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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 577 records · Page 32
All sourcesNASA NTRSDOE OSTI
NASA NTRS2012

Results from Navigator GPS Flight Testing for the Magnetospheric MultiScale Mission

The recent delivery of the first Goddard Space Flight Center (GSFC) Navigator Global Positioning System (GPS) receivers to the Magnetospheric MultiScale (MMS) mission spacecraft is a high water mark crowning a decade of research and development in high-altitude space-based GPS. Preceding MMS delivery, the engineering team had developed receivers to support multiple missions and mission studies, such as Low Earth Orbit (LEO) navigation for the Global Precipitation Mission (GPM), above the constellation navigation for the Geostationary Operational Environmental Satellite (GOES) proof-of-concept studies, cis-Lunar navigation with rapid re-acquisition during re-entry for the Orion Project and an orbital demonstration on the Space Shuttle during the Hubble Servicing Mission (HSM-4).

Lulich, Tyler D.↗
NASA NTRS2016

Aerodynamic Models for the Low Density Supersonic Decelerator (LDSD) Test Vehicles

An overview of aerodynamic models for the Low Density Supersonic Decelerator (LDSD) Supersonic Flight Dynamics Test (SFDT) campaign test vehicle is presented, with comparisons to reconstructed flight data and discussion of model updates. The SFDT campaign objective is to test Supersonic Inflatable Aerodynamic Decelerator (SIAD) and large supersonic parachute technologies at high altitude Earth conditions relevant to entry, descent, and landing (EDL) at Mars. Nominal SIAD test conditions are attained by lifting a test vehicle (TV) to 36 km altitude with a helium balloon, then accelerating the TV to Mach 4 and 53 km altitude with a solid rocket motor. Test flights conducted in June of 2014 (SFDT-1) and 2015 (SFDT-2) each successfully delivered a 6 meter diameter decelerator (SIAD-R) to test conditions and several seconds of flight, and were successful in demonstrating the SFDT flight system concept and SIAD-R technology. Aerodynamic models and uncertainties developed for the SFDT campaign are presented, including the methods used to generate them and their implementation within an aerodynamic database (ADB) routine for flight simulations. Pre- and post-flight aerodynamic models are compared against reconstructed flight data and model changes based upon knowledge gained from the flights are discussed. The pre-flight powered phase model is shown to have a significant contribution to off-nominal SFDT trajectory lofting, while coast and SIAD phase models behaved much as predicted.

Van Norman, John W.↗
NASA NTRS2015

Synthetic Fungal Strains for Solar System Exploration and Colonization

Solar system exploration and eventual colonization efforts are constrained by limits on the mass of material that can embark from Earth. Thus, creative use of the resources available in situ could reduce mission costs and extend the scope of such activities. To that end, we are developing synthetic fungal strains to produce specialized materials from the resources found throughout the solar system. A primary goal is to develop a suite of Saccharomyces cerevisiae strains to serve as generic production chassis for synthetic metabolic pathways. These strains must perform consistently upon challenge by unique conditions including exposure to microgravity, cosmic radiation, the rigors of launch and re-entry, and long-term stasis. Presently, we are establishing systematic datasets profiling epigenetic, transcriptional, translational and metabolic states of S. cerevisiae under relevant operating conditions. These will deepen our understanding of the physiological changes associated with space travel and enable rational engineering of optimal production strains.

