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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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NASA NTRS2015

Planning for Crew Exercise for Future Deep Space Mission Scenarios

Providing the necessary exercise capability to protect crew health for deep space missions will bring new sets of engineering and research challenges. Exercise has been found to be a necessary mitigation for maintaining crew health on‐orbit and preparing the crew for return to earth's gravity. Health and exercise data from Apollo, Space Lab, Shuttle, and International Space Station missions have provided insight into crew deconditioning and the types of activities that can minimize the impacts of microgravity on the physiological systems. The hardware systems required to implement exercise can be challenging to incorporate into spaceflight vehicles. Exercise system design requires encompassing the hardware required to provide mission specific anthropometrical movement ranges, desired loads, and frequencies of desired movements as well as the supporting control and monitoring systems, crew and vehicle interfaces, and vibration isolation and stabilization subsystems. The number of crew and operational constraints also contribute to defining the what exercise systems will be needed. All of these features require flight vehicle mass and volume integrated with multiple vehicle systems. The International Space Station exercise hardware requires over 1,800 kg of equipment and over 24 m3 of volume for hardware and crew operational space. Improvements towards providing equivalent or better capabilities with a smaller vehicle impact will facilitate future deep space missions. Deep space missions will require more understanding of the physiological responses to microgravity, understanding appropriate mitigations, designing the exercise systems to provide needed mitigations, and integrating effectively into vehicle design with a focus to support planned mission scenarios. Recognizing and addressing the constraints and challenges can facilitate improved vehicle design and exercise system incorporation.

Moore, Cherice↗
NASA NTRS2015

High Ice Water Content at Low Radar Reflectivity near Deep Convection: Consistency of In Situ and Remote-Sensing Observations with Stratiform Rain Column Simulations - Part I

Occurrences of jet engine power loss and damage have been associated with flight through fully glaciated deep convection at -10 to -50 degrees Centigrade. Power loss events commonly occur during flight through radar reflectivity (Zeta (sub e)) less than 20-30 decibels relative to Zeta (dBZ - radar returns) and no more than moderate turbulence, often overlying moderate to heavy rain near the surface. During 2010-2012, Airbus carried out flight tests seeking to characterize the highest ice water content (IWC) in such low-radar-reflectivity regions of large, cold-topped storm systems in the vicinity of Cayenne, Darwin, and Santiago. Within the highest IWC regions encountered, at typical sampling elevations (circa 11 kilometers), the measured ice size distributions exhibit a notably narrow concentration of mass over area-equivalent diameters of 100-500 micrometers. Given substantial and poorly quantified measurement uncertainties, here we evaluate the consistency of the Airbus in situ measurements with ground-based profiling radar observations obtained under quasi-steady, heavy stratiform rain conditions in one of the Airbus-sampled locations. We find that profiler-observed radar reflectivities and mean Doppler velocities at Airbus sampling temperatures are generally consistent with those calculated from in situ size-distribution measurements. We also find that column simulations using the in situ size distributions as an upper boundary condition are generally consistent with observed profiles of radar reflectivity (Ze), mean Doppler velocity (MDV), and retrieved rain rate. The results of these consistency checks motivate an examination of the microphysical pathways that could be responsible for the observed size-distribution features in Ackerman et al. (2015).

Rain↗
NASA NTRS1994

Low Power Ground-Based Laser Illumination for Electric Propulsion Applications

A preliminary evaluation of low power, ground-based laser powered electric propulsion systems is presented. A review of available and near-term laser, photovoltaic, and adaptive optic systems indicates that approximately 5-kW of ground-based laser power can be delivered at an equivalent one-sun intensity to an orbit of approximately 2000 km. Laser illumination at the proper wavelength can double photovoltaic array conversion efficiencies compared to efficiencies obtained with solar illumination at the same intensity, allowing a reduction in array mass. The reduced array mass allows extra propellant to be carried with no penalty in total spacecraft mass. The extra propellant mass can extend the satellite life in orbit, allowing additional revenue to be generated. A trade study using realistic cost estimates and conservative ground station viewing capability was performed to estimate the number of communication satellites which must be illuminated to make a proliferated system of laser ground stations economically attractive. The required number of satellites is typically below that of proposed communication satellite constellations, indicating that low power ground-based laser beaming may be commercially viable. However, near-term advances in low specific mass solar arrays and high energy density batteries for LEO applications would render the ground-based laser system impracticable.

