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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 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Technical Characteristics and Accuracy Capabilities of Delta Differential One-Way Ranging (DeltaDOR) as a Spacecraft Navigation Tool

The technique of determining spacecraft angular position by Delta Differential One-Way Ranging (DeltaDOR) is described. The dependence of measurement accuracy on parameters which define the spacecraft signal spectrum is shown. An illustration of DeltaDOR measurement accuracy, using the Mars Observer spacecraft, is presented. Guidelines for specifying spacecraft tone spectra are presented based on typical accuracy and ambiguity resolution needs, on frequency allocation for deep space tracking, on the efficient use of spacecraft signal power, and on the efficient use of ground tracking resources.

VLBI Very Long Baseline Interferometry DOR tones D↗
DOE OSTI2024

Spacecraft surface charging as a function of material properties

Spacecraft material behavior plays a very important role in space missions. Spacecraft immersed in plasma get charged by absorbing plasma particles and by emitting electrons from spacecraft surfaces via photoelectron and secondary electron emission. Spacecraft charging depends heavily on material properties such as work function, secondary electron yield, dielectric constant, and electric conductivity among other. Material properties are typically assumed to be static in charging models. However, it is well known that this is not the case in space. This makes spacecraft charging predictions very challenging. Material properties are well characterized before the spacecraft is put in orbit through characterization in the lab under clean conditions. However, once in space, material properties change due to the harsh and very dynamic space environment. We present a new capability to predict material behavior in space from first-principles modeling. The ongoing effort seeks to couple material models, density functional theory (DFT) and molecular dynamic (MD) codes, with environment models, plasma kinetic codes. This preliminary study will show results of surface charging as a function of material work function, dielectric constant, and conductivity.

36 MATERIALS SCIENCE↗
NASA NTRS2012

Printable Spacecraft: Flexible Electronic Platforms for NASA Missions

Atmospheric confetti. Inchworm crawlers. Blankets of ground penetrating radar. These are some of the unique mission concepts which could be enabled by a printable spacecraft. Printed electronics technology offers enormous potential to transform the way NASA builds spacecraft. A printed spacecraft's low mass, volume and cost offer dramatic potential impacts to many missions. Network missions could increase from a few discrete measurements to tens of thousands of platforms improving areal density and system reliability. Printed platforms could be added to any prime mission as a low-cost, minimum resource secondary payload to augment the science return. For a small fraction of the mass and cost of a traditional lander, a Europa flagship mission might carry experimental printed surface platforms. An Enceladus Explorer could carry feather-light printed platforms to release into volcanic plumes to measure composition and impact energies. The ability to print circuits directly onto a variety of surfaces, opens the possibility of multi-functional structures and membranes such as "smart" solar sails and balloons. The inherent flexibility of a printed platform allows for in-situ re-configurability for aerodynamic control or mobility. Engineering telemetry of wheel/soil interactions are possible with a conformal printed sensor tape fit around a rover wheel. Environmental time history within a sample return canister could be recorded with a printed sensor array that fits flush to the interior of the canister. Phase One of the NIAC task entitled "Printable Spacecraft" investigated the viability of printed electronics technologies for creating multi-functional spacecraft platforms. Mission concepts and architectures that could be enhanced or enabled with this technology were explored. This final report captures the results and conclusions of the Phase One study. First, the report presents the approach taken in conducting the study and a mapping of results against the proposed Phase One objectives. Then an overview of the general field of printed electronics is provided, including manufacturing approaches, commercial drivers, and the current state of integrated systems. The bulk of the report contains the results and findings of Phase One organized into four sections: a survey of components required for a printable spacecraft, technology roadmaps considerations, science mission and engineering applications, and potential risks and challenges of the technology.