Biotechnology↗
NASA NTRS2011

Launch Vehicle Demonstrator Using Shuttle Assets

The Marshall Space Flight Center Advanced Concepts Office (ACO) has the leading role for NASA s preliminary conceptual launch vehicle design and performance analysis. Over the past several years the ACO Earth-to-Orbit Team has evaluated thousands of launch vehicle concept variations for a multitude of studies including agency-wide efforts such as the Exploration Systems Architecture Study (ESAS), Constellation, Heavy Lift Launch Vehicle (HLLV), Heavy Lift Propulsion Technology (HLPT), Human Exploration Framework Team (HEFT), and Space Launch System (SLS). NASA plans to continue human space exploration and space station utilization. Launch vehicles used for heavy lift cargo and crew will be needed. One of the current leading concepts for future heavy lift capability is an inline one and a half stage concept using solid rocket boosters (SRB) and based on current Shuttle technology and elements. Potentially, the quickest and most cost-effective path towards an operational vehicle of this configuration is to make use of a demonstrator vehicle fabricated from existing shuttle assets and relying upon the existing STS launch infrastructure. Such a demonstrator would yield valuable proof-of-concept data and would provide a working test platform allowing for validated systems integration. Using shuttle hardware such as existing RS-25D engines and partial MPS, propellant tanks derived from the External Tank (ET) design and tooling, and four-segment SRB s could reduce the associated upfront development costs and schedule when compared to a concept that would rely on new propulsion technology and engine designs. There are potentially several other additional benefits to this demonstrator concept. Since a concept of this type would be based on man-rated flight proven hardware components, this demonstrator has the potential to evolve into the first iteration of heavy lift crew or cargo and serve as a baseline for block upgrades. This vehicle could also serve as a demonstration and test platform for the Orion Program. Critical spacecraft systems, re-entry and recovery systems, and launch abort systems of Orion could also be demonstrated in early test flights of the launch vehicle demo. Furthermore, an early demonstrator of this type would provide a stop-gap for retaining critical human capital and infrastructure while affording the current emerging generation of young engineers opportunity to work with and capture lessons learned from existing STS program offices and personnel, who were integral in the design and development of the Space Shuttle before these resources are no longer available. The objective of this study is to define candidate launch vehicle demonstration concepts that are based on Space Shuttle assets and determine their performance capabilities and how these demonstration vehicles could evolve to a heavy lift capability to low earth orbit.

Threet, Grady E., Jr.↗
DOE OSTI2020

Coarse-grain cluster analysis of tensors with application to climate biome identification

A tensor provides a concise way to codify the interdependence of complex data. Treating a tensor as a d-way array, each entry records the interaction between the different indices. Clustering provides a way to parse the complexity of the data into more readily understandable information. Clustering methods are heavily dependent on the algorithm of choice, as well as the chosen hyperparameters of the algorithm. However, their sensitivity to data scales is largely unknown. In this work, we apply the discrete wavelet transform to analyze the effects of coarse-graining on clustering tensor data. We are particularly interested in understanding how scale affects clustering of the Earth's climate system. The discrete wavelet transform allows classification of the Earth's climate across a multitude of spatial-temporal scales. The discrete wavelet transform is used to produce an ensemble of classification estimates, as opposed to a single classification. Each element of the ensemble is a clustering at a different spatial-temporal scale. Information theoretic approaches are used to identify important scale lengths in clustering the L15 Climate Datset. We also discover a sub-collection of the ensemble that spans the majority of the variance observed, allowing for efficient consensus clustering techniques that can be used to identify climate biomes.

54 ENVIRONMENTAL SCIENCES↗
NASA NTRS1980

Technologies involved in configuring an advanced earth-to-orbit transport for low structural mass

A tradeoff study for a single-stage, earth-to-orbit transport spacecraft configured with low structural mass as the prime objective is presented. Among the major design elements affecting vehicle mass are: maneuverability and re-entry heating, the compatibility of an irregular cargo bay geometry to future mission requirements, the fabrication and assembly of large honeycomb structure fuselage sections, and the development of LOX/LH2 engine extendable nozzles and dual-propulsion schemes. It is shown that vehicle geometry and materials and structural technologies will be the critical areas in the development of such a system.

