope missions to this capacity, and gives three specific mission concept implementation examples: a 4-m monolithic off-axis telescope, an 8-m monolithic on-axis telescope, and a 12-m segmented on-axis telescope.

  3. Radioisotope Power Systems Reference Book for Mission Designers and Planners

    NASA Technical Reports Server (NTRS)

    Lee, Young; Bairstow, Brian

    2015-01-01

    The RPS Program's Program Planning and Assessment (PPA) Office commissioned the Mission Analysis team to develop the Radioisotope Power Systems (RPS) Reference Book for Mission Planners and Designers to define a baseline of RPS technology capabilities with specific emphasis on performance parameters and technology readiness. The main objective of this book is to provide RPS technology information that could be utilized by future mission concept studies and concurrent engineering practices. A progress summary from the major branches of RPS technology research provides mission analysis teams with a vital tool for assessing the RPS trade space, and provides concurrent engineering centers with a consistent set of guidelines for RPS performance characteristics. This book will be iterated when substantial new information becomes available to ensure continued relevance, serving as one of the cornerstone products of the RPS PPA Office. This book updates the original 2011 internal document, using data from the relevant publicly released RPS technology references and consultations with RPS technologists. Each performance parameter and RPS product subsection has been reviewed and cleared by at least one subject matter representative. A virtual workshop was held to reach consensus on the scope and contents of the book, and the definitions and assumptions that should be used. The subject matter experts then reviewed and updated the appropriate sections of the book. The RPS Mission Analysis Team then performed further updates and crosschecked the book for consistency. Finally, a second virtual workshop was held to ensure all subject matter experts and stakeholders concurred on the contents.

  4. Improved OSC Amtec generator design to meet goals of JPL's candidate Europa Orbiter mission

    SciTech Connect

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

    1998-07-01

    The preceding paper (Paper IECEC.98.244) described OSC's initial designs of AMTEC (Alkali Metal Thermal-to-Electrical Conversion) power systems, consisting of one or two generators, each with 2, 3, or 4 General Purpose Heat Source (GPHS) modules and with 16 refractory AMTEC cells containing 5 Beta Alumina Solid Electrolyte (BASE) tubes; and presented the effect of heat input and voltage output on the generator's BOM evaporator and clad temperatures and on its EOM system efficiency and power output. Comparison of the computed results with JPL's goals for the Europa Orbiter mission showed that all of the initial 16-cell design options yielded either excessive evaporator and clad temperatures or insufficient EOM power to satisfy the JPL-specified mission goals. The present paper describes modified OSC generator designs with different numbers of AMTEC cells, cell diameters, cell lengths, cell materials, BASE tube lengths, and number of tubes per cell. These efforts succeeded in identifying generator designs with only half the number of AMTEC cells which -- for the same assumptions -- can produce EOM power outputs substantially in excess of JPL's goals for NASA's Europa Orbiter mission while operating well below the prescribed BOM limits on evaporator and clad temperature; and revealed that lowering the emissivity of the generator's housing to raise the cells' condenser temperatures can achieve substantial additional performance improvement. Finally, the paper culminates in programmatic recommendations.

  5. Solid breeder blanket option for the ITER conceptual design

    SciTech Connect

    Gohar, Y.; Attaya, H.; Billone, M.C.; Finn, P.; Majumdar, S.; Turner, L.R.; Baker, C.C.; Nelson, B.E.; Raffray, R.; Oak Ridge National Lab., TN; California Univ., Los Angeles, CA )

    1989-10-01

    A solid-breeder water-cooled blanket option was developed for ITER based on a multilayer configuration. The blanket uses beryllium for neutron multiplication and lithium oxide for tritium breeding. The material forms are sintered products for both material with 0.8 density factor. The lithium-6 enrichment is 90%. This blanket has the capability to accommodate a factor of two change in the neutron wall loading without violating the different design guidelines. The design philosophy adopted for the blanket is to produce the necessary tritium required for the ITER operation and to operate at power reactor conditions as much as possible. At the same time, the reliability and the safety aspects of the blanket are enhanced by the use of a low-pressure coolant and the separation of the tritium purge lines from the coolant system. The blanket modules are made by hot vacuum forming and diffusion bonding a double wall structure with integral cooling channels. The different aspects of the blanket design including tritium breeding, nuclear heat deposition, activation analyses, thermal-hydraulics, tritium inventory, structural analyses, and water coolant conditions are summarized in this paper. 12 refs., 2 figs., 1 tab.

  6. An Engineering Design Reference Mission for a Future Large-Aperture UVOIR Space Observatory

    NASA Astrophysics Data System (ADS)

    Thronson, Harley A.; Bolcar, Matthew R.; Clampin, Mark; Crooke, Julie A.; Redding, David; Rioux, Norman; Stahl, H. Philip

    2016-01-01

    From the 2010 NRC Decadal Survey and the NASA Thirty-Year Roadmap, Enduring Quests, Daring Visions, to the recent AURA report, From Cosmic Birth to Living Earths, multiple community assessments have recommended development of a large-aperture UVOIR space observatory capable of achieving a broad range of compelling scientific goals. Of these priority science goals, the most technically challenging is the search for spectroscopic biomarkers in the atmospheres of exoplanets in the solar neighborhood. Here we present an engineering design reference mission (EDRM) for the Advanced Technology Large-Aperture Space Telescope (ATLAST), which was conceived from the start as capable of breakthrough science paired with an emphasis on cost control and cost effectiveness. An EDRM allows the engineering design trade space to be explored in depth to determine what are the most demanding requirements and where there are opportunities for margin against requirements. Our joint NASA GSFC/JPL/MSFC/STScI study team has used community-provided science goals to derive mission needs, requirements, and candidate mission architectures for a future large-aperture, non-cryogenic UVOIR space observatory. The ATLAST observatory is designed to operate at a Sun-Earth L2 orbit, which provides a stable thermal environment and excellent field of regard. Our reference designs have emphasized a serviceable 36-segment 9.2 m aperture telescope that stows within a five-meter diameter launch vehicle fairing. As part of our cost-management effort, this particular reference mission builds upon the engineering design for JWST. Moreover, it is scalable to a variety of launch vehicle fairings. Performance needs developed under the study are traceable to a variety of additional reference designs, including options for a monolithic primary mirror.

  7. A mission design for International Manned Mars Mission - From the 1991 International Space University (ISU) Design Project

    NASA Technical Reports Server (NTRS)

    Mendell, Wendell W.

    1991-01-01

    The International Space University (ISU) conducted a study of an international program to support human exploration of Mars as its annual Design Project activity during its 1991 summer session in Toulouse, France. Although an ISU Design Project strives to produce an in-depth analysis during the intense 10-week summer session, the International Mars Mission (IMM) project was conducted in a manner designed to provide a learning experience for young professionals working in an unusual multidisciplinary and multinational environment. The breadth of the IMM study exceeds that of most Mars mission studies of the past, encompassing political organization for long-term commitment, multinational management structure, cost analysis, mission architecture, vehicle configuration, crew health, life support, Mars surface infrastructure, mission operations, technology evaluation, risk assessment, scientific planning, exploration, communication networks, and Martian resource utilization. The IMM Final Report has particular value for those seeking insight into the choices made by a multinational group working in an apolitical environment on the problems of international cooperation in space.

  8. Automated design of multiphase space missions using hybrid optimal control

    NASA Astrophysics Data System (ADS)

    Chilan, Christian Miguel

    A modern space mission is assembled from multiple phases or events such as impulsive maneuvers, coast arcs, thrust arcs and planetary flybys. Traditionally, a mission planner would resort to intuition and experience to develop a sequence of events for the multiphase mission and to find the space trajectory that minimizes propellant use by solving the associated continuous optimal control problem. This strategy, however, will most likely yield a sub-optimal solution, as the problem is sophisticated for several reasons. For example, the number of events in the optimal mission structure is not known a priori and the system equations of motion change depending on what event is current. In this work a framework for the automated design of multiphase space missions is presented using hybrid optimal control (HOC). The method developed uses two nested loops: an outer-loop that handles the discrete dynamics and finds the optimal mission structure in terms of the categorical variables, and an inner-loop that performs the optimization of the corresponding continuous-time dynamical system and obtains the required control history. Genetic algorithms (GA) and direct transcription with nonlinear programming (NLP) are introduced as methods of solution for the outer-loop and inner-loop problems, respectively. Automation of the inner-loop, continuous optimal control problem solver, required two new technologies. The first is a method for the automated construction of the NLP problems resulting from the use of a direct solver for systems with different structures, including different numbers of categorical events. The method assembles modules, consisting of parameters and constraints appropriate to each event, sequentially according to the given mission structure. The other new technology is for a robust initial guess generator required by the inner-loop NLP problem solver. Two new methods were developed for cases including low-thrust trajectories. The first method, based on GA

  9. Mission Designs for Demonstrating Gravity Tractor Asteroid Deflection

    NASA Astrophysics Data System (ADS)

    Busch, M.; Faber, N.; Eggl, S.; Morrison, D.; Clark, A.; Frost, C.; Jaroux, B. A.; Khetawat, V.

    2015-12-01

    Gravity tractor asteroid deflection relies on the gravitational attraction between the target and a nearby spacecraft; using low-thrust propulsion to change the target's trajectory slowly but continuously. Our team, based at the NASA Ames Mission Design Center, prepared designs for a Gravity Tractor Demonstration Mission (GTDM) for the European Commission's NEOShield initiative. We found five asteroids with well-known orbits and opportunities for efficient stand-alone demonstrations in the 2020s. We selected one object, 2000 FJ10, for a detailed design analysis. Our GTDM design has a 4 kW solar-electric propulsion system and launch mass of 1150 kg. For a nominal asteroid mass of 3 x 109 kg and diameter 150 m, and a hovering altitude 125 m above the asteroid's surface, GTDM would change FJ10's semi-major axis by 10 km over 2 years. To measure the deflection clearly and to permit safe hovering by the spacecraft, several months of survey and characterization are required prior to the active tractoring phase of the mission. Accurate tracking is also required after the tractoring phase, to ensure that the asteroid has indeed been deflected as intended. The GTDM design includes both spacecraft and Earth-based observations of FJ10 to verify the deflection. The estimated cost of GTDM is $280 million. Trajectory analysis for GTDM confirmed that the outcome of a deflection of any asteroid depends on when that deflection is performed. Compared to kinetic impactor deflection, the gradual deflection from a gravity tractor produces comparable results for a given total momentum transfer. However, a gravity tractor can have greater flexibility in the direction in which the target asteroid can be deflected. Asteroid deflection scenarios must be modeled carefully on a case-to-case basis. We will review implications of the results of the GTDM study to other proposed gravity tractor demonstrations, such as that included in NASA's Asteroid Redirect Mission.

  10. FLIP II - Concept Designs to Meet Future Scientific Mission Requirements

    NASA Astrophysics Data System (ADS)

    Laible, D. H.

    2002-12-01

    R/P FLIP has successfully operated for 40 years in support of important oceanographic research missions. The simple platform, which has the unique ability to provide a heave-stable operating location in open ocean environments, has over time been modified and upgraded. Its capability has been extended to the physical limits imposed by buoyancy and stability constraints. Nonetheless, there are oceanographic research operations that can use FLIP's unique characteristics, but which exceed its capabilities. Over the years researchers at the Marine Physical Laboratory of Scripps Institution of Oceanography have led investigations into second generation heave-stable ocean platforms with capabilities substantially exceeding those of R/P FLIP. This paper discusses several design concepts that have been developed. The designs are presented in terms of the ability to meet current and future scientific mission requirements.

  11. KEPLER MISSION DESIGN, REALIZED PHOTOMETRIC PERFORMANCE, AND EARLY SCIENCE

    SciTech Connect

    Koch, David G.; Borucki, William J.; Lissauer, Jack J.; Basri, Gibor; Marcy, Geoffrey; Batalha, Natalie M.; Brown, Timothy M.; Caldwell, Douglas; DeVore, Edna; Jenkins, Jon; Christensen-Dalsgaard, Joergen; Cochran, William D.; Dunham, Edward W.; Gautier, Thomas N.; Gilliland, Ronald L.; Gould, Alan; Kondo, Yoji; Monet, David

    2010-04-20

    The Kepler Mission, launched on 2009 March 6, was designed with the explicit capability to detect Earth-size planets in the habitable zone of solar-like stars using the transit photometry method. Results from just 43 days of data along with ground-based follow-up observations have identified five new transiting planets with measurements of their masses, radii, and orbital periods. Many aspects of stellar astrophysics also benefit from the unique, precise, extended, and nearly continuous data set for a large number and variety of stars. Early results for classical variables and eclipsing stars show great promise. To fully understand the methodology, processes, and eventually the results from the mission, we present the underlying rationale that ultimately led to the flight and ground system designs used to achieve the exquisite photometric performance. As an example of the initial photometric results, we present variability measurements that can be used to distinguish dwarf stars from red giants.

  12. OSIRIS-REx Touch-and-Go (TAG) Mission Design for Asteroid Sample Collection

    NASA Technical Reports Server (NTRS)

    May, Alexander; Sutter, Brian; Linn, Timothy; Bierhaus, Beau; Berry, Kevin; Mink, Ron

    2014-01-01

    The Origins Spectral Interpretation Resource Identification Security Regolith Explorer (OSIRIS-REx) mission is a NASA New Frontiers mission launching in September 2016 to rendezvous with the near-Earth asteroid Bennu in October 2018. After several months of proximity operations to characterize the asteroid, OSIRIS-REx flies a Touch-And-Go (TAG) trajectory to the asteroid's surface to collect at least 60 g of pristine regolith sample for Earth return. This paper provides mission and flight system overviews, with more details on the TAG mission design and key events that occur to safely and successfully collect the sample. An overview of the navigation performed relative to a chosen sample site, along with the maneuvers to reach the desired site is described. Safety monitoring during descent is performed with onboard sensors providing an option to abort, troubleshoot, and try again if necessary. Sample collection occurs using a collection device at the end of an articulating robotic arm during a brief five second contact period, while a constant force spring mechanism in the arm assists to rebound the spacecraft away from the surface. Finally, the sample is measured quantitatively utilizing the law of conservation of angular momentum, along with qualitative data from imagery of the sampling device. Upon sample mass verification, the arm places the sample into the Stardust-heritage Sample Return Capsule (SRC) for return to Earth in September 2023.

  13. Navigation Design and Analysis for the Orion Cislunar Exploration Missions

    NASA Technical Reports Server (NTRS)

    D'Souza, Christopher; Holt, Greg; Gay, Robert; Zanetti, Renato

    2014-01-01

    This paper details the design and analysis of the cislunar optical navigation system being proposed for the Orion Earth-Moon (EM) missions. In particular, it presents the mathematics of the navigation filter. It also presents the sensitivity analysis that has been performed to understand the performance of the proposed system, with particular attention paid to entry flight path angle constraints and the DELTA V performance

  14. Selection and trajectory design to mission secondary targets

    NASA Astrophysics Data System (ADS)

    Victorino Sarli, Bruno; Kawakatsu, Yasuhiro

    2017-02-01

    Recently, with new trajectory design techniques and use of low-thrust propulsion systems, missions have become more efficient and cheaper with respect to propellant. As a way to increase the mission's value and scientific return, secondary targets close to the main trajectory are often added with a small change in the transfer trajectory. As a result of their large number, importance and facility to perform a flyby, asteroids are commonly used as such targets. This work uses the Primer Vector theory to define the direction and magnitude of the thrust for a minimum fuel consumption problem. The design of a low-thrust trajectory with a midcourse asteroid flyby is not only challenging for the low-thrust problem solution, but also with respect to the selection of a target and its flyby point. Currently more than 700,000 minor bodies have been identified, which generates a very large number of possible flyby points. This work uses a combination of reachability, reference orbit, and linear theory to select appropriate candidates, drastically reducing the simulation time, to be later included in the main trajectory and optimized. Two test cases are presented using the aforementioned selection process and optimization to add and design a secondary flyby to a mission with the primary objective of 3200 Phaethon flyby and 25143 Itokawa rendezvous.

  15. Lunar missions using advanced chemical propulsion: System design issues

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan

    1994-01-01

    To provide the transportation of lunar base elements to the moon, large high-energy propulsion systems will be required. Advanced propulsion systems for lunar missions can provide significant launch mass reductions and payload increases. These mass reductions and added payload masses can be translated into significant launch cost savings for the lunar base missions. The masses in low Earth orbit (LEO) were compared for several propulsion systems: nitrogen tetroxide/monomethyl hydrazine (NTO/MMH), oxygen/methane (O2/CH4), oxygen/hydrogen (O2/H2), and metallized O2/H2/Al propellants. Also, the payload mass increases enabled with O2/H2 and O2/H2/Al systems were addressed. In addition, many system design issues involving the engine thrust levels, engine commonality between the transfer vehicle and the excursion vehicle, and the number of launches to place the lunar mission vehicles into LEO will be discussed. Analyses of small lunar missions launched from a single STS-C flight are also presented.

  16. Design Considerations for Spacecraft Operations During Uncrewed Dormant Phases of Human Exploration Missions

    NASA Technical Reports Server (NTRS)

    Williams-Byrd, Julie; Antol, Jeff; Jefferies, Sharon; Goodliff, Kandyce; Williams, Phillip; Ambrose, Rob; Sylvester, Andre; Anderson, Molly; Dinsmore, Craig; Hoffman, Stephen; Lawrence, James; Seibert, Marc; Schier, Jim; Frank, Jeremy; Alexander, Leslie; Ruff, Gary; Soeder, Jim; Guinn, Joseph; Stafford, Matthew

    2016-01-01

    NASA is transforming human spaceflight. The Agency is shifting from an exploration-based program with human activities in low Earth orbit (LEO) and targeted robotic missions in deep space to a more sustainable and integrated pioneering approach. However, pioneering space involves daunting technical challenges of transportation, maintaining health, and enabling crew productivity for long durations in remote, hostile, and alien environments. Subject matter experts from NASA's Human Exploration and Operations Mission Directorate (HEOMD) are currently studying a human exploration campaign that involves deployment of assets for planetary exploration. This study, called the Evolvable Mars Campaign (EMC) study, explores options with solar electric propulsion as a central component of the transportation architecture. This particular in-space transportation option often results in long duration transit to destinations. The EMC study is also investigating deployed human rated systems like landers, habitats, rovers, power systems and ISRU system to the surface of Mars, which also will involve long dormant periods when these systems are staged on the surface. In order to enable the EMC architecture, campaign and element design leads along with system and capability development experts from HEOMD's System Maturation Team (SMT) have identified additional capabilities, systems and operation modes that will sustain these systems especially during these dormant phases of the mission. Dormancy is defined by the absence of crew and relative inactivity of the systems. For EMC missions, dormant periods could range from several months to several years. Two aspects of uncrewed dormant operations are considered herein: (1) the vehicle systems that are placed in a dormant state and (2) the autonomous vehicle systems and robotic capabilities that monitor, maintain, and repair the vehicle and systems. This paper describes the mission stages of dormancy operations, phases of dormant

  17. Interplanetary mission design handbook. Volume 1, part 4: Earth to Saturn ballistic mission opportunities, 1985-2005

    NASA Technical Reports Server (NTRS)

    Sergeyevsky, A. B.; Snyder, G. C.

    1981-01-01

    Graphical data necessary for the preliminary design of ballistic missions to Saturn are provided. Contours of launch energy requirements as well as many other launch and Saturn arrival parameters, are presented in launch date/arrival date space for all launch opportunities from 1985 through 2005. In addition, an extensive text is included which explains mission design methods, from launch window development to Saturn probe and orbiter arrival design, utilizing the graphical data in this volume as well as numerous equations elating various parameters. This is the first of a planned series of mission design documents which will apply to all planets and some other bodies in the solar system.

  18. Advanced Single-Aisle Transport Propulsion Design Options Revisited

    NASA Technical Reports Server (NTRS)

    Guynn, Mark D.; Berton, Jeffrey J.; Tong, Michael T.; Haller, William J.

    2013-01-01

    Future propulsion options for advanced single-aisle transports have been investigated in a number of previous studies by the authors. These studies have examined the system level characteristics of aircraft incorporating ultra-high bypass ratio (UHB) turbofans (direct drive and geared) and open rotor engines. During the course of these prior studies, a number of potential refinements and enhancements to the analysis methodology and assumptions were identified. This paper revisits a previously conducted UHB turbofan fan pressure ratio trade study using updated analysis methodology and assumptions. The changes incorporated have decreased the optimum fan pressure ratio for minimum fuel consumption and reduced the engine design trade-offs between minimizing noise and minimizing fuel consumption. Nacelle drag and engine weight are found to be key drivers in determining the optimum fan pressure ratio from a fuel efficiency perspective. The revised noise analysis results in the study aircraft being 2 to 4 EPNdB (cumulative) quieter due to a variety of reasons explained in the paper. With equal core technology assumed, the geared engine architecture is found to be as good as or better than the direct drive architecture for most parameters investigated. However, the engine ultimately selected for a future advanced single-aisle aircraft will depend on factors beyond those considered here.

  19. An Integrated Approach for Entry Mission Design and Flight Simulations

    NASA Technical Reports Server (NTRS)

    Lu, Ping; Rao, Prabhakara

    2004-01-01

    An integrated approach for entry trajectory design, guidance, and simulation is proposed. The key ingredients for this approach are an on-line 3 degree-of-freedom entry trajectory planning algorithm and the entry guidance algorithm that generates the guidance gains automatically. When fully developed, such a tool could enable end-bend entry mission design and simulations in 3DOF and 6DOF mode from de-orbit burn to the TAEM interface and beyond, all in one key stroke. Some preliminary examples of such a capability are presented in this paper that demonstrate the potential of this type of integrated environment.

  20. Evaluation of Low-Thrust Propulsion Options for Cargo Missions to Near-Earth Objects

    NASA Astrophysics Data System (ADS)

    Spaun, Micah; Sankaran, Kamesh

    2011-10-01

    A simple method developed to optimize low-thrust trajectories to near-Earth objects is presented. A computational tool developed using this method was tested for a round-trip cargo mission carrying a payload of 1000-2000 kg to the asteroid 1996XB27. Several existing electric thrusters (a Bi Hall thruster at two levels of operation, a Xe Hall thruster, an applied field Li Lorentz force accelerator, the HiPEP ion thruster and the VASIMR) were considered for this mission, at their demonstrated values of performance parameters, to examine the validity of this new computational tool. In the range of power levels considered (150 to 600 kW), increasing the power level had only a small effect on reducing the trip time. It was found that a Bi Hall thruster operating at a high thrust-to-power ratio allowed for the shortest trip time, and a Bi Hall thruster and the HiPEP ion thruster with very high values of exhaust velocity and efficiency required the least initial mass to transport the payload. It was found that this new tool yielded results that were realistic, and insights from the results for thruster development are discussed.

  1. Abort Options for Human Missions to Earth-Moon Halo Orbits

    NASA Technical Reports Server (NTRS)

    Jesick, Mark C.

    2013-01-01

    Abort trajectories are optimized for human halo orbit missions about the translunar libration point (L2), with an emphasis on the use of free return trajectories. Optimal transfers from outbound free returns to L2 halo orbits are numerically optimized in the four-body ephemeris model. Circumlunar free returns are used for direct transfers, and cislunar free returns are used in combination with lunar gravity assists to reduce propulsive requirements. Trends in orbit insertion cost and flight time are documented across the southern L2 halo family as a function of halo orbit position and free return flight time. It is determined that the maximum amplitude southern halo incurs the lowest orbit insertion cost for direct transfers but the maximum cost for lunar gravity assist transfers. The minimum amplitude halo is the most expensive destination for direct transfers but the least expensive for lunar gravity assist transfers. The on-orbit abort costs for three halos are computed as a function of abort time and return time. Finally, an architecture analysis is performed to determine launch and on-orbit vehicle requirements for halo orbit missions.

  2. 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

  3. DICE Mission Design, Development, and Implementation: Success and Challenges

    NASA Astrophysics Data System (ADS)

    Stromberg, E.; Swenson, C.; Fish, C. S.; Crowley, G.; Barjatya, A.; Petersen, J.

    2012-12-01

    Funded by the NSF CubeSat and NASA ELaNa programs, the Dynamic Ionosphere CubeSat Experiment (DICE) mission consists of two 1.5U CubeSats which were launched into an eccentric low Earth orbit on October 28, 2011. Each identical spacecraft carries two Langmuir probes to measure ionospheric in-situ plasma densities, electric field probes to measure in-situ DC and AC electric fields, and a magnetometer to measure in-situ DC and AC magnetic fields. Given the tight integration of these multiple sensors with the CubeSat platforms, each of the DICE spacecraft is effectively a "sensor-sat" capable of comprehensive ionospheric diagnostics. Over time, the sensor-sats will separate relative to each other due to differences in the ejection velocity and enable accurate identification of geospace storm-time features, such as the geomagnetic Storm Enhanced Density (SED) bulge and plume. The use of two identical sensor-sats permits the de-convolution of spatial and temporal ambiguities in the observations of the ionosphere from a moving platform. In addition to demonstrating nanosat constellation science, the DICE mission downlink communications system is operating at 3 Mbit/s. To our knowledge, this transmission rate is a factor of 100 or more greater than previous CubeSat missions to date. This paper will focus on the DICE mission design, implementation, and on-orbit operations successes as well as the challenges faced in implementing a high-return science mission with limited resources. Specifically, it will focus on the lessons learned in integrating, calibrating, and managing a small constellation of sensor-sats for global science measurements.

  4. Design of multihundredwatt DIPS for robotic space missions

    NASA Technical Reports Server (NTRS)

    Bents, D. J.; Geng, S. M.; Schreiber, J. G.; Withrow, C. A.; Schmitz, P. C.; Mccomas, Thomas J.

    1991-01-01

    Design of a dynamic isotope power system (DIPS) general purpose heat source (GPHS) and small free piston Stirling engine (FPSE) is being pursued as a potential lower cost alternative to radioisotope thermoelectric generators (RTG's). The design is targeted at the power needs of future unmanned deep space and planetary surface exploration missions ranging from scientific probes to SEI precursor missions. These are multihundredwatt missions. The incentive for any dynamic system is that it can save fuel which reduces cost and radiological hazard. However, unlike a conventional DIPS based on turbomachinery converions, the small Stirling DIPS can be advantageously scaled to multihundred watt unit size while preserving size and weight competitiveness with RTG's. Stirling conversion extends the range where dynamic systems are competitive to hundreds of watts (a power range not previously considered for dynamic systems). The challenge of course is to demonstrate reliability similar to RTG experience. Since the competative potential of FPSE as an isotope converter was first identified, work has focused on the feasibility of directly integrating GPHS with the Stirling heater head. Extensive thermal modeling of various radiatively coupled heat source/heater head geometries were performed using data furnished by the developers of FPSE and GPHS. The analysis indicates that, for the 1050 K heater head configurations considered, GPHS fuel clad temperatures remain within safe operating limits under all conditions including shutdown of one engine. Based on these results, preliminary characterizations of multihundred watt units were established.

  5. Mars Aerocapture Studies for the Design Reference Mission

    NASA Technical Reports Server (NTRS)

    Lyne, James Evans; Wercinski, Paul; Walberg, Gerald; Jits, Roman

    1997-01-01

    The recent discovery of possible fossilized microbes in a Martian meteorite sample and the spectacular success of the Mars Pathfinder mission have substantially increased public interest and support for future robotic and manned exploration of Mars. NASA is currently refining a plan known as the Design Reference Mission (DRM) in which the first human landing would occur in 2014 after a series of cargo launches which would place surface systems and an Earth return vehicle at Mars two years prior to the crew's arrival. At each subsequent launch opportunity (which occur approximately every twenty-six months), an additional Earth return vehicle, surface facility and crew would depart for Mars, with each crew employing the systems launched during the previous opportunity. The mission design calls for a long-duration surface stay, rapid crew transits, in-situ manufacture of the Mars ascent propellant, nuclear thermal propulsion for the trans-Mars injection burn, and the use of aerocapture for both the cargo and crew vehicles at Mars.

  6. Handling and Emplacement Options for Deep Borehole Disposal Conceptual Design.

    SciTech Connect

    Cochran, John R.; Hardin, Ernest

    2015-07-01

    This report presents conceptual design information for a system to handle and emplace packages containing radioactive waste, in boreholes 16,400 ft deep or possibly deeper. Its intended use is for a design selection study that compares the costs and risks associated with two emplacement methods: drill-string and wireline emplacement. The deep borehole disposal (DBD) concept calls for siting a borehole (or array of boreholes) that penetrate crystalline basement rock to a depth below surface of about 16,400 ft (5 km). Waste packages would be emplaced in the lower 6,560 ft (2 km) of the borehole, with sealing of appropriate portions of the upper 9,840 ft (3 km). A deep borehole field test (DBFT) is planned to test and refine the DBD concept. The DBFT is a scientific and engineering experiment, conducted at full-scale, in-situ, without radioactive waste. Waste handling operations are conceptualized to begin with the onsite receipt of a purpose-built Type B shipping cask, that contains a waste package. Emplacement operations begin when the cask is upended over the borehole, locked to a receiving flange or collar. The scope of emplacement includes activities to lower waste packages to total depth, and to retrieve them back to the surface when necessary for any reason. This report describes three concepts for the handling and emplacement of the waste packages: 1) a concept proposed by Woodward-Clyde Consultants in 1983; 2) an updated version of the 1983 concept developed for the DBFT; and 3) a new concept in which individual waste packages would be lowered to depth using a wireline. The systems described here could be adapted to different waste forms, but for design of waste packaging, handling, and emplacement systems the reference waste forms are DOE-owned high- level waste including Cs/Sr capsules and bulk granular HLW from fuel processing. Handling and Emplacement Options for Deep Borehole Disposal Conceptual Design July 23, 2015 iv ACKNOWLEDGEMENTS This report has

  7. Trajectory Options for a Potential Mars Mission Combining Orbiting Science, Relay and a Sample Return Rendezvous Demonstration

    NASA Technical Reports Server (NTRS)

    Guinn, Joseph R.; Kerridge, Stuart J.; Wilson, Roby S.

    2012-01-01

    Mars sample return is a major scientific goal of the 2011 US National Research Council Decadal Survey for Planetary Science. Toward achievement of this goal, recent architecture studies have focused on several mission concept options for the 2018/2020 Mars launch opportunities. Mars orbiters play multiple roles in these architectures such as: relay, landing site identification/selection/certification, collection of on-going or new measurements to fill knowledge gaps, and in-orbit collection and transportation of samples from Mars to Earth. This paper reviews orbiter concepts that combine these roles and describes a novel family of relay orbits optimized for surface operations support. Additionally, these roles provide an intersection of objectives for long term NASA science, human exploration, technology development and international collaboration.

  8. Moon Age and Regolith Explorer (MARE) Mission Design and Performance

    NASA Technical Reports Server (NTRS)

    Condon, Gerald L.; Lee, David E.

    2016-01-01

    The moon’s surface last saw a controlled landing from a U.S. spacecraft on December 11, 1972 with Apollo 17. Since that time, there has been an absence of methodical in-situ investigation of the lunar surface. In addition to the scientific value of measuring the age and composition of a relatively young portion of the lunar surface near Aristarchus Plateau, the Moon Age and Regolith Explorer (MARE) proposal provides the first U.S. soft lunar landing since the Apollo Program and the first ever robotic soft lunar landing employing an autonomous hazard detection and avoidance system, a system that promises to enhance crew safety and survivability during a manned lunar (or other) landing. This report focuses on the mission design and performance associated with the MARE robotic lunar landing subject to mission and trajectory constraints.

  9. Invariant Manifolds, Lagrangian Trajectories and Space Mission Design

    NASA Astrophysics Data System (ADS)

    Belló, Miguel; Gómez, Gerard; Masdemont, Josep J.

    The last 30 years have produced an explosion in the capabilities of designing and managing libration point missions. The starting point was the ground-breaking mission of the third International Sun-Earth Explorer spacecraft (ISEE-3). The ISEE-3 was launched August 12, 1978 to pursue studies of the Earth-Sun interactions, in a first step of what now is known as Space Weather. After a direct transfer of the ISEE-3 to the vicinity of the Sun-Earth Lagrange point, it was inserted into a nearly-periodic halo orbit, in order to monitor the solar wind about 1 h before it reached the Earth's magneto-sphere as well as the ISEE-1 and 2 spacecraft (which where in an elliptical orbit around the Earth).

  10. Interplanetary mission design handbook. Volume 1, part 2: Earth to Mars ballistic mission opportunities, 1990-2005

    NASA Technical Reports Server (NTRS)

    Sergeyevsky, A. B.; Snyder, G. C.; Cunniff, R. A.

    1983-01-01

    Graphical data necessary for the preliminary design of ballistic missions to Mars are provided. Contours of launch energy requirements, as well as many other launch and Mars arrival parameters, are presented in launch date/arrival date space for all launch opportunities from 1990 through 2005. In addition, an extensive text is included which explains mission design methods, from launch window development to Mars probe and orbiter arrival design, utilizing the graphical data as well as numerous equations relating various parameters.

  11. Interplanetary mission design handbook. Volume 1, part 1: Earth to Venus ballistic mission opportunities, 1991-2005

    NASA Technical Reports Server (NTRS)

    Sergeyevsky, A. B.; Yin, N. H.

    1983-01-01

    Graphical data necessary for the preliminary design of ballistic missions to Venus is presented. Contours of launch energy requirements, as well as many other launch and arrival parameters, are presented in launch data/arrival date space for all launch opportunities from 1991 through 2005. An extensive text is included which explains mission design methods, from launch window development to Venus probe and orbiter arrival design, utilizing the graphical data in this volume as well as numerous equations relating various parameters.

  12. Space Station needs, attributes and architectural options. Volume 2, book 2, part 1: Mission implementation concepts

    NASA Technical Reports Server (NTRS)

    1983-01-01

    The overall configuration and modules of the initial and evolved space station are described as well as tended industrial and polar platforms. The mass properties that are the basis for costing are summarized. User friendly attributes (interfaces, resources, and facilities) are identified for commercial; science and applications; industrial park; international participation; national security; and the external tank option. Configuration alternates studied to determine a baseline are examined. Commonality for clustered 3-man and 9-man stations are considered as well as the use of tethered platforms. Requirements are indicated for electrical, communication and tracking; data management Subsystem requirements for electrical, data management, communication and tracking, environment control/life support system; and guidance navigation and control subsystems are identified.

  13. Design of Spacecraft Missions to Remove Multiple Orbital Debris Objects

    NASA Technical Reports Server (NTRS)

    Barbee, Brent W.; Alfano, Salvatore; Pinon, Elfego; Gold, Kenn; Gaylor, David

    2012-01-01

    The amount of hazardous debris in Earth orbit has been increasing, posing an evergreater danger to space assets and human missions. In January of 2007, a Chinese ASAT test produced approximately 2600 pieces of orbital debris. In February of 2009, Iridium 33 collided with an inactive Russian satellite, yielding approximately 1300 pieces of debris. These recent disastrous events and the sheer size of the Earth orbiting population make clear the necessity of removing orbital debris. In fact, experts from both NASA and ESA have stated that 10 to 20 pieces of orbital debris need to be removed per year to stabilize the orbital debris environment. However, no spacecraft trajectories have yet been designed for removing multiple debris objects and the size of the debris population makes the design of such trajectories a daunting task. Designing an efficient spacecraft trajectory to rendezvous with each of a large number of orbital debris pieces is akin to the famous Traveling Salesman problem, an NP-complete combinatorial optimization problem in which a number of cities are to be visited in turn. The goal is to choose the order in which the cities are visited so as to minimize the total path distance traveled. In the case of orbital debris, the pieces of debris to be visited must be selected and ordered such that spacecraft propellant consumption is minimized or at least kept low enough to be feasible. Emergent Space Technologies, Inc. has developed specialized algorithms for designing efficient tour missions for near-Earth asteroids that may be applied to the design of efficient spacecraft missions capable of visiting large numbers of orbital debris pieces. The first step is to identify a list of high priority debris targets using the Analytical Graphics, Inc. SOCRATES website and then obtain their state information from Celestrak. The tour trajectory design algorithms will then be used to determine the itinerary of objects and v requirements. These results will shed light

  14. Interplanetary mission design handbook. Volume 1, Part 5: Mars-to-Earth ballistic mission opportunities, 1992-2007

    NASA Technical Reports Server (NTRS)

    Sergeyevsky, Andrey; Cunniff, Ross

    1987-01-01

    This document contains graphical data necessary for the preliminary design of ballistic missions returning from Mars. Contours of Mars-departure energy requirements, as well as many other launch and Earth-arrival parameters are presented in arrival-date/launch-date space for all departure opportunities from 1992 through 2007. In addition, an extensive companion document (Part 2) is available; it contains Earth-Mars graphical data and explains mission design methods, using the graphical data as well as numerous equations relating various parameters. This is one of a planned series of mission design handbooks.

  15. SP-100 planetary mission/system preliminary design study. Final report, technical information report

    SciTech Connect

    Jones, R.M.

    1986-02-01

    This report contains a discussion on many aspects of a nuclear electric propulsion planetary science mission and spacecraft using the proposed SP-100 nuclear power subsystem. A review of the science rationale for such missions is included. A summary of eleven nuclear electric propulsion planetary missions is presented. A conceptual science payload, mission design, and spacecraft design is included for the Saturn Ring Rendezvous mission. Spacecraft and mission costs have been estimated for two potential sequences of nuclear electric propulsion planetary missions. The integration issues and requirements on the proposed SP-100 power subsystems are identified.