Spacecraft↗
NASA NTRS2016

Spacecraft Fire Safety Research at NASA Glenn Research Center

Appropriate design of fire detection systems requires knowledge of both the expected fire signature and the background aerosol levels. Terrestrial fire detection systems have been developed based on extensive study of terrestrial fires. Unfortunately there is no corresponding data set for spacecraft fires and consequently the fire detectors in current spacecraft were developed based upon terrestrial designs. In low gravity, buoyant flow is negligible which causes particles to concentrate at the smoke source, increasing their residence time, and increasing the transport time to smoke detectors. Microgravity fires have significantly different structure than those in 1-g which can change the formation history of the smoke particles. Finally the materials used in spacecraft are different from typical terrestrial environments where smoke properties have been evaluated. It is critically important to detect a fire in its early phase before a flame is established, given the fixed volume of air on any spacecraft. Consequently, the primary target for spacecraft fire detection is pyrolysis products rather than soot. Experimental investigations have been performed at three different NASA facilities which characterize smoke aerosols from overheating common spacecraft materials. The earliest effort consists of aerosol measurements in low gravity, called the Smoke Aerosol Measurement Experiment (SAME), and subsequent ground-based testing of SAME smoke in 55-gallon drums with an aerosol reference instrument. Another set of experiments were performed at NASAs Johnson Space Center White Sands Test Facility (WSTF), with additional fuels and an alternate smoke production method. Measurements of these smoke products include mass and number concentration, and a thermal precipitator was designed for this investigation to capture particles for microscopic analysis. The final experiments presented are from NASAs Gases and Aerosols from Smoldering Polymers (GASP) Laboratory, with selected results focusing on realistic fuel preparations and heating profiles with regards to early detection of smoke. SAFFIRE is the upcoming large-scale fire experiment which will be executed in a Cygnus vehicle after it undocks from the ISS.

Aerosol↗
NASA NTRS2021

BUMPER: A Tool for Analyzing Spacecraft Micrometeoroid and Orbital Debris Risk

“Bumper” is a computer program for analyzing spacecraft micrometeoroid and orbital debris (MMOD) risk. Bumper was developed in the late-1980s and has been continuously maintained and used since. The user base has grown from a few government entities now include numerous commercial entities as well. The National Aeronautics and Space Administration (NASA) Johnson Space Center (JSC) Hypervelocity Impact Technology (HVIT) group is responsible for all aspects of the Bumper software. Bumper has been used to characterize MMOD risk on many spacecraft. All of the International Space Station (ISS) modules, visiting vehicles and numerous external components and systems have been analyzed. Bumper was used to analyze the Space Shuttle, Orion, and many space probes, telescopes and satellites. Bumper is also being used to analyze future spacecraft such as the Deep Space Gateway (DSG) and Mars Sample Return (MSR) missions. The Bumper Configuration Control Board (CCB) ensures that all changes to the code are approved, reviewed, and documented. The current Bumper version – “Bumper 3” – is a Fortran executable that utilizes a 64-bit architecture. Bumper has numerous features that make it a powerful tool for analyzing spacecraft MMOD risk. Bumper uses the latest orbital debris and meteoroidenvironment models. Bumper also has a large library of ballistic limit “damage” equations available that can be used for a wide variety of MMOD shielding configurations. Bumper can also handle large spacecraft finite elementmodels (FEMs) and conducts checks of the model. This paper introduces Bumper and the MMOD risk analysis process using a simplified cube-shaped spacecraft model

hypervelocity↗
NASA NTRSDate not supplied

Capabilities of Single Launch NTP and Chemical Spacecrafts for Cis-Lunar Tug Missions

Renewed interest in the region between the Earth and the moon suggests that a tug operating between these two destinations will be most likely required in the near future. This report details a parametric spacecraft model which sizes a nuclear thermal propulsion or chemical propulsion spacecraft and compares their payload transport capabilities for missions to the moon. A spacecraft sized to fit in a single launch vehicle is assumed, and six launch vehicles are investigated. Maximum spacecraft payload mass and minimum mission transit time are compared between nuclear and chemical propulsion spacecrafts. For vehicles sized to the maximum capability of a launch vehicle, payloads of at least five metric tons are possible with either propulsion option; with nuclear thermal propulsion able to provide additional mass or reduced transit time when the spacecraft is sized to the maximum capability of the heaviest-lift launch vehicles. Transit times to the moon with a five-ton payload could be reduced by at least 0.2 days with a nuclear propulsion system.