Macconochie, I. O.↗
NASA NTRS1993

Aerothermodynamic performance and thermal protection design for blunt re-entry bodies at L/D = 0.3

Aerodynamic heating and thermal protection design analyses were performed for three blunt re-entry bodies at an L/D = 0.3 returning from low earth orbit. These configurations consisted of a scaled up Apollo command module, a Viking re-entry vehicle, and an Aeroassist Flight Experiment (AFE) aerobrake, each with a maximum diameter of 4.42 m. The aerothermodynamic analysis determined the equilibrium stagnation point heating rate and heat load for nominal and 3-sigma re-entry trajectories and the distribution of heating along the pitch and yaw planes for each of the vehicles at the time of highest heat flux. Using the predicted heating rates and heating distributions, a Thermal Protection System (TPS) design with flight certified materials was tailored for each of the configurations. Results indicated that the heating to the corner of the Viking aeroshell would exceed current limits of reusable tile material. Also, the maximum heating for the AFE would be 15 percent greater than the maximum heating for the Apollo flying the same trajectory. TPS designs showed no significant advantage in TPS weight between the different vehicles; however, heat-shield areal density comparisons showed the Apollo configuration to be the most efficient in terms of TPS weight.

Caram, Jose M.↗
NASA NTRS1999

Visual-vestibular integration motion perception reporting

Self-orientation and self/surround-motion perception derive from a multimodal sensory process that integrates information from the eyes, vestibular apparatus, proprioceptive and somatosensory receptors. Results from short and long duration spaceflight investigations indicate that: (1) perceptual and sensorimotor function was disrupted during the initial exposure to microgravity and gradually improved over hours to days (individuals adapt), (2) the presence and/or absence of information from different sensory modalities differentially affected the perception of orientation, self-motion and surround-motion, (3) perceptual and sensorimotor function was initially disrupted upon return to Earth-normal gravity and gradually recovered to preflight levels (individuals readapt), and (4) the longer the exposure to microgravity, the more complete the adaptation, the more profound the postflight disturbances, and the longer the recovery period to preflight levels. While much has been learned about perceptual and sensorimotor reactions and adaptation to microgravity, there is much remaining to be learned about the mechanisms underlying the adaptive changes, and about how intersensory interactions affect perceptual and sensorimotor function during voluntary movements. During space flight, SMS and perceptual disturbances have led to reductions in performance efficiency and sense of well-being. During entry and immediately after landing, such disturbances could have a serious impact on the ability of the commander to land the Orbiter and on the ability of all crew members to egress from the Orbiter, particularly in a non-nominal condition or following extended stays in microgravity. An understanding of spatial orientation and motion perception is essential for developing countermeasures for Space Motion Sickness (SMS) and perceptual disturbances during spaceflight and upon return to Earth. Countermeasures for optimal performance in flight and a successful return to Earth require the development of preflight and in-flight training to help astronauts acquire and maintain a dual adaptive state. Despite the considerable experience with, and use of, an extensive set of countermeasures in the Russian space program, SMS and perceptual disturbances remain an unresolved problem on long-term flights. Reliable, valid perceptual reports are required to develop and refine stimulus rearrangements presented in the PAT devices currently being developed as countermeasures for the prevention of motion sickness and perceptual disturbances during spaceflight, and to ensure a less hazardous return to Earth. Prior to STS-8, crew member descriptions of their perceptual experiences were, at best, anecdotal. Crew members were not schooled in the physiology or psychology of sensory perception, nor were they exposed to the appropriate professional vocabulary. However, beginning with the STS-8 Shuttle flight, a serious effort was initiated to teach astronauts a systematic method to classify and quantify their perceptual responses in space, during entry, and after flight. Understanding, categorizing, and characterizing perceptual responses to spaceflight has been greatly enhanced by implementation of that training system.

