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At least 973 records · Page 54
All sourcesNASA NTRSDOE OSTI
NASA NTRS2007

Biomedical Support of U.S. Extravehicular Activity

The world's first extravehicular activity (EVA) was performed by A. A. Leonov on March 18, 1965 during the Russian Voskhod-2 mission. The first US EVA was executed by Gemini IV astronaut Ed White on June 3, 1965, with an umbilical tether that included communications and an oxygen supply. A hand-held maneuvering unit (HHMU) also was used to test maneuverability during the brief EVA; however the somewhat stiff umbilical limited controlled movement. That constraint, plus difficulty returning through the vehicle hatch, highlighted the need for increased thermal control and improved EVA ergonomics. Clearly, requirements for a useful EVA were interrelated with the vehicle design. The early Gemini EVAs generated requirements for suits providing micro-meteor protection, adequate visual field and eye protection from solar visual and infrared radiation, gloves optimized for dexterity while pressurized, and thermal systems capable of protecting the astronaut while rejecting metabolic heat during high workloads. Subsequent Gemini EVAs built upon this early experience and included development of a portable environmental control and life support systems (ECLSS) and an astronaut maneuvering unit. The ECLSS provided a pressure vessel and controller with functional control over suit pressure, oxygen flow, carbon dioxide removal, humidity, and temperature control. Gemini EVA experience also identified the usefulness of underwater neutral buoyancy and altitude chamber task training, and the importance of developing reliable task timelines. Improved thermal management and carbon dioxide control also were required for high workload tasks. With the Apollo project, EVA activity was primarily on the lunar surface; and suit durability, integrated liquid cooling garments, and low suit operating pressures (3.75 pounds per square inch absolute [psia] or 25.8 kilopascal [kPa],) were required to facilitate longer EVAs with ambulation and significant physical workloads with average metabolic rates of 1000 BTU/hr and peaks of up to 2200 BTU/hr. Mobility was further augmented with the Lunar Roving Vehicle. The Apollo extravehicular mobility unit (EMU) was made up of over 15 components, ranging from a biomedical belt for capturing and transmitting biomedical data, urine and fecal containment systems, a liquid cooling garment, communications cap, a modular portable life support system (PLSS), a boot system, thermal overgloves, and a bubble helmet with eye protection. Apollo lunar astronauts performed successful EVAs on the lunar surface from a 5 psia (34.4 kPa) 100% oxygen environment in the Lunar Lander. A maximum of three EVAs were performed on any mission. For Skylab a modified A7LB suit, used for Apollo 15, was selected. The Skylab astronaut life support assembly (ALSA) provided umbilical support through the life support umbilical (LSU) and used open loop oxygen flow, rather than closed-loop as in Apollo missions. Thermal control was provided by liquid water circulated by spacecraft pumps and electrical power also was provided from the spacecraft via the umbilical. The cabin atmosphere of 5 psia (34.4 kPa), 70% oxygen, provided a normoxic atmosphere and because of the very low nitrogen partial pressures, no special protocols were required to protect against decompression sickness (DCS) as was the case with the Apollo spacecraft with a 5 psi, 100% oxygen environment.

Gernhardt, Michael L.↗
NASA NTRS2004

JPL Facilities and Software for Collaborative Design: 1994 - Present

The viewgraph presentation provides an overview of the history of the JPL Project Design Center (PDC) and, since 2000, the Center for Space Mission Architecture and Design (CSMAD). The discussion includes PDC objectives and scope; mission design metrics; distributed design; a software architecture timeline; facility design principles; optimized design for group work; CSMAD plan view, facility design, and infrastructure; and distributed collaboration tools.

project design center↗
NASA NTRS2001

Quantitative risk assessment system (QRAS)

A quantitative risk assessment system (QRAS) builds a risk model of a system for which risk of failure is being assessed, then analyzes the risk of the system corresponding to the risk model. The QRAS performs sensitivity analysis of the risk model by altering fundamental components and quantifications built into the risk model, then re-analyzes the risk of the system using the modifications. More particularly, the risk model is built by building a hierarchy, creating a mission timeline, quantifying failure modes, and building/editing event sequence diagrams. Multiplicities, dependencies, and redundancies of the system are included in the risk model. For analysis runs, a fixed baseline is first constructed and stored. This baseline contains the lowest level scenarios, preserved in event tree structure. The analysis runs, at any level of the hierarchy and below, access this baseline for risk quantitative computation as well as ranking of particular risks. A standalone Tool Box capability exists, allowing the user to store application programs within QRAS.