Cis-lunar tug↗
NASA NTRSDate not supplied

Capabilities of Single Launch NTP and Chemical Spacecrafts for Cis-Lunar Tug Missions

Renewed interest in the region between the Earth and the moon suggests that a tug operating between these two destinations will be most likely required in the near future. This report details a parametric spacecraft model which sizes a nuclear thermal propulsion or chemical propulsion spacecraft and compares their payload transport capabilities for missions to the moon. A spacecraft sized to fit in a single launch vehicle is assumed, and six launch vehicles are investigated. Maximum spacecraft payload mass and minimum mission transit time are compared between nuclear and chemical propulsion spacecrafts. For vehicles sized to the maximum capability of a launch vehicle, payloads of at least five metric tons are possible with either propulsion option; with nuclear thermal propulsion able to provide additional mass or reduced transit time when the spacecraft is sized to the maximum capability of the heaviest-lift launch vehicles. Transit times to the moon with a five-ton payload could be reduced by at least 0.2 days with a nuclear propulsion system.

Cis-lunar tug↗
DOE OSTI2020

Hypervelocity impact flash expansion geometry under various spacecraft surface electrical conditions

A more congested space environment due to increasing space activities introduces new threats to spacecraft operators. Among these threats include hypervelocity impacts from microparticles (< 1 µg) in the form of meteoroids and space debris. These microparticles travel at speeds between 11.2 and 72.8 km/s with respect to the Earth and can impact spacecraft forming a small (~1 µm) and dense (~ 10 26 m -3 ) plasma. This plasma can generate a strong optical emission (impact flash) and a radio frequency (RF) emission. These emissions can lead to spacecraft electrical anomalies when the impacted spacecraft surface (target) carries high electrical potential due to various space weather effects. To understand the microparticle hypervelocity impact plasma and its associated threat to spacecraft electronics, we studied the hypervelocity impact flash expansion geometry under various spacecraft surface electrical conditions. Three high-speed broadband photomultiplier tubes (PMT) are used to capture the temporal evolution of the impact flash and the temporal evolution of the angular distribution of the impact flash expansion geometry in a 3 MV electrostatic dust accelerator during a ground-based experiment. Overall, results indicate that the impact surface electrical condition has a statistically significant influence on the impact flash geometry, and a strong similarity between the impact plasma and flash expansion geometry. Thus, this paper provides new experimental support for the impact flash origin from and connection with the impact plasma, and serves as the first impact flash expansion geometry characterization under various impact surface electrical conditions in a collisionless environment (< 10 -5 Torr).

42 ENGINEERING↗
DOE OSTI2022

Spacecraft Geometry Effects on Kinetic Impactor Missions

Abstract The DART (Double Asteroid Redirection Test) mission will impact a spacecraft on the secondary (Dimorphos) of the binary asteroid system Didymos in 2022 September, with the goal of altering the orbital period of Dimorphos about Didymos sufficiently to be observed from ground-based observations. Numerical impact modeling is a crucial component in understanding the outcome of the DART experiment, and while many have investigated the effects of target properties, such as material strength and porosity (which remain unknown), an often overlooked factor is the importance of accurately representing the spacecraft itself in such models. Most impact modeling to date has considered simple impactor geometries such as a solid uniform sphere, but in reality the spacecraft is a complex shape full of different components, open spaces, and thin walled structures. At a minimum, a simple solid representation underestimates the surface area of the impact: for a small body such as Dimorphos (approximately 160 m in diameter), the difference between a spacecraft spanning 20 m (including solar arrays) impacting and a sub-1 m idealized shape may be important. In this paper, we compare models impacting high-fidelity models of the spacecraft based on the CAD geometry with various simplified impactors, in order to assess the potential importance of this effect. We find that the difference between the simplest impactor geometries (such as a uniform sphere) and the real spacecraft is measurable, and has an interesting dependence on the material properties of the asteroid itself.