Lapointe, Michael R.↗
NASA NTRS2013

Intravenous Fluid Generation System

The ability to stabilize and treat patients on exploration missions will depend on access to needed consumables. Intravenous (IV) fluids have been identified as required consumables. A review of the Space Medicine Exploration Medical Condition List (SMEMCL) lists over 400 medical conditions that could present and require treatment during ISS missions. The Intravenous Fluid Generation System (IVGEN) technology provides the scalable capability to generate IV fluids from indigenous water supplies. It meets USP (U.S. Pharmacopeia) standards. This capability was performed using potable water from the ISS; water from more extreme environments would need preconditioning. The key advantage is the ability to filter mass and volume, providing the equivalent amount of IV fluid: this is critical for remote operations or resource- poor environments. The IVGEN technology purifies drinking water, mixes it with salt, and transfers it to a suitable bag to deliver a sterile normal saline solution. Operational constraints such as mass limitations and lack of refrigeration may limit the type and volume of such fluids that can be carried onboard the spacecraft. In addition, most medical fluids have a shelf life that is shorter than some mission durations. Consequently, the objective of the IVGEN experiment was to develop, design, and validate the necessary methodology to purify spacecraft potable water into a normal saline solution, thus reducing the amount of IV fluids that are included in the launch manifest. As currently conceived, an IVGEN system for a space exploration mission would consist of an accumulator, a purifier, a mixing assembly, a salt bag, and a sterile bag. The accumulator is used to transfer a measured amount of drinking water from the spacecraft to the purifier. The purifier uses filters to separate any air bubbles that may have gotten trapped during the drinking water transfer from flowing through a high-quality deionizing cartridge that removes the impurities in the water before entering the salt bag and mixing with the salt to create a normal saline solution.

McQuillen, John↗
NASA NTRS1988

Conceptual design of a lunar base solar power plant lunar base systems study task 3.3

The best available concepts for a 100 kW Solar Lunar Power Plant based on static and dynamic conversion concepts have been examined. The two concepts which emerged for direct comparison yielded a difference in delivered mass of 35 MT, the mass equivalent of 1.4 lander payloads, in favor of the static concept. The technologies considered for the various elements are either state-of-the-art or near-term. Two photovoltaic cell concepts should receive high priority for development: i.e., amorphous silicon and indium phosphide cells. The amorphous silicon, because it can be made so light weight and rugged; and the indium phosphide, because it shows very high efficiency potential and is reportedly not degraded by radiation. Also the amorphous silicon cells may be mounted on flexible backing that may roll up much like a carpet for compact storage, delivery, and ease of deployment at the base. The fuel cell and electrolysis cell technology is quite well along for lunar base applications, and because both the Shuttle and the forthcoming Space Station incorporate these devices, the status quo will be maintained. Early development of emerging improvements should be implemented so that essential life verification test programs may commence.

Source record↗
NASA NTRSDate not supplied

Advanced Electric Propulsion System (AEPS) Enabling a Sustainable Return to the Lunar Surface through NASA Gateway

NASA continues to evolve a human exploration approach for beyond low-Earth orbit. The center of this approach is NASA’s Gateway that is envisioned to provide a maneuverable outpost in lunar orbit to extend human presence in deep space and expand on NASA exploration goals. The Gateway represents the initial step in NASA’s architecture for human cislunar operations, lunar surface access and missions to Mars. NASA announced at the May 2020 NASA Advisory Council’s Human Explorations and Operations Committee a new plan that calls for launching the first two elements of Gateway as a co-manifested mission in the late 2023 timeframe [2]. Launching the Power and Propulsion Element (PPE) and the Habitation and Logistics Outpost (HALO) together reduces mission risk, utilizes the PPE high-powered Electric Propulsion (EP) system to transport both elements to the lunar orbit, and reduces overall cost. NASA and Maxar Technologies have a commercial partnership to develop and demonstration a high-powered Solar Electric Propulsion (SEP) spacecraft [3, 4]. The PPE is baselined to include three 12.5-kW Advanced Electric Propulsion Systems (AEPS) and four 6-kW Hall thrusters, currently under development by Maxar, for a total beginning of life propulsion power of over 48-kW [5]. High-power solar electric propulsion is one of the key technologies that has been prioritized because of its significant exploration benefits, specifically, for missions beyond low Earth orbit. Spacecraft size and mass are currently dominated by onboard chemical propulsion systems and propellants that may constitute more than 50 percent of spacecraft mass. This impact can be substantially reduced through the utilization of SEP, due to its higher specific impulse and lower propellant load required to meet the equivalent mission delta-V. Studies performed for NASA’s HEOMD and Science Mission Directorate (SMD) have demonstrated that 40-kW-class SEP provides the necessary capabilities that would enable near term and future architectures, and science missions [6]. Accordingly, NASA has been developing a 12 kW Hall thruster electric propulsion thruster that can serve as the building block for a 40-kW-class SEP capability. The AEPS development, led by the NASA Glenn Research Center (GRC) and the Jet Propulsion Laboratory (JPL), began with the maturation of the high-power Hall thruster. The technology development work has transitioned to AR via a competitive procurement selection for the AEPS contract in May 2016. Management of the AEPS contract is being led by NASA GRC with funding from NASA’s Science Technology Mission Directorate (STMD) under the Technology Demonstration Missions (TDM) program. NASA continues to support the AEPS development leveraging in-house expertise, plasma modeling capability, and world-class test facilities.