Weinstock, Robert M↗
NASA NTRS2008

In-situ Resource Utilization (ISRU) to Support the Lunar Outpost and the Rationale for Precursor Missions

One of the ways that the Constellation Program can differ from Apollo is to employ a live-off-the-land or In-Situ Resource Utilization (ISRU) supported architecture. The options considered over the past decades for using indigenous materials have varied considerably in terms of what resources to attempt to acquire, how much to acquire, and what the motivations are to acquiring these resources. The latest NASA concepts for supporting the lunar outpost have considered many of these plans and compared these options to customers requirements and desires. Depending on the architecture employed, ISRU technologies can make a significant contribution towards a sustainable and affordable lunar outpost. While extensive ground testing will reduce some mission risk, one or more flight demonstrations prior to the first crew's arrival will build confidence and increase the chance that outpost architects will include ISRU as part of the early outpost architecture. This presentation includes some of the options for using ISRU that are under consideration for the lunar outpost, the precursor missions that would support these applications, and a notional timeline to allow the lessons learned from the precursor missions to support outpost hardware designs.

Simon, Thomas M.↗
NASA NTRS2008

A Brief History of Meteoroid and Orbital Debris Shielding Technology for US Manned Spacecraft

Meteoroid and orbital debris shielding has played an important role from the beginning of manned spaceflight. During the early 60 s, meteoroid protection drove requirements for new meteor and micrometeoroid impact science. Meteoroid protection also stimulated advances in the technology of hypervelocity impact launchers and impact damage assessment methodologies. The first phase of meteoroid shielding assessments closed in the early 70 s with the end of the Apollo program. The second phase of meteoroid protection technology began in the early 80 s when it was determined that there is a manmade Earth orbital debris belt that poses a significant risk to LEO manned spacecraft. The severity of the Earth orbital debris environment has dictated changes in Space Shuttle and ISS operations as well as driven advances in shielding technology and assessment methodologies. A timeline of shielding technology and assessment methodology advances is presented along with a summary of risk assessment results.

Bjorkman, Michael D.↗
NASA NTRS2007

Mission Analysis for the Don Quijote Phase-A Study

The Don Quijote Phase-A study is a definition study funded by ESA and devoted to the analysis of the possibilities to deflect a Near Earth Object (NEO) in the range of 300-800 m diameter. DEIMOS Space S.L. and EADS Astrium have teamed up within this study to form one of the three consortia that have analyzed these aspects for ESA. Target asteroids for the mission are 1989 ML, 2002 AT4 and Apophis. This paper presents the mission analysis activities within the consortium providing: low-thrust interplanetary rendezvous Orbiter trajectories to the target asteroids, ballistic interplanetary trajectories for the Impactor, Orbiter arrival description at the asteroids, Orbiter stable orbits characterization at the asteroid, deflection determination by means of a Radio Science Experiment (RSE) as well as the mission timelines and overall mission scenarios.