Harm, Deborah L.↗
NASA NTRS2020

NASA’s Space Launch System: Launch Capability for Lunar Exploration and Transformative Science

Excitement is building for the first launch of NASA’s Space Launch System (SLS), a unique exploration asset for the agency’s Artemis lunar program as well as for a new generation of science missions. SLS is designed for an array of missions beyond Earth’s orbit. The flexible system, which can be configured for Orion, cargo or Orion with co-manifested payload missions, offers high escape velocities to send more mass to deep space destinations. When configured with an 8.4 m-diameter fairing, SLS offers unmatched payload volume for human exploration and science missions. The initial Block 1 variant will insert at least 26 metric tons (t) to trans-lunar injection (TLI) and the more powerful Block 1B vehicle will launch 34-37 t to TLI using a new-development upper stage. Much of the initial SLS Block 1 vehicle is complete, including the upper stage and payload section, the core stage, engines and the solid rocket boosters. The first mission, Artemis I, launching from modernized and upgraded facilities at Kennedy Space Center (KSC), will be an uncrewed test flight of SLS, Orion and ground processing, with a primary objective of testing Orion’s heat shield at lunar re-entry velocity. Artemis I will have accommodations for 13 6U CubeSat payloads. These CubeSat missions will be deployed along the upper stage disposal trajectory after Orion separates from the vehicle. A rare opportunity for CubeSats to be deployed beyond low Earth orbit (LEO), Artemis I CubeSat missions range from searching for hydrogen and other volatiles on the lunar South Pole to studying the acceleration mechanisms of solar and interplanetary particles from a heliocentric trajectory. With manufacturing of the initial vehicle complete, fabrication and procurement is progressing for the second flight of SLS and Orion, Artemis II. Also an SLS Block 1 and Orion flight launching from KSC, Artemis II will mark the return of American astronauts to deep space with a lunar flyby-free return trajectory mission. With the Artemis III flight, NASA has the goal to land the first woman and the next man on the Moon. Infrastructure beyond SLS will be required for this effort, including elements of the lunar Gateway as well as lunar rovers, landers and additional commercially supplied launch services. SLS, as the only vehicle with the capability to lift 26 t of mass to TLI in its initial Block 1 variant, will remain a key component of this new-era exploration program. Future variants – Block 1B and Block 2 –will lift 34-45 t to TLI. This paper will discuss the status of testing and integration for the Artemis I vehicle, manufacturing progress for the second vehicle and the manifest outlook for primary, co-manifested and secondary payloads in the current deep space exploration environment.

Creech, Stephen D.↗
NASA NTRS1961

ANALYSIS OF QUARTZ AND TEFLON SHIELDS FOR A PARTICULAR RE-ENTRY MISSION

The transient performance of ablation type heat protection shields is treated herein for the surface of a vehicle returning from outer space to the earth. The vehicle weighs 8640 kg, has a ballistic factor of 500 lb/ft2, re-enters with a speed of 11 km/sec at 120 km altitude, has a lift over drag ratio of -0.5, and is subjected to a maximum deceleration of 7.7 times the gravity constant. By use of well-known equations for the heat transfer and the mass transfer at a heated surface, a numerical calculation method is derived which, for the investigated ablation processes, yields exact transient solutions of the fundamental system of partial differential equations. The method is applied to various quartz shields and to one Teflon shield, which all evaporate so readily under the conditions of the problem at hand that practically no flow of molten shield material exists. The solutions also show comparatively small temperature changes parallel to the surface. The results show that the nose of the vehicle is cooled predominantly by the evaporation of the quartz or the Teflon; the rest of the vehicle's surface is cooled by radiation of the quartz or evaporation of the Teflon. The large mass transfer effects on the nose of the vehicle are detrimental since the resulting low surface temperatures prevent the radiative heat transfer out of the shield, which does not involve any mass loss, from being the desirable governing cooling factor.

ABLATING MATERIAL↗
NASA NTRS1987

The compressional ULF foreshock boundary of Venus - Observations by the PVO magnetometer

Entries and exits of the compressional magnetic wave segments of the Pioneer Venus Orbiter data are identified as crossings of a foreshock boundary. When these crossings are plotted in an appropriate coordinate system, they form scatter diagrams representing the geometry of a composite boundary covering all solar wind conditions of the sample. When the orientation of the boundary for the average interplanetary magnetic field direction at Venus is extracted, the boundary is similar, although not identical within present uncertainties, to that found for the analogous foreshock boundary of the earth.