AEPS↗
NASA NTRS2004

Evaluating and optimizing horticultural regimes in space plant growth facilities

In designing innovative space plant growth facilities (SPGF) for long duration space flight, various limitations must be addressed including onboard resources: volume, energy consumption, heat transfer and crew labor expenditure. The required accuracy in evaluating on board resources by using the equivalent mass methodology and applying it to the design of such facilities is not precise. This is due to the uncertainty of the structure and not completely understanding the properties of all associated hardware, including the technology in these systems. We present a simple criteria of optimization for horticultural regimes in SPGF: Qmax = max [M x (EBI)2/(V x E x T], where M is the crop harvest in terms of total dry biomass in the plant growth system; EBI is the edible biomass index (harvest index), V is volume occupied by the crop; E is the crop light energy supply during growth; T is the crop growth duration. The criterion reflects directly on the consumption of onboard resources for crop production. c2004 COSPAR. Published by Elsevier Ltd. All rights reserved.

NASA Discipline Life Support Systems↗
NASA NTRS2003

Nanolaminate Mirrors With "Piston" Figure-Control Actuators

Efforts are under way to develop a special class of thin-shell curved mirrors for high-resolution imaging in visible and infrared light in a variety of terrestrial or extraterrestrial applications. These mirrors can have diameters of the order of a meter and include metallic film reflectors on nanolaminate substrates supported by multiple distributed piezoceramic gpiston h-type actuators for micron-level figure control. Whereas conventional glass mirrors of equivalent size and precision have areal mass densities between 50 and 150 kg/sq m, the nanolaminate mirrors, including not only the reflector/ shell portions but also the actuators and the backing structures needed to react the actuation forces, would have areal mass densities that may approach .5 kg/m2. Moreover, whereas fabrication of a conventional glass mirror of equivalent precision takes several years, the reflector/shell portion of a nanolaminate mirror can be fabricated in less than a week, and its actuation system can be fabricated in 1 to 2 months. The engineering of these mirrors involves a fusion of the technological heritage of multisegmented adaptive optics and deformable mirrors with more recent advances in metallic nanolaminates and in mathematical modeling of the deflections of thin, curved shells in response to displacements by multiple, distributed actuators. Because a nanolaminate shell is of the order of 10 times as strong as an otherwise identical shell made of a single, high-strength, non-nanolaminate metal suitable for mirror use, a nanolaminate mirror can be made very thin (typically between 100 and 150 m from the back of the nanolaminate substrate to the front reflecting surface). The thinness and strength of the nanolaminate are what make it possible to use distributed gpiston h-type actuators for surface figure control with minimal local concentrated distortion (called print-through in the art) at the actuation points.

Lowman, Andrew↗
NASA NTRS1987

A photometric and spectroscopic comparison of the cataclysmic variables on the 2 sides of the period gap and at a specific orbital period

New and previously existing photometric colors and spectroscopic H and He line strengths are used to compare CV disk systems below the period gap (P=80-130 min) with those directly above (P=190-240 min). Significant differences are found in the mean U-B, V-J colors and in the H-beta equivalent widths on the two sides of the gap. A detailed comparison from UV-IR of 3 high inclination novalike systems with periods near 3.3 hr (PG1012-029, PG1030 + 590 and V1315 Aql) shows similar continuum distributions (and implied mass accretion rates) but large differences in high excitation lines (He III 640 A and 4686 A; C IV 1550 A).