  16. Design options for automotive batteries in advanced car electrical systems

    NASA Astrophysics Data System (ADS)

    Peters, K.

    The need to reduce fuel consumption, minimize emissions, and improve levels of safety, comfort and reliability is expected to result in a much higher demand for electric power in cars within the next 5 years. Forecasts vary, but a fourfold increase in starting power to 20 kW is possible, particularly if automatic stop/start features are adopted to significantly reduce fuel consumption and exhaust emissions. Increases in the low-rate energy demand are also forecast, but the use of larger alternators may avoid unacceptable high battery weights. It is also suggested from operational models that the battery will be cycled more deeply. In examining possible designs, the beneficial features of valve-regulated lead-acid batteries made with compressed absorbent separators are apparent. Several of their attributes are considered. They offer higher specific power, improved cycling capability and greater vibration resistance, as well as more flexibility in packaging and installation. Optional circuits considered for dual-voltage supplies are separate batteries for engine starting (36 V) and low-power duties (12 V), and a universal battery (36 V) coupled to a d.c.-d.c. converter for a 12-V equipment. Battery designs, which can be made on commercially available equipment with similar manufacturing costs (per W h and per W) to current products, are discussed. The 36-V battery, made with 0.7 mm thick plates, in the dual-battery system weighs 18.5 kg and has a cold-cranking amp (CCA) rating of 790 A at -18°C to 21.6 V (1080 W kg -1 at a mean voltage of 25.4 V). The associated, cycleable 12-V battery, provides 1.5 kW h and weighs 24.6 kg. Thus, the combined battery weight is 43.1 kg. The single universal battery, with cycling capability, weighs 45.4 kg, has a CCA rating of 810 A (441 W kg -1 at a mean voltage of 24.7 V), and when connected to the d.c.-d.c. converter at 75% efficiency provides a low-power capacity of 1.5 kW h.

  17. Exo-C: Mission and Science Payload Design

    NASA Astrophysics Data System (ADS)

    Dekens, Frank G.; Stapelfeldt, Karl R.; Warfield, Keith; Unwin, Stephen C.; Exo-C Science; Technology Definition Team, Exo-C JPL Study Design Team

    2015-01-01

    We present NASA's Exoplanet Coronagraph (Exo-C) mission design and science payload completed as part of a probe-class concept study under consideration for launch following JWST. The payload consists of an unobscured Cassegrain telescope with a 1.4-m clear aperture, a barrel assembly, and an internal coronagraph instrument. The mission has a 3 year lifetime and is in a highly stable Earth-trailing orbit. The coronagraph instrument is mounted laterally on the anti-Sun side of the telescope, obviating the need for high incidence reflections and better isolating it from spacecraft disturbances. The instrument has both an Imaging Camera and an Integral Field Spectrograph (IFS). The former obtains filter imaging with 1e-9 raw contrast from 2 - 20 λ/D in radius, while the IFS delivers the same contrast with spectral resolution of R = 70 from 450 to 1000 nm, but with a reduced outer working angle.The Exo-C science performance requirements are achieved with a specialized observatory design enabled by several new technologies. The telescope is designed for precision pointing and high stability to maintain a slowly evolving speckle pattern. Vibration isolation is achieved with two stages between the reaction wheels and the science payload. The solar arrays and high gain antenna are body-fixed, and a stiff barrel assembly is used as the telescope metering structure. Telescope pointing is updated at a high rate by monitoring the bright science target star with a low order wavefront sensor and driving a fine steering mirror for compensation. Active thermal control is used to minimize thermal drifts of the telescope, instrument, and barrel assemblies. Stability analyses via modeling of the structural, thermal, and optical performance of this configuration show that the proposed mission configuration would enable unprecedented exoplanet and circumstellar disk science with direct imaging.

  18. Design and application of electromechanical actuators for deep space missions

    NASA Technical Reports Server (NTRS)

    Haskew, Tim A.; Wander, John

    1994-01-01

    This progress report documents research and development efforts performed from August 16, 1993 through August 15, 1994 on NASA Grant NAG8-240, 'Design and Application of Electromechanical Actuators for Deep Space Missions.' Since the submission of our last progress report in February 1994, our efforts have been almost entirely focused on final construction of the test stand and experiment design. Hence, this report is dedicated solely to these topics. However, updates on our research personnel and our health monitoring and fault management efforts are provided in this summary. Following this executive summary are two report sections. The first is devoted to the motor drive being constructed for the test stand. The thrust of the next section is the mechanical and hydraulic design and construction based on the planned experimental requirements. Following both major sections are three appendices.

  19. Design of an afocal telescope for the ARIEL mission

    NASA Astrophysics Data System (ADS)

    Da Deppo, Vania; Middleton, Kevin; Focardi, Mauro; Morgante, Gianluca; Pace, Emanuele; Claudi, Riccardo; Micela, Giuseppina

    2016-07-01

    ARIEL (Atmospheric Remote-sensing Infrared Exoplanet Large-survey) is one of the three candidates for the next ESA medium-class science mission (M4) expected to be launched in 2026. This mission will be devoted to observe spectroscopically in the infrared (IR) a large population of known transiting planets in our Galaxy. ARIEL is based on a 1-m class telescope ahead of two spectrometer channels covering the band 1.95 to 7.8 microns. In addition there are four photometric channels: two wide band, also used as fine guidance sensors, and two narrow band. During its 3.5 years operations from L2 orbit, ARIEL will continuously observe exoplanets transiting their host star. The ARIEL design is conceived as a fore-module common afocal telescope that will feed the spectrometer and photometric channels. The telescope optical design is an off-axis portion of a two-mirror classic telescope coupled to a tertiary off-axis paraboloidal mirror providing a collimating output beam. The telescope and optical bench operating temperatures, as well as those of some subsystems, will be monitored and fine tuned/stabilised mainly by means of a thermal control subsystem (TCU - Telescope Control Unit) working in closed-loop feedback and hosted by the main Payload electronics unit, i.e. the Instrument Control Unit (ICU). In this paper the telescope requirements will be given together with the foreseen design. The technical solution chosen to passively cool the telescope unit will be detailed discussed.

  20. 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.

  1. Maneuver Design for the Juno Mission: Inner Cruise

    NASA Technical Reports Server (NTRS)

    Pavlak, Thomas A.; Frauenholz, Raymond B.; Bordi, John J.; Kangas, Julie A.; Helfrich, Clifford E.

    2014-01-01

    The Juno spacecraft launched in August 2011 and, following a successful Earth flyby in October 2013, is on course for a nominal orbit insertion at Jupiter in July 2016. This paper examines the design and execution of deterministic and statistical trajectory correction maneuvers during the first approximately 27 months of post-launch operations that defined the "Inner Cruise" phase of the Juno mission. Topics of emphasis include the two deep space maneuvers, Earth flyby altitude biasing strategy, and the sequence of trajectory correction maneuvers executed in the weeks prior to the successful Earth gravity assist.

  2. Design and application of electromechanical actuators for deep space missions

    NASA Technical Reports Server (NTRS)

    Haskew, Tim A.; Wander, John

    1995-01-01

    This third semi-annual progress report covers the reporting period from August 16, 1994 through February 15, 1995 on NASA Grant NAG8-240, 'Design and Application of Electromechanical Actuators for Deep Space Missions'. There are two major report sections: Motor Control Status/Electrical Experiment Planning and Experiment Planning and Initial Results. The primary emphasis of our efforts during the reporting period has been final construction and testing of the laboratory facilities. As a result, this report is dedicated to that topic.

  3. Design of a Flush Airdata System (FADS) for the Hypersonic Air Launched Option (HALO) Vehicle

    NASA Technical Reports Server (NTRS)

    Whitmore, Stephen A.; Moes, Timothy R.; Deets, Dwain A. (Technical Monitor)

    1994-01-01

    This paper presents a design study for a pressure based Flush airdata system (FADS) on the Hypersonic Air Launched Option (HALO) Vehicle. The analysis will demonstrate the feasibility of using a pressure based airdata system for the HALO and provide measurement uncertainty estimates along a candidate trajectory. The HALO is a conceived as a man-rated vehicle to be air launched from an SR-71 platform and is proposed as a testbed for an airbreathing hydrogen scramjet. A feasibility study has been performed and indicates that the proposed trajectory is possible with minimal modifications to the existing SR71 vehicle. The mission consists of launching the HALO off the top of an SR-71 at Mach 3 and 80,000 ft. A rocket motor is then used to accelerate the vehicle to the test condition. After the scramjet test is completed the vehicle will glide to a lakebed runway landing. This option provides reusability of the vehicle and scramjet engine. The HALO design will also allow for various scramjet engine and flowpath designs to be flight tested. For the HALO flights, measurements of freestream airdata are considered to be a mission critical to perform gain scheduling and trajectory optimization. One approach taken to obtaining airdata involves measurement of certain parameters such as external atmospheric winds, temperature, etc to estimate the airdata quantities. This study takes an alternate approach. Here the feasibility of obtaining airdata using a pressure-based flush airdata system (FADS) methods is assessed. The analysis, although it is performed using the HALO configuration and trajectory, is generally applicable to other hypersonic vehicles. The method to be presented offers the distinct advantage of inferring total pressure, Mach number, and flow incidence angles, without stagnating the freestream flow. This approach allows for airdata measurements to be made using blunt surfaces and significantly diminishes the heating load at the sensor. In the FADS concept a

  4. Mission Design Overview for Mars 2003/2005 Sample Return Mission

    NASA Technical Reports Server (NTRS)

    Lee, Wayne J.; DAmario, Louis A.; Roncoli, Ralph B.; Smith, John C.

    2000-01-01

    In May 2003, a new and exciting chapter in Mars exploration will begin with the launch of the first of three spacecraft that will collectively contribute toward the goal of delivering samples from the Red Planet to Earth. This mission is called Mars Sample Return (MSR) and will utilize both the 2003 and 2005 launch opportunities with an expected sample return in October 2008. NASA and CNES are major partners in this mission. The baseline mission mode selected for MSR is Mars orbit rendezvous (MOR), analogous in concept to the lunar orbit rendezvous (LOR) mode used for Apollo in the 1960s. Specifically, MSR will employ two NASA-provided landers of nearly identical design and one CNES-provided orbiter carrying a NASA payload of rendezvous sensors, orbital capture mechanisms, and an Earth entry vehicle (EEV). The high-level concept is that the landers will launch surface samples into Mars orbit, and the orbiter will retrieve the samples in orbit and then carry them back to Earth. The first element to depart for Mars will be one of the two landers. Currently, it is proposed that an intermediate class launch vehicle, such as the Boeing Delta 3 or Lockheed Martin Atlas 3A, will launch this 1800-kg lander from Cape Canaveral during the May 2003 opportunity. The lander will utilize a Type-1 transfer trajectory with an arrival at Mars in mid-December 2003. Landing will be aided by precision approach navigation and a guided hypersonic entry to achieve a touchdown accuracy of 10 km or better. Although the exact landing site has not yet been determined, it is estimated that lander resource constraints will limit the site to between 15 degrees north and south latitudes. Following touchdown, the lander will deploy a six-wheeled, 60-kg rover carrying an extensive suite of instruments designed to aid in the analysis of the local terrain and collection of core samples from selected rocks. The surface mission is currently designed around a concept called the surface traverse. Each

  5. Magnetospheric Multiscale (MMS) Mission Attitude Ground System Design

    NASA Technical Reports Server (NTRS)

    Sedlak, Joseph E.; Superfin, Emil; Raymond, Juan C.

    2010-01-01

    This paper describes the attitude ground system (AGS) design to be used for support of the Magnetospheric MultiScale (MMS) mission. The AGS exists as one component of the mission operations control center. It has responsibility for validating the onboard attitude and accelerometer bias estimates, calibrating the attitude sensors and the spacecraft inertia tensor, and generating a definitive attitude history for use by the science teams. NASA's Goddard Space Flight Center (GSFC) in Greenbelt, Maryland is responsible for developing the MMS spacecraft, for the overall management of the MMS mission, and for mission operations. MMS is scheduled for launch in 2014 for a planned two-year mission. The MMS mission consists of four identical spacecraft flying in a tetrahedral formation in an eccentric Earth orbit. The relatively tight formation, ranging from 10 to 400 km, will provide coordinated observations giving insight into small-scale magnetic field reconnection processes. By varying the size of the tetrahedron and the orbital semi-major axis and eccentricity, and making use of the changing solar phase, this geometry allows for the study of both bow shock and magnetotail plasma physics, including acceleration, reconnection, and turbulence. The mission divides into two phases for science; these phases will have orbit dimensions of 1.2 x 12 Earth radii in the first phase and 1.2x25 Earth radii in the second in order to study the dayside magnetopause and the nightside magnetotail, respectively. The orbital periods are roughly one day and three days for the two mission phases. Each of the four MMS spacecraft will be spin stabilized at 3 revolutions per minute (rpm), with the spin axis oriented near the ecliptic north pole but tipped approximately 2.5 deg towards the Sun line. The main body of each spacecraft will be an eight-sided platform with diameter of 3.4 m and height of 1.2 m. Several booms are attached to this central core: two axial booms of 14.9 m length, two

  6. Design of a High Sensitivity GNSS receiver for Lunar missions

    NASA Astrophysics Data System (ADS)

    Musumeci, Luciano; Dovis, Fabio; Silva, João S.; da Silva, Pedro F.; Lopes, Hugo D.

    2016-06-01

    This paper presents the design of a satellite navigation receiver architecture tailored for future Lunar exploration missions, demonstrating the feasibility of using Global Navigation Satellite Systems signals integrated with an orbital filter to achieve such a scope. It analyzes the performance of a navigation solution based on pseudorange and pseudorange rate measurements, generated through the processing of very weak signals of the Global Positioning System (GPS) L1/L5 and Galileo E1/E5 frequency bands. In critical scenarios (e.g. during manoeuvres) acceleration and attitude measurements from additional sensors complementing the GNSS measurements are integrated with the GNSS measurement to match the positioning requirement. A review of environment characteristics (dynamics, geometry and signal power) for the different phases of a reference Lunar mission is provided, focusing on the stringent requirements of the Descent, Approach and Hazard Detection and Avoidance phase. The design of High Sensitivity acquisition and tracking schemes is supported by an extensive simulation test campaign using a software receiver implementation and navigation results are validated by means of an end-to-end software simulator. Acquisition and tracking of GPS and Galileo signals of the L1/E1 and L5/E5a bands was successfully demonstrated for Carrier-to-Noise density ratios as low as 5-8 dB-Hz. The proposed navigation architecture provides acceptable performances during the considered critical phases, granting position and velocity errors below 61.4 m and 3.2 m/s, respectively, for the 99.7% of the mission time.

  7. Space station needs, attributes and architectural options. Volume 3, attachment 1, task 1: Mission requirements

    NASA Technical Reports Server (NTRS)

    1983-01-01

    The development and systems architectural requirements of the space station program are described. The system design is determined by user requirements. Investigated topics include physical and life science experiments, commercial utilization, U.S. national security, and remote space operations. The economic impact of the space station program is analyzed.

  8. Mission and Navigation Design for the 2009 Mars Science Laboratory Mission

    NASA Technical Reports Server (NTRS)

    D'Amario, Louis A.

    2008-01-01

    NASA s Mars Science Laboratory mission will launch the next mobile science laboratory to Mars in the fall of 2009 with arrival at Mars occurring in the summer of 2010. A heat shield, parachute, and rocket-powered descent stage, including a sky crane, will be used to land the rover safely on the surface of Mars. The direction of the atmospheric entry vehicle lift vector will be controlled by a hypersonic entry guidance algorithm to compensate for entry trajectory errors and counteract atmospheric and aerodynamic dispersions. The key challenges for mission design are (1) develop a launch/arrival strategy that provides communications coverage during the Entry, Descent, and Landing phase either from an X-band direct-to-Earth link or from a Ultra High Frequency link to the Mars Reconnaissance Orbiter for landing latitudes between 30 deg North and 30 deg South, while satisfying mission constraints on Earth departure energy and Mars atmospheric entry speed, and (2) generate Earth-departure targets for the Atlas V-541 launch vehicle for the specified launch/arrival strategy. The launch/arrival strategy employs a 30-day baseline launch period and a 27-day extended launch period with varying arrival dates at Mars. The key challenges for navigation design are (1) deliver the spacecraft to the atmospheric entry interface point (Mars radius of 3522.2 km) with an inertial entry flight path angle error of +/- 0.20 deg (3 sigma), (2) provide knowledge of the entry state vector accurate to +/- 2.8 km (3 sigma) in position and +/- 2.0 m/s (3 sigma) in velocity for initializing the entry guidance algorithm, and (3) ensure a 99% probability of successful delivery at Mars with respect to available cruise stage propellant. Orbit determination is accomplished via ground processing of multiple complimentary radiometric data types: Doppler, range, and Delta-Differential One-way Ranging (a Very Long Baseline Interferometry measurement). The navigation strategy makes use of up to five

  9. Design Concepts for the Generation-X Mission

    NASA Astrophysics Data System (ADS)

    Lillie, Charles F.; Dailey, D.; Danner, R.; Pearson, D.; Shropshire, D.

    2010-03-01

    The Generation-X mission, proposed by Roger Brissenden at SAO, is one of the Advanced Strategic Mission Concepts that NASA is considering for development in the post-2020 time period. As currently conceived Gen-X would be a follow-on to the International X-ray Observatory (IXO), with a collecting area ≥ 50 m2, 60-m focal length and 0.1 arc-second spatial resolution, which would be launched in 2030 with an Ares V Cargo Launch Vehicle to an L2 orbit. Our design concept assumes an Ares V with a 10-m diameter, 1,400 m3 volume fairing (or an equivalent launch vehicle) will be developed for NASA's exploration program. The key features of this design include a 16-m diameter deployable x-ray mirror provides a collecting area of 136 m2 a 60-m deployable optical bench which utilizes a Tensegrity structure to achieve high stiffness with low mass; and adaptive grazing incidence optics. Gen-X's combination of large collecting area and high spatial resolution will provide 4 to 5 orders of magnitude greater sensitivity than IXO, enabling scientists to study the formation and growth of the first black holes at z ≈ 8-15 with 0.1 to 10 keV fluxes of ≈ 10-20 erg cm-2s-1.

  10. Design Concepts for the Generation-X Mission

    NASA Astrophysics Data System (ADS)

    Lillie, Charles F.; Dailey, D.; Danner, R.; Shropshire, D.; Pearson, D.

    2009-09-01

    The Generation-X mission, proposed by Roger Brissenden at SAO, is one of the Advanced Strategic Mission Concepts that NASA is considering for development in the post-2020 time period. As currently conceived Gen-X would be a follow-on to the International X-ray Observatory (IXO), with a collecting area ≥ 50 m^2, 60-m focal length and 0.1 arc-second spatial resolution, which would be launched in ˜2030 with an Ares V Cargo Launch Vehicle to an L2 orbit. Our design concept assumes an Ares V with a 10-m diameter, 1,400 m^3 volume fairing (or an equivalent launch vehicle) will be developed for NASA's exploration program. The key features of this design include a 16-m diameter deployable x-ray mirror provides a collecting area of 136 m^2; a 60-m deployable optical bench which utilizes a Tensegrity structure to achieve high stiffness with low mass; and adaptive grazing incidence optics. Gen-X's combination of large collecting area and high spatial resolution will provide 4 to 5 orders of magnitude greater sensitivity than IXO, enabling scientists to study the formation and growth of the first black holes at z ≈ 8-15 with 0.1 to 10 keV fluxes of ≈ 10-20 erg cm^{-2}s^{-1}.

  11. A Review of High Thrust, High Delta-V Options for Microsatellite Missions

    DTIC Science & Technology

    2009-06-25

    millinewtons of thrust. Pushing the limits of microsatellite capability is the Hall thruster design of Berti, et al.23 and Biagioni , et al.,24...of thrust with an Isp greater than 1000 s. Biagioni , et al. further specify that their thruster weighs 0.6 kg and that the power and flow control...Sept. 2002, AIAA-2002-5714. 23Berti, M., Biagioni , L., Cesari, U., Saverdi, M., and Andrenucci, M., “Development and Preliminary Characterization of a

  12. Soil Moisture Active Passive Mission: Fault Management Design Analyses

    NASA Technical Reports Server (NTRS)

    Meakin, Peter; Weitl, Raquel

    2013-01-01

    As a general trend, the complexities of modern spacecraft are increasing to include more ambitious mission goals with tighter timing requirements and on-board autonomy. As a byproduct, the protective features that monitor the performance of these systems have also increased in scope and complexity. Given cost and schedule pressures, there is an increasing emphasis on understanding the behavior of the system at design time. Formal test-driven verification and validation (V&V) is rarely able to test the significant combinatorics of states, and often finds problems late in the development cycle forcing design changes that can be costly. This paper describes the approach the SMAP Fault Protection team has taken to address some of the above-mentioned issues.

  13. Space station needs, attributes and architectural options. Volume 4, attachment 1: Task 2 and 3 mission implementation and cost

    NASA Technical Reports Server (NTRS)

    1983-01-01

    Mission scenario analysis and architectural concepts, alternative systems concepts, mission operations and architectural development, architectural analysis trades, evolution, configuration, and technology development are assessed.

  14. Validation of a Low-Thrust Mission Design Tool Using Operational Navigation Software

    NASA Technical Reports Server (NTRS)

    Englander, Jacob A.; Knittel, Jeremy M.; Williams, Ken; Stanbridge, Dale; Ellison, Donald H.

    2017-01-01

    Design of flight trajectories for missions employing solar electric propulsion requires a suitably high-fidelity design tool. In this work, the Evolutionary Mission Trajectory Generator (EMTG) is presented as a medium-high fidelity design tool that is suitable for mission proposals. EMTG is validated against the high-heritage deep-space navigation tool MIRAGE, demonstrating both the accuracy of EMTG's model and an operational mission design and navigation procedure using both tools. The validation is performed using a benchmark mission to the Jupiter Trojans.

  15. Design and Analysis of RTGs for Solar and Martian Exploration Missions

    SciTech Connect

    Schock, Alfred

    1990-05-01

    The paper described the results of design, analysis and spacecraft integration studies of Radioisotope Thermoelectric Generators (RTGs) for three unmanned space exploration missions. The three missions, consisting of the Mars Rover and Sample Return (MRSR) mission, the Solar Probe mission, and the Mars Global Net work (MGN) mission, are under study by the Jet Propulsion Laboratory (JPL) for the U.S. National Aeronautics and Space Administration (NASA). The NASA/JPL mission studies are supported by the U.S. Department of Energy's Office of Special Applications (DOE/OSA), which has commissioned Fairchild Space Company to carry out the required RTG design studies.

  16. Tether-mission design for multiple flybys of moon Europa

    NASA Astrophysics Data System (ADS)

    Sanmartin, J. R. S.; Charro, M. C.; Sanchez-Arriaga, G. S. A.; Sanchez-Torres, A. S. T.

    2015-10-01

    A tether mission to carry out multiple flybys of Jovian moon Europa is here presented. There is general agreement on elliptic-orbit flybys of Europa resulting in cost to attain given scientific goals lower than if actually orbiting the moon, tethers being naturally fit to fly-by rather than orbit moons1. The present mission is similar in this respect to the Clipper mission considered by NASA, the basic difference lying in location of periapsis, due to different emphasis on mission-challenge metrics. Clipper minimizes damaging radiation-dose by avoiding the Jupiter neighborhood and its very harsh environment; periapsis would be at Europa, apoapsis as far as moon Callisto. As in all past outer-planet missions, Clipper faces, however, critical power and propulsion needs. On the other hand, tethers can provide both propulsion and power, but must reach near the planet to find high plasma density and magnetic field values, leading to high induced tether current, and Lorentz drag and power. The bottom line is a strong radiation dose under the very intense Radiation Belts of Jupiter. Mission design focuses on limiting dose. Perijove would be near Jupiter, at about 1.2-1.3 Jovian radius, apojove about moon Ganymede, corresponding to 1:1 resonance with Europa, so as to keep dose down: setting apojove at Europa, for convenient parallel flybys, would require two perijove passes per flyby (the Ganymede apojove, resulting in high eccentricity, about 0.86, is also less requiring on tether operations). Mission is designed to attain reductions in eccentricity per perijove pass as high as Δe ≈ - 0.04. Due the low gravity-gradient, tether spinning is necessary to keep it straight, plasma contactors placed at both ends taking active turns at being cathodic. Efficiency of capture of the incoming S/C by the tether is gauged by the ratio of S/C mass to tether mass; efficiency is higher for higher tape-tether length and lower thickness and perijove. Low tether bowing due to the Lorentz

  17. LUGH, the Proposed Mercury Express Mission, as an Ideal, Current, Low-Cost, Low-Risk Option for Mercury Exploration

    NASA Technical Reports Server (NTRS)

    Clark, P. E.; Lawlor, S. McKenna; Curtis, S.; Marr, G.; Giles, B.

    2000-01-01

    We propose an ESA Flexi Mission, LUGH, Mercury Express Mission, an extremely fast, low cost, low risk, high return, three-platform, multiple flyby mission which would provide data which are unique and complimentary to recently selected long lead time Mercury missions.

  18. Space Station needs, attributes and architectural options. Volume 2, book 1, part 2, task 1: Mission requirements

    NASA Technical Reports Server (NTRS)

    1983-01-01

    Mission areas analyzed for input to the baseline mission model include: (1) commercial materials processing, including representative missions for producing metallurgical, chemical and biological products; (2) commercial Earth observation, represented by a typical carry-on mission amenable to commercialization; (3) solar terrestrial and resource observations including missions in geoscience and scientific land observation; (4) global environment, including representative missions in meteorology, climatology, ocean science, and atmospheric science; (5) materials science, including missions for measuring material properties, studying chemical reactions and utilizing the high vacuum-pumping capacity of space; and (6) life sciences with experiments in biomedicine and animal and plant biology.

  19. Mission to the Trojan asteroids: Lessons learned during a JPL Planetary Science Summer School mission design exercise

    NASA Astrophysics Data System (ADS)

    Diniega, Serina; Sayanagi, Kunio M.; Balcerski, Jeffrey; Carande, Bryce; Diaz-Silva, Ricardo A.; Fraeman, Abigail A.; Guzewich, Scott D.; Hudson, Jennifer; Nahm, Amanda L.; Potter-McIntyre, Sally; Route, Matthew; Urban, Kevin D.; Vasisht, Soumya; Benneke, Bjoern; Gil, Stephanie; Livi, Roberto; Williams, Brian; Budney, Charles J.; Lowes, Leslie L.

    2013-02-01

    The 2013 Planetary Science Decadal Survey identified a detailed investigation of the Trojan asteroids occupying Jupiter's L4 and L5 Lagrange points as a priority for future NASA missions. Observing these asteroids and measuring their physical characteristics and composition would aid in identification of their source and provide answers about their likely impact history and evolution, thus yielding information about the makeup and dynamics of the early Solar System. We present a conceptual design for a mission to the Jovian Trojan asteroids: the Trojan ASteroid Tour, Exploration, and Rendezvous (TASTER) mission, that is consistent with the NASA New Frontiers candidate mission recommended by the Decadal Survey and the final result of the 2011 NASA-JPL Planetary Science Summer School. Our proposed mission includes visits to two Trojans in the L4 population: a 500 km altitude fly-by of 1999 XS143, followed by a rendezvous with and detailed observations of 911 Agamemnon at orbital altitudes of 1000-100 km over a 12 month nominal science data capture period. Our proposed instrument payload - wide- and narrow-angle cameras, a visual and infrared mapping spectrometer, and a neutron/gamma ray spectrometer - would provide unprecedented high-resolution, regional-to-global datasets for the target bodies, yielding fundamental information about the early history and evolution of the Solar System. Although our mission design was completed as part of an academic exercise, this study serves as a useful starting point for future Trojan mission design studies. In particular, we identify and discuss key issues that can make large differences in the complex trade-offs required when designing a mission to the Trojan asteroids.

  20. Automated Design of Multiphase Space Missions Using Hybrid Optimal Control

    ERIC Educational Resources Information Center

    Chilan, Christian Miguel

    2009-01-01

    A modern space mission is assembled from multiple phases or events such as impulsive maneuvers, coast arcs, thrust arcs and planetary flybys. Traditionally, a mission planner would resort to intuition and experience to develop a sequence of events for the multiphase mission and to find the space trajectory that minimizes propellant use by solving…

  1. Interplanetary mission design handbook. Volume 1, part 3: Earth to Jupiter ballistic mission opportunities, 1985-2005

    NASA Technical Reports Server (NTRS)

    Sergeyevsky, A. B.; Snyder, G. C.

    1982-01-01

    Graphical data necessary for the preliminary design of ballistic missions to Jupiter are provided. Contours of launch energy requirements, as well as many other launch and Jupiter arrival parameters, are presented in launch date/arrival date space for all launch opportunities from 1985 through 2005. In addition, an extensive text is included which explains mission design methods, from launch window development to Jupiter probe and orbiter arrival design, utilizing the graphical data in this volume as well as numerous equations relating various parameters.

  2. Design Options to Reduce Development Cost of First Generation Surface Reactors

    SciTech Connect

    Poston, David I.; Marcille, Thomas F.

    2006-01-20

    Low-power surface reactors have the potential to have the lowest development cost of any space reactor application, primarily because system alpha (mass/kg) is not of utmost importance and mission lifetimes do not have to be a decade or more. Even then, the development cost of a surface reactor can vary substantially depending on the performance requirements (e.g. mass, power, lifetime, reliability) and technical development risk deemed acceptable by the end-user. It is important for potential users to be aware of these relationships before they determine their future architecture (i.e. decide what they need). Generally, the greatest potential costs of a space reactor program are a nuclear-powered ground test and extensive material development campaigns, so it is important to consider options that can minimize the need for or complexity of such tasks. The intended goal of this paper is to inform potential surface reactor users of the potential sensitivities of surface reactor development cost to design requirements, and areas where technical risk can be traded with development cost.

  3. Mars Surface Habitability Options

    NASA Technical Reports Server (NTRS)

    Howe, A. Scott; Simon, Matthew; Smitherman, David; Howard, Robert; Toups, Larry; Hoffman, Stephen J.

    2015-01-01

    This paper reports on current habitability concepts for an Evolvable Mars Campaign (EMC) prepared by the NASA Human Spaceflight Architecture Team (HAT). For many years NASA has investigated alternative human Mars missions, examining different mission objectives, trajectories, vehicles, and technologies; the combinations of which have been referred to as reference missions or architectures. At the highest levels, decisions regarding the timing and objectives for a human mission to Mars continue to evolve while at the lowest levels, applicable technologies continue to advance. This results in an on-going need for assessments of alternative system designs such as the habitat, a significant element in any human Mars mission scenario, to provide meaningful design sensitivity characterizations to assist decision-makers regarding timing, objectives, and technologies. As a subset of the Evolvable Mars Campaign activities, the habitability team builds upon results from past studies and recommends options for Mars surface habitability compatible with updated technologies.

  4. Investigation of design options for improving the energy efficiency of conventionally designed refrigerator-freezers

    SciTech Connect

    Sand, J.R.; Vineyard, E.A.; Bohman, R.H.

    1993-11-01

    Several design options for improving the energy efficiency of conventionally-designed, domestic refrigerator freezers (RFs) were incorporated into two 1990 production RF cabinets and refrigeration systems. The baseline performance of the original units and unit components were extensively documented to provide a firm basis for experimentally measured energy savings. A detailed refrigerator system computer model which could simulate cycling behavior was used to evaluate the daily energy use impacts for each modification, and modeled versus experimental results are compared. The model was shown to track measured RF performance improvement sufficiently well that it was used with some confidence to investigate additional options that could not be experimentally investigated. Substantial improvements in RF efficiency were demonstrated with relatively minor changes in system components and refrigeration circuit design. However, each improvement exacts a penalty in terms of increased cost or system complexity/reliability. For RF sizes typically sold in the United States (18-22 ft{sup 3} [510--620 1]), alternative, more-elaborate, refrigeration cycles may be required to achieve the program goal (1.00 Kilowatt-hour per day for a 560 l, top mount RF.

  5. Spacecraft Station-Keeping Trajectory and Mission Design Tools

    NASA Technical Reports Server (NTRS)

    Chung, Min-Kun J.

    2009-01-01

    Two tools were developed for designing station-keeping trajectories and estimating delta-v requirements for designing missions to a small body such as a comet or asteroid. This innovation uses NPOPT, a non-sparse, general-purpose sequential quadratic programming (SQP) optimizer and the Two-Level Differential Corrector (T-LDC) in LTool (Libration point mission design Tool) to design three kinds of station-keeping scripts: vertical hovering, horizontal hovering, and orbiting. The T-LDC is used to differentially correct several trajectory legs that join hovering points. In a vertical hovering, the maximum and minimum range points must be connected smoothly while maintaining the spacecrafts range from a small body, all within the law of gravity and the solar radiation pressure. The same is true for a horizontal hover. A PatchPoint is an LTool class that denotes a space-time event with some extra information for differential correction, including a set of constraints to be satisfied by T-LDC. Given a set of PatchPoints, each with its own constraint, the T-LDC differentially corrects the entire trajectory by connecting each trajectory leg joined by PatchPoints while satisfying all specified constraints at the same time. Vertical and horizontal hover both are needed to minimize delta-v spent for station keeping. A Python I/F to NPOPT has been written to be used from an LTool script. In vertical hovering, the spacecraft stays along the line joining the Sun and a small body. An instantaneous delta-v toward the anti- Sun direction is applied at the closest approach to the small body for station keeping. For example, the spacecraft hovers between the minimum range (2 km) point and the maximum range (2.5 km) point from the asteroid 1989ML. Horizontal hovering buys more time for a spacecraft to recover if, for any reason, a planned thrust fails, by returning almost to the initial position after some time later via a near elliptical orbit around the small body. The mapping or

  6. Revised Point of Departure Design Options for Nuclear Thermal Propulsion

    NASA Technical Reports Server (NTRS)

    Fittje, James E.; Schnitzler, Bruce G.; Borowoski, Stanley

    2015-01-01

    Four Revised Point of Departure NTR Engines were Designed and Analyzed using MCNP and NESS. All Four Engines Have Thermodynamically Closed Cycles at Nominal Chamber Pressures. 111 kilonewton (25 kip-force) Cermet Design Required Dedicated Heater Elements to Close the Cycle. Cermet Based Designs had Slightly Higher TW Ratios, but Required Substantially More U-235. NERVA Derived Criticality Limited Engine Could Operate at Lower Power and Thrust Levels Compared to the Criticality Limited Cermet Design.

  7. Navigation system design for a Halley Flyby/Tempel 2 Rendezvous mission using ion drive

    NASA Technical Reports Server (NTRS)

    Wood, L. J.; Hast, S. L.

    1979-01-01

    A dual comet (Hall Flyby/Tempel 2 Rendezvous) mission, making use of the solar electric propulsion system, is under consideration for a 1985 launch. This paper describes the preliminary navigation system design for this mission. Orbit determination and guidance strategies for each mission phase are discussed. Navigation accuracy analyses and parametric senstivity studies for the Tempel 2 rendezvous approach phase are presented.

  8. Rapid Preliminary Design of Interplanetary Trajectories Using the Evolutionary Mission Trajectory Generator

    NASA Technical Reports Server (NTRS)

    Englander, Jacob

    2016-01-01

    This set of tutorial slides is an introduction to the Evolutionary Mission Trajectory Generator (EMTG), NASA Goddard Space Flight Center's autonomous tool for preliminary design of interplanetary missions. This slide set covers the basics of creating and post-processing simple interplanetary missions in EMTG using both high-thrust chemical and low-thrust electric propulsion along with a variety of operational constraints.

  9. MAIUS-1- Vehicle, Subsystems Design and Mission Operations

    NASA Astrophysics Data System (ADS)

    Stamminger, A.; Ettl, J.; Grosse, J.; Horschgen-Eggers, M.; Jung, W.; Kallenbach, A.; Raith, G.; Saedtler, W.; Seidel, S. T.; Turner, J.; Wittkamp, M.

    2015-09-01

    In November 2015, the DLR Mobile Rocket Base will launch the MAIUS-1 rocket vehicle at Esrange, Northern Sweden. The MAIUS-A experiment is a pathfinder atom optics experiment. The scientific objective of the mission is the first creation of a BoseEinstein Condensate in space and performing atom interferometry on a sounding rocket [3]. MAIUS-1 comprises a two-stage unguided solid propellant VSB-30 rocket motor system. The vehicle consists of a Brazilian 53 1 motor as 1 st stage, a 530 motor as 2nd stage, a conical motor adapter, a despin module, a payload adapter, the MAIUS-A experiment consisting of five experiment modules, an attitude control system module, a newly developed conical service system, and a two-staged recovery system including a nosecone. In contrast to usual payloads on VSB-30 rockets, the payload has a diameter of 500 mm due to constraints of the scientific experiment. Because of this change in design, a blunted nosecone is necessary to guarantee the required static stability during the ascent phase of the flight. This paper will give an overview on the subsystems which have been built at DLR MORABA, especially the newly developed service system. Further, it will contain a description of the MAIUS-1 vehicle, the mission and the unique requirements on operations and attitude control, which is additionally required to achieve a required attitude with respect to the nadir vector. Additionally to a usual microgravity environment, the MAIUS-l payload requires attitude control to achieve a required attitude with respect to the nadir vector.

