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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 NTRS1989

Artificial gravity Mars spaceship

Experience gained in the study of artificial gravity for a manned trip to Mars is reviewed, and a snowflake-configured interplanetary vehicle cluster of habitat modules, descent vehicles, and propulsion systems is presented. An evolutionary design is described which permits sequential upgrading from five to nine crew members, an increase of landers from one to as many a three per mission, and an orderly, phased incorporation of advanced technologies as they become available.

Clark, Benton C.↗

NASA NTRS1985

Ion propulsion technology requirements for planetary mission applications

Integrating a clustered, four engine ion propulsion system to the Mariner Mark II spacecraft is shown to reduce total trip time for the five proposed Mariner Mark II missions by fifteen years. In addition, this advanced Solar Electric Propulsion (SEP) system enables rendezvous encounters with five main belt asteroids rather than one. A precursor Advanced Capability Explorer (ACE) technology demonstration vehicle is proposed to investigate SEP spacecraft interaction and control requirements in support of these future mission opportunities. It is shown that the ACE spacecraft could be assembled from existing residual hardware, protoflight equipment and commercially available flight tested components. The ion propulsion system requirements to support ambitious Nuclear Electric Propulsion (NEP) missions are also examined and specific technology development objectives presented.

Aston, G.↗

NASA NTRS2012

Near Earth Asteroid Human Mission Possibilities Using Nuclear Thermal Rocket (NTR) Propulsion

The NTR is a proven technology that generates high thrust and has a specific impulse (Isp (is) approximately 900 s) twice that of today's best chemical rockets. During the Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) programs, twenty rocket reactors were designed, built and ground tested. These tests demonstrated: (1) a wide range of thrust; (2) high temperature carbide-based nuclear fuel; (3) sustained engine operation; (4) accumulated lifetime; and (5) restart capability - all the requirements needed for a human mission to Mars. Ceramic metal fuel was also evaluated as a backup option. In NASA's recent Mars Design reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, versatile vehicle design, simple assembly, and growth potential. In contrast to other advanced propulsion options, NTP requires no large technology scale-ups. In fact, the smallest engine tested during the Rover program - the 25 klbf 'Pewee' engine is sufficient for a human Mars mission when used in a clustered engine configuration. The 'Copernicus crewed NTR Mars transfer vehicle design developed for DRA 5.0 has significant capability that can enable reusable '1-year' round trip human missions to candidate near Earth asteroids (NEAs) like 1991 JW in 2027, or 2000 SG344 and Apophis in 2028. A robotic precursor mission to 2000 SG344 in late 2023 could provide an attractive Flight Technology Demonstration of a small NTR engine that is scalable to the 25 klbf-class engine used for human missions 5 years later. In addition to the detailed scientific data gathered from on-site inspection, human NEA missions would also provide a valuable 'check out' function for key elements of the NTR transfer vehicle (its propulsion module, TransHab and life support systems, etc.) in a 'deep space' environment prior to undertaking the longer duration Mars orbital and landing missions that would follow. The initial mass in low Earth orbit required for a mission to Apophis is approximately 323 t consisting of the NTR propulsion module ((is) approximately 138 t), the integrated saddle truss and LH2 drop tank assembly ((is) approximately 123 t), and the 6-crew payload element ((is) approximately 62 t). The later includes a multi-mission Space Excursion Vehicle (MMSEV) used for close-up examination and sample gathering. The total burn time and required restarts on the three 25 klbf 'Pewee-class' engines operating at Isp (is) approximately 906 s, are approximately 76.2 minutes and 4, respectively, well below the 2 hours and 27 restarts demonstrated on the NERVA eXperimental Engine, the NRX-XE. The paper examines the benefits, requirements and characteristics of using NTP for the above NEA missions. The impacts on vehicle design of HLV payload volume and lift capability, crew size, and reusability are also quantified.

