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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS2002

Piloted Simulation Assessment of a High-Speed Civil Transport Configuration

An assessment of a proposed configuration of a high-speed civil transport was conducted by using NASA and industry research pilots. The assessment was conducted to evaluate operational aspects of the configuration from a pilot's perspective, with the primary goal being to identify potential deficiencies in the configuration. The configuration was evaluated within and at the limits of the design operating envelope to determine the suitability of the configuration to maneuver in a typical mission as well as in emergency or envelope-limit conditions. The Cooper-Harper rating scale was used to evaluate the flying qualities of the configuration. A summary flying qualities metric was also calculated. The assessment was performed in the Langley six-degree-of-freedom Visual Motion Simulator. The effect of a restricted cockpit field-of-view due to obstruction by the vehicle nose was not included in this study. Tasks include landings, takeoffs, climbs, descents, overspeeds, coordinated turns, and recoveries from envelope limit excursions. Emergencies included engine failures, loss of stability augmentation, engine inlet unstarts, and emergency descents. Minimum control speeds and takeoff decision, rotation, and safety speeds were also determined.

Jackson, E. Bruce↗
NASA NTRS2006

A Case for Developing a Ground Based Replication of the Earth, Moon and Mars Spaceflight Infrastructure

When the systems are developed and in place to provide the services needed to operate en route and on the Lunar and Martian surfaces, an Earth based replication will need to be in place for the safety and protection of mission success. The replication will entail all aspects of the flight configuration end to end but will not include any closed loop systems. This would replicate the infrastructure from Lunar and Martian robots, manned surface excursions, through man and unmanned terrestrial bases, through the various types of communication systems and technologies, manned and un-manned space vehicles (large and small), to Earth based systems and control centers. An Earth based replicated infrastructure will enable checkout and test of new technologies, hardware, software updates and upgrades and procedures without putting humans and missions at risk. Analysis of events, what ifs and trouble resolution could be played out on the ground to remove as much risk as possible from any type of proposed change to flight operational systems. With adequate detail, it is possible that failures could be predicted with a high probability and action taken to eliminate failures. A major factor in any mission to the Moon and to Mars is the complexity of systems, interfaces, processes, their limitations, associated risks and the factor of the unknown including the development by many contractors and NASA centers. The need to be able to introduce new technologies over the life of the program requires an end to end test bed to analyze and evaluate these technologies and what will happen when they are introduced into the flight system. The ability to analyze system behaviors end to end under varying conditions would enhance safety e.g. fault tolerances. This analysis along with the ability to mine data from the development environment (e.g. test data), flight ops and modeling/simulations data would provide a level of information not currently available to operations and astronauts. In this paper we will analyze the beginnings of such a replication and what it could do in terms of reducing risk in the near term for development. We will analyze the Space Shuttle Main Engine (SSME) test lab which has to a large extent accomplished this replication for the SSME and has been highly successful in analyzing hardware and software problems and changes. The cost of replicating the flight system as proposed here could be very high if attempted as an afterthought. We will describe the initial steps for the development of a replication of this infrastructure starting with the communication infrastructure. The Constellation of Labs (CofL) under the Command, Control, Communication and Information (C3I) project for the NASA Exploration Initiative will provide the initial foundation upon which to base this replication. Simply put, there is very little margin for error in high latency situations e.g. en-route to/from Mars or in an autonomous process on the Lunar far side. Any thought out approach to reduce risk and increase safety needs to be accomplished end to end with the actual systems configuration.