  10. Hayabusa Re-Entry: Trajectory Analysis and Observation Mission Design

    NASA Technical Reports Server (NTRS)

    Cassell, Alan M.; Winter, Michael W.; Allen, Gary A.; Grinstead, Jay H.; Antimisiaris, Manny E.; Albers, James; Jenniskens, Peter

    2011-01-01

    On June 13th, 2010, the Hayabusa sample return capsule successfully re-entered Earth s atmosphere over the Woomera Prohibited Area in southern Australia in its quest to return fragments from the asteroid 1998 SF36 Itokawa . The sample return capsule entered at a super-orbital velocity of 12.04 km/sec (inertial), making it the second fastest human-made object to traverse the atmosphere. The NASA DC-8 airborne observatory was utilized as an instrument platform to record the luminous portion of the sample return capsule re-entry (60 sec) with a variety of on-board spectroscopic imaging instruments. The predicted sample return capsule s entry state information at 200 km altitude was propagated through the atmosphere to generate aerothermodynamic and trajectory data used for initial observation flight path design and planning. The DC- 8 flight path was designed by considering safety, optimal sample return capsule viewing geometry and aircraft capabilities in concert with key aerothermodynamic events along the predicted trajectory. Subsequent entry state vector updates provided by the Deep Space Network team at NASA s Jet Propulsion Laboratory were analyzed after the planned trajectory correction maneuvers to further refine the DC-8 observation flight path. Primary and alternate observation flight paths were generated during the mission planning phase which required coordination with Australian authorities for pre-mission approval. The final observation flight path was chosen based upon trade-offs between optimal viewing requirements, ground based observer locations (to facilitate post-flight trajectory reconstruction), predicted weather in the Woomera Prohibited Area and constraints imposed by flight path filing deadlines. To facilitate sample return capsule tracking by the instrument operators, a series of two racetrack flight path patterns were performed prior to the observation leg so the instruments could be pointed towards the region in the star background where

  11. Review of a Spoke-Cavity Design Option for the RIA Driver Linac

    SciTech Connect

    Petr Ostroumov; Kenneth Shepard; Jean Delayen

    2005-05-16

    A design option for the 1.4 GV, multiple-charge-state driver linac required for the U. S. Rare Isotope Accelerator Project based on 345 MHz, 3-cell spoke-loaded cavities has been previously discussed [1]. This paper updates consideration of design options for the RIA driver, including recent results from numerically-modeling the multi-charge-state beam dynamics and also cold test results for prototype superconducting niobium 3-cell spoke-loaded cavities.

  12. Review of a spoke-cavity design option for the RIA driver linac.

    SciTech Connect

    Ostroumov, P. N.; Shepard, K. W.; Delayen, J. R.; Physics; Thomas Jefferson National Accel. Facility

    2005-01-01

    A design option for the 1.4 GV, multiple-charge-state driver linac required for the U.S. Rare Isotope Accelerator Project based on 345 MHz, 3-cell spoke-loaded cavities has been previously discussed [1]. This paper updates consideration of design options for the RIA driver, including recent results from numerically-modeling the multi-charge-state beam dynamics and also cold test results for prototype superconducting niobium three-spoke-loaded cavities.

  13. Development of On-Orbit Servicing Concepts, Technology Option and Roadmap (Part II) - Technical Design

    NASA Astrophysics Data System (ADS)

    Ulrich, S.; Schwartz, J.; Abbasi, V.; Cree, D.; Daae, M.; Grover, D.; Hay, J.; He, W.; Huang, X.; Jacobs, S.; Jun, Z.; Kearney, S.; Kuwahara, T.; Lenzi, F.; Mirahmetoglu, H.; Morley, S.; Otani, M.; Pastena, M.; Pinni, M.; Shala, K.; Shi, J.-F.; Singh Koral, J.; Steinkellner, M.; Treat, D.; Verheyden, P.; Wang, X. Y.; Weeden, C.

    On-Orbit-Servicing (OOS) can enable valuable assets to be repaired and maintained in space, thus lowering spacecraft development costs and timelines. As space activities grow and ever more spacecraft are placed in orbit, such servicing technologies will become increasingly important. This paper proposes a system architecture design for an Orbital Replacement Unit (ORU) exchange service mission in Geostationary Orbit (GEO). It describes in detail the mission and spacecraft design, thereby demonstrating its feasibility in the midterm future.

  14. Design options for an ITER ion cyclotron system

    SciTech Connect

    Swain, D.W.; Baity, F.W.; Bigelow, T.S.; Ryan, P.M.; Goulding, R.H.; Carter, M.D.; Stallings, D.C.; Batchelor, D.B.; Hoffman, D.J.

    1995-09-01

    Recent changes have occurred in the design requirements for the ITER ion cyclotron system, requiring in-port launchers in four main horizontal ports to deliver 50 MW of power to the plasma. The design is complicated by the comparatively large antenna-separatrix distance of 10--20 cm. Designs of a conventional strap launcher and a folded waveguide launcher than can meet the new requirements are presented.

  15. Mission analysis and guidance, navigation, and control design for rendezvous and docking phase of advanced reentry vehicle mission

    NASA Astrophysics Data System (ADS)

    Strippoli, L.; Colmenarejo, P.; Strauch, H.

    2013-12-01

    Advanced Reentry Vehicle (ARV) belongs to the family of vehicles designed to perform rendezvous and docking (RvD) with the International space station (ISS) [1]. Differently from its predecessor ATV (Automated Transfer Vehicle), the ARV will transport a reentry capsule, equipped with a heatshield and able to bring back cargo, experiments, or, as a possible future development, even crew, being this latter scenario very attracting in view of the Space Shuttle retirement. GMV, as subcontractor of EADS-Astrium Germany, is in charge of the RvD and departure mission analysis and GNC (Guidance, Navigation, and Control) design of ARV mission. This paper will present the main outcomes of the study.

  16. Mission-based Scenario Research: Experimental Design And Analysis

    DTIC Science & Technology

    2012-01-01

    loudspeakers. One of the pre-recorded voices was a simulated tactical operating commander ( TOC ) who provided the mission directives a Commander would...expect on a patrol mission. One of the experimenters also operated a soundboard with controls to activate pre-recorded TOC responses, facilitating...simulated interactions between the TOC and Commander; for example, one button allowed the TOC to respond “Roger” when the Commander called in mission

  17. A study of space station needs, attributes and architectural options. Volume 2: Technical. Book 1: Mission requirements

    NASA Technical Reports Server (NTRS)

    Steinbronn, O.

    1983-01-01

    The following types of space missions were evaluated to determine those that require, or will be benefited materially, by a manned space station: (1) science and applications, (2) commercial, (3) technology development, (4) space operations, and (5) national security. Integrated mission requirements for man-operated and man-tended free-flying missions were addressed. A manned space station will provide major performance and economic benefits to a wide range of missions planned for the 1990s.

  18. Jovian Tour Design for Orbiter and Lander Missions to Europa

    NASA Technical Reports Server (NTRS)

    Campagnola, Stefano; Buffington, Brent B.; Petropoulos, Anastassios E.

    2013-01-01

    Europa is one of the most interesting targets for solar system exploration, as its ocean of liquid water could harbor life. Following the recommendation of the Planetary Decadal Survey, NASA commissioned a study for a flyby mission, an orbiter mission, and a lander mission. This paper presents the moon tours for the lander and orbiter concepts. The total delta v and radiation dose would be reduced by exploiting multi-body dynamics and avoiding phasing loops in the Ganymede-to- Europa transfer. Tour 11-O3, 12-L1 and 12-L4 are presented in details and their performaces compared to other tours from previous Europa mission studies.

  19. New window design options for CEBAF energy upgrade

    SciTech Connect

    Phillips, L.; Mammosser, J.; Nguyen, V.

    1997-07-01

    As the Jefferson Laboratory upgrades the existing CEBAF electron accelerator to operate at higher energies, the fundamental power coupler windows will be required to operate with lower RF dissipation and increased immunity to radiation from cavity field emission. New designs and modifications to existing designs which can achieve these goals are described.

  20. Planning Coverage Campaigns for Mission Design and Analysis: CLASP for DESDynl

    NASA Technical Reports Server (NTRS)

    Knight, Russell L.; McLaren, David A.; Hu, Steven

    2013-01-01

    Mission design and analysis presents challenges in that almost all variables are in constant flux, yet the goal is to achieve an acceptable level of performance against a concept of operations, which might also be in flux. To increase responsiveness, automated planning tools are used that allow for the continual modification of spacecraft, ground system, staffing, and concept of operations, while returning metrics that are important to mission evaluation, such as area covered, peak memory usage, and peak data throughput. This approach was applied to the DESDynl mission design using the CLASP planning system, but since this adaptation, many techniques have changed under the hood for CLASP, and the DESDynl mission concept has undergone drastic changes. The software produces mission evaluation products, such as memory highwater marks, coverage percentages, given a mission design in the form of coverage targets, concept of operations, spacecraft parameters, and orbital parameters. It tries to overcome the lack of fidelity and timeliness of mission requirements coverage analysis during mission design. Previous techniques primarily use Excel in ad hoc fashion to approximate key factors in mission performance, often falling victim to overgeneralizations necessary in such an adaptation. The new program allows designers to faithfully represent their mission designs quickly, and get more accurate results just as quickly.

  1. Key risk attributes in the perception of engineering design options

    SciTech Connect

    Grindrod, P.; Waters, D.J.; Yousaf, F.A.; Takase, H.

    1996-12-01

    The design of an engineered barrier system (EBS) for the containment of radioactive waste buried at depth incorporates a wide range of decisions based on quantitative engineering science, site specific hydrological information and expert judgement. Even at the concept design and planning stage of the EBS, there may be some key alternatives or choices which, though usually considered from an executive engineering perspective, may have a large impact upon the success of the programme as a whole. Therefore it is of interest to ask {open_quotes}what are the key attributes?{close_quotes} of the design process from the perspective of those experts working in the perception/communication fields, as well as the supporting research assessments and programmes. This involves the consideration of subjective expert opinions in various disciplines, and the identification of differences in the structure of their cognitive reasoning regarding the EBS. This report describes how a group of experts responded to a range of EBS designs.

  2. Multi-Agent Modeling and Simulation Approach for Design and Analysis of MER Mission Operations

    NASA Technical Reports Server (NTRS)

    Seah, Chin; Sierhuis, Maarten; Clancey, William J.

    2005-01-01

    A space mission operations system is a complex network of human organizations, information and deep-space network systems and spacecraft hardware. As in other organizations, one of the problems in mission operations is managing the relationship of the mission information systems related to how people actually work (practices). Brahms, a multi-agent modeling and simulation tool, was used to model and simulate NASA's Mars Exploration Rover (MER) mission work practice. The objective was to investigate the value of work practice modeling for mission operations design. From spring 2002 until winter 2003, a Brahms modeler participated in mission systems design sessions and operations testing for the MER mission held at Jet Propulsion Laboratory (JPL). He observed how designers interacted with the Brahms tool. This paper discussed mission system designers' reactions to the simulation output during model validation and the presentation of generated work procedures. This project spurred JPL's interest in the Brahms model, but it was never included as part of the formal mission design process. We discuss why this occurred. Subsequently, we used the MER model to develop a future mission operations concept. Team members were reluctant to use the MER model, even though it appeared to be highly relevant to their effort. We describe some of the tool issues we encountered.

  3. Internal Control Rod Drive Mechanisms, Design Options for IRIS

    SciTech Connect

    Conway, Lawrence E.; Petrovic, Bojan

    2004-07-01

    IRIS (International Reactor Innovative and Secure) is a medium-power (335 MWe) PWR with an integral, primary circuit configuration, where all the reactor coolant system components are contained within the reactor vessel. This integral configuration is a key reason for the success of IRIS' 'safety-by-design' approach, whereby accident initiators are eliminated or the accident consequences and/or frequency are reduced. The most obvious example of the IRIS safety by design approach is the elimination of large LOCA's, since the integral reactor coolant system has no large loop piping. Another serious accident scenario that is being addressed in IRIS is the postulated ejection of a reactor control cluster assembly (RCCA). This accident initiator can be eliminated by locating the RCCA drive mechanisms (CRDMs) inside the reactor vessel. This eliminates the mechanical drive rod penetration between the RCCA and the external CRDM, eliminating the potential for differential pressure across the pressure boundary, and thus eliminating 'by design' the possibility for rod ejection accident. Moreover, the elimination of the 'large' drive-rod penetrations and the external CRDM pressure housings decreases the likelihood of boric acid leakage and subsequent corrosion of the reactor pressure boundary (like the Davis-Besse incident). This paper will discuss the IRIS top level design requirements and objectives for internal CRDMs, and provide examples candidate designs and their specific performance characteristics. (authors)

  4. A mission-oriented orbit design method of remote sensing satellite for region monitoring mission based on evolutionary algorithm

    NASA Astrophysics Data System (ADS)

    Shen, Xin; Zhang, Jing; Yao, Huang

    2015-12-01

    Remote sensing satellites play an increasingly prominent role in environmental monitoring and disaster rescue. Taking advantage of almost the same sunshine condition to same place and global coverage, most of these satellites are operated on the sun-synchronous orbit. However, it brings some problems inevitably, the most significant one is that the temporal resolution of sun-synchronous orbit satellite can't satisfy the demand of specific region monitoring mission. To overcome the disadvantages, two methods are exploited: the first one is to build satellite constellation which contains multiple sunsynchronous satellites, just like the CHARTER mechanism has done; the second is to design non-predetermined orbit based on the concrete mission demand. An effective method for remote sensing satellite orbit design based on multiobjective evolution algorithm is presented in this paper. Orbit design problem is converted into a multi-objective optimization problem, and a fast and elitist multi-objective genetic algorithm is utilized to solve this problem. Firstly, the demand of the mission is transformed into multiple objective functions, and the six orbit elements of the satellite are taken as genes in design space, then a simulate evolution process is performed. An optimal resolution can be obtained after specified generation via evolution operation (selection, crossover, and mutation). To examine validity of the proposed method, a case study is introduced: Orbit design of an optical satellite for regional disaster monitoring, the mission demand include both minimizing the average revisit time internal of two objectives. The simulation result shows that the solution for this mission obtained by our method meet the demand the users' demand. We can draw a conclusion that the method presented in this paper is efficient for remote sensing orbit design.

  5. Design options for improving protective gloves for industrial assembly work.

    PubMed

    Dianat, Iman; Haslegrave, Christine M; Stedmon, Alex W

    2014-07-01

    The study investigated the effects of wearing two new designs of cotton glove on several hand performance capabilities and compared them against the effects of barehanded, single-layered and double cotton glove conditions when working with hand tools (screwdriver and pliers). The new glove designs were based on the findings of subjective hand discomfort assessments for this type of work and aimed to match the glove thickness to the localised pressure and sensitivity in different areas of the hand as well as to provide adequate dexterity for fine manipulative tasks. The results showed that the first prototype glove and the barehanded condition were comparable and provided better dexterity and higher handgrip strength than double thickness gloves. The results support the hypothesis that selective thickness in different areas of the hand could be applied by glove manufacturers to improve the glove design, so that it can protect the hands from the environment and at the same time allow optimal hand performance capabilities.

  6. Design and application of electromechanical actuators for deep space missions

    NASA Technical Reports Server (NTRS)

    Haskew, Tim A.; Wander, John

    1994-01-01

    This progress report documents research and development efforts performed from August 16, 1993 through February 15, 1994 on NASA Grant NAG8-240, 'Design and Application of Electromechanical Actuators for Deep Space Missions.' Following the executive summary are four report sections: Motor Selection, Tests Stand Development, Health Monitoring and Fault Management, and Experiment Planning. Three specific motor types have been considered as prime movers for TVC EMA applications: the brushless dc motor, the permanent magnet synchronous motor, and the induction motor. The fundamental finding was that, in general, the primary performance issues were energy efficiency and thermal dissipation (rotor heating). In terms of all other issues, the three motor types were found to compare quite equally. Among the design changes made to the test stand since the last progress report is the addition of more mounting holes in the side beams. These additional holes allow the movable end beam to be attached in a greater number of positions than previously. With this change the movable end beam can move from full forward to full back in three inch increments. Specific mathematical details on the approach that have been employed for health monitoring and fault management (HMFM) have been reported previously. This approach is based on and adaptive Kalman filter strategy. In general, a bank of filters can be implemented for each primary fault type. Presently under consideration for the brushless dc machine are the following faults: armature winding open-circuits, armature winding short-circuits (phase-to-phase and phase-to-ground), bearing degradation, and rotor flux weakening. The mechanically oriented experiments include transient loading experiments, transverse loading experiment, friction experiment, motor performance experiment, and HMFM experiment.

  7. Options Study Documenting the Fast Reactor Fuels Innovative Design Activity

    SciTech Connect

    Jon Carmack; Kemal Pasamehmetoglu

    2010-07-01

    This document provides presentation and general analysis of innovative design concepts submitted to the FCRD Advanced Fuels Campaign by nine national laboratory teams as part of the Innovative Transmutation Fuels Concepts Call for Proposals issued on October 15, 2009 (Appendix A). Twenty one whitepapers were received and evaluated by an independent technical review committee.

  8. Designing Undergraduate Research Experiences: A Multiplicity of Options

    NASA Astrophysics Data System (ADS)

    Manduca, C. A.

    2001-12-01

    Research experiences for undergraduate students can serve many goals including: developing student understanding of the process of science; providing opportunities for students to develop professional skills or test career plans; completing publishable research; enabling faculty professional development; or enhancing the visibility of a science program. The large range of choices made in the design of an undergraduate research program or opportunity must reflect the goals of the program, the needs and abilities of the students and faculty, and the available resources including both time and money. Effective program design, execution, and evaluation can all be enhanced if the goals of the program are clearly articulated. Student research experiences can be divided into four components: 1) defining the research problem; 2) developing the research plan or experiment design; 3) collecting and interpreting data, and 4) communicating results. In each of these components, the program can be structured in a wide variety of ways and students can be given more or less guidance or freedom. While a feeling of ownership of the research project appears to be very important, examples of successful projects displaying a wide range of design decisions are available. Work with the Keck Geology Consortium suggests that four strategies can enhance the likelihood of successful student experiences: 1) students are well-prepared for research experience (project design must match student preparation); 2) timelines and events are structured to move students through intermediate goals to project completion; 3) support for the emotional, financial, academic and technical challenges of a research project is in place; 4) strong communications between students and faculty set clear expectations and enable mid-course corrections in the program or project design. Creating a research culture for the participants or embedding a project in an existing research culture can also assist students in

  9. Design Considerations of Help Options in Computer-Based L2 Listening Materials Informed by Participatory Design

    ERIC Educational Resources Information Center

    Cárdenas-Claros, Mónica Stella

    2015-01-01

    This paper reports on the findings of two qualitative exploratory studies that sought to investigate design features of help options in computer-based L2 listening materials. Informed by principles of participatory design, language learners, software designers, language teachers, and a computer programmer worked collaboratively in a series of…

  10. Mission-based guidance system design for autonomous UAVs

    NASA Astrophysics Data System (ADS)

    Moon, Jongki

    The advantages of UAVs in the aviation arena have led to extensive research activities on autonomous technology of UAVs to achieve specific mission objectives. This thesis mainly focuses on the development of a mission-based guidance system. Among various missions expected for future needs, autonomous formation flight (AFF) and obstacle avoidance within safe operation limits are investigated. In the design of an adaptive guidance system for AFF, the leader information except position is assumed to be unknown to a follower. Thus, the only measured information related to the leader is the line-of-sight (LOS) range and angle. Adding an adaptive element with neural networks into the guidance system provides a capability to effectively handle leader's velocity changes. Therefore, this method can be applied to the AFF control systems that use a passive sensing method. In this thesis, an adaptive velocity command guidance system and an adaptive acceleration command guidance system are developed and presented. Since relative degrees of the LOS range and angle are different depending on the outputs from the guidance system, the architecture of the guidance system changes accordingly. Simulations and flight tests are performed using the Georgia Tech UAV helicopter, the GTMax, to evaluate the proposed guidance systems. The simulation results show that the neural network (NN) based adaptive element can improve the tracking performance by effectively compensating for the effect of unknown dynamics. It has also been shown that the combination of an adaptive velocity command guidance system and the existing GTMax autopilot controller performs better than the combination of an adaptive acceleration command guidance system and the GTMax autopilot controller. The successful flight evaluation using an adaptive velocity command guidance system clearly shows that the adaptive guidance control system is a promising solution for autonomous formation flight of UAVs. In addition, an

  11. Mars Relays Satellite Orbit Design Considerations for Global Support of Robotic Surface Missions

    NASA Technical Reports Server (NTRS)

    Hastrup, Rolf; Cesarone, Robert; Cook, Richard; Knocke, Phillip; McOmber, Robert

    1993-01-01

    This paper discusses orbit design considerations for Mars relay satellite (MRS)support of globally distributed robotic surface missions. The orbit results reported in this paper are derived from studies of MRS support for two types of Mars robotic surface missions: 1) the mars Environmental Survey (MESUR) mission, which in its current definition would deploy a global network of up to 16 small landers, and 2)a Small Mars Sample Return (SMSR) mission, which included four globally distributed landers, each with a return stage and one or two rovers, and up to four additional sets of lander/rover elements in an extended mission phase.

  12. High Energy Astronomy Observatory, Mission C, Phase A. Volume 2: Preliminary analyses and conceptual design

    NASA Technical Reports Server (NTRS)

    1972-01-01

    An analysis and conceptual design of a baseline mission and spacecraft are presented. Aspects of the HEAO-C discussed include: baseline experiments with X-ray observations of space, analysis of mission requirements, observatory design, structural analysis, thermal control, attitude sensing and control system, communication and data handling, and space shuttle launch and retrieval of HEAO-C.

  13. Mission design for a ballistic slow flyby Comet Encke 1980

    NASA Technical Reports Server (NTRS)

    Farquhar, R. W.; Mccarthy, D. K.; Muhonen, D. P.; Yeomans, D. K.

    1974-01-01

    Preliminary mission analyses for a proposed 1980 slow flyby (7-9 km/s) of comet Encke are presented. Among the topics covered are science objectives, Encke's physical activity and ephemeris accuracy, trajectory and launch-window analysis, terminal guidance, and spacecraft concepts. The nominal mission plan calls for a near-perihelion intercept with two spacecraft launched on a single launch vehicle. Both spacecraft will arrive at the same time, one passing within 500 km from Encke's nucleus on its sunward side, the other cutting through the tail region. By applying a small propulsive correction about three weeks after the encounter, it is possible to retarget both spacecraft for a second Encke intercept in 1984. The potential science return from the ballistic slow flyby is compared with other proposed mission modes for the 1980 Encke flyby mission, including the widely advocated slow flyby using solar-electric propulsion. It is shown that the ballistic slow flyby is superior in every respect.

  14. NASA Now Minute: Engineering Design: Curiosity Mission to Mars

    NASA Video Gallery

    Meet Nagin Cox from the Mission Operations team for the Mars ScienceLaboratory, or Curiosity. Getting to Mars doesn’t happen by chance! Learnabout some of the scientific, technological and engi...

  15. Design of a Mars Airplane Propulsion System for the Aerial Regional-Scale Environmental Survey (ARES) Mission Concept

    NASA Technical Reports Server (NTRS)

    Kuhl, Christopher A.

    2008-01-01

    The Aerial Regional-Scale Environmental Survey (ARES) is a Mars exploration mission concept that utilizes a rocket propelled airplane to take scientific measurements of atmospheric, surface, and subsurface phenomena. The liquid rocket propulsion system design has matured through several design cycles and trade studies since the inception of the ARES concept in 2002. This paper describes the process of selecting a bipropellant system over other propulsion system options, and provides details on the rocket system design, thrusters, propellant tank and PMD design, propellant isolation, and flow control hardware. The paper also summarizes computer model results of thruster plume interactions and simulated flight performance. The airplane has a 6.25 m wingspan with a total wet mass of 185 kg and has to ability to fly over 600 km through the atmosphere of Mars with 45 kg of MMH / MON3 propellant.

  16. Preliminary Design Options for Meteor Burst Communications Systems Buoy Relays

    DTIC Science & Technology

    1986-12-01

    the lithium - thionyl chloride cell exhibit specific energies of the order of 500 watt hours per kilogram, more than 50 percent higher than previous...Supply Buoy Design Type 90 Day Storage Weight Type Energy Type Size (lb) Remote Lithium 2.2 kWh Deployable 8" x 8" x 4’ 200 Battery Pendulous Master...however, that there are various typcs of lithium batteries presently being developed that have energy densities equal to’fuel cell power systems. It is

  17. Current Status of the EJSM Jupiter Europa Orbiter Flagship Mission Design

    NASA Astrophysics Data System (ADS)

    Clark, K.; Pappalardo, R.; Greeley, R.; Hendrix, A.; Boldt, J.; van Houten, T.; Jun, I.; Lock, R.; Ludwinski, J.; Rasmussen, R.; Tan-Wang, G.

    2008-12-01

    NASA and ESA have embarked on a joint study of a mission to Europa and the Jupiter system with orbiters developed by NASA, ESA, and possibly JAXA. An international Joint Jupiter Science Definition Team (JJSDT) is defining the science content for the Jupiter Europa Orbiter (JEO) mission study run by NASA and for the Jupiter Ganymede Orbiter (JGO) mission study run by ESA. Engineering teams for both missions are working closely with the JJSDT to define mission concepts that optimize science, cost, and risk. The NASA-led JEO mission addresses a scientifically rich subset of the complete EJSM science goals and is designed to stand alone or in conjunction with the ESA-led JGO. This paper focuses on the NASA-led JEO mission and will describe the concept in the context of a standalone mission. An orbital mission to Europa is driven by the desire to investigate an astrobiological archetype for icy satellite habitability, with a putative warm, salty, water ocean with plausible energy sources. Additionally, JEO will explore the Jupiter system to better understand how Europa's possible habitability is related to the formation scenario of the other Jovian satellites. The JEO mission will perform 2.5-3 years of Jupiter system science, including encounters with Io, Ganymede and Callisto, before insertion into orbit around Europa for a comprehensive set of science campaigns lasting for up to one year. This paper will highlight the JEO mission design and implementation concept. The work reported was sponsored by the National Aeronautics and Space Administration.

  18. Current Status of the EJSM Jupiter Europa Orbiter: Mission Design and Architecture

    NASA Astrophysics Data System (ADS)

    Grunthaner, Paula; Clark, K.; Pappalardo, R.; Greeley, R.; Hendrix, A.; Boldt, J.; Van Houten, T.; Jun, I.; Lock, R.; Ludwinski, J.; Rasmussen, R.; Tan-Wang, G.

    2008-09-01

    NASA and ESA have embarked on a joint study of a mission to Europa and the Jupiter system with orbiters developed by NASA, ESA, and possibly JAXA. An international Jupiter Joint Science Definition Team (JJSDT) is defining the science content for the Europa Orbiter (JEO) mission study run by NASA and for the Jupiter Ganymede Orbiter (JGO) mission study run by ESA. Engineering teams for both missions are working closely with the JJSDT to define mission concepts that optimize science, cost, and risk. The NASA-led JEO mission addresses a scientifically rich subset of the complete EJSM science goals and is designed to stand alone or in conjunction with the ESA-led JGO. This paper focuses on the NASA-led JEO mission and will describe the concept in the context of a standalone mission. An orbital mission to Europa is driven by the desire to investigate an astrobiological archetype for icy satellite habitability, with a warm, salty, water ocean with plausible energy sources. Additionally, JEO will explore the Jupiter system to better understand how Europa's possible habitability is related to the formation scenario of the other Jovian satellites. The JEO mission will perform 2.5 to 3 years of Jupiter system science, including encounters with Io, Ganymede and Callisto, before insertion into orbit around Europa for a comprehensive set of science campaigns lasting for up to one year. This paper will highlight the JEO mission design and implementation concept. The work reported was sponsored by the National Aeronautics and Space Administration.

  19. From neighborhood design and food options to residents' weight status.

    PubMed

    Cerin, Ester; Frank, Lawrence D; Sallis, James F; Saelens, Brian E; Conway, Terry L; Chapman, James E; Glanz, Karen

    2011-06-01

    This study examined associations of accessibility, availability, price, and quality of food choices and neighborhood urban design with weight status and utilitarian walking. To account for self-selection bias, data on adult residents of a middle-to-high-income neighborhood were used. Participants kept a 2-day activity/travel diary and self-reported socio-demographics, height, and weight. Geographic Information Systems data were used to objectively quantify walking-related aspects of urban design, and number of and distance to food outlets within respondents' 1km residential buffers. Food outlets were audited for availability, price, and quality of healthful food choices. Number of convenience stores and in-store healthful food choices were positively related to walking for errands which, in turn, was predictive of lower risk of being overweight/obese. Negative associations with overweight/obesity unexplained by walking were found for number of grocery stores and healthful food choices in sit-down restaurants. Aspects of urban form and food environment were associated with walking for eating purposes which, however, was not predictive of overweight/obesity. Access to diverse destinations, food outlets and healthful food choices may promote pedestrian activity and contribute to better weight regulation. Accessibility and availability of healthful food choices may lower the risk of overweight/obesity by providing opportunities for healthier dietary patterns.

  20. On options for interdisciplinary analysis and design optimization

    NASA Technical Reports Server (NTRS)

    Haftka, R. T.; Sobieszczanski-Sobieski, J.; Padula, S. L.

    1992-01-01

    The interdisciplinary optimization of engineering systems is discussed from the standpoint of the computational alternatives available to the designer. The analysis of such systems typically requires the solution of coupled systems of nonlinear algebraic equations. The solution procedure is necessarily iterative in nature. It is shown that the system can be solved by fixed point iteration, by Newton's method, or by a combination of the two. However, the need for sensitivity analysis may affect the choice of analysis solution method. Similarly, the optimization of the system can be formulated in several ways that are discussed in the paper. It is shown that the effect of the topology of the interaction between disciplines is a key factor in the choice of analysis, sensitivity and optimization methods. Several examples are presented to illustrate the discussion.

  1. Optical design options for hypertelescopes and prototype testing

    NASA Astrophysics Data System (ADS)

    Bondoux, Erick; Bosio, Sandra; Chakraborty, Rijuparna; Dali-Ali, Wassila; Labeyrie, Antoine; Lacamp, Bruno; Maillot, Jerome; Mourard, Denis; Nunez, Paul D.; Pijoan, Jordi; Prudhomme, Rémi; Riaud, Pierre; Roussel, Martine; Surya, Arun; Tregon, Bernard; Houllier, Thomas; Lepine, Thierry; Rabou, Patrick; Rondi, André; Bresson, Yves; Vernet, David

    2016-08-01

    Hypertelescopes are large optical interferometric arrays, employing many small mirrors and a miniature pupildensifier before the focal camera, expected to produce direct images of celestial sources at high resolution. Their peculiar imaging properties, initially explored through analytical derivations, had been verified with simulations before testing a full-size testbed instrument. We describe several architectures and optical design solutions and present recent progress made on the Ubaye hypertelescope experiment. Arecibo-like versions with a fixed spherical primary meta-mirror, or an active aspheric one, have a suspended focal beam combiner equipped for pupil-drift accommodation, with a field-mosaic arrangement for observing multiple sources such as exoplanetary systems, globular clusters or active galactic nuclei. We have developed a cable suspension and drive system with tracking accuracy reaching a millimeter at 100m above ground.

  2. The Integrated Mission Design Center (IMDC) at NASA Goddard Space Flight Center

    NASA Technical Reports Server (NTRS)

    Karpati, Gabriel; Martin, John; Steiner, Mark; Reinhardt, K.

    2002-01-01

    NASA Goddard has used its Integrated Mission Design Center (IMDC) to perform more than 150 mission concept studies. The IMDC performs rapid development of high-level, end-to-end mission concepts, typically in just 4 days. The approach to the studies varies, depending on whether the proposed mission is near-future using existing technology, mid-future using new technology being actively developed, or far-future using technology which may not yet be clearly defined. The emphasis and level of detail developed during any particular study depends on which timeframe (near-, mid-, or far-future) is involved and the specific needs of the study client. The most effective mission studies are those where mission capabilities required and emerging technology developments can synergistically work together; thus both enhancing mission capabilities and providing impetus for ongoing technology development.

  3. Options and obstacles for designing a universal influenza vaccine.

    PubMed

    Jang, Yo Han; Seong, Baik Lin

    2014-08-18

    Since the discovery of antibodies specific to a highly conserved stalk region of the influenza virus hemagglutinin (HA), eliciting such antibodies has been considered the key to developing a universal influenza vaccine that confers broad-spectrum protection against various influenza subtypes. To achieve this goal, a prime/boost immunization strategy has been heralded to redirect host immune responses from the variable globular head domain to the conserved stalk domain of HA. While this approach has been successful in eliciting cross-reactive antibodies against the HA stalk domain, protective efficacy remains relatively poor due to the low immunogenicity of the domain, and the cross-reactivity was only within the same group, rather than among different groups. Additionally, concerns are raised on the possibility of vaccine-associated enhancement of viral infection and whether multiple boost immunization protocols would be considered practical from a clinical standpoint. Live attenuated vaccine hitherto remains unexplored, but is expected to serve as an alternative approach, considering its superior cross-reactivity. This review summarizes recent advancements in the HA stalk-based universal influenza vaccines, discusses the pros and cons of these approaches with respect to the potentially beneficial and harmful effects of neutralizing and non-neutralizing antibodies, and suggests future guidelines towards the design of a truly protective universal influenza vaccine.

  4. Magnetospheric Multiscale (MMS) Mission Attitude Ground System Design

    NASA Technical Reports Server (NTRS)

    Sedlak, Joseph E.; Superfin, Emil; Raymond, Juan C.

    2011-01-01

    This paper presents an overview of the attitude ground system (AGS) currently under development for the Magnetospheric Multiscale (MMS) mission. The primary responsibilities for the MMS AGS are definitive attitude determination, validation of the onboard attitude filter, and computation of certain parameters needed to improve maneuver performance. For these purposes, the ground support utilities include attitude and rate estimation for validation of the onboard estimates, sensor calibration, inertia tensor calibration, accelerometer bias estimation, center of mass estimation, and production of a definitive attitude history for use by the science teams. Much of the AGS functionality already exists in utilities used at NASA's Goddard Space Flight Center with support heritage from many other missions, but new utilities are being created specifically for the MMS mission, such as for the inertia tensor, accelerometer bias, and center of mass estimation. Algorithms and test results for all the major AGS subsystems are presented here.

  5. Leveraging Knowledge: Impact on Low Cost Planetary Mission Design.

    ERIC Educational Resources Information Center

    Momjian, Jennifer

    This paper discusses innovations developed by the Jet Propulsion Laboratory (JPL) librarians to reduce the information query cycle time for teams planning low-cost, planetary missions. The first section provides background on JPL and its library. The second section addresses the virtual information environment, including issues of access, content,…

  6. Trajectory Design for the Europa Clipper Mission Concept

    NASA Technical Reports Server (NTRS)

    Buffington, Brent

    2014-01-01

    Europa is one of the most scientifically intriguing targets in planetary science due to its potential suitability for extant life. As such, NASA has funded the California Institute of Technology Jet Propulsion Laboratory and the Johns Hopkins University Applied Physics Laboratory to jointly determine and develop the best mission concept to explore Europa in the near future. The result of nearly 4 years of work--the Europa Clipper mission concept--is a multiple Europa flyby mission that could efficiently execute a number of high caliber science investigations to meet Europa science priorities specified in the 2011 NRC Decadal Survey, and is capable of providing reconnaissance data to maximize the probability of both a safe landing and access to surface material of high scientific value for a future Europa lander. This paper will focus on the major enabling component for this mission concept--the trajectory. A representative trajectory, referred to as 13F7-A21, would obtain global-regional coverage of Europa via a complex network of 45 flybys over the course of 3.5 years while also mitigating the effects of the harsh Jovian radiation environment. In addition, 5 Ganymede and 9 Callisto flybys would be used to manipulate the trajectory relative to Europa. The tour would reach a maximum Jovicentric inclination of 20.1 deg. have a deterministic (Delta)V of 164 m/s (post periapsis raise maneuver), and a total ionizing dose of 2.8 Mrad (Si).

  7. Astronomy sortie missions definition study. Volume 3, book 1: Design analysis and trade studies

    NASA Technical Reports Server (NTRS)

    1972-01-01

    A study to define the astronomy sortie missions was conducted. The design analyses and tradeoff studies conducted for candidate concepts are presented. The subjects discussed are: (1) system and subsystem requirements, (2) space shuttle interfaces, (3) infrared telescope development, and (4) experiments to be conducted during the mission.

  8. Preliminary design polymeric materials experiment. [for space shuttles and Spacelab missions

    NASA Technical Reports Server (NTRS)

    Mattingly, S. G.; Rude, E. T.; Marshner, R. L.

    1975-01-01

    A typical Advanced Technology Laboratory mission flight plan was developed and used as a guideline for the identification of a number of experiment considerations. The experiment logistics beginning with sample preparation and ending with sample analysis are then overlaid on the mission in order to have a complete picture of the design requirements. The results of this preliminary design study fall into two categories. First specific preliminary designs of experiment hardware which is adaptable to a variety of mission requirements. Second, identification of those mission considerations which affect hardware design and will require further definition prior to final design. Finally, a program plan is presented which will provide the necessary experiment hardware in a realistic time period to match the planned shuttle flights. A bibliography of all material reviewed and consulted but not specifically referenced is provided.