Borowski, Stanley↗

NASA NTRS2012

Modular Growth NTR Space Transportation System for Future NASA Human Lunar, NEA and Mars Exploration Missions

The nuclear thermal rocket (NTR) is a proven, high thrust propulsion technology that has twice the specific impulse (I(sub sp) approx.900 s) of today's best chemical rockets. During the Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) programs, twenty rocket reactors were designed, built and ground tested. These tests demonstrated: (1) a wide range of thrust; (2) high temperature carbide-based nuclear fuel; (3) sustained engine operation; (4) accumulated lifetime; and (5) restart capability - everything required for affordable human missions beyond LEO. In NASA's recent Mars Design Reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower IMLEO, versatile vehicle design, and growth potential. Furthermore, the NTR requires no large technology scale-ups since the smallest engine tested during the Rover program - the 25 klb(sub f) "Pewee" engine is sufficient for human Mars missions when used in a clustered engine configuration. The "Copernicus" crewed Mars transfer vehicle developed for DRA 5.0 was an expendable design sized for fast-conjunction, long surface stay Mars missions. It therefore has significant propellant capacity allowing a reusable "1-year" round trip human mission to a large, high energy near Earth asteroid (NEA) like Apophis in 2028. Using a "split mission" approach, Copernicus and its two key elements - a common propulsion stage and integrated "saddle truss" and LH2 drop tank assembly - configured as an Earth Return Vehicle / propellant tanker, can also support a short round trip (approx.18 month) / short orbital stay (60 days) Mars reconnaissance mission in the early 2030's before a landing is attempted. The same short stay orbital mission can be performed with an "all-up" vehicle by adding an "in-line" LH2 tank to Copernicus to supply the extra propellant needed for this higher energy, opposition-class mission. To transition to a reusable Mars architecture, Copernicus' saddle truss / drop tank assembly is replaced by an in-line tank and "star truss" assembly with paired modular drop tanks to further increase the vehicle's propellant capacity. Shorter "1-way" transit time fast-conjunction Mars missions are another possibility using this vehicle configuration but, as with reusability, increased launch mass is required. "Scaled down" versions of Copernicus (sized to a SLS lift capability of approx.70 t - 100 t) can be developed initially allowing reusable lunar cargo delivery and crewed landing missions, easy NEA missions (e.g., 2000 SG344 also in 2028) or an expendable mission to Apophis. Mission scenario descriptions, key vehicle features and operational characteristics are provided along with a brief discussion of NASA's current activities and its "pre-decisional" plans for future NTR development.

Borowski, Stanley K.↗

NASA NTRS2012

Nuclear Thermal Rocket (Ntr) Propulsion: A Proven Game-Changing Technology for Future Human Exploration Missions

The NTR represents the next evolutionary step in high performance rocket propulsion. It generates high thrust and has a specific impulse (Isp) of approx.900 seconds (s) or more V twice that of today s best chemical rockets. The technology is also proven. During the previous Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) nuclear rocket programs, 20 rocket reactors were designed, built and ground tested. These tests demonstrated: (1) a wide range of thrust; (2) high temperature carbide-based nuclear fuel; (3) sustained engine operation; (4) accumulated lifetime; and (5) restart capability V all the requirements needed for a human mission to Mars. Ceramic metal cermet fuel was also pursued, as a backup option. The NTR also has significant growth and evolution potential. Configured as a bimodal system, it can generate electrical power for the spacecraft. Adding an oxygen afterburner nozzle introduces a variable thrust and Isp capability and allows bipropellant operation. In NASA s recent Mars Design Reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, simple assembly and mission operations. In contrast to other advanced propulsion options, NTP requires no large technology scale-ups. In fact, the smallest engine tested during the Rover program V the 25,000 lbf (25 klbf) Pewee engine is sufficient for human Mars missions when used in a clustered engine arrangement. The Copernicus crewed spacecraft design developed in DRA 5.0 has significant capability and a human exploration strategy is outlined here that uses Copernicus and its key components for precursor near Earth asteroid (NEA) and Mars orbital missions prior to a Mars landing mission. Initially, the basic Copernicus vehicle can enable reusable 1-year round trip human missions to candidate NEAs like 1991 JW and Apophis in the late 2020 s to check out vehicle systems. Afterwards, the Copernicus spacecraft and its 2 key components, now configured as an Earth Return Vehicle / propellant tanker, would be used for a short round trip (approx.18 - 20 months)/short orbital stay (60 days) Mars / Phobos survey mission in 2033 using a split mission approach. The paper also discusses NASA s current Foundational Technology Development activities and its pre-decisional plans for future system-level Technology Demonstrations that include ground testing a small (approx.7.5 klbf) scalable NTR before the decade is out with a flight test shortly thereafter.