36 MATERIALS SCIENCE↗
NASA NTRS1971

Simulation of Mariner Mars 1971 spacecraft

Preparation for the Mariner Mars 1971 mission is reported, including an extensive training program for operations personnel during which the primary source of spacecraft data was a computer program simulating the spacecraft. The objectives of a simulation model for training purposes differed from objectives appropriate to a design or analysis model. Model subsystems were designed to provide realistic telemetry data reflecting changes due both to commands and environmental parameters affecting the spacecraft at various times during the mission. The spacecraft is modeled along two separate functional lines. Boolean operations are concentrated in the spacecraft logic model, which determines the spacecraft state or mode, while mathematical operations or algorithms are executed in computational subsystem models. Although logic parameters are interrogated as a part of each computational pass, actual logic model processing occurs only when a change-of-state input is generated by the operations organization. The program design, some of the special characteristics of each of the modeled subsystems, and how the model was used in support of mission operations training are presented.

Ausman, N. E., Jr.↗
NASA NTRS1972

Predicting spacecraft self-contamination in space and in a test chamber

The self-contamination of spacecraft (defined as the return and deposition of outgassed molecules on its critical surfaces, either in orbit or while undergoing vacuum) is considered. Theoretical relations for the flux, density, and pressure of the emitted gas as a function of altitude, radius, and distance from the spacecraft surface are developed. The flux of the outgassed molecules that return to the emitting surface is also obtained and shown to be dependent on altitude, spacecraft dimensions, and the magnitude of outgassing. The rate of condensation and the time for the formation of a monolayer of the returning molecules can be calculated. The self-contamination of spacecraft undergoing vacuum chamber test is also theoretically examined and compared with the equivalent parameters for orbit conditions. It is concluded that, depending on the dimensions of the spacecraft relative to those of the chamber and the wall capture coefficient, ground tests conducted in the more usual space simulation chambers can provide returning fluxes and self-contamination comparable to those occurring in space up to an altitude of about 400 km. For higher altitudes and return fluxes less than 0.001 of those emitted, the chamber test can produce a greater contamination. In this case, the ground results can be related to those obtained in space, provided that the wall capture coefficient is known or if the ratio of returned to emitted flux at the spacecraft surface is measured.

Scialdone, J. J.↗
NASA NTRS1973

Standardization and economics of nuclear spacecraft: Executive summary

Feasibility and cost benefits of nuclear-powered standardized spacecraft were investigated. The study indicates that two shuttle-launched nuclear-powered spacecraft should be able to serve the majority of unmanned NASA missions anticipated for the 1980's. The standard spacecraft include structure, thermal control, power, attitude control, some propulsion capability and tracking, telemetry, and command subsystems. One spacecraft design, powered by the radioisotope thermoelectric generator, can serve missions requiring up to 450 watts. The other spacecraft design, powered by similar nuclear heat sources in a Brayton-cycle generator, can serve missions requiring up to 2200 watts. Design concepts and trade-offs are discussed. The conceptual designs selected are presented and successfully tested against a variety of missions. The thermal design is such that both spacecraft are capable of operating in any earth orbit and any orientation without modification.

Source record↗
NASA NTRS1972

Tracking and data relay satellite system configuration and tradeoff study. Volume 5: TDRS spacecraft design, part 1

A dual spin stabilized TDR spacecraft design is presented for low data rate (LDR) and medium data rate (MDR) user spacecraft telecommunication relay service. The relay satellite provides command and data return channels for unmanned users together with duplex voice and data communication channels for manned user spacecraft. TDRS/ground links are in the Ku band. Command links are provided at UHF for LDR users and S band for MDR users. Voice communication channels are provided at UHF/VHF for LDR users and at S band for MDR users. The spacecraft is designed for launch on the Delta 2914 with system deployment planned for 1978. This volume contains a description of the overall TDR spacecraft configuration, a detailed description of the spacecraft subsystems, a reliability analysis, and a product effectiveness plan.