  9. Overview: Solar Electric Propulsion Concept Designs for SEP Technology Demonstration Mission

    NASA Technical Reports Server (NTRS)

    Mcguire, Melissa L.; Hack, Kurt J.; Manzella, David; Herman, Daniel

    2014-01-01

    JPC presentation of the Concept designs for NASA Solar Electric Propulsion Technology Demonstration mission paper. Multiple Solar Electric Propulsion Technology Demonstration Missions were developed to assess vehicle performance and estimated mission cost. Concepts ranged from a 10,000 kg spacecraft capable of delivering 4000 kg of payload to one of the Earth Moon Lagrange points in support of future human-crewed outposts to a 180 kg spacecraft capable of performing an asteroid rendezvous mission after launched to a geostationary transfer orbit as a secondary payload.

  10. A study of space station needs, attributes and architectural options. Volume 2: Technical. Book 1: Mission requirements. Appendixes 1 and 2

    NASA Technical Reports Server (NTRS)

    1983-01-01

    The space station mission requirements data base consists of 149 attached and free-flying missions each of which is documented by a set of three interrelated documents: (1) NASA LaRC Data Sheets - with three sheets comprising a set for each payload element described. These sheets contain user payload element data necessary to drive Space Station architectural options. (2) GDC-derived operations descriptions that supplement the LaRC payload element data in the operations areas such as further descriptions of crew involvement, EVA, etc. (3) Payload elements synthesis sheets used by GDC to provide requirements traceability to data sources and to provide a narrative describing the basis for formulating the payload element requirements.

  11. The Gevaltig: An inertial fusion powered manned spacecraft design for outer solar system missions

    SciTech Connect

    Murray, K.A.

    1989-10-01

    The Gevaltig is an inertial fusion powered rocket engine capable of manned missions to other planets with round trip mission times as low as 100 days. The Gevaltig design was previously described for a mission to Mars. This effort defines the spacecraft components in terms of mass and presents a mission analysis for a manned trip to Titan, a moon of Saturn. The Gevaltig component masses are provided as a function of fuel pellet ignition frequency. These variable mass components include the fuel tanks, radiators, structure and EM pumps. Fixed mass components include the drivers, coil, coil shield, power processing system, payload, crew shield and laser mirrors. A 6 MW nuclear reactor is included in the design for startup purposes. Various combinations of thrust, mission duration and specific impulse were evaluated to determine a reasonable mission scenario for the Titan mission. The mission analysis yielded several viable mission scenarios, with round trip durations of 370 to 500 days and initial (launch) masses from lunar orbit of 2500 to 20,000 metric tons. 15 refs., 13 figs., 14 tabs.

  12. Conceptual Design of a Communications Relay Satellite for a Lunar Sample Return Mission

    NASA Technical Reports Server (NTRS)

    Brunner, Christopher W.

    2005-01-01

    In 2003, NASA solicited proposals for a robotic exploration of the lunar surface. Submissions were requested for a lunar sample return mission from the South Pole-Aitken Basin. The basin is of interest because it is thought to contain some of the oldest accessible rocks on the lunar surface. A mission is under study that will land a spacecraft in the basin, collect a sample of rock fragments, and return the sample to Earth. Because the Aitken Basin is on the far side of the Moon, the lander will require a communications relay satellite (CRS) to maintain contact with the Earth during its surface operation. Design of the CRS's orbit is therefore critical. This paper describes a mission design which includes potential transfer and mission orbits, required changes in velocity, orbital parameters, and mission dates. Several different low lunar polar orbits are examined to compare their availability to the lander versus the distance over which they must communicate. In addition, polar orbits are compared to a halo orbit about the Earth-Moon L2 point, which would permit continuous communication at a cost of increased fuel requirements and longer transmission distances. This thesis also examines some general parameters of the spacecraft systems for the mission under study. Mission requirements for the lander dictate the eventual choice of mission orbit. This mission could be the first step in a period of renewed lunar exploration and eventual human landings.

  13. [Interior] Configuration options, habitability and architectural aspects of the transfer habitat module (THM) and the surface habitat on Mars (SHM)/ESA's AURORA human mission to Mars (HMM) study

    NASA Astrophysics Data System (ADS)

    Imhof, Barbara

    2007-02-01

    This paper discusses the findings for [Interior] configuration options, habitability and architectural aspects of a first human spacecraft to Mars. In 2003 the space architecture office LIQUIFER was invited by the European Space Agency's (ESA) AURORA Program committee to consult the scientists and engineers from the European Space and Technology Center (ESTEC) and other European industrial communities with developing the first human mission to Mars, which will take place in 2030, regarding the architectural issues of crewed habitats. The task was to develop an interior configuration for a transfer vehicle (TV) to Mars, especially a transfer habitation module (THM) and a surface habitat module (SHM) on Mars. The total travel time Earth—Mars and back for a crew of six amounts to approximately 900 days. After a 200-day-flight three crewmembers will land on Mars in the Mars excursion vehicle (MEV) and will live and work in the SHM for 30 days. For 500 days before the 200-day journey back the spacecraft continues to circle the Martian orbit for further exploration. The entire mission program is based on our present knowledge of technology. The project was compiled during a constant feedback-design process and trans-disciplinary collaboration sessions in the ESA-ESTEC concurrent design facility. Long-term human space flight sets new spatial conditions and requirements to the design concept. The guidelines were developed from relevant numbers and facts of recognized standards, interviews with astronauts/cosmonauts and from analyses about habitability, sociology, psychology and configuration concepts of earlier space stations in combination with the topics of the individual's perception and relation of space. Result of this study is the development of a prototype concept for the THM and SHM with detailed information and complete plans of the interior configuration, including mass calculations. In addition the study contains a detailed explanation of the development of

  14. 17 CFR 33.4 - Designation as a contract market for the trading of commodity options.

    Code of Federal Regulations, 2012 CFR

    2012-04-01

    ... 17 Commodity and Securities Exchanges 1 2012-04-01 2012-04-01 false Designation as a contract market for the trading of commodity options. 33.4 Section 33.4 Commodity and Securities Exchanges... value in normal cash marketing channels; (B) There exists an accurate and widely-disseminated...

  15. Interaction-Region Design Options for a Linac-Ring LHeC

    SciTech Connect

    Zimmermann, Frank; Bettoni, Simona; Bruning, Oliver; Holzer, Bernhard; Russenschuck, Stephan; Schulte, Daniel; Tomas, Rogelio; Aksakal, Husnu; Appleby, Robert; Chattopadhyay, Swapan; Korostelev, Maxim; Ciftci, Abbas; Ciftci, Rena; Zengin, Kahraman; Dainton, John; Klein, Max; Eroglu, Emre; Tapan, Ilhan; Kostka, Peter; Litvinenko, Vladimir; Paoloni, Eugenio; /INFN, Pisa /INFN, Bologna /DESY /SLAC

    2012-06-21

    The interaction-region design for a linac-ring electron-proton collider based on the LHC ('LR-LHeC') poses numerous challenges related to collision scheme, synchrotron radiation, aperture, magnet technology, and optics. We report a first assessment and various options.

  16. Design of RF systems for the RTD mission VASIMR

    SciTech Connect

    Baity, F. W.; Barber, G. C.; Carter, M. D.; Goulding, R. H.; Sparks, D. O.; Chang-Diaz, F. R.; McCaskill, G. E.; Squire, J. P.

    1999-09-20

    The first flight test of the variable specific impulse magnetoplasma rocket (VASIMR) is tentatively scheduled for the Radiation and Technology Demonstration (RTD) in 2003. This mission to map the radiation environment out to several earth radii will employ both a Hall thruster and a VASIMR during its six months duration, beginning from low earth orbit. The mission will be powered by a solar array providing 12 kW of direct current electricity at 50 V. The VASIMR utilizes radiofrequency (RF) power both to generate a high-density plasma in a helicon source and to accelerate the plasma ions to high velocity by ion cyclotron resonance heating (ICRH). The VASIMR concept is being developed by the National Aeronautics and Space Administration (NASA) in collaboration with national laboratories and universities. Prototype plasma sources, RF amplifiers, and antennas are being developed in the experimental facilities of the Advanced Space Propulsion Laboratory (ASPL). (c) 1999 American Institute of Physics.

  17. Design of RF Systems for the RTD Mission VASIMR

    SciTech Connect

    Baity, F.W.; Barber, G.C.; Carter, M.D.; Chang-Diaz, F.R.; Goulding, R.H.; McCaskill, G.E.; Sparks, D.O.; Squire, J.P.

    1999-04-12

    The first flight test of the variable specific impulse magnetoplasma rocket (VASIMR) is tentatively scheduled for the Radiation and Technology Demonstration (RTD) in 2003. This mission to map the radiation environment out to several earth radii will employ both a Hall thruster and a VASIMR during its six months duration, beginning from low earth orbit. The mission will be powered by a solar array providing 12 kW of direct current electricity at 50 V. The VASIMR utilizes radiofrequency (RF) power both to generate a high-density plasma in a helicon source and to accelerate the plasma ions to high velocity by ion cyclotron resonance heating (ICRH). The VASIMR concept is being developed by the National Aeronautics and Space Administration (NASA) in collaboration with national laboratories and universities. Prototype plasma sources, RF amplifiers, and antennas are being developed in the experimental facilities of the Advanced Space Propulsion Laboratory (ASPL).

  18. Calipso's Mission Design: Sun-Glint Avoidance Strategies

    NASA Technical Reports Server (NTRS)

    Mailhe, Laurie M.; Schiff, Conrad; Stadler, John H.

    2004-01-01

    CALIPSO will fly in formation with the Aqua spacecraft to obtain a coincident image of a portion of the Aqua/MODIS swath. Since MODIS pixels suffering sun-glint degradation are not processed, it is essential that CALIPSO only co- image the glint h e portion of the MODIS instrument swath. This paper presents sun-glint avoidance strategies for the CALIPSO mission. First, we introduce the Aqua sun-glint geometry and its relation to the CALIPSO-Aqua formation flying parameters. Then, we detail our implementation of the computation and perform a cross-track trade-space analysis. Finally, we analyze the impact of the sun-glint avoidance strategy on the spacecraft power and delta-V budget over the mission lifetime.

  19. Aircraft design for mission performance using nonlinear multiobjective optimization methods

    NASA Technical Reports Server (NTRS)

    Dovi, Augustine R.; Wrenn, Gregory A.

    1990-01-01

    A new technique which converts a constrained optimization problem to an unconstrained one where conflicting figures of merit may be simultaneously considered was combined with a complex mission analysis system. The method is compared with existing single and multiobjective optimization methods. A primary benefit from this new method for multiobjective optimization is the elimination of separate optimizations for each objective, which is required by some optimization methods. A typical wide body transport aircraft is used for the comparative studies.

  20. Design and Analysis of a Formation Flying System for the Cross-Scale Mission Concept

    NASA Technical Reports Server (NTRS)

    Cornara, Stefania; Bastante, Juan C.; Jubineau, Franck

    2007-01-01

    The ESA-funded "Cross-Scale Technology Reference Study has been carried out with the primary aim to identify and analyse a mission concept for the investigation of fundamental space plasma processes that involve dynamical non-linear coupling across multiple length scales. To fulfill this scientific mission goal, a constellation of spacecraft is required, flying in loose formations around the Earth and sampling three characteristic plasma scale distances simultaneously, with at least two satellites per scale: electron kinetic (10 km), ion kinetic (100-2000 km), magnetospheric fluid (3000-15000 km). The key Cross-Scale mission drivers identified are the number of S/C, the space segment configuration, the reference orbit design, the transfer and deployment strategy, the inter-satellite localization and synchronization process and the mission operations. This paper presents a comprehensive overview of the mission design and analysis for the Cross-Scale concept and outlines a technically feasible mission architecture for a multi-dimensional investigation of space plasma phenomena. The main effort has been devoted to apply a thorough mission-level trade-off approach and to accomplish an exhaustive analysis, so as to allow the characterization of a wide range of mission requirements and design solutions.

  1. Design Concepts for a Small Space-Based GEO Relay Satellite for Missions Between Low Earth and near Earth Orbits

    NASA Technical Reports Server (NTRS)

    Bhasin, Kul B.; Warner, Joseph D.; Oleson, Steven; Schier, James

    2014-01-01

    The main purpose of the Small Space-Based Geosynchronous Earth orbiting (GEO) satellite is to provide a space link to the user mission spacecraft for relaying data through ground networks to user Mission Control Centers. The Small Space Based Satellite (SSBS) will provide services comparable to those of a NASA Tracking Data Relay Satellite (TDRS) for the same type of links. The SSBS services will keep the user burden the same or lower than for TDRS and will support the same or higher data rates than those currently supported by TDRS. At present, TDRSS provides links and coverage below GEO; however, SSBS links and coverage capability to above GEO missions are being considered for the future, especially for Human Space Flight Missions (HSF). There is also a rising need for the capability to support high data rate links (exceeding 1 Gbps) for imaging applications. The communication payload on the SSBS will provide S/Ka-band single access links to the mission and a Ku-band link to the ground, with an optical communication payload as an option. To design the communication payload, various link budgets were analyzed and many possible operational scenarios examined. To reduce user burden, using a larger-sized antenna than is currently in use by TDRS was considered. Because of the SSBS design size, it was found that a SpaceX Falcon 9 rocket could deliver three SSBSs to GEO. This will greatly reduce the launch costs per satellite. Using electric propulsion was also evaluated versus using chemical propulsion; the power system size and time to orbit for various power systems were also considered. This paper will describe how the SSBS will meet future service requirements, concept of operations, and the design to meet NASA users' needs for below and above GEO missions. These users' needs not only address the observational mission requirements but also possible HSF missions to the year 2030. We will provide the trade-off analysis of the communication payload design in terms of

  2. Conceptual Design of an Electric Sail Technology Demonstration Mission Spacecraft

    NASA Technical Reports Server (NTRS)

    Wiegmann, Bruce M.

    2017-01-01

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

  3. Multi-Objective Hybrid Optimal Control for Multiple-Flyby Low-Thrust Mission Design

    NASA Technical Reports Server (NTRS)

    Englander, Jacob A.; Vavrina, Matthew A.; Ghosh, Alexander R.

    2015-01-01

    Preliminary design of low-thrust interplanetary missions is a highly complex process. The mission designer must choose discrete parameters such as the number of flybys, the bodies at which those flybys are performed, and in some cases the final destination. In addition, a time-history of control variables must be chosen that defines the trajectory. There are often many thousands, if not millions, of possible trajectories to be evaluated. The customer who commissions a trajectory design is not usually interested in a point solution, but rather the exploration of the trade space of trajectories between several different objective functions. This can be a very expensive process in terms of the number of human analyst hours required. An automated approach is therefore very desirable. This work presents such an approach by posing the mission design problem as a multi-objective hybrid optimal control problem. The method is demonstrated on a hypothetical mission to the main asteroid belt.

  4. Rapid Preliminary Design of Interplanetary Trajectories Using the Evolutionary Mission Trajectory Generator

    NASA Technical Reports Server (NTRS)

    Englander, Jacob

    2016-01-01

    Preliminary design of interplanetary missions is a highly complex process. The mission designer must choose discrete parameters such as the number of flybys, the bodies at which those flybys are performed, and in some cases the final destination. In addition, a time-history of control variables must be chosen that defines the trajectory. There are often many thousands, if not millions, of possible trajectories to be evaluated. This can be a very expensive process in terms of the number of human analyst hours required. An automated approach is therefore very desirable. This work presents such an approach by posing the mission design problem as a hybrid optimal control problem. The method is demonstrated on notional high-thrust chemical and low-thrust electric propulsion missions. In the low-thrust case, the hybrid optimal control problem is augmented to include systems design optimization.

  5. Advanced methods of low cost mission design for Jovian moons exploration

    NASA Astrophysics Data System (ADS)

    Grushevskii, Alexey; Koryanov, Victor; Tuchin, Andrey; Golubev, Yury; Tuchin, Denis

    2016-07-01

    DeltaV-low-cost gravity assists tours mission design of for the Jovian Moons exploration is considered (orbiters and probes around Io, Europa, Ganymede, Callisto), taking radiation hazard into account. Limited dynamic opportunities of using flybys require multiple gravity assists. Relevance of regular creation of optimum scenarios - sequences of passing of celestial bodies with definition of conditions of their execution is obvious. This work is devoted to the description of criteria for creation of such chains. New Multi-Tisserand coordinates [1,2] for this purpose are introduced for the best study of features for the radiation hazard decrease and the spacecraft asymptotic velocity reduction. One of main problems of the Jovian system mission design (JIMO, JUICE, Laplas P) is that the reduction of the asymptotic velocity of the spacecraft with respect to the satellite for the Jovian moon's capture is impossible. A valid reason is in the invariance of Jacobi integral and Tisserand parameter in a restricted three-body model (RTBP) [3]. Furthermore, the same-body flybys tour falls into the hazard radiation zone according the Tisserand-Poincaré graph. Formalized beam's algorithm to overcome this "problem of the ballistic destiny" with using full ephemeris model and with several coupled RTBP engaging has been implemented. Withal low-cost reduction of the spacecraft asymptotic velocity for the capture of the moon is required. The corresponding numerical scheme was developed with using Tisserand-Poincaré graph and the simulation of tens of millions of options. The Delta V-low cost searching was utilized also with help of the modeling of the multiple rebounds (cross gravity assists) of the beam of trajectories. The techniques are developed by the authors specifically to the needs of the mission "Laplas P" of Roscosmos. If we have answers to the questions "what kind of gravity assists", we need answer on the question "when". New Multi-Tisserand coordinates for this

  6. Conceptual spacecraft and arcjet propulsion system design for the SP-100 interim reference mission

    NASA Technical Reports Server (NTRS)

    Zafran, S.; Bell, M. W. J.

    1990-01-01

    An arcjet propulsion system, delivering 7.5 N thrust, was defined for the SP-100 Space Reactor Power System Interim Reference Mission. Conceptual design trades and configuration studies of a spacecraft suitable for the mission were performed to the extent necessary to define propulsion system requirements and interfaces. The propulsion system design is based on the use of 30-kW, constricted arc, ammonia arcjet engines operating in parallel during orbit boost from low earth to geosynchronous orbit.

  7. An analytic model for footprint dispersions and its application to mission design

    NASA Technical Reports Server (NTRS)

    Rao, J. R. Jagannatha; Chen, Yi-Chao

    1992-01-01

    This is the final report on our recent research activities that are complementary to those conducted by our colleagues, Professor Farrokh Mistree and students, in the context of the Taguchi method. We have studied the mathematical model that forms the basis of the Simulation and Optimization of Rocket Trajectories (SORT) program and developed an analytic method for determining mission reliability with a reduced number of flight simulations. This method can be incorporated in a design algorithm to mathematically optimize different performance measures of a mission, thus leading to a robust and easy-to-use methodology for mission planning and design.

  8. Shock-Tolerant Low-Power Generator Design for Landed Missions

    NASA Astrophysics Data System (ADS)

    Gelderloos, Carl J.; Decino, Jim; Lock, Jennifer; Miller, Dan D.; Taylor, Robert

    2004-02-01

    A shock-tolerant thermal enclosure has been designed for use in distributed landed missions. Missions such as Pascal and the Mars Long-Lived Landed Network require low power sources capable of surviving an omnidirectional load at impact and delivering reliable power for several Martian years. With the use of a radioisotope heat source and a thermoelectric converter, power can be generated reliably, but the challenge of developing an insulating canister that delivers sufficient power at end of life and is shock tolerant has been elusive. We describe a manufacturable design using conventional materials that meets mission requirements and show preliminary analysis of impact load response.

  9. Shuttle/IUS performance for planetary missions. [Interim Upper Stage

    NASA Technical Reports Server (NTRS)

    Cork, M. J.; Driver, J. M.; Wright, J. L.

    1975-01-01

    Potential requirements for planetary missions in the 1980s, capabilities of the Interim Upper Stage (IUS) candidates to perform those missions, and Shuttle/IUS mission profile options for performance enhancement are examined. The most demanding planetary missions are the Pioneer Saturn/Uranus/Titan Probe and the Mariner-class orbiters of Mercury, Jupiter, and Saturn. Options available to designers of these missions will depend on the specific IUS selected for development and the programmatic phasing of the IUS and the NASA Tug. Use of Shuttle elliptic orbits as initial conditions for IUS ignition offers significant performance improvements; specific values are mission dependent.

  10. A study of space station needs, attributes, and architectural options, volume 2, technical. Book 2: Mission implementation concepts

    NASA Technical Reports Server (NTRS)

    1983-01-01

    Space station systems characteristics and architecture are described. A manned space station operational analysis is performed to determine crew size, crew task complexity and time tables, and crew equipment to support the definition of systems and subsystems concepts. This analysis is used to select and evaluate the architectural options for development.

  11. Jupiter Orbiter and Probe project - Synthesis of the mission design

    NASA Technical Reports Server (NTRS)

    Beckman, J. C.; Roberts, P. H., Jr.

    1977-01-01

    The Jupiter Orbiter Probe (JOP), scheduled to be launched by the Shuttle IUS in 1981, is described in terms of its scientific mission objectives. These include: analysis of the chemical composition and physical state of Jupiter's atmosphere, the chemical composition and physical state of Ganymede and Callisto, and the topology and behavior of the magnetic field and energetic particle fluxes. Attention is given to an atmospheric probe which will be launched from the main probe, and to the navigation requirements necessary to 'bounce' the JOP around the Jovian moons.

  12. Probabilistic Risk Assessment for Concurrent, Conceptual Design of Space Missions

    NASA Technical Reports Server (NTRS)

    Meshkat, Leila

    2005-01-01

    NASA is expanding its capability to perform PRA. This capability gives insight into the links of a suggested design and drives the refinement of the design by identifying optimal areas for investments. Clearly, it is more viable and less expensive to refine a design at the time that it is being conceived. Hence the utility of conducting PRA at the conceptual design phase. Concurrent engineering teams greatly reduce the design time and costs. However, there is currently no standardized means for building probabilistic risk models to assess risks associated with a design produced by such teams. The capability to produce a consistent and valid risk metric associated with such designs would greatly enhance the value of such design teams. This paper explains the experimental results obtained to date from building probabilistic risk models for sample studies conducted at the concurrent engineering design team at the Jet Propulsion Laboratory (TeamX).

  13. The mechanical design of an imaging photopolarimeter for the Jupiter missions (Pioneer 10 and 11)

    NASA Technical Reports Server (NTRS)

    Kodak, J. C.

    1975-01-01

    The mechanical design and fabrication are discussed of the imaging photopolarimeter (IPP), a multifunction space-qualified instrument used on the Jupiter Pioneer missions. The extreme environmental requirements for the structural design, optical system, and mechanisms are described with detailed discussion of some of the design and fabrication problems encountered.

  14. Design of a Slowed-Rotor Compound Helicopter for Future Joint Service Missions

    NASA Technical Reports Server (NTRS)

    Silva, Christopher; Yeo, Hyeonsoo; Johnson, Wayne R.

    2010-01-01

    A slowed-rotor compound helicopter has been synthesized using the NASA Design and Analysis of Rotorcraft (NDARC) conceptual design software. An overview of the design process and the capabilities of NDARC are presented. The benefits of trading rotor speed, wing-rotor lift share, and trim strategies are presented for an example set of sizing conditions and missions.

  15. Planetary mission requirements, technology and design considerations for a solar electric propulsion stage

    NASA Technical Reports Server (NTRS)

    Cork, M. J.; Hastrup, R. C.; Menard, W. A.; Olson, R. N.

    1979-01-01

    High energy planetary missions such as comet rendezvous, Saturn orbiter and asteroid rendezvous require development of a Solar Electric Propulsion Stage (SEPS) for augmentation of the Shuttle-IUS. Performance and functional requirements placed on the SEPS are presented. These requirements will be used in evolution of the SEPS design, which must be highly interactive with both the spacecraft and the mission design. Previous design studies have identified critical SEPS technology areas and some specific design solutions which are also presented in the paper.

  16. Monte Carlo Analysis as a Trajectory Design Driver for the TESS Mission

    NASA Technical Reports Server (NTRS)

    Nickel, Craig; Lebois, Ryan; Lutz, Stephen; Dichmann, Donald; Parker, Joel

    2016-01-01

    The Transiting Exoplanet Survey Satellite (TESS) will be injected into a highly eccentric Earth orbit and fly 3.5 phasing loops followed by a lunar flyby to enter a mission orbit with lunar 2:1 resonance. Through the phasing loops and mission orbit, the trajectory is significantly affected by lunar and solar gravity. We have developed a trajectory design to achieve the mission orbit and meet mission constraints, including eclipse avoidance and a 30-year geostationary orbit avoidance requirement. A parallelized Monte Carlo simulation was performed to validate the trajectory after injecting common perturbations, including launch dispersions, orbit determination errors, and maneuver execution errors. The Monte Carlo analysis helped identify mission risks and is used in the trajectory selection process.

  17. Monte Carlo Analysis as a Trajectory Design Driver for the Transiting Exoplanet Survey Satellite (TESS) Mission

    NASA Technical Reports Server (NTRS)

    Nickel, Craig; Parker, Joel; Dichmann, Don; Lebois, Ryan; Lutz, Stephen

    2016-01-01

    The Transiting Exoplanet Survey Satellite (TESS) will be injected into a highly eccentric Earth orbit and fly 3.5 phasing loops followed by a lunar flyby to enter a mission orbit with lunar 2:1 resonance. Through the phasing loops and mission orbit, the trajectory is significantly affected by lunar and solar gravity. We have developed a trajectory design to achieve the mission orbit and meet mission constraints, including eclipse avoidance and a 30-year geostationary orbit avoidance requirement. A parallelized Monte Carlo simulation was performed to validate the trajectory after injecting common perturbations, including launch dispersions, orbit determination errors, and maneuver execution errors. The Monte Carlo analysis helped identify mission risks and is used in the trajectory selection process.

  18. Decoding the "CoDe": A Framework for Conceptualizing and Designing Help Options in Computer-Based Second Language Listening

    ERIC Educational Resources Information Center

    Cardenas-Claros, Monica Stella; Gruba, Paul A.

    2013-01-01

    This paper proposes a theoretical framework for the conceptualization and design of help options in computer-based second language (L2) listening. Based on four empirical studies, it aims at clarifying both conceptualization and design (CoDe) components. The elements of conceptualization consist of a novel four-part classification of help options:…

  19. Target selection and comparison of mission design for space debris removal by DLR's advanced study group

    NASA Astrophysics Data System (ADS)

    van der Pas, Niels; Lousada, Joao; Terhes, Claudia; Bernabeu, Marc; Bauer, Waldemar

    2014-09-01

    Space debris is a growing problem. Models show that the Kessler syndrome, the exponential growth of debris due to collisions, has become unavoidable unless an active debris removal program is initiated. The debris population in LEO with inclination between 60° and 95° is considered as the most critical zone. In order to stabilize the debris population in orbit, especially in LEO, 5 to 10 objects will need to be removed every year. The unique circumstances of such a mission could require that several objects are removed with a single launch. This will require a mission to rendezvous with a multitude of objects orbiting on different altitudes, inclinations and planes. Removal models have assumed that the top priority targets will be removed first. However this will lead to a suboptimal mission design and increase the ΔV-budget. Since there is a multitude of targets to choose from, the targets can be selected for an optimal mission design. In order to select a group of targets for a removal mission the orbital parameters and political constraints should also be taken into account. Within this paper a number of the target selection criteria are presented. The possible mission targets and their order of retrieval is dependent on the mission architecture. A comparison between several global mission architectures is given. Under consideration are 3 global missions of which a number of parameters are varied. The first mission launches multiple separate deorbit kits. The second launches a mother craft with deorbit kits. The third launches an orbital tug which pulls the debris in a lower orbit, after which a deorbit kit performs the final deorbit burn. A RoM mass and cost comparison is presented. The research described in this paper has been conducted as part of an active debris removal study by the Advanced Study Group (ASG). The ASG is an interdisciplinary student group working at the DLR, analyzing existing technologies and developing new ideas into preliminary

  20. Architectural design proposal for a Martian base to continue NASA Mars Design Reference Mission

    NASA Astrophysics Data System (ADS)

    Kozicki, Janek

    The issue of extraterrestrial bases has recently been a very vivid one. There are orbital stations currently existing and humans will travel to Mars around 2030. They will need stations established there, which will provide them the proper living conditions. Firstly, it might be a small module brought from Earth (e.g. NASA Mars Design Reference Mission module (DRM)), in later stages equivalents of Earth houses may be built from local resources. The goal of this paper is to propose an architectural design for an intermediate stage — for a larger habitable unit transported from Earth. It is inspired by terrestrial portable architecture ideas. A pneumatic structure requires small volume during transportation. However, it provides large habitable space after deployment. It is designed for transport by DRM transportation module and its deployment is considerable easy and brief. An architectural solution analogous to a terrestrial house with a studio and a workshop was assumed. Its form was a result of technical and environmental limitations, and the need for an ergonomic interior. The spatial placement of following zones was carefully considered: residential, agricultural and science, as well as a garage with a workshop, transportation routes, and a control and communication center. The issues of Life Support System, energy, food, water and waste recycling were also discussed. This Martian base was designed to be crewed by a team of eight people to stay on Mars for at least 1.5 year. An Open Plan architectural solution was assumed in pneumatic modules, with a high level of modularity. Walls of standardized sizes with zip-fasteners allow free rearrangement of the interior to adapt to a new situation (e.g. damage of one of the pneumatic modules or a psychological ,,need of a change"). The architectural design focuses on ergonomic and psychological aspects of longer stay in hostile Martian environment. This solution provides Martian crew with a comfortable habitable

  1. An evaluation of LMR design options for reduction of sodium void worth

    SciTech Connect

    Hill, R N; Khalil, H

    1989-01-01

    In this study, we analyze the relationship between the sodium void worth ({rho}{sub NA}) and other important performance characteristics for various design options which reduce {rho}{sub NA}. Our objective was to identify a preferred design option for reducing {rho}{sub NA} based on an overall consideration of performance tradeoffs. The focus of this study is on core designs of recent interest in the US LMR program, i.e. designs in the 450 to 1200 MWt size range that make use of metal alloy fuel. A key objective of the LMR development program in the US has been to design cores that can passively avoid damage when the control rods fail to scram in response to postulated accident initiators (e.g. inadvertent reactivity insertion or loss of coolant flow). Analyses and experimental tests of such unprotected events have demonstrated that the physical properties of metallic fuel alloys and the neutronic feedback characteristics of metal-fueled cores can be exploited to obtain favorable relations among the power, power/flow, and inlet temperature coefficients of reactivity and, consequently, large margins to sodium boiling and fuel damage under accident conditions. Since the reactivity effects of sodium density variation during postulated accidents are effectively compensated by other feedback effects, reduction of the sodium void worth has not been a primary design objective for recent LMR concepts; relatively large values ($4 to $6) are predicted for current core designs. 23 refs., 11 figs.

  2. Design of a Formation of Earth Orbiting Satellites: The Auroral Lites Mission

    NASA Technical Reports Server (NTRS)

    Hametz, Mark E.; Conway, Darrel J.; Richon, Karen

    1999-01-01

    The National Aeronautics and Space Administration (NASA) Goddard Space Flight Center (GSFC) has proposed a set of spacecraft flying in close formation around the Earth in order to measure the behavior of the auroras. The mission, named Auroral Lites, consists of four spacecraft configured to start at the vertices of a tetrahedron, flying over three mission phases. During the first phase, the distance between any two spacecraft in the formation is targeted at 10 kilometers (km). The second mission phase is much tighter, requiring satellite interrange spacing targeted at 500 meters. During the final phase of the mission, the formation opens to a nominal 100-km interrange spacing. In this paper, we present the strategy employed to initialize and model such a close formation during each of these phases. The analysis performed to date provides the design and characteristics of the reference orbit, the evolution of the formation during Phases I and II, and an estimate of the total mission delta-V budget. AI Solutions' mission design tool, FreeFlyer(R), was used to generate each of these analysis elements. The tool contains full force models, including both impulsive and finite duration maneuvers. Orbital maintenance can be fully modeled in the system using a flexible, natural scripting language built into the system. In addition, AI Solutions is in the process of adding formation extensions to the system facilitating mission analysis for formations like Auroral Lites. We will discuss how FreeFlyer(R) is used for these analyses.

  3. Design of a Formation of Earth-Orbiting Satellites: The Auroral Lites Mission

    NASA Technical Reports Server (NTRS)

    Hametz, Mark E.; Conway, Darrel J.; Richon, Karen

    1999-01-01

    The National Aeronautics and Space Administration (NASA) Goddard Space Flight Center (GSFC) has proposed a set of spacecraft flying in close formation around the Earth in order to measure the behavior of the auroras. The mission, named Auroral Lites, consists of four spacecraft configured to start at the vertices of a tetrahedron, flying over three mission phases. During the first phase, the distance between any two spacecraft in the formation is targeted at 10 kilometers (km). The second mission phase is much tighter, requiring satellite interrange spacing targeted at 500 meters. During the final phase of the mission, the formation opens to a nominal 100-km interrange spacing. In this paper, we present the strategy employed to initialize and model such a close formation during each of these phases. The analysis performed to date provides the design and characteristics of the reference orbit, the evolution of the formation during Phases I and II, and an estimate of the total mission delta-V budget. AI Solutions' mission design tool, FreeFlyer, was used to generate each of these analysis elements. The tool contains full force models, including both impulsive and finite duration maneuvers. Orbital maintenance can be fully modeled in the system using a flexible, natural scripting language built into the system. In addition, AI Solutions is in the process of adding formation extensions to the system facilitating mission analysis for formations like Auroral Lites. We will discuss how FreeFlyer is used for these analyses.

  4. Cryogenic Propulsion Stage (CPS) Configuration in Support of NASA's Multiple Design Reference Missions (DRMs)

    NASA Technical Reports Server (NTRS)

    Hanna, Stephen G.; Jones, David L.; Creech, Stephen D.; Lawrence, Thomas D.

    2012-01-01

    In support of the National Aeronautics and Space Administration's (NASA) Human Exploration and Operations Mission Directorate (HEOMD), the Space Launch System (SLS) is being designed for safe, affordable, and sustainable human and scientific exploration missions beyond Earth's or-bit (BEO). The SLS Team is tasked with developing a system capable of safely and repeatedly lofting a new fleet of spaceflight vehicles beyond Earth orbit. The Cryogenic Propulsion Stage (CPS) is a key enabler for evolving the SLS capability for BEO missions. This paper reports on the methodology and initial recommendations relative to the CPS, giving a brief retrospective of early studies on this promising propulsion hardware. This paper provides an overview of the requirements development and CPS configuration in support of NASA's multiple Design Reference Missions (DRMs).

  5. An orbital radar mapper of Venus in the 1980's - Mission design and analysis

    NASA Technical Reports Server (NTRS)

    Asnin, S. K.

    1973-01-01

    A reasonable approach to the examination of Venus topography, obscured for photographic imaging, is available in the application of airborne radar mapping systems to an orbiter mission about the planet. Extrapolating the improving capabilities of earth-based radar study of Venus into the 1980's suggests that only a non-uniform, poorly resolved surface profile will be possible relative to the potential for 100% coverage at 100 meter resolution with an orbital radar. The intent of this paper is to define mission opportunities favorable for a Venus orbital mapper during the 1980's, to examine orbit design problem associated with mapping radar systems, to establish what flexibility exists for an adaptive mapping strategy, to contribute to the sizing of particular spacecraft systems, to suggest a reference mission design and demonstrate mission feasibility.

  6. Lunar Cube Transfer Trajectory Options

    NASA Technical Reports Server (NTRS)

    Folta, David; Dichmann, Donald James; Clark, Pamela E.; Haapala, Amanda; Howell, Kathleen

    2015-01-01

    Numerous Earth-Moon trajectory and lunar orbit options are available for Cubesat missions. Given the limited Cubesat injection infrastructure, transfer trajectories are contingent upon the modification of an initial condition of the injected or deployed orbit. Additionally, these transfers can be restricted by the selection or designs of Cubesat subsystems such as propulsion or communication. Nonetheless, many trajectory options can b e considered which have a wide range of transfer duration, fuel requirements, and final destinations. Our investigation of potential trajectories highlights several options including deployment from low Earth orbit (LEO) geostationary transfer orbits (GTO) and higher energy direct lunar transfer and the use of longer duration Earth-Moon dynamical systems. For missions with an intended lunar orbit, much of the design process is spent optimizing a ballistic capture while other science locations such as Sun-Earth libration or heliocentric orbits may simply require a reduced Delta-V imparted at a convenient location along the trajectory.

  7. Lunar Cube Transfer Trajectory Options

    NASA Technical Reports Server (NTRS)

    Folta, David; Dichmann, Donald J.; Clark, Pamela; Haapala, Amanda; Howell, Kathleen

    2015-01-01

    Numerous Earth-Moon trajectory and lunar orbit options are available for Cubesat missions. Given the limited Cubesat injection infrastructure, transfer trajectories are contingent upon the modification of an initial condition of the injected or deployed orbit. Additionally, these transfers can be restricted by the selection or designs of Cubesat subsystems such as propulsion or communication. Nonetheless, many trajectory options can be considered which have a wide range of transfer durations, fuel requirements, and final destinations. Our investigation of potential trajectories highlights several options including deployment from low Earth orbit (LEO), geostationary transfer orbits (GTO), and higher energy direct lunar transfers and the use of longer duration Earth-Moon dynamical systems. For missions with an intended lunar orbit, much of the design process is spent optimizing a ballistic capture while other science locations such as Sun-Earth libration or heliocentric orbits may simply require a reduced Delta-V imparted at a convenient location along the trajectory.