Borowski, Stanley K.↗

NASA NTRS2019

Application of Machine Learning Techniques to Delay Tolerant Network Routing

This dissertation discusses several machine learning techniques to improve routing in delay tolerant networks (DTNs). These are networks in which there may be long one-way trip times, asymmetric links, high error rates, and deterministic as well as non-deterministic loss of contact between network nodes, such as interplanetary satellite networks, mobile ad hoc networks and wireless sensor networks. This work uses historical network statistics to train a multi-label classifier to predict reliable paths through the network. In addition, a clustering technique is used to predict future mobile node locations. Both of these techniques are used to reduce the consumption of resources such as network bandwidth, memory and data storage that is required by replication routing methods often used in opportunistic DTN environments. Thesis contributions include: an emulation tool chain developed to create a DTN test bed for machine learning, the network and software architecture for a machine learning based routing method, the development and implementation of classification and clustering techniques and performance evaluation in terms of machine learning and routing metrics.

Dudukovich, Rachel M.↗

NASA NTRS2009

Aerodynamics of a Gulfstream G550 Nose Landing Gear Model

In this paper we discuss detailed steady and unsteady aerodynamic measurements of a Gulfstream G550 nose landing gear model. The quarter-scale, high-fidelity model includes part of the lower fuselage and the gear cavity. The full model configuration allowed for removal of various gear components (e.g. light cluster, steering mechanism, hydraulic lines, etc.) in order to document their effects on the local flow field. The measurements were conducted at a Reynolds number of 7.3 x 10(exp 4) based on the shock strut (piston) diameter and a freestream Mach number of 0.166. Additional data were also collected at lower Mach numbers of 0.12 and 0.145 and correspondingly lower Reynolds numbers. The boundary layer on the piston was tripped to enable turbulent flow separation, so as to better mimic the conditions encountered during flight. Steady surface pressures were gathered from an extensive number of static ports on the wheels, door, fuselage, and within the gear cavity. To better understand the resultant flow interactions between gear components, surface pressure fluctuations were collected via sixteen dynamic pressure sensors strategically placed on various subcomponents of the gear. Fifteen of the transducers were flush mounted on the gear surface at fixed locations, while the remaining one was a mobile transducer that could be placed at numerous varying locations. The measured surface pressure spectra are mainly broadband in nature, lacking any local peaks associated with coherent vortex shedding. This finding is in agreement with off-surface flow measurements using PIV that revealed the flow field to be a collection of separated shear layers without any dominant vortex shedding processes.