Bradford, Robert N.↗
NASA NTRSDate not supplied

Predicting Orion Launch Abort Acoustics

Computational fluid dynamics (CFD) scale-resolving simulations were performed to help assess the Orion launch abort system’s vibro-acoustic environment in collaboration with the Orion Loads and Dynamics team at the Johnson Space Center (JSC). The primary objectives are to establish best practices and to validate CFD predictions for surface fluctuating pressure spectra with ground, wind tunnel, and flight test data. Excellent agreement is obtained between CFD predictions by the Cartesian adaptive mesh refinement (AMR) module of the Launch, Ascent, and Vehicle Aerodynamics (LAVA) solver and measurements from a ground test representative of the start of a pad abort scenario, and with measurements from a wind tunnel test case corresponding to a transonic, high total angle of attack ascent abort scenario. A simplified (static and at rest) simulation of a pad abort flight test shows good agreement with the flight data, whereas a more costly simulation where the vehicle is accelerating relative to the ground from rest to Mach 0.4 shows consistent over-prediction of the spectra. Simulation results for a supersonic ascent abort flight test are in reasonable agreement with the measurements, but show larger excursions in regions where non-linear interactions between the plume turbulence, shocks, and turbulent boundary layer are strongest – likely due to CFD modeling assumptions, the effect of which are discussed. Overall, the CFD results demonstrate that scale-resolving simulation methodologies like the one presented can help characterize full-scale launch abort vehicle near-field noise and can complement wind tunnel, ground and flight tests by providing trends and reducing uncertainty for scenarios that cannot or are not planned to be tested.

Orion↗
NASA NTRS2008

NASA Ares I Crew Launch Vehicle Upper Stage Overview

By incorporating rigorous engineering practices, innovative manufacturing processes and test techniques, a unique multi-center government/contractor partnership, and a clean-sheet design developed around the primary requirements for the International Space Station (ISS) and Lunar missions, the Upper Stage Element of NASA's Crew Launch Vehicle (CLV), the "Ares I," is a vital part of the Constellation Program's transportation system. Constellation's exploration missions will include Ares I and Ares V launch vehicles required to place crew and cargo in low-Earth orbit (LEO), crew and cargo transportation systems required for human space travel, and transportation systems and scientific equipment required for human exploration of the Moon and Mars. Early Ares I configurations will support ISS re-supply missions. A self-supporting cylindrical structure, the Ares I Upper Stage will be approximately 84' long and 18' in diameter. The Upper Stage Element is being designed for increased supportability and increased reliability to meet human-rating requirements imposed by NASA standards. The design also incorporates state-of-the-art materials, hardware, design, and integrated logistics planning, thus facilitating a supportable, reliable, and operable system. With NASA retiring the Space Shuttle fleet in 2010, the success of the Ares I Project is essential to America's continued leadership in space. The first Ares I test flight, called Ares I-X, is scheduled for 2009. Subsequent test flights will continue thereafter, with the first crewed flight of the Crew Exploration Vehicle (CEV), "Orion," planned for no later than 2015. Crew transportation to the ISS will follow within the same decade, and the first Lunar excursion is scheduled for the 2020 timeframe.

McArthur, J. Craig↗
NASA NTRS2008

NASA Ares I Crew Launch Vehicle Upper Stage Overview

By incorporating rigorous engineering practices, innovative manufacturing processes and test techniques, a unique multi-center government/contractor partnership, and a clean-sheet design developed around the primary requirements for the International Space Station (ISS) and Lunar missions, the Upper Stage Element of NASA's Crew Launch Vehicle (CLV), the "Ares I," is a vital part of the Constellation Program's transportation system. Constellation's exploration missions will include Ares I and Ares V launch vehicles required to place crew and cargo in low-Earth orbit (LEO), crew and cargo transportation systems required for human space travel, and transportation systems and scientific equipment required for human exploration of the Moon and Mars. Early Ares I configurations will support ISS re-supply missions. A self-supporting cylindrical structure, the Ares I Upper Stage will be approximately 84' long and 18' in diameter. The Upper Stage Element is being designed for increased supportability and increased reliability to meet human-rating requirements imposed by NASA standards. The design also incorporates state-of-the-art materials, hardware, design, and integrated logistics planning, thus facilitating a supportable, reliable, and operable system. With NASA retiring the Space Shuttle fleet in 2010, the success of the Ares I Project is essential to America's continued leadership in space. The first Ares I test flight, called Ares 1-X, is scheduled for 2009. Subsequent test flights will continue thereafter, with the first crewed flight of the Crew Exploration Vehicle (CEV), "Orion," planned for no later than 2015. Crew transportation to the ISS will follow within the same decade, and the first Lunar excursion is scheduled for the 2020 timeframe.