Cano, Juan L.↗
NASA NTRS2007

Automated Target Planning for FUSE Using the SOVA Algorithm

The SOVA algorithm was originally developed under the Resilient Systems and Operations Project of the Engineering for Complex Systems Program from NASA s Aerospace Technology Enterprise as a conceptual framework to support real-time autonomous system mission and contingency management. The algorithm and its software implementation were formulated for generic application to autonomous flight vehicle systems, and its efficacy was demonstrated by simulation within the problem domain of Unmanned Aerial Vehicle autonomous flight management. The approach itself is based upon the precept that autonomous decision making for a very complex system can be made tractable by distillation of the system state to a manageable set of strategic objectives (e.g. maintain power margin, maintain mission timeline, and et cetera), which if attended to, will result in a favorable outcome. From any given starting point, the attainability of the end-states resulting from a set of candidate decisions is assessed by propagating a system model forward in time while qualitatively mapping simulated states into margins on strategic objectives using fuzzy inference systems. The expected return value of each candidate decision is evaluated as the product of the assigned value of the end-state with the assessed attainability of the end-state. The candidate decision yielding the highest expected return value is selected for implementation; thus, the approach provides a software framework for intelligent autonomous risk management. The name adopted for the technique incorporates its essential elements: Strategic Objective Valuation and Attainability (SOVA). Maximum value of the approach is realized for systems where human intervention is unavailable in the timeframe within which critical control decisions must be made. The Far Ultraviolet Spectroscopic Explorer (FUSE) satellite, launched in 1999, has been collecting science data for eight years.[1] At its beginning of life, FUSE had six gyros in two IRUs and four reaction wheels. Over time through various failures, the satellite has been left with one reaction wheel on the vehicle skew axis and two gyros. To remain operational, a control scheme has been implemented using the magnetic torque rods and the remaining momentum wheel.[2] As a consequence, there are attitude regions where there is insufficient torque authority to overcome environmental disturbances (e.g. gravity gradient torques). The situation is further complicated by the fact that these attitude regions shift inertially with time as the spacecraft moves through earth s magnetic field during the course of its orbit. Under these conditions, the burden of planning targets and target-to-target slew maneuvers has increased significantly since the beginning of the mission.[3] Individual targets must be selected so that the magnetic field remains roughly aligned with the skew wheel axis to provide enough control authority to the other two orthogonal axes. If the field moves too far away from the skew axis, the lack of control authority allows environmental torques to pull the satellite away from the target and can potentially cause it to tumble. Slew maneuver planning must factor the stability of targets at the beginning and end, and the torque authority at all points along the slew. Due to the time varying magnetic field geometry relative to any two inertial targets, small modifications in slew maneuver timing can make large differences in the achievability of a maneuver.

Heatwole, Scott↗
NASA NTRS2004

NASA STARDUST Sample Return Mission

This viewgraph presentation reviews the Stardust mission. The objectives of the mission is to rendezvous with the Wild 2 comet, and using a specially designed aerogel capture some of the cometary material, and then return a capsule with the material to Earth. Other objectives are to provide images of the comet, and capture interstellar dust. There is a description of the aerogel, the trajectory, and views of the Stardust Sample Return Capsule (SRC), Cometary and Interstellar Dust Analyzer (CIDA) instrument, and the JPL designed camera. Also included is a timeline of the projected release of data to the Planetary Data System.

Wild 2↗
NASA NTRS2008

The Mars Climate Sounder In-Flight Positioning Anomaly

The paper discusses the Mars Climate Sounder (MCS) instrument s in-flight positioning errors and presents background material about it. A short overview of the instrument s science objectives and data acquisition techniques is provided. The brief mechanical description familiarizes the reader with the MCS instrument. Several key items of the flight qualification program, which had a rigorous joint drive test program but some limitations in overall system testing, are discussed. Implications this might have had for the flight anomaly, which began after several months of flawless space operation, are mentioned. The detection, interpretation, and instrument response to the errors is discussed. The anomaly prompted engineering reviews, renewed ground, and some in-flight testing. A summary of these events, including a timeline, is included. Several items of concern were uncovered during the anomaly investigation, the root cause, however, was never found. The instrument is now used with two operational constraints that work around the anomaly. It continues science gathering at an only slightly diminished pace that will yield approximately 90% of the originally intended science.

Jau, Bruno M.↗
NASA NTRS2008

Launching to the Moon, Mars, and Beyond

This viewgraph presentation discusses NASA's mission and addresses the following questions: 1) What is NASA's mission? 2) Why do we explore? 3) What is out timeline? 4) why the Moon first? 5) What will the vehicles look like? 6) What progress have we made? who will be doing the work? and 7) What are the benefits of space exploration?

Shivers, C. Herbert↗
NASA NTRS2008

Launching to the Moon, Mars, and Beyond

This viewgraph presentation reviews NASA's mission to launch to the Moon, Mars, and Beyond. The following questions will be answered: 1) What is NASA's mission? 2) Why do we explore? 3) What is our timeline? 4) Why the Moon first? 5) What will the vehicles look like? 5) What progress have we made? 6) Who will be doing the work? and 7) What are the benefits of space exploration?