Greenstadt, E. W.↗
NASA NTRS2000

NASA Participation in the ISAS MUSES C Asteroid Sample Return Mission

NASA and Japan's Institute of Space and Astronautical Science (ISAS) have agreed to cooperate on the first mission to collect samples from the surface of an asteroid and return them to Earth for in-depth study. The MUSES-C mission will be launched on a Japanese MV launch vehicle in January 2002 from Kagoshima Space Center, Japan, toward a touchdown on the asteroid Nereus in September 2003. A NASA-provided miniature rover will conduct in-situ measurements on the surface. The asteroid samples will be returned to Earth by MUSES-C via a parachute-borne recovery capsule in January 2006. NASA and ISAS will cooperate on several aspects of the mission, including mission support and scientific analysis. In addition to providing the rover, NASA will arrange for the testing of the MUSES-C re-entry heat shield at NASA/Ames Research Center, provide supplemental Deep Space Network tracking of the spacecraft, assist in navigating the spacecraft and provide arrangements for the recovery of the sample capsule at a landing site in the U. S. Scientific coinvestigators from the U.S. and Japan will share data from the instruments on the rover and the spacecraft. They will also collaborate on the investigations of the returned samples. With a mass of about I kg, the rover experiment will be a direct descendant of the technology used to build the Sojourner rover. The rover will carry three science instruments: a visible imaging camera, a near-infrared point spectrometer and an alpha X ray spectrometer. The solarpowered rover will move around the surface of Nereus collecting imagery data which are complimentary to the spacecraft investigation. The imaging system will be capable of making surface texture, composition, and morphology measurements at resolutions better than 1 cm. The rover will transmit this data to the spacecraft for relay back to Earth. Due to the microgravity environment on Nereus, the rover has been designed to right itself in case it flips over. Solar panels on all sides of the rover will ensure that enough power will always be available to the rover to activate the motors needed to turn over. Posable struts will allow the rover to position its chassis such that the camera can be pointed straight down at the surface or straight up at the sky. The presentation will describe the mission, scientific objectives and current state of the rover in detail.

Jones, Ross↗
NASA NTRS2004

Atmospheric Models for Aeroentry and Aeroassist

Eight destinations in the Solar System have sufficient atmosphere for aeroentry, aeroassist, or aerobraking/aerocapture: Venus, Earth, Mars, Jupiter, Saturn, Uranus, and Neptune, plus Saturn's moon Titan. Engineering-level atmospheric models for Earth, Mars, Titan, and Neptune have been developed for use in NASA s systems analysis studies of aerocapture applications. Development has begun on a similar atmospheric model for Venus. An important capability of these models is simulation of quasi-random perturbations for Monte Carlo analyses in developing guidance, navigation and control algorithms, and for thermal systems design. Characteristics of these atmospheric models are compared, and example applications for aerocapture are presented. Recent Titan atmospheric model updates are discussed, in anticipation of applications for trajectory and atmospheric reconstruct of Huygens Probe entry at Titan. Recent and planned updates to the Mars atmospheric model, in support of future Mars aerocapture systems analysis studies, are also presented.

Justus, C. G.↗
NASA NTRS2005

Atmospheric Models for Aeroentry and Aeroassist

Eight destinations in the Solar System have sufficient atmosphere for aeroentry, aeroassist, or aerobraking/aerocapture: Venus, Earth, Mars, Jupiter, Saturn, Uranus, and Neptune, plus Saturn's moon Titan. Engineering-level atmospheric models for Earth, Mars, Titan, and Neptune have been developed for use in NASA's systems analysis studies of aerocapture applications. Development has begun on a similar atmospheric model for Venus. An important capability of these models is simulation of quasi-random perturbations for Monte Carlo analyses in developing guidance, navigation and control algorithms, and for thermal systems design. Characteristics of these atmospheric models are compared, and example applications for aerocapture are presented. Recent Titan atmospheric model updates are discussed, in anticipation of applications for trajectory and atmospheric reconstruct of Huygens Probe entry at Titan. Recent and planned updates to the Mars atmospheric model, in support of future Mars aerocapture systems analysis studies, are also presented.