Szkody, P.↗
NASA NTRS1992

Solar electric propulsion for the orbital transfer of large spacecraft to geosynchronous orbit

Low thrust trajectories utilizing electric propulsion and photovoltaic power sources are derived for the transfer of a large spacecraft from assembly in low earth orbit to operation at geosynchronous orbit. Conventional silicon photovoltaic devices are assumed. A 1 MeV equivalent electron flux model is used. The solar panel mass and size are taken from Advanced Photovoltaic Solar Array design concepts with 3-mil shielding. Shadowing of the vehicle and low earth orbit atmospheric drag are also modeled. The power system is sized so that end of life output is 30 percent of initial output. Steering of the thrust vector is defined by a three term series in which the coefficients are varied as controls to minimize the total power loss with cumulative exposure in the Van Allen region.

Garrison, James L.↗
NASA NTRS2015

Swarm Flyby Gravimetry

This study describes a new technology for discerning the gravity fields and mass distribution of a solar system small body, without requiring dedicated orbiters or landers. Instead of a lander, a spacecraft releases a collection of small, simple probes during a flyby past an asteroid or comet. By tracking those probes from the host spacecraft, one can estimate the asteroid's gravity field and infer its underlying composition and structure. This approach offers a diverse measurement set,equivalent to planning and executing many independent and unique flyby encounters of a single spacecraft. This report assesses a feasible hardware implementation, derives the underlying models,and analyzes the performance of this concept via simulation.In terms of hardware, a small, low mass, low cost implementation is presented, which consists of a dispenser and probes. The dispenser contains roughly 12 probes in a tube and has a total size commensurate with a 6U P-Pod. The probes are housed in disc shaped sabots. When commanded,the dispenser ejects the top-most probe using a linear motor. The ejected probe separates from its sabots and unfolds using internal springs. There are two types of probes, each designed for a particular tracking modality. The reflective probe type, tracked by a telescope, unfolds to forma diffusely reflective sphere. The retroreflector probe type, tracked by a lidar, unfolds to form a corner-cube retroreflector assembly. Both types are designed to spherical so that their attitude doesn't affect the spacecraft's tracking performance.This analysis indicates that the point-mass term of small bodies larger than roughly 500 m in diameter can be observed from a host spacecraft that tracks locally deployed probes throughout a flyby to an uncertainty of better than 5%. The conditions by which this measurement is possible depends on the characteristics of the asteroid (size, type), the flyby velocity, and the type of tracking available (angles-only or angles+ranging). For most encounters, a few (1-3) well placed probes can be very effective, with marginal improvement for additional probes. Given realistic deployment errors, an encounter may require roughly 10-12 probes to ensure that 1-3 achieve their target. Long duration tracking of probes flying by large asteroids (>5 km diameter) can sometimes provide observability of the gravity field's first spherical harmonic, J( sub 2). In summary, this method offers a feasible, affordable approach to enabling or augmenting flyby science.

Atchison, Justin↗
NASA NTRS1977

Wind tunnel tests of a two bladed model rotor to evaluate the TAMI system in descending forward flight

A research investigation was conducted to assess the potential of the Tip Air Mass Injection system in reducing the noise output during blade vortex interaction in descending low speed flight. In general it was concluded that the noise output due to blade vortex interaction can be reduced by 4 to 6 db with an equivalent power expenditure of approximately 14 percent of installed power.

White, R. P., Jr.↗
NASA NTRSDate not supplied

Reactor Parametric Assessments for Alternative Propellant Nuclear Thermal Propulsion Engines

This work focuses on the implications of alternative working fluids for nuclear thermal propulsion (NTP) reactors. To perform this analysis, NASA’s Testing Reference Design (TRD) is altered using parametric studies with selected alternative propellants to determine the reactor modifications required to enable an operable system. This research investigates ammonia, water, methane, helium, and enriched diborane as feasible options to enable a specific impulse (𝐼 𝑠𝑝 ) greater than the chemical propulsion alternatives. Frozen and dissociated 𝐼 𝑠𝑝 for each propellant is shown for variable chamber temperature conditions. Geometrical sensitivities are performed to observe the impact on the neutron multiplication factor (k eff ), system mass, and HALEU loading. Control drum worth curves are included for 5- degree increments of rotation. The base TRD configuration does not enable a critical system for ammonia and enriched diborane, thus these propellants will be highlighted in the final results. Equivalent study results for hydrogen will be included as a point of comparison to the performance of the other options.