Davis, Daniel J.↗
NASA NTRS1963

Vehicle Design for Mars Landing and Return to Mars Orbit

This paper briefly describes three modes for accomplishing the Mars landing mission and compares them on a gross basis to indicate their probable order of merit and to identify design requirements placed on the Mars-excursion module (MEM) by the choice of mode. The paper shows that a flyby-rendezvous mode requiring low weight in earth orbit requires the MEM to enter the Mars atmosphere at velocities ranging from 20,000 to 30,000 ft/sec. The MEM for the flyby-rendezvous mode is not covered in this paper but merits further study. The MEM for the other modes of mission accomplishment begins its active operational sequence in Mars orbit and need not be greatly influenced by the method of delivery to Mars orbit. Parametric studies of the entry problem for two vehicles typifying a ballistic-type and a lifting-body-type were conducted to identify the problems associated with design of a MEM to accommodate the extremes of Mars atmospheric density presently predicted. This brief study indicates that: (a) the presently predicted density extremes of the Mars atmosphere present no serious design problems for a MEM which can operate across the entire band of predicted densities; (b) details of operational requirements and mission objectives will control the choice of configuration rather than entry requirements; and (c) the ballistic-type MEM is lighter and simpler but has less operational flexibility than a high L/D MEM.

Hammock, D. M.↗
NASA NTRS2008

Determining Exercise Strength Requirements for Astronaut Critical Mission Tasks: Reaching Under G-Load

The critical mission tasks assessments effort seeks to determine the physical performance requirements that astronauts must meet in order to safely and successfully accomplish lunar exploration missions. These assessments will determine astronaut preflight strength, fitness, and flexibility requirements, and the extent to which exercise and other countermeasures must prevent the physical deconditioning associated with prolonged weightlessness. The purpose is to determine the flexibility and strength that crewmembers must possess in order to reach Crew Exploration Vehicle controls during maneuvers that result in sustained acceleration levels ranging from 3.7G to 7.8G. An industry standard multibody dynamics application was used to create human models representing a 5th percentile female, a 50th percentile male, and a 95th percentile male. The additional mass of a space suit sleeve was added to the reaching arm to account for the influence of the suit mass on the reaching effort. The human model was merged with computer models of a pilot seat and control panel for the Crew Exploration Vehicle. Three dimensional paths were created that guided the human models hand from a starting position alongside its thigh to three control targets: a joystick, a keyboard, and an overhead switch panel. The reaching motion to each target was repeated under four vehicle acceleration conditions: nominal ascent (3.7G), two ascent aborts (5.5G and 7.8G) and lunar reentry (4.6G). Elbow and shoulder joint angular excursions were analyzed to assess range of motion requirements. Mean and peak elbow and shoulder joint torques were determined and converted to equivalent resistive exercise loads to assess strength requirements. Angular excursions for the 50th and 95th percentile male models remained within joint range of motion limits. For the 5th percentile female, both the elbow and the shoulder exceeded range of motion limits during the overhead reach. Elbow joint torques ranged from 10 N-m (nominal ascent) to 60 N-m (ascent abort). Shoulder joint torques ranged from 65 N-m (nominal ascent) to 280 N-m (ascent abort). Maximal equivalent exercise loads reached 30 lb in tricep extension, 9 lb in bicep curl, 110 lb in unilateral pullover and unilateral bench press for nominal conditions (lunar reentry), and 188 lb in unilateral pullover and unilateral bench press. The location of the pilot seat was found to be inadequately located to allow a 5th percentile female to reach the switches on the overhead panel. Elbow strength requirements were found to be well within population norms. Shoulder strength was found to be a limiting factor. Reaching under nominal ascent and lunar reentry conditions was found to require near maximal shoulder strength. Reaching under ascent abort conditions requires shoulder strength well beyond population norms. Pilot seats must adjust to accomodate a 5th percentile female. Exercise countermeasures must maintain maximal pullover and bench press strength to allow pilots to reach and operate controls during lunar reentry. Reaching will not be possible during ascent abort conditions. Flight controls should be built into armrests or flight control must be accomplished by autonomous systems during acceleration exceeding 4.6G.