Source record↗
NASA NTRS1973

Telescopic booms for the Hawkeye spacecraft

Several boom arrangements were proposed for the Hawkeye spacecraft. The design chosen to fly consists of one pair of stem roll-out booms and a set of telescopic booms. Some of the advantages of choosing this boom configuration were that the booms could be stored in the belt line area, which is near the center of gravity of the spacecraft, and also with the approach the telescopic booms could be extended at the spacecraft's initial spin rate (180 rpm). By being able to extend these booms at the initial spin rate of the spacecraft the need for a despin system was eliminated, thus weight saving to the spacecraft. With the method of storage and extension employed by the telescopic design, the desirable sensor distance from the spacecraft could also be attained.

Anderson, R. D.↗
NASA NTRS1973

System design of the Pioneer Venus spacecraft. Volume 1: Executive summary

The NASA Ames Research Center Pioneer Venus Project objective is to conduct scientific investigations of the planet Venus using spin stabilized spacecraft. The defined approach to accomplish this goal is to implement a multiprobe spacecraft mission and an orbiter spacecraft mission. Candidate launch vehicles for the Pioneer Venus missions were the Thor/Delta and Atlas/Centaur. The multiprobe spacecraft consists of a probe bus, one large probe, and three small probes. The probes are designed to survive to the surface of Venus, and to make in situ measurements of the Venusian atmosphere; the probe bus enters the atmosphere and makes scientific measurements until it burns out. The orbiter mission uses a spacecraft designed to orbit Venus for 225 days with an orbit period of about 24 hours (h). The probe bus and orbiter designs are to use a common spacecraft bus.

Dorfman, S. D.↗
NASA NTRS1976

Method and apparatus for neutralizing potentials induced on spacecraft surfaces

A potential induced on the surface of an orbiting spacecraft is neutralized to the potential of a plasma through which the spacecraft is traveling by directing charged particles into the plasma from the spacecraft surface. The induced potential occurs in response to bombardment of the spacecraft surface by ambient charged particles which may be negative or positive. The charged particles directed into the plasma from the surface have the same polarity as the induced potential to provide the neutralization. The invention can be utilized to maintain different, electrically isolated segments of a spacecraft surface at the same potential to prevent electric discharges between the different parts and thereby protect electric circuits within the spacecraft.

Hunter, R. E.↗
NASA NTRS1978

A highly accurate method for the determination of mass and center of mass of a spacecraft

An extremely accurate method for the measurement of mass and the lateral center of mass of a spacecraft has been developed. The method was needed for the Voyager spacecraft mission requirement which limited the uncertainty in the knowledge of lateral center of mass of the spacecraft system weighing 750 kg to be less than 1.0 mm (0.04 in.). The method consists of using three load cells symmetrically located at 120 deg apart on a turntable with respect to the vertical axis of the spacecraft and making six measurements for each load cell. These six measurements are taken by cyclic rotations of the load cell turntable and of the spacecraft, about the vertical axis of the measurement fixture. This method eliminates all alignment, leveling, and load cell calibration errors for the lateral center of mass determination, and permits a statistical best fit of the measurement data. An associated data reduction computer program called MASCM has been written to implement this method and has been used for the Voyager spacecraft.

Chow, E. Y.↗
NASA NTRS1978

System design of an ion drive spacecraft

As electric propulsion technology has improved and mission requirements have changed, a series of Ion Propulsion Module (IPM) design concepts have evolved. The most recent iteration occurred in the NASA-sponsored Halley Comet Rendezvous Mission (HCRM) study of ion drive. Spacecraft system design considerations introduced by the integration of such an IPM as the primary propulsion source are described with reference to the synthesis of the HCRM spacecraft and spacecraft design considerations for other interplanetary applications. IPM interactions with the system (especially telecommunications and science) are found to be manageable. The spacecraft design developed for the HCRM indicates the interface simplicity between the IPM and the spacecraft. Methods are shown for readily applying this IPM to a variety of planetary missions. Methods are also described for the IPM to provide up to 5 kW to the spacecraft for increasing the mission science return

Stuart, J.↗