Shivers, Charles Herbert↗
NASA NTRS2008

Spacecraft Formation Control: Managing Line-of-Sight Drift Based on the Dynamics of Relative Motion

In a quest to improve space-based observational capability, an increasing number of investigators are proposing missions with precision formation flying architectures. Typical missions include the Micro- Arcsecond X-ray Imaging Mission (MAXIM), Stellar Imager (SI), and the New Worlds Observer (NWO). Missions designed to explore targets in deep-space generally require holding a formation configuration fixed in inertial space during science observation. Analysis in this paper is specifically aimed at the NWO architecture, characterizing the natural drift of the line-of-sight and the separation range for two spacecraft operating in the vicinity of the Earth/Moon-Sun L(sub 2) libration point. Analysis employs a linear form of the relative dynamics associated with an n-body gravity field. The study is designed to identify favorable observation directions, characterized by minimal line-of-sight drift, along the mission timeline.

Luquette, Richard J.↗
NASA NTRS2008

New Worlds Observer Formation Control Design Based on the Dynamics of Relative Motion

The New Worlds Observer (NWO) mission is designed for the direct detection and characterization of extrasolar planets. The NWO mission concept employs a two spacecraft leader-follower formation on a trajectory around the Earth/Moon-Sun L(sub 2) Libration Point. The leader spacecraft is baselined as a 4 meter optical telescope. The follower, Starshade spacecraft, is designed to suppress light from a central body star permitting direct detection of a surrounding exoplanetary system. The current design requires a nominal leader-follower separation range of 72 Megameters. NWO poses many challenges including formation control. NWO cycles between three principal control modes during the nominal mission timeline: science (fine pointing), realignment and transition. This paper examines formation control strategies in the context of dynamics of relative motion for two spacecraft operating in the vicinity of the Earth/Moon-Sun L(sub 2)libration point. The paper presents an overview of the equations of relative motion followed by a discussion of each of the control modes. Discussion and analysis characterize control strategies for each of the mission control modes, including requirements, implementation challenges and project fuel budgets.

Luquette, Richard J.↗
NASA NTRS2008

Computational Model of Heat Transfer on the ISS

SCRAM Lite (SCRAM signifies Station Compact Radiator Analysis Model) is a computer program for analyzing convective and radiative heat-transfer and heat-rejection performance of coolant loops and radiators, respectively, in the active thermal-control systems of the International Space Station (ISS). SCRAM Lite is a derivative of prior versions of SCRAM but is more robust. SCRAM Lite computes thermal operating characteristics of active heat-transport and heat-rejection subsystems for the major ISS configurations from Flight 5A through completion of assembly. The program performs integrated analysis of both internal and external coolant loops of the various ISS modules and of an external active thermal control system, which includes radiators and the coolant loops that transfer heat to the radiators. The SCRAM Lite run time is of the order of one minute per day of mission time. The overall objective of the SCRAM Lite simulation is to process input profiles of equipment-rack, crew-metabolic, and other heat loads to determine flow rates, coolant supply temperatures, and available radiator heat-rejection capabilities. Analyses are performed for timelines of activities, orbital parameters, and attitudes for mission times ranging from a few hours to several months.

Torian, John G.↗
NASA NTRS2008

Estimating the Need for Medical Intervention due to Sleep Disruption on the International Space Station

During ISS and shuttle missions, difficulties with sleep affect more than half of all US crews. Mitigation strategies to help astronauts cope with the challenges of disrupted sleep patterns can negatively impact both mission planning and vehicle design. The methods for addressing known detrimental impacts for some mission scenarios may have a substantial impact on vehicle specific consumable mass or volume or on the mission timeline. As part of the Integrated Medical Model (IMM) task, NASA Glenn Research Center is leading the development of a Monte Carlo based forecasting tool designed to determine the consumables required to address risks related to sleep disruption. The model currently focuses on the International Space Station and uses an algorithm that assembles representative mission schedules and feeds this into a well validated model that predicts relative levels of performance, and need for sleep (SAFTE Model, IBR Inc). Correlation of the resulting output to self-diagnosed needs for hypnotics, stimulants, and other pharmaceutical countermeasures, allows prediction of pharmaceutical use and the uncertainty of the specified prediction. This paper outlines a conceptual model for determining a rate of pharmaceutical utilization that can be used in the IMM model for comparison and optimization of mitigation methods with respect to all other significant medical needs and interventions.