Schaffner, Grant↗
NASA NTRS1974

Atmosphere Explorer orbit adjust propulsion system

The Atmosphere Explorer-C (AE-C) spacecraft was launched into a highly elliptical orbit by a two-stage Delta 1900 vehicle on December 15, 1973. Its mission is to observe and study phenomena in the earth's atmosphere at altitudes between 150 and 4000 kilometers with several excursions down to 130 kilometers. An Orbit Adjust Propulsion System (OAPS) is included on the spacecraft in order (1) to perform low perigee maneuvers at 130 to 150 kilometers; (2) to periodically restore apogee in order to offset the effect of aerodynamic drag at perigee; and (3) to provide the capability for occasional large attitude maneuvers. This paper describes the configuration of the OAPS and the design characteristics of its individual components. Special emphasis is placed on the valving arrangement, which provides a means of adjusting the center-of-mass of the spacecraft through control of the distribution and usage of the propellant within the system.

Woodruff, W. L.↗
NASA NTRS1987

Analysis and design of a gas-lubricated, sectored, floating ring seal

Helium buffered floating-ring seals used in cryogenic turbopumps are subject to large temperature variations, shaft centrifugal growth, and dynamic shaft excursions. To prevent dangerous contact friction over a wide range of operating conditions generally requires liberal clearances with consequent high leakage. Reduction of helium consumption translates into lower space vehicle helium storage requirements and thus increased payload. A hydrostatic sectored floating-ring seal is described with self-adjusting clearance capabilities that enables the seal to operate with small clearances over the operating range to minimize leakage.

Shapiro, W.↗
NASA NTRS1998

Behavior and Performance Project

Behavior is an overt manifestation of underlying biology. As such, alterations in biological systems that result from spaceflight would be expected to evidence themselves in subtle or even pronounced changes in the behavior of that organism. These alterations in visible behavior may then indicate or even be diagnostic of alterations in the physical well-being of humans and other animals as they adapt to space environments or readapt to Earth--alterations that might not otherwise be detected without relatively invasive and frequently expensive procedures. Moreover, behavior and performance constitute the central standard for evaluating the mission success of spaceflight ventures. The success or failure of any excursion into space is primarily indexed by the ability of astronauts and cosmonauts to perform the tasks and experiments of any particular mission, to land a re-entry vehicle after extended exposure to microgravity (as with the space shuttle), or to make time-critical and life-saving repairs or other decisions while in orbit (as with the recent Mir events). Thus, understanding the effects spaceflight on behavior and performance is inherently important, in addition to those insights that behavior can provide about the physiological consequences of space adaptation.

Washburn, David A.↗
NASA NTRS1990

Use of random Martian atmosphere to evaluate potential entry guidance schemes

A random Martian atmosphere was developed and was used with three guidance schemes to determine the effect of random density variations on the guidance. This random atmosphere was shown to be useful for testing the robustness of guidance schemes for vehicles encountering random disturbances during aerobraking for capture into planetary orbit. Levels of disturbance that could be tolerated and areas where performance could be improved were established. The need for Monte Carlo studies to define the excursion boundaries of capture orbit parameters was indicated.