Justus, C. G.↗
NASA NTRSDate not supplied

Enabling In-Situ Exploration of the Ice Giants Using Aerocapture

Investigation of the Ice Giants, especially Uranus, via orbiter and atmospheric probes, is required to answer pressing science questions that have been raised in the latest Decadal Survey. As the Ice Giants are the farthest planets from Earth, traditional fully-propulsive orbit insertion missions have transit times to the planetary bodies nearing 13- 15 years and require a large amount of propellant, leaving less mass for the scientific payload and a planetary probe (dry mass percentages of around 30- 40%). Aerocapture uses aerodynamic forces generated by flight within a planetary atmosphere to decelerate and achieve orbit insertion. Although, aerocapture has not been used in the past, recent developments in thermal protection systems (TPS), guidance and control, and interplanetary navigation capabilities enable the use of rigid, heritage entry vehicle configurations already flown at other planetary bodies for Ice Giants aerocapture. Aerocapture can robustly deliver spacecraft to Ice Giant orbits, while substantially increasing on-orbit payload mass (more than 40%) that can be used for a robust atmospheric entry probe. Additionally, the aerocapture maneuver would reduce the interplanetary transit time by 2-5 years (15-30%) relative to fully-propulsive orbit insertion.

S Dutta↗
NASA NTRS2019

The Exo-Brake as an Inexpensive Means of Achieving Sample Return from Low Earth Orbit – Recent Flight

The Exo-Brake is a simple, non-propulsive means of deorbiting small pay-loads from orbital platforms such as the Inter-national Space Station (ISS). Recent flight experiments involving the TechEdSat (TES) 6, 7, 8 are discussed in terms of both ‘targeted’ and ‘disposal’ de-orbit techniques. These build on the previous flight experiments with fixed surface areas – and now involve improved uplink/downlink communication and GPS for improved targeting and control. The recent targeting experiments are dis-cussed involving the TechEdSat-6,7,8 nanosatellites. The extension of the concept to a 1-stage, 3-stage, and lifting entry sample re-turn system are discussed.

Murbach, Marcus S.↗
NASA NTRS2009

Aerothermodynamic Design of the Mars Science Laboratory Heatshield

Aerothermodynamic design environments are presented for the Mars Science Laboratory entry capsule heatshield. The design conditions are based on Navier-Stokes flowfield simulations on shallow (maximum total heat load) and steep (maximum heat flux, shear stress, and pressure) entry trajectories from a 2009 launch. Boundary layer transition is expected prior to peak heat flux, a first for Mars entry, and the heatshield environments were defined for a fully-turbulent heat pulse. The effects of distributed surface roughness on turbulent heat flux and shear stress peaks are included using empirical correlations. Additional biases and uncertainties are based on computational model comparisons with experimental data and sensitivity studies. The peak design conditions are 197 W/sq cm for heat flux, 471 Pa for shear stress, 0.371 Earth atm for pressure, and 5477 J/sq cm for total heat load. Time-varying conditions at fixed heatshield locations were generated for thermal protection system analysis and flight instrumentation development. Finally, the aerothermodynamic effects of delaying launch until 2011 are previewed.

Edquist, Karl T.↗
NASA NTRS2013

Entry, Descent, and Landing with Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing and Shock Phenomena

The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet's moon atmospheres for entry, and descent can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retro-propulsion (SRP) rocket system for the final soft landing. A three engine retro-propulsion configuration with a 2.5 inch diameter sphere-cone aeroshell model was tested in the NASA Glenn 1x1 Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70 degree Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a supersonic retro-propulsion system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retro-propulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. In addition, special topics of electromagnetic interference with retro-propulsion induced shock waves and retro-propulsion for Earth launched booster recovery are also addressed.

Palaszewski, Bryan↗