Myers, Jerry G.↗
NASA NTRS2009

Food Mass Reduction Trade Study

Future long duration manned space flights beyond low earth orbit will require the food system to remain safe, acceptable, and nutritious while efficiently balancing appropriate vehicle resources such as mass, volume, power, water, and crewtime. Often, this presents a challenge since maintaining the quality of the food system can result in a higher mass and volume. The Orion vehicle is significantly smaller than the Shuttle vehicle and the International Space Station and the mass and volume available for food is limited. Therefore, the food team has been challenged to reduce the mass of the packaged food from 1.82 kg per person per day to 1.14 kg per person per day. Past work has concentrated on how to reduce the mass of the packaging which contributes to about 15% of the total mass of the packaged food system. Designers have also focused on integrating and optimizing the Orion galley equipment as a system to reduce mass. To date, there has not been a significant effort to determine how to reduce the food itself. The objective of this project is to determine how the mass and volume of the packaged food can be reduced while maintaining caloric and hydration requirements. The following tasks are the key elements to this project: (1) Conduct further analysis of the ISS Standard Menu to determine moisture, protein, carbohydrate, and fat levels. (2) Conduct trade studies to determine how to bring the mass of the food system down. Trade studies may include removing the water of the total food system and/or increasing the fat content. (3) Determine the preferred method for delivery of the new food (e.g. bars, or beverages) and the degree of replacement. (4) Determine whether there are commercially available products that meet the requirements. By the end of this study, an estimate of the mass and volume savings will be provided to the Constellation Program. In addition, if new technologies need to be developed to achieve the mass savings, the technologies, timeline, and budget will be identified at the end of the project.

Perchonok, Michele H.↗
NASA NTRS2006

Mars Science Laboratory: Entry, Descent, and Landing System Performance

In 2010, the Mars Science Laboratory (MSL) mission will pioneer the next generation of robotic Entry, Descent, and Landing (EDL) systems, by delivering the largest and most capable rover to date to the surface of Mars. To do so, MSL will fly a guided lifting entry at a lift-to-drag ratio in excess of that ever flown at Mars, deploy the largest parachute ever at Mars, and perform a novel Sky Crane maneuver. Through improved altitude capability, increased latitude coverage, and more accurate payload delivery, MSL is allowing the science community to consider the exploration of previously inaccessible regions of the planet. The MSL EDL system is a new EDL architecture based on Viking heritage technologies and designed to meet the challenges of landing increasing massive payloads on Mars. In accordance with level-1 requirements, the MSL EDL system is being designed to land an 850 kg rover to altitudes as high as 1 km above the Mars Orbiter Laser Altimeter defined areoid within 10 km of the desired landing site. Accordingly, MSL will enter the largest entry mass, fly the largest 70 degree sphere-cone aeroshell, generate the largest hypersonic lift-to-drag ratio, and deploy the largest Disk-Gap-Band supersonic parachute of any previous mission to Mars. Major EDL events include a hypersonic guided entry, supersonic parachute deploy and inflation, subsonic heatshield jettison, terminal descent sensor acquisition, powered descent initiation, sky crane terminal descent, rover touchdown detection, and descent stage flyaway. Key performance metrics, derived from level-1 requirements and tracked by the EDL design team to indicate performance capability and timeline margins, include altitude and range at parachute deploy, time on radar, and propellant use. The MSL EDL system, which will continue to develop over the next three years, will enable a notable extension in the advancement of Mars surface science by delivering more science capability than ever before to the surface of Mars. This paper describes the current MSL EDL system performance as predicted by end-to-end EDL simulations, highlights the sensitivity of this baseline performance to several key environmental assumptions, and discusses some of the challenges faced in delivering such an unprecedented rover payload to the surface of Mars.

Way, David W.↗
NASA NTRS2009

Rendezvous Integration Complexities of NASA Human Flight Vehicles

Propellant-optimal trajectories, relative sensors and navigation, and docking/capture mechanisms are rendezvous disciplines that receive much attention in the technical literature. However, other areas must be considered. These include absolute navigation, maneuver targeting, attitude control, power generation, software development and verification, redundancy management, thermal control, avionics integration, robotics, communications, lighting, human factors, crew timeline, procedure development, orbital debris risk mitigation, structures, plume impingement, logistics, and in some cases extravehicular activity. While current and future spaceflight programs will introduce new technologies and operations concepts, the complexity of integrating multiple systems on multiple spacecraft will remain. The systems integration task may become more difficult as increasingly complex software is used to meet current and future automation, autonomy, and robotic operation requirements.

Brazzel, Jack P.↗