Tomlinson, Barbara S.↗
NASA NTRS1994

Project ECHO: Electronic Communications from Halo Orbit

The design of a communications relay to provide constant access between the Earth and the far side of the Moon is presented. Placement of the relay in a halo orbit about the L2 Earth-Moon Lagrange point allows the satellite to maintain constant simultaneous communication between Earth and scientific payloads on the far side of the Moon. The requirements of NASA's Discovery-class missions adopted and modified for this design are: total project cost should not exceed $150 million excluding launch costs, launch must be provided by Delta-class vehicle, and the satellite should maintain an operational lifetime of 10 to 15 years. The spacecraft will follow a transfer trajectory to the L2 point, after launch by a Delta II 7925 vehicle in 1999. Low-level thrust is used for injection into a stationkeeping-free halo orbit once the spacecraft reaches the L2 point. The shape of this halo orbit is highly elliptical with the maximum excursion from the L2 point being 35000 km. A spun section and despun section connected through a bearing and power transfer assembly (BAPTA) compose the structure of the spacecraft. Communications equipment is placed on the despun section to provide for a stationary dual parabolic offset-feed array antenna system. The dual system is necessary to provide communications coverage during portions of maximum excursion on the halo orbit. Transmissions to the NASA Deep Space Network 34 m antenna include six channels (color video, two voice, scientific data from lunar payloads, satellite housekeeping and telemetry and uplinked commands) using the S- and X-bands. Four radioisotope thermoelectric generators (RTG's) provide a total of 1360 W to power onboard systems and any two of the four Hughes 13 cm ion thrusters at once. Output of the ion thrusters is approximately 17.8 mN each with xenon as the propellant. Presence of torques generated by solar pressure on the antenna dish require the addition of a 'skirt' extending from the spun section of the satellite for balance. Total mass of the satellite is approximately 900 kg at a cost of $130 million FY99.

Borrelli, Jason↗
NASA NTRS2012

Desert RATS 2011: Near-Earth Asteroid Human Exploration Operations

The Desert Research and Technology Studies (D-RATS) 2011 field test involved the planning and execution of a series of exploration scenarios under operational conditions similar to those that would be expected during a human exploration mission to a near-Earth asteroid (NEA). The focus was on understanding the operations tempo during simulated NEA exploration and the implications of communications latency and limited data bandwidth. Anchoring technologies and sampling techniques were not evaluated due to the immaturity of those technologies and the inability to meaningfully test them at D-RATS. Reduced gravity analogs and simulations are being used to fully evaluate Multi-Mission Space Exploration Vehicle (MMSEV) and extravehicular (EVA) operations and interactions in near-weightlessness at a NEA as part of NASA s integrated analogs program. Hypotheses were tested by planning and performing a series of 1-day simulated exploration excursions comparing test conditions all of which involved a single Deep Space Habitat (DSH) and either zero, one, or two MMSEVs; three or four crewmembers; one of two different communications bandwidths; and a 100-second roundtrip communications latency between the field site and Houston. Excursions were executed at the Black Point Lava Flow test site with a Mission Control Center and Science Support Room at Johnson Space Center (JSC) being operated with 100-second roundtrip communication latency to the field. Crews were composed of astronauts and professional field geologists and teams of Mission Operations, Science, and Education & Public Outreach (EPO) experts also supported the mission simulations each day. Data were collected separately from the Crew, Mission Operations, Science, and EPO teams to assess the test conditions from multiple perspectives. For the operations tested, data indicates practically significant benefits may be realized by including at least one MMSEV and by including 4 versus 3 crewmembers in the NEA exploration architecture as measured by increased Scientific Data Quality, EVA Exploration Time, Capability Assessment Ratings, and Overall Acceptability ratings by Crew, Mission Operations, Science, and Education & Public Outreach teams. A combination of text and voice was used to effectively communicate over the 100-second roundtrip communications latency and increased communication bandwidth yielded a small but practically significant improvement in Overall Acceptability as rated by the Science team, although the impact of bandwidth on scientific strategic planning was not assessed. No effect of increased bandwidth was observed with respect to Crew, Mission Operations, or EPO team ratings of Overall Acceptability.