  8. Design of Photovoltaic Power System for a Precursor Mission for Human Exploration of Mars

    NASA Technical Reports Server (NTRS)

    Mcnatt, Jeremiah; Landis, Geoffrey; Fincannon, James

    2016-01-01

    This project analyzed the viability of a photovoltaic power source for technology demonstration mission to demonstrate Mars in-situ resource utilization (ISRU) to produce propellant for a future human mission, based on technology available within the next ten years. For this assessment, we performed a power-system design study for a scaled ISRU demonstrator lander on the Mars surface based on existing solar array technologies.

  9. Hubble Space Telescope servicing mission scientific instrument protective enclosure design requirements and contamination controls

    NASA Technical Reports Server (NTRS)

    Hansen, Patricia A.; Hughes, David W.; Hedgeland, Randy J.; Chivatero, Craig J.; Studer, Robert J.; Kostos, Peter J.

    1994-01-01

    The Scientific Instrument Protective Enclosures were designed for the Hubble Space Telescope Servicing Missions to provide a beginning environment to a Scientific Instrument during ground and on orbit activities. The Scientific Instruments required very stringent surface cleanliness and molecular outgassing levels to maintain ultraviolet performance. Data from the First Servicing Mission verified that both the Scientific Instruments and Scientific Instrument Protective Enclosures met surface cleanliness level requirements during ground and on-orbit activities.

  10. Design and assembly sequence analysis of option 3 for CETF reference space station

    NASA Technical Reports Server (NTRS)

    Garrett, L. Bernard; Andersen, Gregory C.; Hall, John B., Jr.; Allen, Cheryl L.; Scott, A. D., Jr.; So, Kenneth T.

    1987-01-01

    A design and assembly sequence was conducted on one option of the Dual Keel Space Station examined by a NASA Critical Evaluation Task Force to establish viability of several variations of that option. A goal of the study was to produce and analyze technical data to support Task Force decisions to either examine particular Option 3 variations in more depth or eliminate them from further consideration. An analysis of the phasing assembly showed that use of an Expendable Launch Vehicle in conjunction with the Space Transportation System (STS) can accelerate the buildup of the Station and ease the STS launch rate constraints. The study also showed that use of an Orbital Maneuvering Vehicle on the first flight can significantly benefit Station assembly and, by performing Station subsystem functions, can alleviate the need for operational control and reboost systems during the early flights. In addition to launch and assembly sequencing, the study assessed stability and control, and analyzed node-packaging options and the effects of keel removal on the structural dynamics of the Station. Results of these analyses are presented and discussed.

  11. Space station needs, attributes and architectural options. Volume 4, task 2 and 3: Mission implementation and cost

    NASA Technical Reports Server (NTRS)

    1983-01-01

    An overview of the basic space station infrastructure is presented. A strong case is made for the evolution of the station using the basic Space Transportation System (STS) to achieve a smooth transition and cost effective implementation. The integrated logistics support (ILS) element of the overall station infrastructure is investigated. The need for an orbital transport system capability that is the key to servicing and spacecraft positioning scenarios and associated mission needs is examined. Communication is also an extremely important element and the basic issue of station autonomy versus ground support effects the system and subsystem architecture.

  12. G-Guidance Interface Design for Small Body Mission Simulation

    NASA Technical Reports Server (NTRS)

    Acikmese, Behcet; Carson, John; Phan, Linh

    2008-01-01

    The G-Guidance software implements a guidance and control (G and C) algorithm for small-body, autonomous proximity operations, developed under the Small Body GN and C task at JPL. The software is written in Matlab and interfaces with G-OPT, a JPL-developed optimization package written in C that provides G-Guidance with guaranteed convergence to a solution in a finite computation time with a prescribed accuracy. The resulting program is computationally efficient and is a prototype of an onboard, real-time algorithm for autonomous guidance and control. Two thruster firing schemes are available in G-Guidance, allowing tailoring of the software for specific mission maneuvers. For example, descent, landing, or rendezvous benefit from a thruster firing at the maneuver termination to mitigate velocity errors. Conversely, ascent or separation maneuvers benefit from an immediate firing to avoid potential drift toward a second body. The guidance portion of this software explicitly enforces user-defined control constraints and thruster silence times while minimizing total fuel usage. This program is currently specialized to small-body proximity operations, but the underlying method can be generalized to other applications.

  13. Electric thruster models for multimegawatt nuclear electric propulsion mission design

    NASA Technical Reports Server (NTRS)

    Leifer, Stephanie D.; Blandino, John J.; Sercel, Joel C.

    1991-01-01

    Three types of electric thrusters currently under development at JPL have potential to support future missions which utilize multimegawatt nuclear electric propulsion. These electric thrusters are the electron bombardment ion thruster, the magnetoplasmadynamic (MPD) thruster, and the electron-cyclotron-resonance (ECR) thruster. The electron bombardment ion thruster is a relatively mature technology which has been developed for operation at kilowatt power levels but will require new development for application in the multimegawatt regime. The MPD engine represents a technology which may be very well suited to steady-state multimegawatt applications but which has been limited to sub-scale (100's of kW) and pulsed (MW) testing thus far. The ECR plasma engine represents a class of very promising new concepts which are still in the basic research phase of development, but which may possess important fundamental advantages over other electric thruster technologies. Models of these thrusters are described and used to make projections of thrusters specific mass, efficiency, and power handling capacity for operation in the multimegawatt regime.

  14. Design of modular probes for stratospheric balloon mission: Thermo mechanical aspects and lession learned from SORA mission.

    NASA Astrophysics Data System (ADS)

    Bettanini, Carlo; Friso, Enrico; Colombatti, Giacomo; Aboudan, Alessio; Flamini, Enrico; Pirrotta, Simone; Debei, Stefano

    Stratospheric balloon missions provide a very effective facility for testing instruments in a space-like environment with drastically lower requirements in funding and sensibly shorter timelines than common space mission. Mainly during ascent to operative altitude and parachuted de-scent the flight units face fast changing environmental conditions which may induce issues in the mechanical and thermal behavior of the equipment. A new concept modular gondola was engineered by CISAS "G.Colombo" at University of Padova,to be easily reconfigured to host scientific experiments with different power and thermal requirements thus sensibly reducing development times and costs. The gondola was mechanically designed to withstand dynamic loads related to parachute opening and ground impact and provided a 1 m x 1m x 0.3 m volume for scientific payloads which is pressure regulated with the use of relief valves and thermally controlled by main CDMU.Furthermore the whole system was able to float in case of descent in water thanks to an optmised design of the main aluminium structure and use of hermetic connections. A custom Command and Data Management Unit with hard-real-time control capabilities has been developed to manage sensors acquisition, data storage, and experiments monitoring and control. The gondola was equipped with IMU, GPS, a downward looking cam-era and a set of health check and housekeeping sensors which sample key parameters as attitude, acceleration and temperature in several parts of the structure feeding housekeeping data to the main pc in order to monitor overall system health. The unit was successfully assembled and tested at University of Padova and used in the flight of the SORA mission launched in summer 2009 from Svalbard islands to map with a penetrating radar the stratification of ice and rock above Northern Greenland. Because of unexpected wind directions the mission trajectory was several hundred kilometers southern than predicted terminating with a

  15. Preliminary analysis of long-range aircraft designs for future heavy airlift missions

    NASA Technical Reports Server (NTRS)

    Nelms, W. P., Jr.; Murphy, R.; Barlow, A.

    1976-01-01

    A computerized design study of very large cargo aircraft for the future heavy airlift mission was conducted using the Aircraft Synthesis program (ACSYNT). The study was requested by the Air Force under an agreement whereby Ames provides computerized design support to the Air Force Flight Dynamics Laboratory. This effort is part of an overall Air Force program to study advanced technology large aircraft systems. Included in the Air Force large aircraft program are investigations of missions such as heavy airlift, airborne missile launch, battle platform, command and control, and aerial tanker. The Ames studies concentrated on large cargo aircraft of conventional design with payloads from 250,000 to 350,000 lb. Range missions up to 6500 n.mi. and radius missions up to 3600 n.mi. have been considered. Takeoff and landing distances between 7,000 and 10,000 ft are important constraints on the configuration concepts. The results indicate that a configuration employing conventional technology in all disciplinary areas weighs approximately 2 million pounds to accomplish either a 6500-n.mi. range mission or a 3600-n.mi. radius mission with a 350,000-lb payload.

  16. Pulsed activation analyses of the ITER blanket design options considered in the blanket trade-off study

    SciTech Connect

    Wang, Q.; Henderson, D.L.

    1994-12-31

    The International Thermonuclear Experimental Reactor (ITER) project began a new design phase called the Engineering Design Activity (EDA) which started in July 1992. A variety of blanket designs options were analyzed as a part of the U.S. ITER home team blanket option trade-off study (BOTS) which began in May 1993. The options considered were a self-cooled Li/V blanket, a helium cooled Li/V blanket and a water cooled 316 SS nonbreeding shield option. Detailed activation, dose rate and waste disposal rating calculations have been performed for these different ITER blanket design options based on a fluence of 3.0 MWa/m{sup 2} and an average neutron wall loading of 2.0 MW/m{sup 2}. A continuous operation assumption was utilized in the analysis. The results of this work are presented in this conference.

  17. An Extended Mission to Multiple Asteroid Flybys after Lunar Exploration: Design and analysis

    NASA Astrophysics Data System (ADS)

    Dong, Qiao; Pingyuan, Cui; Hutao, Cui

    Asteroids are key objects for the research of the origin and evolution of the solar system, because they relatively well conserve their past, as they are often said to be the fossil from the early phase of the solar system. Human have carried out many small celestial body missions (such as DS-I, NEAR, Stardust, Deep impact, etc) and gained a plenty of valuable experience. With many new asteroid or comet missions in various stages of development (Dawn, ROSETTA, Hayabusa-II), the situation regarding our understanding of these bodies can improve in the future. With the success of "Chang-Er" program for lunar exploration and the development of "Ying-Huo" program for Mars exploration, it will provide the great opportunity for Chinese development of the deep space exploration. Following up the "Chang-Er" and "Ying-Huo" program, two promising scenarios have been discussed. One of them is a multiple asteroid rendezvous and sample return mission aiming at several spectral known Near-Earth asteroids. The other is a multiple asteroid flyby mission which is an extended "Chang-Er" mission for lunar exploration. Here, an extended mission to multiple asteroid flybys after lunar exploration is proposed and discussed. We investigated feasible mission scenarios. According to the constraints of the mission, we search and present the optimal rendezvous and flyby opportunities for potential asteroids. The preliminary design results of the proposed missions will be reported. Finally, we analyze the trajectory characteristics and gave some key parameters, which would have a direct impact on communication system, power system, thermal control system of spacecraft and the optical instrument for science mission etc.

  18. Design of Spacecraft Missions to Test Kinetic Impact for Asteroid Deflection

    NASA Technical Reports Server (NTRS)

    Hernandez, Sonia; Barbee, Brent W.

    2011-01-01

    There are currently over 8,000 known near-Earth asteroids (NEAs), and more are being discovered on a continual basis. More than 1,200 of these are classified as Potentially Hazardous Asteroids (PHAs) because their Minimum Orbit Intersection Distance (MOID) with Earth's orbit is <= 0.05 AU and their estimated diameters are >= 150 m. To date, 178 Earth impact structures have been discovered, indicating that our planet has previously been struck with devastating force by NEAs and will be struck again. Such collisions are aperiodic events and can occur at any time. A variety of techniques have been proposed to defend our planet from NEA impacts by deflecting the incoming asteroid. However, none of these techniques have been tested. Unless rigorous testing is conducted to produce reliable asteroid deflection systems, we will be forced to deploy completely untested -- and therefore unreliable -- deflection missions when a sizable asteroid on a collision course with Earth is discovered. Such missions will have a high probability of failure. We propose to address this problem with a campaign of deflection technology test missions deployed to harmless NEAs. The objective of these missions is to safely evaluate and refine the mission concepts and asteroid deflection system designs. Our current research focuses on the kinetic impactor, one of the simplest proposed asteroid deflection techniques in which a spacecraft is sent to collide with an asteroid at high relative velocity. By deploying test missions in the near future, we can characterize the performance of this deflection technique and resolve any problems inherent to its execution before needing to rely upon it during a true emergency. In this paper we present the methodology and results of our survey, including lists of NEAs for which safe and effective kinetic impactor test missions may be conducted within the next decade. Full mission designs are also presented for the NEAs which offer the best mission opportunities.

  19. Multi-Objective Hybrid Optimal Control for Multiple-Flyby Interplanetary Mission Design Using Chemical Propulsion

    NASA Technical Reports Server (NTRS)

    Englander, Jacob A.; Vavrina, Matthew A.

    2015-01-01

    Preliminary design of high-thrust interplanetary missions is a highly complex process. The mission designer must choose discrete parameters such as the number of flybys and the bodies at which those flybys are performed. For some missions, such as surveys of small bodies, the mission designer also contributes to target selection. In addition, real-valued decision variables, such as launch epoch, flight times, maneuver and flyby epochs, and flyby altitudes must be chosen. There are often many thousands of possible trajectories to be evaluated. The customer who commissions a trajectory design is not usually interested in a point solution, but rather the exploration of the trade space of trajectories between several different objective functions. This can be a very expensive process in terms of the number of human analyst hours required. An automated approach is therefore very desirable. This work presents such an approach by posing the impulsive mission design problem as a multiobjective hybrid optimal control problem. The method is demonstrated on several real-world problems.

  20. Equal-Curvature X-ray Telescope Designs for Constellation-X Mission

    NASA Technical Reports Server (NTRS)

    Saha, Timo T.; Content, David A.; Zhang, William W.

    2003-01-01

    We study grazing incidence Equal-Curvature telescope designs for the Constellation-X mission. These telescopes have nearly spherical axial surfaces. The telescopes are designed so that the axial curvature is the same on the primary and secondary. The optical performance of these telescopes is for all practical purposes identical to the equivalent Wolter telescopes.

  1. Evolutionary exploitation of design options by the first animals with hard skeletons.

    PubMed

    Thomas, R D; Shearman, R M; Stewart, G W

    2000-05-19

    The set of viable design elements available for animals to use in building skeletons has been fully exploited. Analysis of animal skeletons in relation to the multivariate, theoretical "Skeleton Space" has shown that a large proportion of these options are used in each phylum. Here, we show that structural elements deployed in the skeletons of Burgess Shale animals (Middle Cambrian) incorporate 146 of 182 character pairs defined in this morphospace. Within 15 million years of the appearance of crown groups of phyla with substantial hard parts, at least 80 percent of skeletal design elements recognized among living and extinct marine metazoans were exploited.

  2. The Impact of Information Technology on the Design, Development, and Implementation of a Lunar Exploration Mission

    NASA Technical Reports Server (NTRS)

    Gross, Anthony R.; Sims, Michael H.; Briggs, Geoffrey A.

    1996-01-01

    From the beginning to the present expeditions to the Moon have involved a large investment of human labor. This has been true for all aspects of the process, from the initial design of the mission, whether scientific or technological, through the development of the instruments and the spacecraft, to the flight and operational phases. In addition to the time constraints that this situation imposes, there is also a significant cost associated with the large labor costs. As a result lunar expeditions have been limited to a few robotic missions and the manned Apollo program missions of the 1970s. With the rapid rise of the new information technologies, new paradigms are emerging that promise to greatly reduce both the time and cost of such missions. With the rapidly increasing capabilities of computer hardware and software systems, as well as networks and communication systems, a new balance of work is being developed between the human and the machine system. This new balance holds the promise of greatly increased exploration capability, along with dramatically reduced design, development, and operating costs. These new information technologies, utilizing knowledge-based software and very highspeed computer systems, will provide new design and development tools, scheduling mechanisms, and vehicle and system health monitoring capabilities that have hitherto been unavailable to the mission and spacecraft designer and the system operator. This paper will utilize typical lunar missions, both robotic and crewed, as a basis to describe and illustrate how these new information system technologies could be applied to all aspects such missions. In particular, new system design tradeoff tools will be described along with technologies that will allow a very much greater degree of autonomy of exploration vehicles than has heretofore been possible. In addition, new information technologies that will significantly reduce the human operational requirements will be discussed.

  3. Generic procedure for designing and implementing plan management systems for space science missions operations

    NASA Astrophysics Data System (ADS)

    Chaizy, P. A.; Dimbylow, T. G.; Allan, P. M.; Hapgood, M. A.

    2011-09-01

    This paper is one of the components of a larger framework of activities whose purpose is to improve the performance and productivity of space mission systems, i.e. to increase both what can be achieved and the cost effectiveness of this achievement. Some of these activities introduced the concept of Functional Architecture Module (FAM); FAMs are basic blocks used to build the functional architecture of Plan Management Systems (PMS). They also highlighted the need to involve Science Operations Planning Expertise (SOPE) during the Mission Design Phase (MDP) in order to design and implement efficiently operation planning systems. We define SOPE as the expertise held by people who have both theoretical and practical experience in operations planning, in general, and in space science operations planning in particular. Using ESA's methodology for studying and selecting science missions we also define the MDP as the combination of the Mission Assessment and Mission Definition Phases. However, there is no generic procedure on how to use FAMs efficiently and systematically, for each new mission, in order to analyse the cost and feasibility of new missions as well as to optimise the functional design of new PMS; the purpose of such a procedure is to build more rapidly and cheaply such PMS as well as to make the latter more reliable and cheaper to run. This is why the purpose of this paper is to provide an embryo of such a generic procedure and to show that the latter needs to be applied by people with SOPE during the MDP. The procedure described here proposes some initial guidelines to identify both the various possible high level functional scenarii, for a given set of possible requirements, and the information that needs to be associated with each scenario. It also introduces the concept of catalogue of generic functional scenarii of PMS for space science missions. The information associated with each catalogued scenarii will have been identified by the above procedure and

  4. Distributed Space Mission Design for Earth Observation Using Model-Based Performance Evaluation

    NASA Technical Reports Server (NTRS)

    Nag, Sreeja; LeMoigne-Stewart, Jacqueline; Cervantes, Ben; DeWeck, Oliver

    2015-01-01

    Distributed Space Missions (DSMs) are gaining momentum in their application to earth observation missions owing to their unique ability to increase observation sampling in multiple dimensions. DSM design is a complex problem with many design variables, multiple objectives determining performance and cost and emergent, often unexpected, behaviors. There are very few open-access tools available to explore the tradespace of variables, minimize cost and maximize performance for pre-defined science goals, and therefore select the most optimal design. This paper presents a software tool that can multiple DSM architectures based on pre-defined design variable ranges and size those architectures in terms of predefined science and cost metrics. The tool will help a user select Pareto optimal DSM designs based on design of experiments techniques. The tool will be applied to some earth observation examples to demonstrate its applicability in making some key decisions between different performance metrics and cost metrics early in the design lifecycle.

  5. Kids as Airborne Mission Scientists: Designing PBL To Inspire Kids.

    ERIC Educational Resources Information Center

    Koszalka, Tiffany A.; Grabowski, Barbara L.; Kim, Younghoon

    Problem-based learning (PBL) has great potential for inspiring K-12 learning. KaAMS, a NASA funded project and an example of PBL, was designed to help teachers inspire middle school students to learn science. The students participate as scientists investigating environmental problems using NASA airborne remote sensing data. Two PBL modules were…

  6. Manchester Coding Option for SpaceWire: Providing Choices for System Level Design

    NASA Technical Reports Server (NTRS)

    Rakow, Glenn; Kisin, Alex

    2014-01-01

    This paper proposes an optional coding scheme for SpaceWire in lieu of the current Data Strobe scheme for three reasons. First reason is to provide a straightforward method for electrical isolation of the interface; secondly to provide ability to reduce the mass and bend radius of the SpaceWire cable; and thirdly to provide a means for a common physical layer over which multiple spacecraft onboard data link protocols could operate for a wide range of data rates. The intent is to accomplish these goals without significant change to existing SpaceWire design investments. The ability to optionally use Manchester coding in place of the current Data Strobe coding provides the ability to DC balanced the signal transitions unlike the SpaceWire Data Strobe coding; and therefore the ability to isolate the electrical interface without concern. Additionally, because the Manchester code has the clock and data encoded on the same signal, the number of wires of the existing SpaceWire cable could be optionally reduced by 50. This reduction could be an important consideration for many users of SpaceWire as indicated by the already existing effort underway by the SpaceWire working group to reduce the cable mass and bend radius by elimination of shields. However, reducing the signal count by half would provide even greater gains. It is proposed to restrict the data rate for the optional Manchester coding to a fixed data rate of 10 Megabits per second (Mbps) in order to make the necessary changes simple and still able to run in current radiation tolerant Field Programmable Gate Arrays (FPGAs). Even with this constraint, 10 Mbps will meet many applications where SpaceWire is used. These include command and control applications and many instruments applications with have moderate data rate. For most NASA flight implementations, SpaceWire designs are in rad-tolerant FPGAs, and the desire to preserve the heritage design investment is important for cost and risk considerations. The

  7. A Potential Operational CryoSat Follow-on Mission Concept and Design

    NASA Astrophysics Data System (ADS)

    Cullen, R.

    2015-12-01

    CryoSat was a planned as a 3 year mission with clear mission objectives to allow the assessment rates of change of thickness in the land and marine ice fields with reduced uncertainties with relation to other non-dedicated missions. Although CryoSat suffered a launch failure in Oct 2005, the mission was recovered with a launch in April 2010 of CryoSat-2. The nominal mission has now been completed, all mission requirements have been fulfilled and CryoSat has been shown to be most successful as a dedicated polar ice sheet measurement system demonstrated by nearly 200 peer reviewed publications within the first four years of launch. Following the completion of the nominal mission in Oct 2013 the platform was shown to be in good health and with a scientific backing provided by the ESA Earth Science Advisory Committee (ESAC) the mission has been extended until Feb 2017 by the ESA Programme Board for Earth Observation. Though not designed to provide data for science and operational services beyond its original mission requirements, a number of services have been developed for exploitation and these are expected to increase over the next few years. Services cover a number of aspects of land and marine ice fields in addition to complementary activities covering glacial monitoring, inland water in addition to coastal and open ocean surface topography science that CryoSat has demonstrated world leading advances with. This paper will present the overall concept for a potential low-cost follow-on to the CryoSat mission with the objective to provide both continuity of the existing CryoSat based data sets, i.e., longer term science and operational services that cannot be provided by the existing Copernicus complement of satellites. This is, in part, due to the high inclination (92°) drifting orbit and state of the art Synthetic Aperture Interferometer Radar Altimeter (SIRAL). In addition, further improvements in performance are expected by use of the instrument timing and

  8. Earth Entry Vehicle Design for Sample Return Missions Using M-SAPE

    NASA Technical Reports Server (NTRS)

    Samareh, Jamshid

    2015-01-01

    Most mission concepts that return sample material to Earth share one common element: an Earth entry vehicle (EEV). The primary focus of this paper is the examination of EEV design space for relevant sample return missions. Mission requirements for EEV concepts can be divided into three major groups: entry conditions (e.g., velocity and flight path angle), payload (e.g., mass, volume, and g-load limit), and vehicle characteristics (e.g., thermal protection system, structural topology, and landing concepts). The impacts of these requirements on the EEV design have been studied with an integrated system analysis tool, and the results will be discussed in details. In addition, through sensitivities analyses, critical design drivers that have been identified will be reviewed.

  9. Habitability as a Tier One Criterion in Exploration Mission and Vehicle Design. Part 1; Habitability

    NASA Technical Reports Server (NTRS)

    Adams, Constance M.; McCurdy, Matthew Riegel

    1999-01-01

    Habitability and human factors are necessary criteria to include in the iterative process of Tier I mission design. Bringing these criteria in at the first, conceptual stage of design for exploration and other human-rated missions can greatly reduce mission development costs, raise the level of efficiency and viability, and improve the chances of success. In offering a rationale for this argument, the authors give an example of how the habitability expert can contribute to early mission and vehicle architecture by defining the formal implications of a habitable vehicle, assessing the viability of units already proposed for exploration missions on the basis of these criteria, and finally, by offering an optimal set of solutions for an example mission. In this, the first of three papers, we summarize the basic factors associated with habitability, delineate their formal implications for crew accommodations in a long-duration environment, and show examples of how these principles have been applied in two projects at NASA's Johnson Space Center: the BIO-Plex test facility, and TransHab.

  10. Baseline design of a Mobile Asteroid Surface Scout (MASCOT) for the Hayabusa-2 mission

    NASA Astrophysics Data System (ADS)

    Lange, Caroline; Richter, Lutz; Ho, Tra Mi; Witte, Lars

    The Hayabusa-2 mission is currently being studied by JAXA/JSPEC as a sample return mis-sion to the C-type near-Earth asteroid 1999JU3. Hayabusa-2, with launch planned for 2014, would be the immediate successor to the currently flying Hayabusa mission. Originally in the context of the proposed ESA Cosmic Vision M-class mission Marco Polo, but then following an invitation by JAXA/JSPEC, the Institute of Space Systems of the German Aerospace Center (DLR) led a proposal for a separate lander package MASCOT (Mobile Asteroid Surface Scout) to be carried on the mission. A feasibility study was subsequently carried out that, upon consultation with the planetary science community, assessed different concepts for the lander that converged to a package with 3 kg of P/L, for a total mass of 10-15 kg. Presently, MASCOT is in the preliminary design phase and an Announcement of Opportunity for its payload complement is being prepared. This paper is intended to give an overview over the current system baseline design and dedicated subsystems, such as mobility. The focus will be on the highly demanding constraints that are being imposed on the system due to the general mission scenario, the asteroid environment and the tight budgetary limitations and the means by which these constraints are being fulfilled, while still offering an excellent science potential.

  11. High Earth Orbit Design for Lunar-Assisted Medium Class Explorer Missions

    NASA Technical Reports Server (NTRS)

    McGiffin, Daniel A.; Mathews, Michael; Cooley, Steven

    2001-01-01

    This study investigates the application of high-Earth orbit (HEO) trajectories to missions requiring long on-target integration times, avoidance of the Earth's radiation belt, and minimal effects of Earth and Lunar shadow periods which could cause thermal/mechanical stresses on the science instruments. As used here, a HEO trajectory is a particular solution to the restricted three-body problem in the Earth-Moon system with the orbit period being either 1/2 of, or 1/4 of, the lunar sidereal period. A primary mission design goal is to find HEO trajectories where, for a five-year mission duration, the minimum perigee radius is greater than seven Earth radii (R(sub E)). This minimum perigee radius is chosen so that, for the duration of the mission, the perigee is always above the relatively heavily populated geosynchronous radius of 6.6 R(sub E). A secondary goal is to maintain as high an ecliptic inclination as possible for the duration of the mission to keep the apsis points well out of the Ecliptic plane. Mission design analysis was completed for launch dates in the month of June 2003, using both direct transfer and phasing loop transfer techniques, to a lunar swingby for final insertion into a HEO. Also provided are analysis results of eclipse patterns for the trajectories studied, as well as the effects of launch vehicle errors and launch delays.

  12. A CubeSat Asteroid Mission: Design Study and Trade-Offs

    NASA Technical Reports Server (NTRS)

    Landis, Geoffrey A.; Oleson, Steven R.; McGuire, Melissa; Hepp, Aloysius; Stegeman, James; Bur, Mike; Burke, Laura; Martini, Michael; Fittje, James E.; Kohout, Lisa; Fincannon, James; Packard, Tom

    2014-01-01

    There is considerable interest in expanding the applicability of cubesat spacecraft into lightweight, low cost missions beyond Low Earth Orbit. A conceptual design was done for a 6-U cubesat for a technology demonstration to demonstrate use of electric propulsion systems on a small satellite platform. The candidate objective was a mission to be launched on the SLS test launch EM-1 to visit a Near-Earth asteroid. Both asteroid fly-by and asteroid rendezvous missions were analyzed. Propulsion systems analyzed included cold-gas thruster systems, Hall and ion thrusters, incorporating either Xenon or Iodine propellant, and an electrospray thruster. The mission takes advantage of the ability of the SLS launch to place it into an initial trajectory of C3=0. Targeting asteroids that fly close to earth minimizes the propulsion required for fly-by/rendezvous. Due to mass constraints, high specific impulse is required, and volume constraints mean the propellant density was also of great importance to the ability to achieve the required deltaV. This improves the relative usefulness of the electrospray salt, with higher propellant density. In order to minimize high pressure tanks and volatiles, the salt electrospray and iodine ion propulsion systems were the optimum designs for the fly-by and rendezvous missions respectively combined with a thruster gimbal and wheel system For the candidate fly-by mission, with a mission deltaV of about 400 m/s, the mission objectives could be accomplished with a 800s electrospray propulsion system, incorporating a propellant-less cathode and a bellows salt tank. This propulsion system is planned for demonstration on 2015 LEO and 2016 GEO DARPA flights. For the rendezvous mission, at a ?V of 2000 m/s, the mission could be accomplished with a 50W miniature ion propulsion system running iodine propellant. This propulsion system is not yet demonstrated in space. The conceptual design shows that an asteroid mission is possible using a cubesat

  13. EPS performance predictions support TOPEX/Poseidon spacecraft mission beyond design goal of 5 years

    SciTech Connect

    Chetty, P.R.K.; Doherty, M.; Sherwood, R.

    1995-12-31

    TOPEX/Poseidon is a scientific satellite, launched successfully on August 10,1992 to map the surface topology of the earth`s oceans. The electrical power system for this satellite consists of the following elements: a single wing solar array, a solar array drive and power transfer assembly, a peak power tracker, and storage batteries. The satellite`s electronics are designed for a three-year primary mission. Because of a potential mission extension, the solar array, batteries, and propellant are sized for a five year mission goal. This paper addresses on overview of the EPS and then performance predictions beyond five years are made. The EPS will meet all the mission requirements of the TOPEX/Poseidon Spacecraft for seven years.

  14. Design and application of electromechanical actuators for deep space missions

    NASA Technical Reports Server (NTRS)

    Haskew, Tim A.; Wander, John

    1993-01-01

    During the period 8/16/92 through 2/15/93, work has been focused on three major topics: (1) screw modeling and testing; (2) motor selection; and (3) health monitoring and fault diagnosis. Detailed theoretical analysis has been performed to specify a full dynamic model for the roller screw. A test stand has been designed for model parameter estimation and screw testing. In addition, the test stand is expected to be used to perform a study on transverse screw loading.

  15. Terrestrial Planet Finder Interferometer: Architecture, Mission Design, and Technology Development

    NASA Technical Reports Server (NTRS)

    Henry, Curt

    2004-01-01

    This slide presentation represents an overview progress report about the system design and technology development of two interferometer concepts studied for the Terrestrial Planet Finder (TPF) project. The two concepts are a structurally-connected interferometer (SCI) intended to fulfill minimum TPF science goals and a formation-flying interferometer (FFI) intended to fulfill full science goals. Described are major trades, analyses, and technology experiments completed. Near term plans are also described. This paper covers progress since August 2003

  16. Kepler Mission Design, Realized Photometric Performance, and Early Science

    DTIC Science & Technology

    2010-04-20

    designed with the explicit capability to detect Earth- size planets in the habitable zone of solar-like stars using the transit photometry method...that can be used to distinguish dwarf stars from red giants. Key words: instrumentation: photometers – planetary systems – space vehicles: instruments...characteristics. The method of choice is transit photometry (Pont et al. 2009), which provides the orbital period and size of the planet relative to its

  17. Heat Exchanger Design Options and Tritium Transport Study for the VHTR System

    SciTech Connect

    Chang H. Oh; Eung S. Kim

    2008-09-01

    This report presents the results of a study conducted to consider heat exchanger options and tritium transport in a very high temperature reactor (VHTR) system for the Next Generation Nuclear Plant Project. The heat exchanger options include types, arrangements, channel patterns in printed circuit heat exchangers (PCHE), coolant flow direction, and pipe configuration in shell-and-tube designs. Study considerations include: three types of heat exchanger designs (PCHE, shell-and-tube, and helical coil); single- and two-stage unit arrangements; counter-current and cross flow configurations; and straight pipes and U-tube designs in shell-and-tube type heat exchangers. Thermal designs and simple stress analyses were performed to estimate the heat exchanger options, and the Finite Element Method was applied for more detailed calculations, especially for PCHE designs. Results of the options study show that the PCHE design has the smallest volume and heat transfer area, resulting in the least tritium permeation and greatest cost savings. It is theoretically the most reliable mechanically, leading to a longer lifetime. The two-stage heat exchanger arrangement appears to be safer and more cost effective. The recommended separation temperature between first and second stages in a serial configuration is 800oC, at which the high temperature unit is about one-half the size of the total heat exchanger core volume. Based on simplified stress analyses, the high temperature unit will need to be replaced two or three times during the plant’s lifetime. Stress analysis results recommend the off-set channel pattern configuration for the PCHE because stress reduction was estimated at up to 50% in this configuration, resulting in a longer lifetime. The tritium transport study resulted in the development of a tritium behavior analysis code using the MATLAB Simulink code. In parallel, the THYTAN code, previously performed by Ohashi and Sherman (2007) on the Peach Bottom data, was revived

  18. Design criteria and mitigation options for thermal fatigue effects in ATW blankets.

    SciTech Connect

    Dunn, F. E.

    2000-12-07

    Thermal fatigue due to beam interruptions is an issue that must be addressed in the design of an ATW blanket. Two different approaches can be taken to address this issue. One approach is to analyze current ATW blanket designs in order to set interrupt frequency design limits for the accelerator. The other approach is to assume that accelerator reliability can not be guaranteed before design and construction of the blanket. In this approach the blanket must be designed so as to accommodate an accelerator with a beam interruption frequency significantly higher than current high power accelerators in order to provide a margin of error. Both approaches are considered in this paper. Both a sodium cooled blanket design and a lead-bismuth cooled blanket design are considered. Thermal hydraulic analysis of the blanket for beam interruption transients is carried out with the SASSYS-1 systems analysis code to obtain the time histories of the coolant temperatures in contact with structural components. These coolant temperatures are then used in a detailed structure temperature calculation to obtain structure surface and structure average temperatures. The difference between the average temperature and the surface temperature is used to obtain thermal strains. Low cycle fatigue curves from the American Society of Mechanical Engineers Boiler and Pressure Vessel Code are used to determine the number of cycles that the structural components can endure, based on these strains. Calculations are made for base case designs and for a number of mitigation options. The mitigation options include using two separate accelerators to provide the beam, reducing the thickness of the above core load pads in the subassemblies, increasing the coolant flow rate or reducing power in order to reduce the core temperature rise, and reducing the superheat in the once-through steam generator.

  19. 77 FR 58604 - Self-Regulatory Organizations; BOX Options Exchange LLC; Notice of Designation of a Longer Period...

    Federal Register 2010, 2011, 2012, 2013, 2014

    2012-09-21

    ... COMMISSION Self-Regulatory Organizations; BOX Options Exchange LLC; Notice of Designation of a Longer Period.... On July 25, 2012, BOX Options Exchange LLC (``Exchange'') filed with the Securities and Exchange... Period (``PIP'')) with respect to the execution of quotes and orders that are on the BOX Book prior...

  20. Orbit Design for Phase I and II of the Magnetospheric Multiscale Mission

    NASA Technical Reports Server (NTRS)

    Hughes, Steve P.

    2004-01-01

    The Magnetospheric Multiscale Mission (MMS) is a NASA mission intended to make fundamental advancements in our understanding of the Earth s magnetosphere. There are three processes that MMS is intended to study including magnetic reconnection, charged particle acceleration, and turbulence. There are four phases of the MMS mission and each phase is designed to study a particular region of the Earth's magnetosphere. The mission is composed of a formation of four spacecraft that are nominally in a regular tetrahedron formation. In this work, we present optimal orbit designs for Phase I and II. This entails designing reference orbits such that the spacecraft dwell-time in the region of interest is a maximum. This is non-trivial because the Earth's magnetosphere is dynamic and its shape and position are not constant in inertial space. Optimal orbit design for MMS also entails designing the formation so that the relative motion of the four spacecraft yields the greatest science return. We develop performance metrics that are directly related to the science return, and use Sequential Quadratic Programming (SQP) to determine optimal relative motion solutions. While designing for optimal science return, we also consider practical constraints such as maximum eclipse time and minimum inter-spacecraft separation distances. Data are presented that illustrates how long we can ensure that the formation remains in the relevant region of the Earth's magnetosphere. We also draw general conclusions about where in the orbit acceptable tetrahedron configurations can be provided and for how long.

  1. 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.

  2. Multiagent Modeling and Simulation in Human-Robot Mission Operations Work System Design

    NASA Technical Reports Server (NTRS)

    Sierhuis, Maarten; Clancey, William J.; Sims, Michael H.; Shafto, Michael (Technical Monitor)

    2001-01-01

    This paper describes a collaborative multiagent modeling and simulation approach for designing work systems. The Brahms environment is used to model mission operations for a semi-autonomous robot mission to the Moon at the work practice level. It shows the impact of human-decision making on the activities and energy consumption of a robot. A collaborative work systems design methodology is described that allows informal models, created with users and stakeholders, to be used as input to the development of formal computational models.