Nuclear Thermal Propulsion↗
NASA NTRSDate not supplied

Reactor Parametric Assessments for Alternative Propellant Nuclear Thermal Propulsion Engines

This work focuses on the implications of alternative working fluids for nuclear thermal propulsion (NTP) reactors. To perform this analysis, NASA’s Testing Reference Design (TRD) is altered using parametric studies with selected alternative propellants to determine the reactor modifications required to enable an operable system. This research investigates ammonia, water, methane, helium, and enriched diborane as feasible options to enable a specific impulse (Isp) greater than the chemical propulsion alternatives. Frozen and dissociated Isp for each propellant is shown for variable chamber temperature conditions. Geometrical sensitivities are performed to observe the impact on the neutron multiplication factor (keff), system mass, and HALEU loading. Control drum worth curves are included for 5- degree increments of rotation. The base TRD configuration does not enable a critical system for ammonia and enriched diborane, thus these propellants will be highlighted in the final results. Equivalent study results for hydrogen will be included as a point of comparison to the performance of the other options.

Nuclear Thermal Propulsion↗
NASA NTRS1988

Experimental Determination of Aerodynamic Damping in a Three-Stage Transonic Axial-Flow Compressor

Rotor blade aerodynamic damping is experimentally determined in a three-stage transonic axial flow compressor having design aerodynamic performance goals of 4.5:1 pressure ratio and 65.5 lbm/sec weight flow. The combined damping associated with each mode is determined by a least squares fit of a single degree of freedom system transfer function to the nonsynchronous portion of the rotor blade strain gauge output power spectra. The combined damping consists of aerodynamic and structural and mechanical damping. The aerodynamic damping varies linearly with the inlet total pressure for a given equivalent speed, equivalent mass flow, and pressure ratio while structural and mechanical damping are assumed to be constant. The combined damping is determined at three inlet total pressure levels to obtain the aerodynamic damping. The third stage rotor blade aerodynamic damping is presented and discussed for 70, 80, 90, and 100 percent design equivalent speed. The compressor overall performance and experimental Campbell diagrams for the third stage rotor blade row are also presented.

Newman, Frederick A.↗
NASA NTRS1994

The rotating molecular core in G10.6 - 0.4: Synthesis maps in (12)C(18)O

Using the Owens Valley Radio Observatory millimeter-wave interferometer, the compact molecular cloud core surrounding the H II region G10.60-0.4 has been resolved in the C(18)O J = 1-0 transition. Since this line is most likely optically thin, it is an excellent tracer of the total column density. We find a centrally condensed and flattened (12.6 sec x 4.2 sec ; 0.3 pc x 0.1 pc) core, which is rotating rapidly with a substantial velocity gradient (17 +/- 2 km/s pc) along the major axis. The core is quite bright (greater than 10K), dense (greater than (10(exp 6)/cu cm), massive (10(exp 3) solar mass), and embedded in a more extended (greater than 60 seconds) envelope. There is no evidence for absorption, implying that the C(18)O emitting gas is both optically thin and substantially hotter than the equivalent continuum temperature at 3 mm. The systemic velocity of the rotating core is -2.5 +/- 0.5 km/s, the same as the systemic velocity of the extended cloud.

Ho, Paul T. P.↗
NASA NTRS1983

Orbital motion and mass flow in the interacting binary Be star HR 2142

The discovery of an unusual, periodic, two-component shell phase of short duration in the 'classical' Be star HR2142 (HD41335, MWC133) offered convincing evidence that this object is a mass-transfer binary system. A model based solely on the phase-dependent behavior of the hydrogen shell lines in this 80(d).860 binary was developed by Peters and Polidan (1973) and by Peters (1976). The present investigation is concerned with a refinement to the earlier model, taking into account the utilization of an orbital solution obtained from measurements of the wings of the broad photospheric features observed in the rapidly rotating primary. Velocities and equivalent widths from the sharp 'shell' lines, presumably formed in or near the gas stream, provide additional information on the mass flow in the Balmer-line-formation region.

Peters, G. J.↗