Abercromby, Andrew↗
NASA NTRS2007

Technical Findings, Lessons Learned, and Recommendations Resulting from the Helios Prototype Vehicle Mishap

The Helios Prototype was originally planned to be two separate vehicles, but because of resource limitations only one vehicle was developed to demonstrate two missions. The vehicle consisted of two configurations, one for each mission. One configuration, designated HP01, was designed to operate at extremely high altitudes using batteries and high-efficiency solar cells spread across the upper surface of its 247-foot wingspan. On August 13, 2001, the HP01 configuration reached an altitude of 96,863 feet, a world record for sustained horizontal flight by a winged aircraft. The other configuration, designated HP03, was designed for long-duration flight. The plan was to use the solar cells to power the vehicle's electric motors and subsystems during the day and to use a modified commercial hydrogen-air fuel cell system for use during the night. The aircraft design used wing dihedral, engine power, elevator control surfaces, and a stability augmentation and control system to provide aerodynamic stability and control. At about 30 minutes into the second flight of HP03, the aircraft encountered a disturbance in the way of turbulence and morphed into an unexpected, persistent, high dihedral configuration. As a result of the persistent high dihedral, the aircraft became unstable in a very divergent pitch mode in which the airspeed excursions from the nominal flight speed about doubled every cycle of the oscillation. The aircraft s design airspeed was subsequently exceeded and the resulting high dynamic pressures caused the wing leading edge secondary structure on the outer wing panels to fail and the solar cells and skin on the upper surface of the wing to rip away. As a result, the vehicle lost its ability to maintain lift, fell into the Pacific Ocean within the confines of the U.S. Navy's Pacific Missile Range Facility, and was destroyed. This paper describes the mishap and its causes, and presents the technical recommendations and lessons learned for improving the design, analysis, and testing methods and techniques required for this class of vehicle.

Noll, Thomas E.↗
NASA NTRS1989

Space vehicle thermal testing: Principles, practices and effectiveness

Component qualification and acceptance temperatures are derived from worst case thermal analyses and analytic uncertainty margin subject to certain specified temperature extremes. Temperature requirements are shown for equipment operation within specification and for survival and turn-on (need not operate within specification, but must not experience any degradation when returned to operational range). Temperature excursions for most equipment are seen to be 20 to 50 C above and below room temperature. Components without active electronics which are mounted outboard, such as solar arrays and antennas, are usually designed to withstand wider temperature excursions, particularly at the cold end. Batteries are tightly controlled at cold temperatures to increase life. Payload components such as extremely accurate clocks for precise navigation are controlled over a relatively narrow temperature range.

Gluck, Donald F.↗
NASA NTRS1991

Overpressurized zero pressure balloon system

Typical zero pressure scientific balloon systems experiencing diurnal cycles require on board expendable ballast to maintain design float altitude. In the past, pressurized balloon systems have been employed to overcome the need for ballast while maintaining small diurnal altitude excursions. However, requirements on material characteristics and manufacturing methods have enforced limitations on balloon size and therefore payload and altitude capabilities. This effort was aimed at considering the use of current (or slightly modified) zero pressure balloon systems (design, manufacturing and materials) as candidates for pressurized vehicles. Data gathered through performance analysis, structural analysis and flight testing has shown that the use of current zero pressure designs can be expanded to enhance balloon performance.

Simpson, Joel M.↗
NASA NTRS2007

Protection of Space Vehicles from Micrometeoroid/Orbital Debris (MMOD) Damages

As the environment that puts space vehicles at risk can never be eliminated, space vehicles must implement protection against the MMOD environment. In general, this protection has been implemented on a risk estimate basis, largely focused on estimates of impactor size and estimated flux. However, there is some uncertainty in applying these methods from data gathered in earth orbit to excursions outside. This paper discusses different past thresholds and processes of the past and suggests additional refinement or methods that could be used for future space endeavors.

Barr, Stephanie↗