  3. Design of a Mission Data Storage and Retrieval System for NASA Dryden Flight Research Center

    NASA Technical Reports Server (NTRS)

    Lux, Jessica; Downing, Bob; Sheldon, Jack

    2007-01-01

    The Western Aeronautical Test Range (WATR) at the NASA Dryden Flight Research Center (DFRC) employs the WATR Integrated Next Generation System (WINGS) for the processing and display of aeronautical flight data. This report discusses the post-mission segment of the WINGS architecture. A team designed and implemented a system for the near- and long-term storage and distribution of mission data for flight projects at DFRC, providing the user with intelligent access to data. Discussed are the legacy system, an industry survey, system operational concept, high-level system features, and initial design efforts.

  4. Trajectory Design to Mitigate Risk on the Transiting Exoplanet Survey Satellite (TESS) Mission

    NASA Technical Reports Server (NTRS)

    Dichmann, Donald

    2016-01-01

    The Transiting Exoplanet Survey Satellite (TESS) will employ a highly eccentric Earth orbit, in 2:1 lunar resonance, reached with a lunar flyby preceded by 3.5 phasing loops. The TESS mission has limited propellant and several orbit constraints. Based on analysis and simulation, we have designed the phasing loops to reduce delta-V and to mitigate risk due to maneuver execution errors. We have automated the trajectory design process and use distributed processing to generate and to optimize nominal trajectories, check constraint satisfaction, and finally model the effects of maneuver errors to identify trajectories that best meet the mission requirements.

  5. 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.

  6. Terrestrial Planet Finder Interferometer: Architecture, Mission Design and Technology Development

    NASA Technical Reports Server (NTRS)

    Henry, Curt; Lay, Oliver; Aung, MiMi; Gunter, Steven M.; Dubovitsky, Serge; Blackwood, Gary

    2004-01-01

    This overview paper is a progress report about the system design and technology development of two interferometer concepts studied for the Terrestrial Planet Finder (TPF) project. The two concepts are a structurally-connected interferometer (SCI) intended to fulfill minimum TPF science goals and a formation-flying interferometer (FFI) intended to fulfill full science goals. Described are major trades, analyses, and technology experiments completed. Near term plans are also described. This paper covers progress since August 2003 and serves as an update to a paper presented at that month's SPIE conference, 'Techniques and Instrumentation for Detection of Exoplanets.

  7. Design of a Solar Sail Mission to Mars

    NASA Technical Reports Server (NTRS)

    Eastridge, Richard; Funston, Kerry; Okia, Aminat; Waldrop, Joan; Zimmerman, Christopher

    1989-01-01

    An evaluation of the design of the solar sail includes key areas such as structures, sail deployment, space environmental effects, materials, power systems, telemetry, communications, attitude control, thermal control, and trajectory analysis. Deployment and material constraints determine the basic structure of the sail, while the trajectory of the sail influences the choice of telemetry, communications, and attitude control systems. The thermal control system of the sail for the structures and electronics takes into account the effects of the space environment. Included also are a cost and weight estimate for the sail.

  8. NEXT Lunar Lander Mission - Overview and Challenges of the Lunar Rover Design

    NASA Astrophysics Data System (ADS)

    Allouis, Elie

    Looking ahead at the 2015-2018 timeframe, the European Space Agency (ESA) has recently started the investigation of the Next Exploration Science and Technology missions (NEXT) to demonstrate a number of key technologies for future programmes such as the Mars Sample Return (MSR). This paper provides the first insights into the mobile rover concept investigated as part of the NEXT Lunar Lander Study. Operating at the South Pole of the Moon, the rover will face a very challenging environment. Subjected to 200-hours long cold lunar nights at -200C for an initial mission duration of 1 year, and a total traverse of 20km, the design and operation of the rover requires careful attention. Its design is initially based on the knowledge developed for the ESA ExoMars mission, but the major differences in the environment and operation of the rover, means that most of the systems need a thorough assessment of their capabilities under Lunar condition and, where required, the development of new solutions. From the locomotion system designed to cope with uncertain lunar terrain, the thermal system dealing with gradients of hundreds of degrees, to the navigation through dark shadows, this paper illustrates some of the challenges future missions will face when targeting location such as the south pole on the Moon, but it will also provide details of the enabling technologies leading to the Mars Sample Return Mission and beyond.

  9. Radioisotope thermophotovoltaic system design and its application to an illustrative space mission

    SciTech Connect

    Schock, A.; Kumar, V.

    1995-01-05

    The paper describes the results of a DOE-sponsored design study of a radioisotope thermophotovoltaic generator (RTPV), to complement similar studies of Radioisotope Thermoelectric Generators (RTGs) and Stirling Generators (RSGs) previously published by the author. Instead of conducting a generic study, it was decided to focus the design effort by directing it at a specific illustrative space mission, Pluto Fast Flyby (PFF). That mission, under study by JPL, envisages a direct eight-year flight to Pluto (the only unexplored planet in the solar system), followed by comprehensive mapping, surface composition, and atmospheric structure measurements during a brief flyby of the planet and its moon Charon, and transmission of the recorded science data to Earth during a post-encounter cruise lasting up to one year. Because of Pluto`s long distance from the sun (30--50 A.U.) and the mission`s large energy demand, JPL has baselined the use of a radioisotope power system for the PFF spacecraft. TRGs have been tentatively selected, because they have been successfully flown on many space missions, and have demonstrated exceptional reliability and durability. The only reason for exploring the applicability of the far less mature RTPV systems is their potential for much higher conversion efficiencies, which would greatly reduce the mass and cost of the required radioisotope heat source. Those attributes are particularly important for the PFF mission, which---like all NASA missions under current consideration---is severely mass- and cost-limited. The paper describes the design of the radioisotope heat source, the thermophotovoltaic converter, and the heat rejection system; and depicts its integration with the PFF spacecraft.

  10. Analysis, optimization, and assessment of radioisotope thermophotovoltaic system design for an illustrative space mission

    SciTech Connect

    Schock, A.; Mukunda, M.; Or, C.; Summers, G.

    1995-01-05

    A companion paper presented at this conference described the design of a Radioisotope Thermophotovoltaic (RTPV) Generator for an illustrative space mission (Pluto Fast Flyby). It presented a detailed design of an integrated system consisting of a radioisotope heat source, a thermophotovoltaic converter, and an optimized heat rejection system. The present paper describes the thermal, electrical, and structural analyses which led to that optimized design, and compares the computed RTPV performance to that of a Radioisotope Thermoelectric Generator (RTG) designed for the same mission. RTPVs are of course much less mature than RTGs, but our results indicate that---when fully developed---they could result in a 60% reduction of the heat source`s mass, cost, and fuel loading, a 50% reduction of generator mass, a tripling of the power system`s specific power, and a quadrupling of its efficiency. The paper concludes by briefly summarizing the RTPV`s current technology status and assessing its potential applicability for the PFF mission. For other power systems (e.g., RTGs), demonstrating their flight readiness for a long mission is a very time-consuming process to determine the long-term effect of temperature-induced degradation mechanisms. But for the case of the described RTPV design, the paper lists a number of factors, primarily its cold (0 to 10 {degree}C) converter temperature, that may greatly reduce the need for long-term tests to demonstrate generator lifetime. In any event, our analytical results suggest that the RTPV generator, when developed by DOE and/or NASA, would be quite valuable not only for the Pluto mission but also for other future missions requiring small, long-lived, low-mass generators. {copyright} {ital 1995} {ital American} {ital Institute} {ital of} {ital Physics}.

  11. Analysis, Optimization, and Assessment of Radioisotope Thermophotovoltaic System Design for an Illustrative Space Mission

    SciTech Connect

    Schock, Alfred; Mukunda, Meera; Summers, G.

    1994-06-28

    A companion paper presented at this conference described the design of a Radioisotope Thermophotovoltaic (RTPV) Generator for an illustrative space mission (Pluto Fast Flyby). It presented a detailed design of an integrated system consisting of a radioisotope heat source, a thermophotovoltaic converter, and an optimized heat rejection system. The present paper describes the thermal, electrical, and structural analyses which led to that optimized design, and compares the computed RTPV performance to that of a Radioisotope Thermoelectric Generator (RTG) designed for the same mission. RTPV's are of course much less mature than RTGs, but our results indicate that - when fully developed - they could result in a 60% reduction of the heat source's mass, cost, and fuel loading, a 50% reduction of generator mass, a tripling of the power system's specific power, and a quadrupling of its efficiency. The paper concludes by briefly summarizing the RTPV's current technology status and assessing its potential applicability for the PFF mission. For other power systems (e.g. RTGs), demonstrating their flight readiness for a long mission is a very time-consuming process to determine the long-term effect of temperature-induced degradation mechanisms. But for the case of the described RTPV design, the paper lists a number of factors, primarily its cold (0 to 10 degrees C) converter temperature, that may greatly reduce the need for long-term tests to demonstrate generator lifetime. In any event, our analytical results suggest that the RTPV generator, when developed by DOE and/or NASA, would be quite valuable not only for the Pluto mission but also for other future missions requiring small, long-lived, low mass generators. Another copy is in the Energy Systems files.

  12. Identifying and attaining LDL-C goals: mission accomplished? Next target: new therapeutic options to raise HDL-C levels.

    PubMed

    Athyros, Vasilios G; Mikhailidis, Dimitri P; Kakafika, Anna I; Karagiannis, Asterios; Hatzitolios, Apostolos; Tziomalos, Konstantinos; Ganotakis, Emmanuel S; Liberopoulos, Evangelos N; Elisaf, Moses

    2007-03-01

    Currently, low density lipoprotein cholesterol (LDL-C) levels are the main, if not the only, lipid target in the effort to reduce cardiovascular disease (CVD) morbidity and mortality. Several primary and secondary CVD prevention trials with statins shaped current guidelines and provided detailed targets across a range of CVD risk categories. These targets can be attained using effective statins or combination therapy. However, the net benefit in CVD risk reduction may be improved if we address other lipid risk factors. High density lipoprotein cholesterol (HDL-C) emerges from epidemiological studies as the most promising target. This review links the increase in HDL-C levels with clinical benefit from "old" (e.g. sustained release niacin) and new treatment options. Synthetically produced recombined apolipoprotein A-I Milano administered intravenously seems to have a marked effect in reducing the atheroma burden. The anti-cholesterol ester transfer protein (CETP) vaccine (CETi-1) produces auto-antibodies against CETP thus increasing the cholesterol ester content in HDL particles. CETP inhibitors (e.g. JTT-705 and torcetrapib) seem to be the most promising regimen to increase HDL-C levels. Torcetrapib (already in phase IIIa studies) can substantially increase HDL-C levels (up to 106%), alone or in combination with atorvastatin. HDL-C strategies, in combination with effective statins, are a new drug target aimed at a further reduction in CVD morbidity and mortality compared with statin monotherapy.

  13. Evaluation of MPLM Design and Mission 6A Coupled Loads Analyses

    NASA Technical Reports Server (NTRS)

    Bookout, Paul S.; Ricks, Ed

    1999-01-01

    Through the development of a space shuttle payload, there are usually several coupled loads analyses (CLA) performed: preliminary design, critical design, final design and verification loads analysis (VLA). A final design CLA is the last analysis conducted prior to model delivery to the shuttle program for the VLA. The finite element models used in the final design CLA and the VLA are test verified dynamic math models. Mission 6A is the first of many flights of the Multi-Purpose Logistics Module (MPLM). The MPLM was developed by Alenia Spazio S.p.A. (an Italian aerospace company) and houses the International Standard Payload Racks (ISPR) for transportation to the space station in the shuttle. Marshall Space Flight Center (MSFC), the payload integrator of the MPLM for Mission 6A, performed the final design CLA using the M6.OZC shuttle data for liftoff and landing conditions using the proper shuttle cargo manifest. Alenia performed the preliminary and critical design CLAs for the development of the MPLM. However, these CLAs did not use the current Mission 6A cargo manifest. An evaluation of the preliminary and critical design performed by Alenia and the final design performed by MSFC is presented.

  14. Trajectory design for the Deep Space Program Science Experiment (DSPSE) mission

    NASA Technical Reports Server (NTRS)

    Carrington, D.; Carrico, J.; Jen, J.; Roberts, C.; Seacord, A.; Sharer, P.; Newman, L.; Richon, K.; Kaufman, B.; Middour, J.

    1993-01-01

    In 1994, the Deep Space Program Science Experiment (DSPSE) spacecraft will become the first spacecraft to perform, in succession, both a lunar orbiting mission and a deep-space asteroid encounter mission. The primary mission objective is to perform a long-duration flight-test of various new-technology lightweight components, such as sensors, in a deep-space environment. The mission has two secondary science objectives: to provide high-resolution imaging of the entire lunar surface for mapping purposes and flyby imaging of the asteroid 1620 Geographos. The DSPSE mission is sponsored by the Strategic Defense Initiative Organization (SDIO). As prime contractor, the Naval Research Laboratory (NRL) is building the spacecraft and will conduct mission operations. The Goddard Space Flight Center's (GSFC) Flight Dynamics Division is supporting NRL in the areas of The Deep Space Network (DSN) will provide tracking support. The DSPSE mission will begin with a launch from the Western Test Range in late January 1994. Following a minimum 1.5-day stay in a low-Earth parking orbit, a solid kick motor burn will boost DSPSE into an 18-day, 2.5-revolution phasing orbit transfer trajectory to the Moon. Two burns to insert DSPSE into a lunar polar orbit suitable for the mapping mission will be followed by mapping orbit maintenance and adjustment operations over a period of 2 sidereal months. In May 1994, a lunar orbit departure maneuver, in conjunction with a lunar swingby 26 days later, will propel DSPSE onto a heliocentric transfer that will intercept Geographos on September 1, 1994. This paper presents the characteristics, deterministic delta-Vs, and design details of each trajectory phase of this unique mission, together with the requirements, constraints, and design considerations to which each phase is subject. Numerous trajectory plots and tables of significant trajectory events are included. Following a discussion of the results of a preliminary launch window analysis, a

  15. 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.

  16. Orbit Design for Phase I and II of the Magnetospheric Multiscale Mission (MMS)

    NASA Technical Reports Server (NTRS)

    Hughes, Steve P.

    2003-01-01

    The Magnetospheric Multiscale Mission (MMS) is a NASA mission intended to make fundamental advancements in our understanding of the Earth's Magnetosphere. There are three processes that MMS is intended to study including magnetic recon- nection, charged particle acceleration, and turbulence. There are four phases of the MMS mission and each phase is designed to study a particular region of the Earth's magnetosphere. The mission is composed of a formation of four spacecraft that are nominally in a regular tetrahedron formation. In this work, we present optimal orbit designs for Phase I and II. This entails designing optimal reference orbits so that the spacecraft dwell-time in the region of interest is a maximum. This is non-trivial because the Earth's magnetosphere is dynamic and its shape and position are not constant in inertial space. Optimal orbit design for MMS also entails designing the formation so that the relative motion of the four spacecraft yields the greatest science return. We develop performance metrics that are related to the science return, and use Sequential Quadratic Programming (SQP) to determine optimal relative motion solutions. We also ensure that practical constraints such as maximum eclipse time and minimum inter-spacecraft separation distances are not violated.

  17. Conceptual Design of a Hypervelocity Asteroid Intercept Vehicle (HAIV) Flight Validation Mission

    NASA Technical Reports Server (NTRS)

    Barbee, Brent W.; Wie, Bong; Steiner, Mark; Getzandanner, Kenneth

    2013-01-01

    In this paper we present a detailed overview of the MDL study results and subsequent advances in the design of GNC algorithms for accurate terminal guidance during hypervelocity NEO intercept. The MDL study produced a conceptual con guration of the two-body HAIV and its subsystems; a mission scenario and trajectory design for a notional flight validation mission to a selected candidate target NEO; GNC results regarding the ability of the HAIV to reliably intercept small (50 m) NEOs at hypervelocity (typically greater than 10 km/s); candidate launch vehicle selection; a notional operations concept and cost estimate for the flight validation mission; and a list of topics to address during the remainder of our NIAC Phase II study.

  18. Life-cycle cost analysis of energy efficiency design options for residential furnaces and boilers

    SciTech Connect

    Lutz, James; Lekov, Alex; Whitehead, Camilla Dunham; Chan, Peter; Meyers,Steve; McMahon, James

    2004-01-20

    In 2001, the U.S. Department of Energy (DOE) initiated a rulemaking process to consider whether to amend the existing energy efficiency standards for furnaces and boilers. A key factor in DOE's consideration of new standards is the economic impacts on consumers of possible revisions to energy-efficiency standards. Determining cost-effectiveness requires an appropriate comparison of the additional first cost of energy efficiency design options with the savings in operating costs. DOE's preferred approach involves comparing the total life-cycle cost (LCC) of owning and operating a more efficient appliance with the LCC for a baseline design. This study describes the method used to conduct the LCC analysis and presents the estimated change in LCC associated with more energy-efficient equipment. The results indicate that efficiency improvement relative to the baseline design can reduce the LCC in each of the product classes considered.

  19. Energy Efficiency Design Options for Residential Water Heaters: Economic Impacts on Consumers

    SciTech Connect

    Lekov, Alex; Franco, Victor; Meyers, Steve; Thompson, Lisa; Letschert, Virginie

    2010-11-24

    The U.S. Department of Energy (DOE) recently completed a rulemaking process in which it amended the existing energy efficiency standards for residential water heaters. A key factor in DOE?s consideration of new standards is the economic impacts on consumers. Determining such impacts requires a comparison of the additional first cost of energy efficiency design options with the savings in operating costs. This paper describes the method used to conduct the life-cycle cost (LCC) and payback period analysis for gas and electric storage water heaters. It presents the estimated change in LCC associated with more energy-efficient equipment, including heat pump electric water heaters and condensing gas water heaters, for a representative sample of U.S. homes. The study included a detailed accounting of installation costs for the considered design options, with a focus on approaches for accommodating the larger dimensions of more efficient water heaters. For heat pump water heaters, the study also considered airflow requirements, venting issues, and the impact of these products on the indoor environment. The results indicate that efficiency improvement relative to the baseline design reduces the LCC in the majority of homes for both gas and electric storage water heaters, and heat pump electric water heaters and condensing gas water heaters provide a lower LCC for homes with large rated volume water heaters.

  20. IXV avionics architecture: Design, qualification and mission results

    NASA Astrophysics Data System (ADS)

    Succa, Massimo; Boscolo, Ilario; Drocco, Alessandro; Malucchi, Giovanni; Dussy, Stephane

    2016-07-01

    The paper details the IXV avionics presenting the architecture and the constituting subsystems and equipment. It focuses on the novelties introduced, such as the Ethernet-based protocol for the experiment data acquisition system, and on the synergy with Ariane 5 and Vega equipment, pursued in order to comply with the design-to-cost requirement for the avionics system development. Emphasis is given to the adopted model philosophy in relation to OTS/COTS items heritage and identified activities necessary to extend the qualification level to be compliant with the IXV environment. Associated lessons learned are identified. Then, the paper provides the first results and interpretation from the flight recorders telemetry, covering the behavior of the Data Handling System, the quality of telemetry recording and real-time/delayed transmission, the performance of the batteries and the Power Protection and Distribution Unit, the ground segment coverage during visibility windows and the performance of the GNC sensors (IMU and GPS) and actuators. Finally, some preliminary tracks of the IXV follow on are given, introducing the objectives of the Innovative Space Vehicle and the necessary improvements to be developed in the frame of PRIDE.

  1. Solar array conceptual design for the Halley's Comet ion drive mission, phase 2

    NASA Technical Reports Server (NTRS)

    Rayl, G. J.; Speight, K. M.; Stanhouse, R. W.

    1977-01-01

    Conceptual design studies were performed directed toward a high power, ultralightweight solar array, compatible with the requirements for the Halley's Comet Ion Drive Mission. A planar, rollup array design concept capable of producing 120 kW at 1 AU and 6 kW at 4.5 AU, and a concentrator, rollup array design concept capable of producing 60 kW at 1 AU and 15.5 kW at 4.5 AU evolved. Both arrays make maximum use of thin film, lightweight technology. The Halley's Comet spacecraft and mission requirements developed from preliminary definition to a more finalized and mature design. As solar array requirements were updated, conceptual design iterations were necessary to keep pace with the rapidly changing program objectives and goals. The Halley's Comet Mission program status and design approaches were reviewed and more realistic power requirements at 4.5 AU for the ion engines were established at the 12 to 16 kW range. This higher power necessitated a change from the planar array design to a concentrator array design in order to remain within suitable cost and weight objectives.

  2. Mission Design for Continental-Scale Carbon Cycle Applications

    NASA Astrophysics Data System (ADS)

    Gervin, J. C.; Esper, J.; McClain, C. R.; Hall, F. G.; Middleton, E. M.; Gregg, W. W.; Mannino, A.; Knox, R. G.; Dabney, P. W.; Huemmrich, K. F.; Wood, H. J.; Roberto, M.

    2003-12-01

    constituents. A hyperspectral instrument design capable of observing in the ultraviolet, in addition to the visible and near infrared spectrum, is essential to investigate the variability, dynamics and biogeochemical cycles of the world's coastal and open ocean regions. For both terrestrial and ocean carbon cycle science objectives, a hyperspectral geostationary sensor should enable the development of new remote sensing measurements for important but as yet unobservable variables, and with the overall goal of linking both terrestrial and ocean carbon cycle processes to climate variability. The GSFC Carbon Team has been pursuing a geosynchronous hyperspectral instrument, which would revolutionize our knowledge of biological processes on land, in the ocean, and along the coast by providing multiple, diurnal coverage. Preliminary studies in Goddard's Instrument Synthesis and Analysis Laboratory indicate that we can meet many of our science requirements: full spectral coverage (360-1000 nm); narrow bandwidths (5-10 nm); adequate ground resolution (100-200 m); and continental-scale coverage 4-6 times per day; all the while achieving a signal to noise ratio of between 500 and 1000 to 1. An innovative and bold focal plane design and a large mirror (1.8 meter diameter) will be required.

  3. Using NASA's Space Launch System to Enable Game Changing Science Mission Designs

    NASA Technical Reports Server (NTRS)

    Creech, Stephen D.

    2013-01-01

    NASA's Marshall Space Flight Center is directing efforts to build the Space Launch System (SLS), a heavy-lift rocket that will help restore U.S. leadership in space by carrying the Orion Multi-Purpose Crew Vehicle and other important payloads far beyond Earth orbit. Its evolvable architecture will allow NASA to begin with Moon fly-bys and then go on to transport humans or robots to distant places such as asteroids, Mars, and the outer solar system. Designed to simplify spacecraft complexity, the SLS rocket will provide improved mass margins and radiation mitigation, and reduced mission durations. These capabilities offer attractive advantages for ambitious missions such as a Mars sample return, by reducing infrastructure requirements, cost, and schedule. For example, if an evolved expendable launch vehicle (EELV) were used for a proposed mission to investigate the Saturn system, a complicated trajectory would be required with several gravity-assist planetary fly-bys to achieve the necessary outbound velocity. The SLS rocket, using significantly higher C3 energies, can more quickly and effectively take the mission directly to its destination, reducing trip times and cost. As this paper will report, the SLS rocket will launch payloads of unprecedented mass and volume, such as monolithic telescopes and in-space infrastructure. Thanks to its ability to co-manifest large payloads, it also can accomplish complex missions in fewer launches. Future analyses will include reviews of alternate mission concepts and detailed evaluations of SLS figures of merit, helping the new rocket revolutionize science mission planning and design for years to come.

  4. The Tohoku Medical Megabank Project: Design and Mission

    PubMed Central

    Kuriyama, Shinichi; Yaegashi, Nobuo; Nagami, Fuji; Arai, Tomohiko; Kawaguchi, Yoshio; Osumi, Noriko; Sakaida, Masaki; Suzuki, Yoichi; Nakayama, Keiko; Hashizume, Hiroaki; Tamiya, Gen; Kawame, Hiroshi; Suzuki, Kichiya; Hozawa, Atsushi; Nakaya, Naoki; Kikuya, Masahiro; Metoki, Hirohito; Tsuji, Ichiro; Fuse, Nobuo; Kiyomoto, Hideyasu; Sugawara, Junichi; Tsuboi, Akito; Egawa, Shinichi; Ito, Kiyoshi; Chida, Koichi; Ishii, Tadashi; Tomita, Hiroaki; Taki, Yasuyuki; Minegishi, Naoko; Ishii, Naoto; Yasuda, Jun; Igarashi, Kazuhiko; Shimizu, Ritsuko; Nagasaki, Masao; Koshiba, Seizo; Kinoshita, Kengo; Ogishima, Soichi; Takai-Igarashi, Takako; Tominaga, Teiji; Tanabe, Osamu; Ohuchi, Noriaki; Shimosegawa, Toru; Kure, Shigeo; Tanaka, Hiroshi; Ito, Sadayoshi; Hitomi, Jiro; Tanno, Kozo; Nakamura, Motoyuki; Ogasawara, Kuniaki; Kobayashi, Seiichiro; Sakata, Kiyomi; Satoh, Mamoru; Shimizu, Atsushi; Sasaki, Makoto; Endo, Ryujin; Sobue, Kenji; Yamamoto, Masayuki

    2016-01-01

    The Great East Japan Earthquake (GEJE) and resulting tsunami of March 11, 2011 gave rise to devastating damage on the Pacific coast of the Tohoku region. The Tohoku Medical Megabank Project (TMM), which is being conducted by Tohoku University Tohoku Medical Megabank Organization (ToMMo) and Iwate Medical University Iwate Tohoku Medical Megabank Organization (IMM), has been launched to realize creative reconstruction and to solve medical problems in the aftermath of this disaster. We started two prospective cohort studies in Miyagi and Iwate Prefectures: a population-based adult cohort study, the TMM Community-Based Cohort Study (TMM CommCohort Study), which will recruit 80 000 participants, and a birth and three-generation cohort study, the TMM Birth and Three-Generation Cohort Study (TMM BirThree Cohort Study), which will recruit 70 000 participants, including fetuses and their parents, siblings, grandparents, and extended family members. The TMM CommCohort Study will recruit participants from 2013 to 2016 and follow them for at least 5 years. The TMM BirThree Cohort Study will recruit participants from 2013 to 2017 and follow them for at least 4 years. For children, the ToMMo Child Health Study, which adopted a cross-sectional design, was also started in November 2012 in Miyagi Prefecture. An integrated biobank will be constructed based on the two prospective cohort studies, and ToMMo and IMM will investigate the chronic medical impacts of the GEJE. The integrated biobank of TMM consists of health and clinical information, biospecimens, and genome and omics data. The biobank aims to establish a firm basis for personalized healthcare and medicine, mainly for diseases aggravated by the GEJE in the two prefectures. Biospecimens and related information in the biobank will be distributed to the research community. TMM itself will also undertake genomic and omics research. The aims of the genomic studies are: 1) to construct an integrated biobank; 2) to return genomic

  5. The Tohoku Medical Megabank Project: Design and Mission.

    PubMed

    Kuriyama, Shinichi; Yaegashi, Nobuo; Nagami, Fuji; Arai, Tomohiko; Kawaguchi, Yoshio; Osumi, Noriko; Sakaida, Masaki; Suzuki, Yoichi; Nakayama, Keiko; Hashizume, Hiroaki; Tamiya, Gen; Kawame, Hiroshi; Suzuki, Kichiya; Hozawa, Atsushi; Nakaya, Naoki; Kikuya, Masahiro; Metoki, Hirohito; Tsuji, Ichiro; Fuse, Nobuo; Kiyomoto, Hideyasu; Sugawara, Junichi; Tsuboi, Akito; Egawa, Shinichi; Ito, Kiyoshi; Chida, Koichi; Ishii, Tadashi; Tomita, Hiroaki; Taki, Yasuyuki; Minegishi, Naoko; Ishii, Naoto; Yasuda, Jun; Igarashi, Kazuhiko; Shimizu, Ritsuko; Nagasaki, Masao; Koshiba, Seizo; Kinoshita, Kengo; Ogishima, Soichi; Takai-Igarashi, Takako; Tominaga, Teiji; Tanabe, Osamu; Ohuchi, Noriaki; Shimosegawa, Toru; Kure, Shigeo; Tanaka, Hiroshi; Ito, Sadayoshi; Hitomi, Jiro; Tanno, Kozo; Nakamura, Motoyuki; Ogasawara, Kuniaki; Kobayashi, Seiichiro; Sakata, Kiyomi; Satoh, Mamoru; Shimizu, Atsushi; Sasaki, Makoto; Endo, Ryujin; Sobue, Kenji; Tohoku Medical Megabank Project Study Group, The; Yamamoto, Masayuki

    2016-09-05

    The Great East Japan Earthquake (GEJE) and resulting tsunami of March 11, 2011 gave rise to devastating damage on the Pacific coast of the Tohoku region. The Tohoku Medical Megabank Project (TMM), which is being conducted by Tohoku University Tohoku Medical Megabank Organization (ToMMo) and Iwate Medical University Iwate Tohoku Medical Megabank Organization (IMM), has been launched to realize creative reconstruction and to solve medical problems in the aftermath of this disaster. We started two prospective cohort studies in Miyagi and Iwate Prefectures: a population-based adult cohort study, the TMM Community-Based Cohort Study (TMM CommCohort Study), which will recruit 80 000 participants, and a birth and three-generation cohort study, the TMM Birth and Three-Generation Cohort Study (TMM BirThree Cohort Study), which will recruit 70 000 participants, including fetuses and their parents, siblings, grandparents, and extended family members. The TMM CommCohort Study will recruit participants from 2013 to 2016 and follow them for at least 5 years. The TMM BirThree Cohort Study will recruit participants from 2013 to 2017 and follow them for at least 4 years. For children, the ToMMo Child Health Study, which adopted a cross-sectional design, was also started in November 2012 in Miyagi Prefecture. An integrated biobank will be constructed based on the two prospective cohort studies, and ToMMo and IMM will investigate the chronic medical impacts of the GEJE. The integrated biobank of TMM consists of health and clinical information, biospecimens, and genome and omics data. The biobank aims to establish a firm basis for personalized healthcare and medicine, mainly for diseases aggravated by the GEJE in the two prefectures. Biospecimens and related information in the biobank will be distributed to the research community. TMM itself will also undertake genomic and omics research. The aims of the genomic studies are: 1) to construct an integrated biobank; 2) to return genomic

  6. Applying Strategic Visualization(Registered Trademark) to Lunar and Planetary Mission Design

    NASA Astrophysics Data System (ADS)

    Frassanito, John R.; Cooke, D. R.

    2002-01-01

    NASA teams, such as the NASA Exploration Team (NEXT), utilize advanced computational visualization processes to develop mission designs and architectures for lunar and planetary missions. One such process, Strategic Visualization (trademark), is a tool used extensively to help mission designers visualize various design alternatives and present them to other participants of their team. The participants, which may include NASA, industry, and the academic community, are distributed within a virtual network. Consequently, computer animation and other digital techniques provide an efficient means to communicate top-level technical information among team members. Today,Strategic Visualization(trademark) is used extensively both in the mission design process within the technical community, and to communicate the value of space exploration to the general public. Movies and digital images have been generated and shown on nationally broadcast television and the Internet, as well as in magazines and digital media. In our presentation will show excerpts of a computer-generated animation depicting the reference Earth/Moon L1 Libration Point Gateway architecture. The Gateway serves as a staging corridor for human expeditions to the lunar poles and other surface locations. Also shown are crew transfer systems and current reference lunar excursion vehicles as well as the Human and robotic construction of an inflatable telescope array for deployment to the Sun/Earth Libration Point.

  7. Correlating Online Course Design with the Institutional Mission and Student Learning Outcomes

    ERIC Educational Resources Information Center

    Lang, Charlotte S.

    2013-01-01

    Purpose and Method of Study: The purpose of this study was to examine the quality of the design of the subject university's online general education courses based on the Quality Matters™ (QM) criteria and scoring rubric and correlate this evaluation with the implementation of the university's overall institutional mission and student learning…

  8. Web Consulting for Non-Academic Educational Missions: How Instructional Design Offers a Competitive Advantage

    ERIC Educational Resources Information Center

    Cates, Ward Mitchell; Mattke, Paige Hawkins

    2013-01-01

    Based on a recently completed study of education directors at science museums, this article addresses how design-and-development consultants might use those findings to enhance the way in which they propose and deliver Website services to non-academic organizations with either primary or complementary educational missions. After a very brief…

  9. Astronomy sortie missions definition study. Volume 3, book 2: Appendix: Design analysis and trade studies

    NASA Technical Reports Server (NTRS)

    1972-01-01

    Backup or supporting data for the design analyses and trade studies which defined the astronomy sortie missions are presented. The subjects discussed are: (1) configuration of space shuttle orbiter, (2) electronic subsystems, (3) electric power requirements, and (4) payload requirements. Mathematical models are developed to illustrate the orbital rendezvous capabilities.

  10. Mechanical Analysis of W78/88-1 Life Extension Program Warhead Design Options

    SciTech Connect

    Spencer, Nathan

    2014-09-01

    Life Extension Program (LEP) is a program to repair/replace components of nuclear weapons to ensure the ability to meet military requirements. The W78/88-1 LEP encompasses the modernization of two major nuclear weapon reentry systems into an interoperable warhead. Several design concepts exist to provide different options for robust safety and security themes, maximum non-nuclear commonality, and cost. Simulation is one capability used to evaluate the mechanical performance of the designs in various operational environments, plan for system and component qualification efforts, and provide insight into the survivability of the warhead in environments that are not currently testable. The simulation efforts use several Sandia-developed tools through the Advanced Simulation and Computing program, including Cubit for mesh generation, the DART Model Manager, SIERRA codes running on the HPC TLCC2 platforms, DAKOTA, and ParaView. Several programmatic objectives were met using the simulation capability including: (1) providing early environmental specification estimates that may be used by component designers to understand the severity of the loads their components will need to survive, (2) providing guidance for load levels and configurations for subassembly tests intended to represent operational environments, and (3) recommending design options including modified geometry and material properties. These objectives were accomplished through regular interactions with component, system, and test engineers while using the laboratory's computational infrastructure to effectively perform ensembles of simulations. Because NNSA has decided to defer the LEP program, simulation results are being documented and models are being archived for future reference. However, some advanced and exploratory efforts will continue to mature key technologies, using the results from these and ongoing simulations for design insights, test planning, and model validation.

  11. The design and fabrication of a prototype trash compacting unit. [for long duration space missions

    NASA Technical Reports Server (NTRS)

    1973-01-01

    A prototype trash compactor, that is compatible with the anticipated requirements of future long-term space missions, is described. Preliminary problem definition studies were conducted to identify typical types and quantities of waste materials to be expected from a typical mission. Bench-scale compaction tests were then conducted on typical waste materials to determine force/compaction curves. These data were used to design a boilerplate compactor that was fabricated to prove the feasibility of the basic design concept. A final design was then prepared from which the deliverable unit was fabricated. Design concepts are presented for suggested further development of the compactor, including a version that is capable of handling wet biodegradable wastes.

  12. Preliminary Assessment of a Neptune Aerocapture Mission Using an Integrated Design Tool

    NASA Technical Reports Server (NTRS)

    Gage, Peter J.; Wercinski, Paul F.

    1998-01-01

    Aerocapture is an efficient orbit insertion technique that uses the planet's atmosphere to decelerate an arriving spacecraft. With current technology and for vehicles of reasonable mass, it is the only technique that might deliver the high delta-V's required for insertion to orbits around the outer planets. Preliminary design studies for outer planet orbital missions must evaluate aerocapture strategies, and must therefore consider the coupling between vehicle geometry, aerodynamics, aerocapture trajectory, heating and thermal protection system mass. The analyses have been linked into an integrated design environment, with the critical parameters grouped in a global database. The designer is free to use single point evaluations, parametric variation, and numerical optimization to evaluate a range of vehicle shapes and insertion trajectories. The application of this design tool to a preliminary study for Neptune aerocapture has implications for the use of such computational environments for any atmospheric entry mission.

  13. Particle-hemodynamics simulations and design options for surgical reconstruction of diseased carotid artery bifurcations.

    PubMed

    Hyun, S; Kleinstreuer, C; Longest, P W; Chen, C

    2004-04-01

    Based on the hypothesis that aggravating hemodynamic factors play a key role in the onset of arterial diseases, the methodology of "virtual prototyping" of branching blood vessels was applied to diseased external carotid artery (ECA) segments. The goals were to understand the underlying particle-hemodynamics and to provide various geometric design options for improved surgical reconstruction based on the minimization of critical hemodynamic wall parameters (HWPs). First, a representative carotid artery bifurcation (CAB) and then CABs with stenosed ECAs, i.e., a distally occluded ECA and an ECA stump, were analyzed based on transient three-dimensional blood flow solutions, employing a user-enhanced commercial finite volume code. Specifically, the HWPs, i.e., oscillatory shear index, wall shear stress angle gradient, near-wall residence time of monocytes, and near-wall helicity angle difference were evaluated to compare the merits of each design option, including a reconstructed near-optimal junction which generates the lowest HWP-values. The results provide physical insight to the biofluid dynamics of branching blood vessels and guide vascular surgeons as well as stent manufacturers towards interventions leading to high sustained patency rates.

  14. Transfer Trajectory Design for the Mars Atmosphere and Volatile Evolution (MAVEN) Mission

    NASA Technical Reports Server (NTRS)

    Folta, David; Demcak, Stuart; Young, Brian; Berry, Kevin

    2013-01-01

    The Mars Atmosphere and Volatile Evolution (MAVEN) mission will determine the history of the loss of volatiles from the Martian atmosphere from a highly inclined elliptical orbit. MAVEN will launch from Cape Canaveral Air Force Station on an Atlas-V 401 during an extended 36-day launch period opening November 18, 2013. The MAVEN Navigation and Mission Design team performed a Monte Carlo analysis of the Type-II transfer to characterize; dispersions of the arrival B-Plane, trajectory correction maneuvers (TCMs), and the probability of Mars impact. This paper presents detailed analysis of critical MOI event coverage, maneuver constraints, deltaV-99 budgets, and Planetary Protection requirements.

  15. 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.

  16. Design of Spacecraft Missions to Test Kinetic Impact for Asteroid Deflection

    NASA Technical Reports Server (NTRS)

    Barbee, Brent W.; Hernandez, Sonia

    2012-01-01

    Earth has previously been struck with devastating force by near-Earth asteroids (NEAs) and will be struck again. Telescopic search programs aim to provide advance warning of such an impact, but no techniques or systems have yet been tested for deflecting an incoming NEA. To begin addressing this problem, we have analyzed the more than 8000 currently known NEAs to identify those that offer opportunities for safe and meaningful near-term tests of the proposed kinetic impact asteroid deflection technique. In this paper we present our methodology and results, including complete mission designs for the best kinetic impactor test mission opportunities.

  17. Human Health and Performance Aspects of Mars Design Reference Mission of July, 1997

    NASA Technical Reports Server (NTRS)

    Charles, John B.

    1999-01-01

    The human element is the most complex element of the mission design Mars missions will pose significant physiological and psychological challenges to crew members Some challenges (human engineering, life support) must be overcome (potential "non-starters") Some challenges (bone, radiation) may be show-stoppers ISS will only Indirectly address Mars questions before any "Go/No Go" decision Significant amount of ground-based and specialized flight research will be required -- Critical Path Roadmap project will direct HSLSPO's research toward Mars exploration objectives

  18. Habitability in Advanced Space Mission Design. Part 2; Evaluation of Habitation Elements

    NASA Technical Reports Server (NTRS)

    Adams, Constance M.; McCurdy, Matthew R.

    2000-01-01

    Habitability is a fundamental component of any long-duration human habitat. Due to the pressures on the crew and the criticality of their performance, this is particularly true of habitats or vehicles proposed for use in any human space mission of duration over 30 days. This paper, the second of three on this subject, will focus on evaluating all the vehicles currently under consideration for the Mars Design Reference Mission through application of metrics for habitability (proposed in a previous paper, see references Adams/McCurdy 1999).

  19. Multi-Objective Hybrid Optimal Control for Multiple-Flyby Interplanetary Mission Design Using Chemical Propulsion

    NASA Technical Reports Server (NTRS)

    Englander, Jacob; Vavrina, Matthew

    2015-01-01

    The customer (scientist or project manager) most often does not want just one point solution to the mission design problem Instead, an exploration of a multi-objective trade space is required. For a typical main-belt asteroid mission the customer might wish to see the trade-space of: Launch date vs. Flight time vs. Deliverable mass, while varying the destination asteroid, planetary flybys, launch year, etcetera. To address this question we use a multi-objective discrete outer-loop which defines many single objective real-valued inner-loop problems.

  20. Mission design software development at the University of Texas at Austin

    NASA Technical Reports Server (NTRS)

    Fowler, Wallace T.

    1993-01-01

    This paper describes the development process, the contents, the update process, and the various uses of a space mission planning FORTRAN subroutine library. This document is written by graduate (and undergraduate) students at the University of Texas at Austin and is used by students in several courses, primarily design courses. The library has been made available to faculty and students at several schools and was provided to students at the 1991 International Space University in Toulouse, France. This paper describes the mission library, its creation, its checking, its update procedure, and the teaching philosophy and procedures involved in its use.

  1. Designing Phase 2 for the Double-Lunar Swingby of the Magnetospheric Multiscale Mission (MMS)

    NASA Technical Reports Server (NTRS)

    Edery, Ariel

    2002-01-01

    The Magnetospheric Multiscale (MMS) mission is a formation flying mission that consists of four distinct phases: phases 1 and 2 are low-inclination highly elliptical orbits (HEO), phase 3 is a double-lunar swingby which transfers phase 2 to phase 4, a high-inclination HE0 orbit. Phase 2 is designed to reach the first lunar swingby of phase 3 in the most efficient fashion. It is shown that when the orientation of the line of apsides of phase 2 is properly chosen, no extra Delta-V is required beyond what is typically needed to raise apogee to lunar distance.

  2. Trajectory Design for MoonRise: A Proposed Lunar South Pole Aitken Basin Sample Return Mission

    NASA Astrophysics Data System (ADS)

    Parker, Jeffrey S.; McElrath, Timothy P.; Anderson, Rodney L.; Sweetser, Theodore H.

    2015-03-01

    This paper presents the mission design for the proposed MoonRise New Frontiers mission: a lunar far side lander and return vehicle, with an accompanying communication satellite. Both vehicles are launched together, but fly separate low-energy transfers to the Moon. The communication satellite enters lunar orbit immediately upon arrival at the Moon, whereas the lander enters a staging orbit about the lunar Lagrange points. The lander descends and touches down on the surface 17 days after the communication satellite enters orbit. The lander remains on the surface for nearly two weeks before lifting off and returning to Earth via a low-energy return.

  3. Proactive Integration of Planetary Protection Needs Into Early Design Phases of Human Exploration Missions

    NASA Astrophysics Data System (ADS)

    Race, Margaret; Conley, Catharine

    yet been developed. Looking ahead, it is recognized that these planetary protection policies will apply to both governmental and non-governmental entities for the more than 100 countries that are signatories to the Outer SpaceTreaty. Fortunately, planetary protection controls for human missions are supportive of many other important mission needs, such as maximizing closed-loop and recycling capabilities to minimize mass required, minimizing exposure of humans to planetary materials for multiple health reasons, and minimizing contamination of planetary samples and environments during exploration and science activities. Currently, there is progress on a number of fronts in translating the basic COSPAR PP Principles and Implementation Guidelines into information that links with early engineering and process considerations. For example, an IAA Study Group on Planetary Protection and Human Missions is engaging robotic and human mission developers and scientists in exploring detailed technical, engineering and operational approaches by which planetary protection objectives can be accomplished for human missions in synergism with robotic exploration and in view of other constraints. This on-going study aims to highlight important information for the early stages of planning, and identify key research and technology development (R&TD) areas for further consideration and work. Such R&TD challenges provide opportunities for individuals, institutions and agencies of emerging countries to be involved in international exploration efforts. In January 2014, the study group presented an Interim Report to the IAA Heads of Agencies Summit in Washington DC. Subsequently, the group has continued to work on expanding the initial technical recommendations and findings, focusing especially on information useful to mission architects and designers as they integrate PP considerations in their varied plans-- scientific, commercial and otherwise. Already the findings and recommendations

  4. Conceptual Design of a Flight Validation Mission for a Hypervelocity Asteroid Intercept Vehicle

    NASA Technical Reports Server (NTRS)

    Barbee, Brent W.; Wie, Bong; Steiner, Mark; Getzandanner, Kenneth

    2013-01-01

    Near-Earth Objects (NEOs) are asteroids and comets whose orbits approach or cross Earth s orbit. NEOs have collided with our planet in the past, sometimes to devastating effect, and continue to do so today. Collisions with NEOs large enough to do significant damage to the ground are fortunately infrequent, but such events can occur at any time and we therefore need to develop and validate the techniques and technologies necessary to prevent the Earth impact of an incoming NEO. In this paper we provide background on the hazard posed to Earth by NEOs and present the results of a recent study performed by the NASA/Goddard Space Flight Center s Mission Design Lab (MDL) in collaboration with Iowa State University s Asteroid Deflection Research Center (ADRC) to design a flight validation mission for a Hypervelocity Asteroid Intercept Vehicle (HAIV) as part of a Phase 2 NASA Innovative Advanced Concepts (NIAC) research project. The HAIV is a two-body vehicle consisting of a leading kinetic impactor and trailing follower carrying a Nuclear Explosive Device (NED) payload. The HAIV detonates the NED inside the crater in the NEO s surface created by the lead kinetic impactor portion of the vehicle, effecting a powerful subsurface detonation to disrupt the NEO. For the flight validation mission, only a simple mass proxy for the NED is carried in the HAIV. Ongoing and future research topics are discussed following the presentation of the detailed flight validation mission design results produced in the MDL.

  5. Conceptual design of a flight validation mission for a Hypervelocity Asteroid Intercept Vehicle

    NASA Astrophysics Data System (ADS)

    Barbee, Brent W.; Wie, Bong; Steiner, Mark; Getzandanner, Kenneth

    2015-01-01

    Near-Earth Objects (NEOs) are asteroids and comets whose orbits approach or cross Earth's orbit. NEOs have collided with our planet in the past, sometimes to devastating effect, and continue to do so today. Collisions with NEOs large enough to do significant damage to the ground are fortunately infrequent, but such events can occur at any time and we therefore need to develop and validate the techniques and technologies necessary to prevent the Earth impact of an incoming NEO. In this paper we provide background on the hazard posed to Earth by NEOs and present the results of a recent study performed by the NASA/Goddard Space Flight Center's Mission Design Lab (MDL) in collaboration with Iowa State University's Asteroid Deflection Research Center (ADRC) to design a flight validation mission for a Hypervelocity Asteroid Intercept Vehicle (HAIV) as part of a Phase 2 NASA Innovative Advanced Concepts (NIAC) research project. The HAIV is a two-body vehicle consisting of a leading kinetic impactor and trailing follower carrying a Nuclear Explosive Device (NED) payload. The HAIV detonates the NED inside the crater in the NEO's surface created by the lead kinetic impactor portion of the vehicle, effecting a powerful subsurface detonation to disrupt the NEO. For the flight validation mission, only a simple mass proxy for the NED is carried in the HAIV. Ongoing and future research topics are discussed following the presentation of the detailed flight validation mission design results produced in the MDL.

  6. The optimisation, design and verification of feed horn structures for future Cosmic Microwave Background missions

    NASA Astrophysics Data System (ADS)

    McCarthy, Darragh; Trappe, Neil; Murphy, J. Anthony; O'Sullivan, Créidhe; Gradziel, Marcin; Doherty, Stephen; Huggard, Peter G.; Polegro, Arturo; van der Vorst, Maarten

    2016-05-01

    In order to investigate the origins of the Universe, it is necessary to carry out full sky surveys of the temperature and polarisation of the Cosmic Microwave Background (CMB) radiation, the remnant of the Big Bang. Missions such as COBE and Planck have previously mapped the CMB temperature, however in order to further constrain evolutionary and inflationary models, it is necessary to measure the polarisation of the CMB with greater accuracy and sensitivity than before. Missions undertaking such observations require large arrays of feed horn antennas to feed the detector arrays. Corrugated horns provide the best performance, however owing to the large number required (circa 5000 in the case of the proposed COrE+ mission), such horns are prohibitive in terms of thermal, mechanical and cost limitations. In this paper we consider the optimisation of an alternative smooth-walled piecewise conical profiled horn, using the mode-matching technique alongside a genetic algorithm. The technique is optimised to return a suitable design using efficient modelling software and standard desktop computing power. A design is presented showing a directional beam pattern and low levels of return loss, cross-polar power and sidelobes, as required by future CMB missions. This design is manufactured and the measured results compared with simulation, showing excellent agreement and meeting the required performance criteria. The optimisation process described here is robust and can be applied to many other applications where specific performance characteristics are required, with the user simply defining the beam requirements.

  7. OSIRIS-REx Touch-And-Go (TAG) Mission Design and Analysis

    NASA Technical Reports Server (NTRS)

    Berry, Kevin; Sutter, Brian; May, Alex; Williams, Ken; Barbee, Brent W.; Beckman, Mark; Williams, Bobby

    2013-01-01

    The Origins Spectral Interpretation Resource Identification Security Regolith Explorer (OSIRIS-REx) mission is a NASA New Frontiers mission launching in 2016 to rendezvous with the near-Earth asteroid (101955) 1999 RQ36 in late 2018. After several months in formation with and orbit about the asteroid, OSIRIS-REx will fly a Touch-And-Go (TAG) trajectory to the asteroid s surface to obtain a regolith sample. This paper describes the mission design of the TAG sequence and the propulsive maneuvers required to achieve the trajectory. This paper also shows preliminary results of orbit covariance analysis and Monte-Carlo analysis that demonstrate the ability to arrive at a targeted location on the surface of RQ36 within a 25 meter radius with 98.3% confidence.

  8. Human Health and Performance Aspects of the Mars Design Reference Mission

    NASA Technical Reports Server (NTRS)

    Charles, John B.

    2000-01-01

    This paper will describe the current planning for exploration-class missions, emphasizing the medical, and human factors aspects of such expeditions. The details of mission architecture are still under study, but a typical Mars design reference mission comprises a six-month transit from Earth to Mar, eighteen months in residence on Mars, and a six-month transit back to Earth. Physiological stressors will include environmental factors such as prolonged exposure to radiation, weightlessness in transit, and hypogravity and a toxic atmosphere while on Mars. Psychological stressors will include remoteness from Earth, confinement, and potential interpersonal conflicts, all complicated by circadian alterations. Medical risks including trauma must also be considered. Results of planning for assuring human health and performance will be presented.

  9. Design Option of Heat Exchanger for the Next Generation Nuclear Plant

    SciTech Connect

    Eung Soo Kim; Chang Oh

    2008-09-01

    The Next Generation Nuclear Plant (NGNP), a very High temperature Gas-Cooled Reactor (VHTGRS) concept, will provide the first demonstration of a closed-loop Brayton cycle at a commercial scale of a few hundred megawatts electric and hydrogen production. The power conversion system (PCS) for the NGNP will take advantage of the significantly higher reactor outlet temperatures of the VHTGRS to provide higher efficiencies than can be achieved in the current generation of light water reactors. Besides demonstrating a system design that can be used directly for subsequent commercial deployment, the NGNP will demonstrate key technology elements that can be used in subsequent advanced power conversion systems for other Generation IV reactors. In anticipation of the design, development and procurement of an advanced power conversion system for the NGNP, the system integration of the NGNP and hydrogen plant was initiated to identify the important design and technology options that must be considered in evaluating the performance of the proposed NGNP. As part of the system integration of the VHTGRS and hydrogen production plant, the intermediate heat exchanger is used to transfer the process heat from VHTGRS to hydrogen plant. Therefore, the design and configuration of the intermediate heat exchanger are very important. This paper will include analysis of one stage versus two stage heat exchanger design configurations and thermal stress analyses of a printed circuit heat exchanger, helical coil heat exchanger, and shell/tube heat exchanger.

  10. BWR fuel design options for self-sustainable Th-{sup 233}U fuel cycle

    SciTech Connect

    Shaposhnik, Y.; Shwageraus, E.; Elias, E.

    2012-07-01

    In this work, we investigate a number of fuel assembly design options for a BWR core operating in a closed self-sustainable Th-{sup 233}U fuel cycle. The designs rely on axially heterogeneous fuel assembly structure in order to improve fertile to fissile conversion ratio. One of the main assumptions of the current study was to restrict the fuel assembly geometry to a single axial fissile zone 'sandwiched' between two fertile blanket zones. The main objective was to study the effect of the most important design parameters, such as dimensions of fissile and fertile zones and average void fraction, on the net breeding of {sup 233}U. The main design challenge in this respect is that the fuel breeding potential is at odds with axial power peaking and therefore limits the maximum achievable core power rating. The calculations were performed with BGCore system, which consists of MCNP code coupled with fuel depletion and thermo-hydraulic feedback modules. A single 3-dimensional fuel assembly with reflective radial boundaries was modeled applying simplified restrictions on maximum central line fuel temperature and Critical Power Ratio. It was found that axially heterogeneous fuel assembly design with single fissile zone can potentially achieve net breeding. In this case however, the achievable core power density is roughly one third of the reference BWR core. (authors)

  11. Advanced Spacecraft Designs in Support of Human Missions to Earth's Neighborhood

    NASA Technical Reports Server (NTRS)

    Fletcher, David

    2002-01-01

    NASA's strategic planning for technology investment draws on engineering studies of potential future missions. A number of hypothetical mission architectures have been studied. A recent study completed by The NASA/JSC Advanced Design Team addresses one such possible architecture strategy for missions to the moon. This conceptual study presents an overview of each of the spacecraft elements that would enable such missions. These elements include an orbiting lunar outpost at lunar L1 called the Gateway, a lunar transfer vehicle (LTV) which ferries a crew of four from the ISS to the Gateway, a lunar lander which ferries the crew from the Gateway to the lunar surface, and a one-way lunar habitat lander capable of supporting the crew for 30 days. Other supporting elements of this architecture discussed below include the LTV kickstage, a solar-electric propulsion (SEP) stage, and a logistics lander capable of re-supplying the 30-day habitat lander and bringing other payloads totaling 10.3 mt in support of surface mission activities. Launch vehicle infrastructure to low-earth orbit includes the Space Shuttle, which brings up the LTV and crew, and the Delta-IV Heavy expendable launch vehicle which launches the landers, kickstage, and SEP.

  12. Water Recovery System Design to Accommodate Dormant Periods for Manned Missions

    NASA Technical Reports Server (NTRS)

    Tabb, David; Carter, Layne

    2015-01-01

    Future manned missions beyond lower Earth orbit may include intermittent periods of extended dormancy. Under the NASA Advanced Exploration System (AES) project, NASA personnel evaluated the viability of the ISS Water Recovery System (WRS) to support such a mission. The mission requirement includes the capability for life support systems to support crew activity, followed by a dormant period of up to one year, and subsequently for the life support systems to come back online for additional crewed missions. Dormancy could be a critical issue due to concerns with microbial growth or chemical degradation that might prevent water systems from operating properly when the crewed mission began. As such, it is critical that the water systems be designed to accommodate this dormant period. This paper details the results of this evaluation, which include identification of dormancy issues, results of testing performed to assess microbial stability of pretreated urine during dormancy periods, and concepts for updating to the WRS architecture and operational concepts that will enable the ISS WRS to support the dormancy requirement.

  13. Design, qualification and operation of nuclear rockets for safe Mars missions

    SciTech Connect

    Buden, D.; Madsen, W.W.; Olson, T.S. ); Redd, L.R. )

    1993-01-01

    Nuclear thermal propulsion modules planned for use on crew missions to Mars improve mission reliability and overall safety of the mission. This, as well as all other systems, are greatly enhanced if the system specifications take into account safety from design initiation, and operational considerations are well thought through and applied. For instance, the use of multiple engines in the propulsion module can lead to very high system safety and reliability. Operational safety enhancements may include: the use of multiple perigee burns, thus allowing time to ensure that all systems are functioning properly prior to departure from Earth orbit; the ability to perform all other parts of the mission in a degraded mode with little or no degradation of the mission; and the safe disposal of the nuclear propulsion module in a heliocentric orbit out of the ecliptic plane. The standards used to qualify nuclear rockets are one of the main cost drivers of the program. Concepts and systems that minimize cost and risk will rely on use of the element and component levels to demonstrate technology readiness and validation. Subsystem or systems testing then is only needed for verification of performance. Also, these will be the safest concepts because they will be more thoroughly understood and the safety margins will be well established and confirmed by tests.

  14. Design, qualification and operation of nuclear rockets for safe Mars missions

    SciTech Connect

    Buden, D.; Madsen, W.W.; Olson, T.S.; Redd, L.R.

    1993-04-01

    Nuclear thermal propulsion modules planned for use on crew missions to Mars improve mission reliability and overall safety of the mission. This, as well as all other systems, are greatly enhanced if the system specifications take into account safety from design initiation, and operational considerations are well thought through and applied. For instance, the use of multiple engines in the propulsion module can lead to very high system safety and reliability. Operational safety enhancements may include: the use of multiple perigee burns, thus allowing time to ensure that all systems are functioning properly prior to departure from Earth orbit; the ability to perform all other parts of the mission in a degraded mode with little or no degradation of the mission; and the safe disposal of the nuclear propulsion module in a heliocentric orbit out of the ecliptic plane. The standards used to qualify nuclear rockets are one of the main cost drivers of the program. Concepts and systems that minimize cost and risk will rely on use of the element and component levels to demonstrate technology readiness and validation. Subsystem or systems testing then is only needed for verification of performance. Also, these will be the safest concepts because they will be more thoroughly understood and the safety margins will be well established and confirmed by tests.

  15. Venus, Mars, and the ices on Mercury and the moon: astrobiological implications and proposed mission designs.

    PubMed

    Schulze-Makuch, Dirk; Dohm, James M; Fairén, Alberto G; Baker, Victor R; Fink, Wolfgang; Strom, Robert G

    2005-12-01

    Venus and Mars likely had liquid water bodies on their surface early in the Solar System history. The surfaces of Venus and Mars are presently not a suitable habitat for life, but reservoirs of liquid water remain in the atmosphere of Venus and the subsurface of Mars, and with it also the possibility of microbial life. Microbial organisms may have adapted to live in these ecological niches by the evolutionary force of directional selection. Missions to our neighboring planets should therefore be planned to explore these potentially life-containing refuges and return samples for analysis. Sample return missions should also include ice samples from Mercury and the Moon, which may contain information about the biogenic material that catalyzed the early evolution of life on Earth (or elsewhere). To obtain such information, science-driven exploration is necessary through varying degrees of mission operation autonomy. A hierarchical mission design is envisioned that includes spaceborne (orbital), atmosphere (airborne), surface (mobile such as rover and stationary such as lander or sensor), and subsurface (e.g., ground-penetrating radar, drilling, etc.) agents working in concert to allow for sufficient mission safety and redundancy, to perform extensive and challenging reconnaissance, and to lead to a thorough search for evidence of life and habitability.

  16. Solar Orbiter- Solar Array- Thermal Design for an Extreme Temperature Mission

    NASA Astrophysics Data System (ADS)

    Muller, Jens; Paarmann, Carola; Lindner, Anton; Kreutz, Martin; Oberhuttinger, Carola; Costello, Ian; Icardi, Lidia

    2014-08-01

    The Solar Orbiter mission is an interdisciplinary mission to the sun, carried out by ESA in collaboration with NASA. The spacecraft will approach the sun close to 0.28 AU. At this distance, the solar array has to be operated under high solar array inclination angles to limit the temperatures to a maximum qualification temperature of 200°C for the photo voltaic assembly (PVA). Nevertheless, extreme temperatures appear at specific locations of the solar array which require purpose-built temperature protection measures. A very specific thermal protection is needed to keep the PVA and its supporting structures within the qualified temperature range and simultaneously minimize the thermal flux into the spacecraft.This paper describes the Solar Orbiter solar array design in general and its specific thermal design in particular. It describes the interdisciplinary steps between thermal- and mechanical analysis as well as design engineering necessary to result to the different shielding methods.

  17. Concurrent Mission and Systems Design at NASA Glenn Research Center: The Origins of the COMPASS Team

    NASA Technical Reports Server (NTRS)

    McGuire, Melissa L.; Oleson, Steven R.; Sarver-Verhey, Timothy R.

    2012-01-01

    Established at the NASA Glenn Research Center (GRC) in 2006 to meet the need for rapid mission analysis and multi-disciplinary systems design for in-space and human missions, the Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) team is a multidisciplinary, concurrent engineering group whose primary purpose is to perform integrated systems analysis, but it is also capable of designing any system that involves one or more of the disciplines present in the team. The authors were involved in the development of the COMPASS team and its design process, and are continuously making refinements and enhancements. The team was unofficially started in the early 2000s as part of the distributed team known as Team JIMO (Jupiter Icy Moons Orbiter) in support of the multi-center collaborative JIMO spacecraft design during Project Prometheus. This paper documents the origins of a concurrent mission and systems design team at GRC and how it evolved into the COMPASS team, including defining the process, gathering the team and tools, building the facility, and performing studies.

  18. The initial design of LAPAN's IR micro bolometer using mission analysis process

    NASA Astrophysics Data System (ADS)

    Bustanul, A.; Irwan, P.; M. T., Andi; Firman, B.

    2016-11-01

    As new player in Infra Red (IR) sector, uncooled, small, and lightweight IR Micro Bolometer has been chosen as one of payloads for LAPAN's next micro satellite project. Driven the desire to create our own IR Micro Bolometer, mission analysis design procedure has been applied. After tracing all possible missions, the Planck's and Wien's Law for black body, Temperature Responsivity (TR), and sub-pixel response had been utilized in order to determine the appropriate spectral radiance. The 3.8 - 4 μm wavelength were available to detect wild fire (forest fire) and active volcanoes, two major problems faced by Indonesia. In order to strengthen and broaden the result, iteration process had been used throughout the process. The analysis, then, were continued by calculating Ground pixel size, IFOV pixel, swath width, and focus length. Meanwhile, regarding of resolution, at least it is 400 m. The further procedure covered the integrated of optical design, wherein we combined among optical design software, Zemax, with mechanical analysis software (structure and thermal analysis), such as Nastran and Thermal Desktop / Sinda Fluint. The integration process was intended to produce high performance optical system of our IR Micro Bolometer that can be used under extreme environment. The results of all those analysis, either in graphs or in measurement, show that the initial design of LAPAN'S IR Micro Bolometer meets the determined requirement. However, it needs the further evaluation (iteration). This paper describes the initial design of LAPAN's IR Micro Bolometer using mission analysis process

  19. Design considerations for a dedicated gravity recovery satellite mission consisting of two pairs of satellites

    NASA Astrophysics Data System (ADS)

    Wiese, D. N.; Nerem, R. S.; Lemoine, F. G.

    2012-02-01

    Future satellite missions dedicated to measuring time-variable gravity will need to address the concern of temporal aliasing errors; i.e., errors due to high-frequency mass variations. These errors have been shown to be a limiting error source for future missions with improved sensors. One method of reducing them is to fly multiple satellite pairs, thus increasing the sampling frequency of the mission. While one could imagine a system architecture consisting of dozens of satellite pairs, this paper explores the more economically feasible option of optimizing the orbits of two pairs of satellites. While the search space for this problem is infinite by nature, steps have been made to reduce it via proper assumptions regarding some parameters and a large number of numerical simulations exploring appropriate ranges for other parameters. A search space originally consisting of 15 variables is reduced to two variables with the utmost impact on mission performance: the repeat period of both pairs of satellites (shown to be near-optimal when they are equal to each other), as well as the inclination of one of the satellite pairs (the other pair is assumed to be in a polar orbit). To arrive at this conclusion, we assume circular orbits, repeat groundtracks for both pairs of satellites, a 100-km inter-satellite separation distance, and a minimum allowable operational satellite altitude of 290 km based on a projected 10-year mission lifetime. Given the scientific objectives of determining time-variable hydrology, ice mass variations, and ocean bottom pressure signals with higher spatial resolution, we find that an optimal architecture consists of a polar pair of satellites coupled with a pair inclined at 72°, both in 13-day repeating orbits. This architecture provides a 67% reduction in error over one pair of satellites, in addition to reducing the longitudinal striping to such a level that minimal post-processing is required, permitting a substantial increase in the spatial

  20. Design Considerations for a Dedicated Gravity Recovery Satellite Mission Consisting of Two Pairs of Satellites

    NASA Technical Reports Server (NTRS)

    Wiese, D. N.; Nerem, R. S.; Lemoine, F. G.

    2011-01-01

    Future satellite missions dedicated to measuring time-variable gravity will need to address the concern of temporal aliasing errors; i.e., errors due to high-frequency mass variations. These errors have been shown to be a limiting error source for future missions with improved sensors. One method of reducing them is to fly multiple satellite pairs, thus increasing the sampling frequency of the mission. While one could imagine a system architecture consisting of dozens of satellite pairs, this paper explores the more economically feasible option of optimizing the orbits of two pairs of satellites. While the search space for this problem is infinite by nature, steps have been made to reduce it via proper assumptions regarding some parameters and a large number of numerical simulations exploring appropriate ranges for other parameters. A search space originally consisting of 15 variables is reduced to two variables with the utmost impact on mission performance: the repeat period of both pairs of satellites (shown to be near-optimal when they are equal to each other), as well as the inclination of one of the satellite pairs (the other pair is assumed to be in a polar orbit). To arrive at this conclusion, we assume circular orbits, repeat groundtracks for both pairs of satellites, a 100-km inter-satellite separation distance, and a minimum allowable operational satellite altitude of 290 km based on a projected 10-year mission lifetime. Given the scientific objectives of determining time-variable hydrology, ice mass variations, and ocean bottom pressure signals with higher spatial resolution, we find that an optimal architecture consists of a polar pair of satellites coupled with a pair inclined at 72deg, both in 13-day repeating orbits. This architecture provides a 67% reduction in error over one pair of satellites, in addition to reducing the longitudinal striping to such a level that minimal post-processing is required, permitting a substantial increase in the spatial

  1. Mission and Instrument Design Trades for a Space-based Gravitational Wave Observatory to Maximize Science Return

    NASA Astrophysics Data System (ADS)

    Livas, Jeffrey; Baker, John; Stebbins, Robin; Thorpe, James; Larson, Shane; Sesana, Alberto

    2016-03-01

    A space-based gravitational wave observatory is required to access the rich array of astrophysical sources expected at frequencies between 0.0001 and 0.1 Hz. The European Space Agency (ESA) chose the Gravitational Universe as the science theme of its L3 launch opportunity. A call for mission proposals will be released soon after the completion of the LISA Pathfinder (LPF) mission. LPF is scheduled to start science operations in March 2016, and finish by the end of the year, so an optimized mission concept is needed now. There are a number of possible design choices for both the instrument and the mission. One of the goals for a good mission design is to maximize the science return while minimizing risk and keeping costs low. This presentation will review some of the main design choices for a LISA-like laser interferometry mission and the impact of these choices on cost, risk, and science return.

  2. The genetics of age-related macular degeneration (AMD)--Novel targets for designing treatment options?

    PubMed

    Grassmann, Felix; Fauser, Sascha; Weber, Bernhard H F

    2015-09-01

    Age-related macular degeneration (AMD) is a progressive disease of the central retina and the main cause of legal blindness in industrialized countries. Risk to develop the disease is conferred by both individual as well as genetic factors with the latter being increasingly deciphered over the last decade. Therapeutically, striking advances have been made for the treatment of the neovascular form of late stage AMD while for the late stage atrophic form of the disease, which accounts for almost half of the visually impaired, there is currently no effective therapy on the market. This review highlights our current knowledge on the genetic architecture of early and late stage AMD and explores its potential for the discovery of novel, target-guided treatment options. We reflect on current clinical and experimental therapies for all forms of AMD and specifically note a persisting lack of efficacy for treatment in atrophic AMD. We further explore the current insight in AMD-associated genes and pathways and critically question whether this knowledge is suited to design novel treatment options. Specifically, we point out that known genetic factors associated with AMD govern the risk to develop disease and thus may not play a role in its severity or progression. Treatments based on such knowledge appear appropriate rather for prevention than treatment of manifest disease. As a consequence, future research in AMD needs to be greatly focused on approaches relevant to the patients and their medical needs.

  3. Mission design and operations of a constellation of small satellites for remote sensing

    NASA Astrophysics Data System (ADS)

    Sorensen, Trevor C.; Pilger, Eric J.; Wood, Mark S.; Nunes, Miguel A.; Yoneshige, Lance K.

    2013-05-01

    The Hawaii Space Flight Laboratory (HSFL) at the University of Hawaii at Manoa is developing the capabilities to design, build, and operate constellations of small satellites than can be tailored to efficiently execute a variety of remote sensing missions. With the Operationally Responsive Space (ORS) Office, HSFL is developing the Super Strypi launch vehicle that on its initial mission in 2013 will launch the HSFL 55-kg HawaiiSat-1 into a near polar orbit, providing the first deployment of these technologies. This satellite will be carrying a miniature hyperspectral thermal imager developed by the Hawaii Institute of Geophysics and Planetology (HIGP). HSFL has also developed a method to efficiently deploy a constellation of small satellites using a minimal number of launch vehicles. Under a three-year NASA grant, HSFL is developing a Comprehensive Open-architecture Space Mission Operations System (COSMOS) to support these types of missions. COSMOS is being designed as a System of Systems (SoS) software integrator, tying together existing elements from different technological domains. This system should be easily adaptable to new architectures and easily scalable. It will be provided as Open Source to qualified users, so will be adoptable by even universities with very restricted budgets. In this paper we present the use of COSMOS as a System of Systems integrator for satellite constellations of up to 100 satellites and numerous ground stations and/or contact nodes, including a fully automated "lights out" satellite contact capability.

  4. Space Technology 5: Changing the Mission Design without Changing the Hardware

    NASA Technical Reports Server (NTRS)

    Carlisle, Candace C.; Webb, Evan H.; Slavin, James A.

    2005-01-01

    The Space Technology 5 (ST-5) Project is part of NASA's New Millennium Program. The validation objectives are to demonstrate the research-quality science capability of the ST-5 spacecraft; to operate the three spacecraft as a constellation; and to design, develop, test and flight-validate three capable micro-satellites with new technologies. A three-month flight demonstration phase is planned, beginning in March 2006. This year, the mission was re-planned for a Pegasus XL dedicated launch into an elliptical polar orbit (instead of the Originally-planned Geosynchronous Transfer Orbit.) The re-plan allows the mission to achieve the same high-level technology validation objectives with a different launch vehicle. The new mission design involves a revised science validation strategy, a new orbit and different communication strategy, while minimizing changes to the ST-5 spacecraft itself. The constellation operations concepts have also been refined. While the system engineers, orbit analysts, and operations teams were re-planning the mission, the implementation team continued to make progress on the flight hardware. Most components have been delivered, and the first spacecraft is well into integration and test.

  5. Design of impulsive Earth-Moon Halo transfers: lunar proximity and direct options

    NASA Astrophysics Data System (ADS)

    Zeng, Hao; Zhang, Jingrui

    2016-10-01

    Techniques associated with stable manifold and lunar flyby have been applied to the construction of optimal transfers to Earth-Moon L1 /L2 libration point orbits. Compared with traditional design methods and to reduce maneuver cost, the design process presents a detailed analysis on the effect of lunar proximity with multiple constraints. An accurate and fast design strategy for seeking an insertion point and modifying the stable manifold to satisfy these constraints is proposed. Combined this strategy with the differential correction algorithm, the optimal transfer trajectory can be determined from a low-Earth orbit to a halo orbit around the L1 /L2 libration point within a little computational time. Different amplitudes and insertion points of halo orbit in conjunction with various constraint conditions about lunar flyby are considered to deeply examine the efficiency and reliability of the design algorithm. Preliminary results indicate that the required mission cost has a significant correlation with lunar proximity constraints, and demonstrate that the method of constructing impulsive lunar halo transfer trajectories with multiple constraints is feasible.

  6. Designing the STS-134 Re-Rendezvous: A Preparation for Future Crewed Rendezvous Missions

    NASA Technical Reports Server (NTRS)

    Stuit, Timothy D.

    2011-01-01

    In preparation to provide the capability for the Orion spacecraft, also known as the Multi-Purpose Crew Vehicle (MPCV), to rendezvous with the International Space Station (ISS) and future spacecraft, a new suite of relative navigation sensors are in development and were tested on one of the final Space Shuttle missions to ISS. The National Aeronautics and Space Administration (NASA) commissioned a flight test of prototypes of the Orion relative navigation sensors on STS-134, in order to test their performance in the space environment during the nominal rendezvous and docking, as well as a re-rendezvous dedicated to testing the prototype sensors following the undocking of the Space Shuttle orbiter at the end of the mission. Unlike the rendezvous and docking at the beginning of the mission, the re-rendezvous profile replicates the newly designed Orion coelliptic approach trajectory, something never before attempted with the shuttle orbiter. Therefore, there were a number of new parameters that needed to be conceived of, designed, and tested for this rerendezvous to make the flight test successful. Additionally, all of this work had to be integrated with the normal operations of the ISS and shuttle and had to conform to the constraints of the mission and vehicles. The result of this work is a separation and rerendezvous trajectory design that would not only prove the design of the relative navigation sensors for the Orion vehicle, but also would serve as a proof of concept for the Orion rendezvous trajectory itself. This document presents the analysis and decision making process involved in attaining the final STS-134 re-rendezvous design.

  7. Conceptual design study for the use of COBE rocket engines on the Tropical Rainfall Measuring Mission

    NASA Technical Reports Server (NTRS)

    1992-01-01

    The objective of this conceptual design study is to verify that the Cosmic Background Explorer (COBE) Hydrazine Propulsion Subsystem (HPS) Rocket Engine Assembly (REA) will satisfy the Tropical Rainfall Measuring Mission (TRMM) requirements and to develop a preliminary thruster module design using the existing REAs. The performance of the COBE HPS 5 lbf thrusters meet the TRMM mission requirements. The preliminary design consists of a single 5 lbf REA REM which is isolation mounted to a spacecraft interface angle bracket (5 or 10 deg angle). The REM incorporates a catalyst bed heater and sensor assembly, and propellant thermal control is achieved by thermostatically controlled heaters on the thruster valves. A ROM cost of approx. $950 K has been estimated for the phase 2 program to finalize the design, fabricate, and test the hardware using mechanical thermostats for thermal control. In the event that solid state thermostats are used, the cost is estimated to be $160 K higher. A ROM cost is approx. $145 K is estimated to study the effects of using Japanese manufactured hydrazine for the TRMM mission.

  8. Rapid model-based inter-disciplinary design of a CubeSat mission

    NASA Astrophysics Data System (ADS)

    Lowe, C. J.; Macdonald, M.

    2014-12-01

    With an increase in the use of small, modular, resource-limited satellites for Earth orbiting applications, the benefit to be had from a model-based architecture that rapidly searches the mission trade-space and identifies near-optimal designs is greater than ever. This work presents an architecture that identifies trends between conflicting objectives (e.g. lifecycle cost and performance) and decision variables (e.g. orbit altitude and inclination) such that informed assessment can be made as to which design/s to take on for further analysis. The models within the architecture exploit analytic methods where possible, in order avoid computationally expensive numerical propagation, and achieve rapid convergence. Two mission cases are studied; the first is an Earth observation satellite and presents a trade-off between ground sample distance and revisit time over a ground target, given altitude as the decision variable. The second is a satellite with a generic scientific payload and shows a more involved trade-off, between data return to a ground station and cost of the mission, given variations in the orbit altitude, inclination and ground station latitude. Results of each case are presented graphically and it is clear that non-intuitive results are captured that would typically be missed using traditional, point-design methods, where only discrete scenarios are examined.

  9. Multi-Objective Hybrid Optimal Control for Multiple-Flyby Interplanetary Mission Design using Chemical Propulsion

    NASA Technical Reports Server (NTRS)

    Englander, Jacob A.; Vavrina, Matthew A.

    2015-01-01

    Preliminary design of high-thrust interplanetary missions is a highly complex process. The mission designer must choose discrete parameters such as the number of flybys and the bodies at which those flybys are performed. For some missions, such as surveys of small bodies, the mission designer also contributes to target selection. In addition, real-valued decision variables, such as launch epoch, flight times, maneuver and flyby epochs, and flyby altitudes must be chosen. There are often many thousands of possible trajectories to be evaluated. The customer who commissions a trajectory design is not usually interested in a point solution, but rather the exploration of the trade space of trajectories between several different objective functions. This can be a very expensive process in terms of the number of human analyst hours required. An automated approach is therefore very desirable. This work presents such an approach by posing the impulsive mission design problem as a multi-objective hybrid optimal control problem. The method is demonstrated on several real-world problems. Two assumptions are frequently made to simplify the modeling of an interplanetary high-thrust trajectory during the preliminary design phase. The first assumption is that because the available thrust is high, any maneuvers performed by the spacecraft can be modeled as discrete changes in velocity. This assumption removes the need to integrate the equations of motion governing the motion of a spacecraft under thrust and allows the change in velocity to be modeled as an impulse and the expenditure of propellant to be modeled using the time-independent solution to Tsiolkovsky's rocket equation [1]. The second assumption is that the spacecraft moves primarily under the influence of the central body, i.e. the sun, and all other perturbing forces may be neglected in preliminary design. The path of the spacecraft may then be modeled as a series of conic sections. When a spacecraft performs a close

  10. Design of a satellite end-to-end mission performance simulator for imaging spectrometers and its application to the ESA's FLEX/Sentinel-3 tandem mission

    NASA Astrophysics Data System (ADS)

    Vicent, Jorge; Sabater, Neus; Tenjo, Carolina; Acarreta, Juan R.; Manzano, María.; Rivera, Juan P.; Jurado, Pedro; Franco, Raffaella; Alonso, Luis; Moreno, Jose

    2015-09-01

    The performance analysis of a satellite mission requires specific tools that can simulate the behavior of the platform; its payload; and the acquisition of scientific data from synthetic scenes. These software tools, called End-to-End Mission Performance Simulators (E2ES), are promoted by the European Space Agency (ESA) with the goal of consolidating the instrument and mission requirements as well as optimizing the implemented data processing algorithms. Nevertheless, most developed E2ES are designed for a specific satellite mission and can hardly be adapted to other satellite missions. In the frame of ESA's FLEX mission activities, an E2ES is being developed based on a generic architecture for passive optical missions. FLEX E2ES implements a state-of-the-art synthetic scene generator that is coupled with dedicated algorithms that model the platform and instrument characteristics. This work will describe the flexibility of the FLEX E2ES to simulate complex synthetic scenes with a variety of land cover classes, topography and cloud cover that are observed separately by each instrument (FLORIS, OLCI and SLSTR). The implemented algorithms allows modelling the sensor behavior, i.e. the spectral/spatial resampling of the input scene; the geometry of acquisition; the sensor noises and non-uniformity effects (e.g. stray-light, spectral smile and radiometric noise); and the full retrieval scheme up to Level-2 products. It is expected that the design methodology implemented in FLEX E2ES can be used as baseline for other imaging spectrometer missions and will be further expanded towards a generic E2ES software tool.

  11. 17 CFR 33.4 - Designation as a contract market for the trading of commodity options.

    Code of Federal Regulations, 2010 CFR

    2010-04-01

    ... and disposition of option customer complaints, and, where applicable, the futures commission merchant..., notification thereof be given to the option grantor. (4) Require, with respect to all written option customer complaints, that each member futures commission merchant which engages in the offer or sale of...

  12. Mariner Missions

    NASA Astrophysics Data System (ADS)

    Snyder, C.; Murdin, P.

    2000-11-01

    Mariner was the name given to the earliest set of American space missions to explore the planets and to the spacecraft developed to carry them out. The missions were planned and executed by the JET PROPULSION LABORATORY (JPL) of the California Institute of Technology, which had been designated by the National Aeronautics and Space Administration (NASA) as its lead center for planetary missions....

  13. Design and Performance of Tropical Rainfall Measuring Mission (TRMM) Super NiCd Batteries

    NASA Technical Reports Server (NTRS)

    Ahmad, Anisa J.; Rao, Gopalakrishna M.; Jallice, Doris E.; Moran Vickie E.

    1999-01-01

    The Tropical Rainfall Measuring Mission (TRMM) is a joint mission between NASA and the National Space Development Agency (NASDA) of Japan. The observatory is designed to monitor and study tropical rainfall and the associated release of energy that helps to power the global atmospheric circulation shaping both weather and climate around the globe. The spacecraft was launched from Japan on November 27,1997 via the NASDA H-2 launch vehicle. The TRMM Power Subsystem is a Peak Power Tracking system that can support the maximum TRMM load of 815 watts at the end of its three year life. The Power Subsystem consists of two 50 Ampere Hour Super NiCd batteries, Gallium Arsenide Solar Array and the Power System Electronics. This paper describes the TRMM Power Subsystem, battery design, cell and battery ground test performance, and in-orbit battery operations and performance.

  14. Development of a Plastic Melt Waste Compactor for Space Missions Experiments and Prototype Design

    NASA Technical Reports Server (NTRS)

    Pace, Gregory; Wignarajah, Kanapathipillai; Pisharody, Suresh; Fisher, John

    2004-01-01

    This paper describes development at NASA Ames Research Center of a heat melt compactor that can be used on both near term and far term missions. Experiments have been performed to characterize the behavior of composite wastes that are representative of the types of wastes produced on current and previous space missions such as International Space Station, Space Shuttle, MIR and Skylab. Experiments were conducted to characterize the volume reduction, bonding, encapsulation and biological stability of the waste composite and also to investigate other key design issues such as plastic extrusion, noxious off-gassing and removal of the of the plastic waste product from the processor. The experiments provided the data needed to design a prototype plastic melt waste processor, a description of which is included in the paper.

  15. Navigation Design and Analysis for the Orion Earth-Moon Mission

    NASA Technical Reports Server (NTRS)

    DSouza, Christopher; Zanetti, Renato

    2014-01-01

    This paper details the design of the cislunar optical navigation system being proposed for the Orion Earth-Moon (EM) missions. In particular, it presents the mathematics of the navigation filter. The unmodeled accelerations and their characterization are detailed. It also presents the analysis that has been performed to understand the performance of the proposed system, with particular attention paid to entry flight path angle constraints and the delta-V performance.

  16. The L-/C-band feed design for the DSS 14 70-meter antenna (Phobos mission)

    NASA Technical Reports Server (NTRS)

    Stanton, P. H.; Reilly, H. F., Jr.

    1991-01-01

    A dual-frequency (1.668 and 5.01 GHz) feed was designed for the Deep Space Station (DSS) 14 70-m antenna to support the Soviet Phobos Mission. This antenna system was capable of supporting telemetry, two-way Doppler, and very long baseline interferometry (VLBI). VLBI and two-way Doppler information on the Phobos spacecraft was acquired with this antenna in 1989.

  17. The Mercury Thermal Environment As A Design Driver and A Scientific Objective of The Bepicolombo Mission

    NASA Astrophysics Data System (ADS)

    Perotto, V.; Malosti, T.; Martino, R.; Briccarello, M.; Anselmi, A.

    The thermal environment of Mercury is extremely severe and a strong design driver for any mission to the planet. The main factors are the large amount of energy both di- rectly received from the sun and reflected/re-emitted from the planet, and the variation of such energy with time. The total thermal flux received by an object in orbit or on the surface of Mercury is a combination of the above-mentioned contributions, weighted according to the orbit characteristics, or the morphology of the surface. For a lander mission, the problems are compounded by the uncertainty in the a-priori knowledge of the surface properties and morphology. The thermal design of the orbiting and land- ing elements of the BepiColombo mission has a major role in the Definition Study being carried out under ESA contract by a team led by Alenia Spazio. The project en- compasses a spacecraft in low, near-circular, polar orbit (Mercury Planetary Orbiter, MPO), a spacecraft in high-eccentricity, polar orbit (Mercury Magnetospheric Orbiter, MMO, provided by ISAS, Japan) and a lander (Mercury Surface Element, MSE). The approach to a feasible mission design must rely on several provisions. For the orbiting elements, the orientation of the orbit plane with respect to the line of apsides of the or- bit of Mercury is found to have a major effect on the achievable orbiter temperatures. The spacecraft configuration, and its attitude with respect to the planet and the sun, drive the accommodation of the scientific instruments. Once the optimal orientation, attitude and configuration are determined, specific thermal control solutions must be elaborated, to maintain all components including the instruments in the required tem- perature range. The objective is maximizing the scientific return under constraints such as the available on-board resources and the project budget. A major outcome of the study so far has been the specification of requirements for improved thermal con- trol technologies, which are

  18. Evolutionary Design of an X-Band Antenna for NASA's Space Technology 5 Mission

    NASA Technical Reports Server (NTRS)

    Lohn, Jason D.; Hornby, Gregory S.; Rodriguez-Arroyo, Adan; Linden, Derek S.; Kraus, William F.; Seufert, Stephen E.

    2003-01-01

    We present an evolved X-band antenna design and flight prototype currently on schedule to be deployed on NASA s Space Technology 5 spacecraft in 2004. The mission consists of three small satellites that wall take science measurements in Earth s magnetosphere. The antenna was evolved to meet a challenging set of mission requirements, most notably the combination of wide beamwidth for a circularly-polarized wave and wide bandwidth. Two genetic algorithms were used: one allowed branching an the antenna arms and the other did not. The highest performance antennas from both algorithms were fabricated and tested. A handdesigned antenna was produced by the contractor responsible for the design and build of the mission antennas. The hand-designed antenna is a quadrifilar helix, and we present performance data for comparison to the evolved antennas. As of this writing, one of our evolved antenna prototypes is undergoing flight qualification testing. If successful, the resulting antenna would represent the first evolved hardware in space, and the first deployed evolved antenna.

  19. 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.

  20. Performance and fuel-cycle cost analysis of one JANUS 30 conceptual design for several fuel-element-design options

    SciTech Connect

    Nurdin, M.; Matos, J.E.; Freese, K.E.

    1982-01-01

    The performance and fuel cycle costs for a 25 MW, JANUS 30 reactor conceptual design by INTERATOM, Federal Republic of Germany, for BATAN, Republic of Indonesia have been studied using 19.75% enriched uranium in four fuel element design options. All of these fuel element designs have either been proposed by INTERATOM for various reactors or are currently in use with 93% enriched uranium in reactors in the Federal Republic of Germany. Aluminide, oxide, and silicide fuels were studied for selected designs using the range of uranium densities that are either currently qualified or are being developed and demonstrated internationally. To assess the long-term fuel adaptation strategy as well as the present fuel acceptance, reactor performance and annual fuel cycle costs were computed for seventeen cases based on a representative end-of-cycle excess reactivity and duty factor. In addition, a study was made to provide data for evaluating the trade-off between the increased safety associated with thicker cladding and the economic penalty due to increased fuel consumption.

  1. Conceptual design of a crewed reusable space transportation system aimed at parabolic flights: stakeholder analysis, mission concept selection, and spacecraft architecture definition

    NASA Astrophysics Data System (ADS)

    Fusaro, Roberta; Viola, Nicole; Fenoglio, Franco; Santoro, Francesco

    2016-06-01

    This paper proposes a methodology to derive architectures and operational concepts for future earth-to-orbit and sub-orbital transportation systems. In particular, at first, it describes the activity flow, methods, and tools leading to the generation of a wide range of alternative solutions to meet the established goal. Subsequently, the methodology allows selecting a small number of feasible options among which the optimal solution can be found. For the sake of clarity, the first part of the paper describes the methodology from a theoretical point of view, while the second part proposes the selection of mission concepts and of a proper transportation system aimed at sub-orbital parabolic flights. Starting from a detailed analysis of the stakeholders and their needs, the major objectives of the mission have been derived. Then, following a system engineering approach, functional analysis tools as well as concept of operations techniques allowed generating a very high number of possible ways to accomplish the envisaged goals. After a preliminary pruning activity, aimed at defining the feasibility of these concepts, more detailed analyses have been carried out. Going on through the procedure, the designer should move from qualitative to quantitative evaluations, and for this reason, to support the trade-off analysis, an ad-hoc built-in mission simulation software has been exploited. This support tool aims at estimating major mission drivers (mass, heat loads, manoeuverability, earth visibility, and volumetric efficiency) as well as proving the feasibility of the concepts. Other crucial and multi-domain mission drivers, such as complexity, innovation level, and safety have been evaluated through the other appropriate analyses. Eventually, one single mission concept has been selected and detailed in terms of layout, systems, and sub-systems, highlighting also logistic, safety, and maintainability aspects.

  2. Conceptual design of a crewed reusable space transportation system aimed at parabolic flights: stakeholder analysis, mission concept selection, and spacecraft architecture definition

    NASA Astrophysics Data System (ADS)

    Fusaro, Roberta; Viola, Nicole; Fenoglio, Franco; Santoro, Francesco

    2017-03-01

    This paper proposes a methodology to derive architectures and operational concepts for future earth-to-orbit and sub-orbital transportation systems. In particular, at first, it describes the activity flow, methods, and tools leading to the generation of a wide range of alternative solutions to meet the established goal. Subsequently, the methodology allows selecting a small number of feasible options among which the optimal solution can be found. For the sake of clarity, the first part of the paper describes the methodology from a theoretical point of view, while the second part proposes the selection of mission concepts and of a proper transportation system aimed at sub-orbital parabolic flights. Starting from a detailed analysis of the stakeholders and their needs, the major objectives of the mission have been derived. Then, following a system engineering approach, functional analysis tools as well as concept of operations techniques allowed generating a very high number of possible ways to accomplish the envisaged goals. After a preliminary pruning activity, aimed at defining the feasibility of these concepts, more detailed analyses have been carried out. Going on through the procedure, the designer should move from qualitative to quantitative evaluations, and for this reason, to support the trade-off analysis, an ad-hoc built-in mission simulation software has been exploited. This support tool aims at estimating major mission drivers (mass, heat loads, manoeuverability, earth visibility, and volumetric efficiency) as well as proving the feasibility of the concepts. Other crucial and multi-domain mission drivers, such as complexity, innovation level, and safety have been evaluated through the other appropriate analyses. Eventually, one single mission concept has been selected and detailed in terms of layout, systems, and sub-systems, highlighting also logistic, safety, and maintainability aspects.

  3. Manned Mars Explorer project: Guidelines for a manned mission to the vicinity of Mars using Phobos as a staging outpost; schematic vehicle designs considering chemical and nuclear electric propulsion

    NASA Technical Reports Server (NTRS)

    Nolan, Sean; Neubek, Deb; Baxmann, C. J.

    1988-01-01

    The Manned Mars Explorer (MME) project responds to the fundamental problems of sending human beings to Mars in a mission scenario and schematic vehicle designs. The mission scenario targets an opposition class Venus inbound swingby for its trajectory with concentration on Phobos and/or Deimos as a staging base for initial and future Mars vicinity operations. Optional vehicles are presented as a comparison using nuclear electric power/propulsion technology. A Manned Planetary Vehicle and Crew Command Vehicle are used to accomplish the targeted mission. The Manned Planetary Vehicle utilizes the mature technology of chemical propulsion combined with an advanced aerobrake, tether and pressurized environment system. The Crew Command Vehicle is the workhorse of the mission performing many different functions including a manned Mars landing, and Phobos rendezvous.

  4. Parallel satellite orbital situational problems solver for space missions design and control

    NASA Astrophysics Data System (ADS)

    Atanassov, Atanas Marinov

    2016-11-01

    Solving different scientific problems for space applications demands implementation of observations, measurements or realization of active experiments during time intervals in which specific geometric and physical conditions are fulfilled. The solving of situational problems for determination of these time intervals when the satellite instruments work optimally is a very important part of all activities on every stage of preparation and realization of space missions. The elaboration of universal, flexible and robust approach for situation analysis, which is easily portable toward new satellite missions, is significant for reduction of missions' preparation times and costs. Every situation problem could be based on one or more situation conditions. Simultaneously solving different kinds of situation problems based on different number and types of situational conditions, each one of them satisfied on different segments of satellite orbit requires irregular calculations. Three formal approaches are presented. First one is related to situation problems description that allows achieving flexibility in situation problem assembling and presentation in computer memory. The second formal approach is connected with developing of situation problem solver organized as processor that executes specific code for every particular situational condition. The third formal approach is related to solver parallelization utilizing threads and dynamic scheduling based on "pool of threads" abstraction and ensures a good load balance. The developed situation problems solver is intended for incorporation in the frames of multi-physics multi-satellite space mission's design and simulation tools.

  5. Formation Tetrahedron Design for Phase I of the Magnetospheric Multiscale Mission

    NASA Technical Reports Server (NTRS)

    Hughes, Steven P.

    2003-01-01

    The Magnetospheric Multiscale Mission (MMS) is a NASA mission intended to make fundamental advancements in our understanding of the Earth's Magnetosphere. There are three processes that MMS will study including magnetic reconnection, charged particle acceleration, and turbulence. There are four phases in the nominal mission and this work addresses some of the outstanding issues in phase I. The nominal phase I orbit is 1.2 x 12 R(sub e) highly elliptic orbit with four spacecraft nominally forming a regular tetrahedron. In this paper we investigate the relative dynamics of the four MMS spacecraft about an assumed reference orbit. There are several tetrahedron dimensions required in Phase I of the mission and in this work we design optimal tetrahedrons for the 10 km baseline. The performance metric used in the optimization process is directly related to the science return, and is based on an extension of previous work performed by Glassmeier. The optimizer we use is a commercially available Sequential Quadratic Programming (SQP) routine. Multiple optimal solutions are found, and we characterize how the performance of the formation varies between different regions of the reference orbit.

  6. Towards Design of Sustainable Energy Systems in Developing Countries: Centralized and Localized Options

    NASA Astrophysics Data System (ADS)

    Kursun, Berrin

    Energy use in developing countries is projected to equal and exceed the demand in developed countries in the next five years. Growing concern about environmental problems, depletion and price fluctuation of fossil fuels pushes the efforts for meeting energy demand in an environmentally friendly and sustainable way. Hence, it is essential to design energy systems consisting of centralized and localized options that generate the optimum energy mix to meet this increasing energy demand in a sustainable manner. In this study, we try to answer the question, "How can the energy demand in Rampura village be met sustainably?" via two centralized clean coal (CCC) technology and three localized energy technology options analyzed. We perform the analysis of these energy technologies through joint use of donor-side analysis technique emergy analysis (EA) and user-side analysis technique life cycle assessment (LCA). Sustainability of such an energy combination depends on its reliance on renewable inputs rather than nonrenewable or purchased inputs. CCC technologies are unsustainable energy systems dependent on purchased external inputs almost 100%. However, increased efficiency and significantly lower environmental impacts of CCC technologies can lead to more environmentally benign utilization of coal as an energy source. CCC technologies supply electricity at a lower price compared to the localized energy options investigated. Localized energy options analyzed include multi-crystalline solar PV, floating drum biogas digester and downdraft biomass gasifier. Solar PV has the lowest water and land use, however, solar electricity has the highest price with a high global warming potential (GWP). Contrary to general opinion, solar electricity is highly non-renewable. Although solar energy is a 100% renewable natural resource, materials utilized in the production of solar panels are mostly non-renewable purchased inputs causing the low renewability of solar electricity. Best

  7. Requirements and design reference mission for the WFIRST/AFTA coronagraph instrument

    NASA Astrophysics Data System (ADS)

    Demers, Richard T.; Dekens, Frank; Calvet, Rob; Chang, Zensheu; Effinger, Robert; Ek, Eric; Hovland, Larry; Jones, Laura; Loc, Anthony; Nemati, Bijan; Noecker, Charley; Neville, Timothy; Pham, Hung; Rud, Mike; Tang, Hong; Villalvazo, Juan

    2015-09-01

    The WFIRST-AFTA coronagraph instrument takes advantage of AFTAs 2.4-meter aperture to provide novel exoplanet imaging science at approximately the same instrument cost as an Explorer mission. The AFTA coronagraph also matures direct imaging technologies to high TRL for an Exo-Earth Imager in the next decade. The coronagraph Design Reference Mission (DRM) optical design is based on the highly successful High Contrast Imaging Testbed (HCIT), with modifications to accommodate the AFTA telescope design, service-ability, volume constraints, and the addition of an Integral Field Spectrograph (IFS). In order to optimally satisfy the three science objectives of planet imaging, planet spectral characterization and dust debris imaging, the coronagraph is designed to operate in two different modes: Hybrid Lyot Coronagraph or Shaped Pupil Coronagraph. Active mechanisms change pupil masks, focal plane masks, Lyot masks, and bandpass filters to shift between modes. A single optical beam train can thus operate alternatively as two different coronagraph architectures. Structural Thermal Optical Performance (STOP) analysis predicts the instrument contrast with the Low Order Wave Front Control loop closed. The STOP analysis was also used to verify that the optical/structural/thermal design provides the extreme stability required for planet characterization in the presence of thermal disturbances expected in a typical observing scenario. This paper describes the instrument design and the flow down from science requirements to high level engineering requirements.

  8. Mission Design Evaluation Using Automated Planning for High Resolution Imaging of Dynamic Surface Processes from the ISS

    NASA Technical Reports Server (NTRS)

    Knight, Russell; Donnellan, Andrea; Green, Joseph J.

    2013-01-01

    A challenge for any proposed mission is to demonstrate convincingly that the proposed systems will in fact deliver the science promised. Funding agencies and mission design personnel are becoming ever more skeptical of the abstractions that form the basis of the current state of the practice with respect to approximating science return. To address this, we have been using automated planning and scheduling technology to provide actual coverage campaigns that provide better predictive performance with respect to science return for a given mission design and set of mission objectives given implementation uncertainties. Specifically, we have applied an adaptation of ASPEN and SPICE to the Eagle-Eye domain that demonstrates the performance of the mission design with respect to coverage of science imaging targets that address climate change and disaster response. Eagle-Eye is an Earth-imaging telescope that has been proposed to fly aboard the International Space Station (ISS).

  9. A Probabilistic Approach of Incorporating Safety and Reliability in System Designs for a Manned Mission to Mars

    NASA Technical Reports Server (NTRS)

    Railsback, Jan W.; Simion, George P.; Himel, Malcolm (Technical Monitor)

    1999-01-01

    Conceptual stages in mission design often lack the input of quantitative safety and reliability assessments, simply because failure rates or other data are not yet available for systems that have not yet been designed. Absence of such data should not, however, prevent the development of a quantitative risk models with placeholders for missing data. Functions (that is, actions the systems must perform) in mission design will eventually require system probabilities of success, and there could be much learned from surrogate data, adequately bounded in uncertainty, used in a large event tree model of a complex mission.

  10. Design of a Four Degree_of_Freedom Manipulator for Northern Light Mars Mission

    NASA Astrophysics Data System (ADS)

    Lee, Regina; Quine, Brendan; Sathiyanathan, Kartheephan; Roberts, Caroline

    Northern Light is a Canadian mission to Mars, currently developed by a team of engineers, scientists and industrial organizations. The mission objectives include scientific goals such as the search for life and water, preparation for a sample return and engineering goals including the demonstration of interplanetary travel, an entry, descent and landing system, a rover design, a manipulator/drilling system, and semi-autonomous control in remote operations. The Northern Light team at York University is developing a four degree-of-freedom manipulator system, specifically for this remote operation. The Northern Light manipulator system will be mounted directly on the lander (not on the rover), providing an opportunity to perform scientific missions directly from the lander. The drilling instrument, to be mounted on the manipulator, is currently under development by Dr. Tze Chuen Ng now with the help of Hong Kong's Polytechnics University. The operation concept is based on a “single command cycle” approach. The operation plans are designed to handle exceptions, failures and unforeseen events using local intelligence and a contingency planner.

  11. Exploration-Related Research on the International Space Station: Connecting Science Results to the Design of Future Missions

    NASA Technical Reports Server (NTRS)

    Rhatigan, Jennifer L.; Robinson, Julie A.; Sawin, Charles F.; Ahlf, Peter R.

    2005-01-01

    In January, 2004, the US President announced a vision for space exploration, and charged NASA with utilizing the International Space Station (ISS) for research and technology targeted at supporting the US space exploration goals. This paper describes: 1) what we have learned from the first four years of research on ISS relative to the exploration mission, 2) the on-going research being conducted in this regard, 3) our current understanding of the major exploration mission risks that the ISS can be used to address, and 4) current progress in realigning NASA s research portfolio for ISS to support exploration missions. Specifically, we discuss the focus of research on solving the perplexing problems of maintaining human health on long-duration missions, and the development of countermeasures to protect humans from the space environment, enabling long duration exploration missions. The interchange between mission design and research needs is dynamic, where design decisions influence the type of research needed, and results of research influence design decisions. The fundamental challenge to science on ISS is completing experiments that answer key questions in time to shape design decisions for future exploration. In this context, exploration-relevant research must do more than be conceptually connected to design decisions-it must become a part of the mission design process.

  12. Mission and system optimization of nuclear electric propulsion vehicles for lunar and Mars missions

    NASA Technical Reports Server (NTRS)

    Gilland, James H.

    1991-01-01

    The detailed mission and system optimization of low thrust electric propulsion missions is a complex, iterative process involving interaction between orbital mechanics and system performance. Through the use of appropriate approximations, initial system optimization and analysis can be performed for a range of missions. The intent of these calculations is to provide system and mission designers with simple methods to assess system design without requiring access or detailed knowledge of numerical calculus of variations optimizations codes and methods. Approximations for the mission/system optimization of Earth orbital transfer and Mars mission have been derived. Analyses include the variation of thruster efficiency with specific impulse. Optimum specific impulse, payload fraction, and power/payload ratios are calculated. The accuracy of these methods is tested and found to be reasonable for initial scoping studies. Results of optimization for Space Exploration Initiative lunar cargo and Mars missions are presented for a range of power system and thruster options.

  13. Environmental Controls and Life Support System (ECLSS) Design for a Multi-Mission Space Exploration Vehicle (MMSEV)

    NASA Technical Reports Server (NTRS)

    Stambaugh, Imelda; Baccus, Shelley; Buffington, Jessie; Hood, Andrew; Naids, Adam; Borrego, Melissa; Hanford, Anthony J.; Eckhardt, Brad; Allada, Rama Kumar; Yagoda, Evan

    2013-01-01

    Engineers at Johnson Space Center (JSC) are developing an Environmental Control and Life Support System (ECLSS) design for the Multi-Mission Space Exploration Vehicle (MMSEV). The purpose of the MMSEV is to extend the human exploration envelope for Lunar, Near Earth Object (NEO), or Deep Space missions by using pressurized exploration vehicles. The MMSEV, formerly known as the Space Exploration Vehicle (SEV), employs ground prototype hardware for various systems and tests it in manned and unmanned configurations. Eventually, the system hardware will evolve and become part of a flight vehicle capable of supporting different design reference missions. This paper will discuss the latest MMSEV ECLSS architectures developed for a variety of design reference missions, any work contributed toward the development of the ECLSS design, lessons learned from testing prototype hardware, and the plan to advance the ECLSS toward a flight design.

  14. Environmental Controls and Life Support System (ECLSS) Design for a Multi-Mission Space Exploration Vehicle (MMSEV)

    NASA Technical Reports Server (NTRS)

    Stambaugh, Imelda; Baccus, Shelley; Naids, Adam; Hanford, Anthony

    2012-01-01

    Engineers at Johnson Space Center (JSC) are developing an Environmental Control and Life Support System (ECLSS) design for the Multi-Mission Space Exploration Vehicle (MMSEV). The purpose of the MMSEV is to extend the human exploration envelope for Lunar, Near Earth Object (NEO), or Deep Space missions by using pressurized exploration vehicles. The MMSEV, formerly known as the Space Exploration Vehicle (SEV), employs ground prototype hardware for various systems and tests it in manned and unmanned configurations. Eventually, the system hardware will evolve and become part of a flight vehicle capable of supporting different design reference missions. This paper will discuss the latest MMSEV ECLSS architectures developed for a variety of design reference missions, any work contributed toward the development of the ECLSS design, lessons learned from testing prototype hardware, and the plan to advance the ECLSS toward a flight design.

  15. The results of the critical design of the mission instruments of GOSAT-2

    NASA Astrophysics Data System (ADS)

    Yajima, Yukie; Suto, Hiroshi; Yotsumoto, Kazuhiko; Miyakawa, Takehiro; Hashimoto, Makiko; Shiomi, Kei; Nakajima, Masakatsu; Hirabayashi, Takeshi

    2016-04-01

    The GOSAT-2 is the successor satellite to the GOSAT which is the satellite dedicated to the measurements of the greenhouse gases such as carbon dioxide and methane. GOSAT was launched in January of 2009 and has been operated for about seven years. The development of the GOSAT-2 has been continued for two years, and through the preliminary and critical design phase the detail of the design of the mission instruments were fixed as well as the bus system design. The mission instruments of the GOSAT-2 are TANSO-FTS-2 and TANSO-CAI-2. TANSO-FTS-2 is the Fourier Transform Spectrometer observing greenhouse gases such as Carbon Dioxide and Methane and TANSO-CAI-2 is the imager observing the aerosols and clouds to compensate the TANSO-FTS-2 data and to grasp the movements of the aerosols such as PM2.5. The mission instruments will adopt the same kinds of instruments as GOSAT. But some improvements will be carried. Based on the results of the preliminary design, the design had been refined in the critical design phase and the results of the design meets all of the requirements on the mission instruments derived from the mission requirements to understand CO2 and CH4 sources and sinks and carbon cycle precisely. To improve the measurement accuracy, the signal to noise ratio will be increased by the extension of the aperture size from 64mm to 73mm and cooling the after optics as well as the thermal detectors. And to increase the number of the useful data, GOSAT-2 will equip the function to avoid the clouds during the observation using the images obtained by the monitor camera in FTS. To observe the carbon monoxide, the 2.3μm observation channel will be added. This function will be realized by the extension of the 2.0μm observation band to 2.3μm. The pointing angle in the along track direction will be extend from 20 degrees of GOSAT to 40 degrees to expand the observation area over the ocean where the sun glint is observed. This will make it possible to increase the number

  16. Techniques for Assuring NASA Mission Success Using Redundancy and Multi-Functionality Designs

    NASA Technical Reports Server (NTRS)

    Shivers, Herb

    2010-01-01

    Topics include NASA centers around the country; 2009 highlights of significant successes in space transportation, exploration, and science; significant accomplishments; places to explore include Lagrange points, near-Earth objects, Mars and the Moon, and International Space Station research; Marshall's missions include propulsion and transportation systems, life support systems, and earth and space science spacecraft, systems, and operations; project lifecycle management model; motivation of avionics fault-tolerance, redundancy needs and concerns, redundancy versus reliability; parallel-series configurations; effect of adding redundancy on mission success; example of rules-based approach where reliability and safety interaction impacts design; impact of common cause failure; approach ot bottom-up reliability analysis; three factors that lead to redundant system failure; Apollo 13 multi-functional reliability and example; and mitigating the risk of single string spacecraft architecture;.

  17. Mission Analysis and Conceptual Design of a Space Ambulance Utilizing the Boeing X-37B Platform

    NASA Astrophysics Data System (ADS)

    Blaesser, Nathaniel James

    This paper discusses the concept of operations and modifications for a crew return vehicle from the International Space Station based on the Boeing X-37B platform. The modifications necessary include accommodating a three person crew, adding an environmental control and life support system and modifying the propulsion system to be more conducive to a fast evacuation. Two crew seating configurations were considered and traded for vehicle mass, crew envelope and ease of loading and unloading the injured patient. Computer models were generated to determine key parameters, and the models were validated using physical mock-ups. The mission is considerably shorter than the design reference mission, meaning that there are significant reductions in propellant mass and volume. The results showed that by utilizing the space saved from smaller propellant tanks, and combining the crew's oxygen supply with the propulsion system, the vehicle can be retooled without modifying the outer mold line..

  18. The design and realisation of the IXV Mission Analysis and Flight Mechanics

    NASA Astrophysics Data System (ADS)

    Haya-Ramos, Rodrigo; Blanco, Gonzalo; Pontijas, Irene; Bonetti, Davide; Freixa, Jordi; Parigini, Cristina; Bassano, Edmondo; Carducci, Riccardo; Sudars, Martins; Denaro, Angelo; Angelini, Roberto; Mancuso, Salvatore

    2016-07-01

    The Intermediate eXperimental Vehicle (IXV) is a suborbital re-entry demonstrator successfully launched in February 2015 focusing on the in-flight demonstration of a lifting body system with active aerodynamic control surfaces. This paper presents an overview of the Mission Analysis and Flight Mechanics of the IXV vehicle, which comprises computation of the End-to-End (launch to splashdown) design trajectories, characterisation of the Entry Corridor, assessment of the Mission Performances through Monte Carlo campaigns, contribution to the aerodynamic database, analysis of the Visibility and link budget from Ground Stations and GPS, support to safety analyses (off nominal footprints), specification of the Centre of Gravity box, selection of the Angle of Attack trim line to be flown and characterisation of the Flying Qualities performances. An initial analysis and comparison with the raw flight data obtained during the flight will be discussed and first lessons learned derived.

  19. TASTER: Trojan ASteroid Tour, Exploration and Rendezvous, a NASA Planetary Science Summer School Mission Design Exercise

    NASA Astrophysics Data System (ADS)

    Diaz-silva, R.; Sayanagi, K. M.; Gil, S.; Diniega, S.; Balcerski, J.; Benneke, B.; Carande, B.; Fraeman, A. A.; Hudson, J. S.; Guzewich, S. D.; Livi, R.; Nahm, A.; Potter, S.; Route, M.; Urban, K. D.; Vasisht, S.; Williams, B.; Budney, C. J.; Lowes, L. L.

    2011-12-01

    A detailed investigation of the Trojan asteroids occupying Jupiter's L4 and L5 Lagrangian points has been identified as a priority for future missions by the 2011 Planetary Science Decadal Survey. Observing these asteroids and getting clear measurements of their physical characteristics and composition may yield answers to fundamental questions relating to the early Solar System. In particular, Trojan asteroids are believed to harbor primordial material dating from the time of its formation. The source region for Trojans is still unknown; the Nice model predicts that some bodies may have originated in the primordial Kuiper belt and were subsequently scattered inward during the migration of Neptune and Uranus and settled in their current location. In alternative models, less radial scattering of small bodies would imply Trojans formed from material at a similar orbital distance to Jupiter. Determination of Trojan composition and structure will help identify their birth location, provide information about the impact history and subsequent evolution. Earth-based observations of size and surface characteristics are sparse; spectral measurements are unable to resolve composition (and show a puzzling lack of volatile signatures), indicating that close-range observation is needed. We present a mission design for a Trojan Tour and Rendezvous mission that is consistent with NASA's New Frontiers candidate recommended by the Decadal Survey, and which is the final result of the 2011 NASA-JPL Planetary Science Summer School Mission Design Exercise. Our proposed mission includes a tour phase that features a 500 km altitude fly-by of 1999 XS143. The spacecraft will then orbit and make detailed observations of 1919FD Agamemnon, a 167 km diameter asteroid located in the leading Lagrangian point (L4), from orbital altitudes of 1000 - 100 km over a 12 month nominal science data capture period. The mission's planned primary observations aim to (1) detect and identify volatile species

  20. TOPEX satellite option study

    NASA Technical Reports Server (NTRS)

    1982-01-01

    The basic design of the fleet satellite communication spacecraft (FLTSATCOM) can easily accommodate any of the three payload options for the ocean dynamic topography experiment (TOPEX). The principal mission requirements as well as the payload accommodations and communications systems needed for launching this payload are reviewed. The existing FLTSATCOM satellite design is identified and the approaches for the proposed propulsion system are described in addition to subsystems for mechanical; power; attitude and velocity control; and telemetry, tracking and control are described. The compatability of FLTSATCOM with the launch vehicle is examined and its capabilities vs TOPEX requirements are summarized. Undetermined changes needed to meet data storage, thermal control, and area to mass ratio requirements are discussed. Cost estimates are included for budgetary and planning purposes. The availability of the described design is assessed based on the continuing production of FLTSATCOM spacecraft during the schedule span planned for TOPEX.