Neuhart, Dan H.↗

NASA NTRS2000

Power Systems Evaluated for Solar Electric Propulsion Vehicles

Solar electric propulsion (SEP) mission architectures are applicable to a wide range NASA missions including the robotic exploration of the outer planets in the next decade and the human exploration of Mars within the next 2 decades. SEP enables architectures that are very mass efficient with reasonable power levels (1-MW class) aerobrake and cryogenic upper-stage transportation technologies are utilized. In this architecture, the efficient SEP stage transfers the payload from low Earth orbit (LEO) High Energy Elliptical Parking Orbit (HEEPO) within a period of 6 to 12 months. highthrust, cryogenic upper stage and payload then separate from the SEP vehicle for injection to the planetary target, allowing for fast heliocentric trip times. This mission architecture offers a potential reduction in mass to LEO in comparison to alternative all-chemical nuclear propulsion schemes. Mass reductions may allow launch vehicle downsizing enable missions that would have been grounded because of cost constraints. The preceding figure illustrates a conceptual SEP stage design for a human Mars mission. Researchers at the NASA Glenn Research Center at Lewis Field designed conceptual SEP vehicle, conceived the mission architecture to use this vehicle, and analyzed the vehicle s performance. This SEP stage has a dry mass of 35 metric tons (MT), 40 MT of xenon propellant, and a photovoltaic array that spans 110 m, providing power to a cluster of eight 100-kW Hall thrusters. The stage can transfer an 80-MT payload and upper stage to the desired HEEPO. Preliminary packaging studies show this space-station-class SEP vehicle meets the proposed "Magnum" launch vehicle and volume requirements with considerable margin. An SEP vehicle for outer planetary missions, such as the Europa Mapper Mission, would be dramatically smaller than human Mars mission SEP stage. In this mission architecture, the SEP power system with the payload to provide spacecraft power throughout the mission. Several photovoltaic array design concepts were considered for the SEP vehicle power system for the human mission to Mars. These include a space station derivative, a SCARLET (Solar Concentrator Arrays with Refractive Linear Element Technology) derivative, and a hybrid inflatable-deployable thin polymer membrane array with thin-film solar cells (as shown in the concept illustration). This concept is based on a design developed for the Next Generation Space Telescope Sun shield. The array is divided into 16 independent electrical sections with 500-V, negative-grounded solar cell strings. The power system employs a channelized, 500-Vdc power management and distribution (PMAD) architecture with lithium ion batteries for energy storage for vehicle and payload secondary loads (the high-power Hall thrusters do not operate in eclipse periods). The 500-V PMAD voltage permits "direct-drive" thruster operation, greatly reducing the power processing unit size, complexity, and power loss. Similar power system architecture, designs, and technology are assumed for the Europa Mapper Mission SEP vehicle. The primary exceptions are that the photovoltaic array is assumed to consist of two rectangular wings and that the power system rating is 15 kW in Earth orbit and 200 W at Europa. To size the SEP vehicle power system, a dedicated Fortran code was developed to predict detailed power system performance, mass, and thermal control requirements. This code also modeled all the relevant Earth orbit environments; that is, the particulate radiation, plasma, meteoroids and debris, ultraviolet radiation, contamination, and thermal conditions. Analysis results for the Human Mars Mission SEP vehicle show a power system mass of 9-MT and photovoltaic array area of 5800-square meters for the thin-membrane design concept with CuInS2 thin-film cells. Power processing unit input power for a thin-membrane array design with three-junction, amorphous SiGe solar cells is shown in the graph. Power falls off rapidly inhe first weeks of the mission because of light-induced (Staebler-Wronksi) solar cell losses. During the next 200 days, power decreases steadily as the SEP stage spirals through the proton belts and sustains the bulk of the mission radiation damage. Once the vehicle apogee is above approximately four Earth radii, little additional degradation is incurred. From 400 to 800 days, a 1100-km "parking" orbit is maintained to await the next payload transfer opportunity. This orbit is below the main proton belt, and thus, little radiation dose is accumulated during this time period. During the second LEO-to-HEEPO transfer, power degrades somewhat further, but power requirements are still met. In comparison, the Europa Mapper SEP vehicle power system had a mass of 150 kg and a thin membrane array area of 100 square meters.

Kerslake, Thomas W.↗

NASA NTRS1995

Flight Schedule and the Circadian Clock Influence on Sleep Loss During Overnight Cargo Operations

Thirty-four flight crew members were monitored before, during, and after two 8-day overnight cargo duty patterns which involved multiple flights at night crossing no more than one time zone per 24 h. Rectal temperature, heart rate, and wrist activity were recorded every 2 min. Sleep quantity and quality, and nap timing, were noted in a logbook. To reduce the masking effects of physical activity on temperature, 0.28 C was added to each subject's raw temperature data whenever he reported being asleep. For both masked and unmasked data, daily temperature minima were estimated from the multiple complex demodulated waveform. The temperature minima did not show a progressive adaptation to night duty, which was interrupted by a night off after 5 nights on one trip pattern and after 3 nights on the other. On duty days, the average temperature minimum delayed by about 3 h, occurring near the end of the duty period. Daytime sleep episodes averaged 2.9 h shorter than nighttime sleep episodes, and were rated as lighter, less restorative, and poorer overall. Fifty-three percent of subjects slept more than once per 24 h while they were on night duty, compared to 17% when able to sleep at night. The total sleep per 24 h on duty days averaged 1.2 h less than pretrip. Twenty-nine percent of subjects lost more than 2 h of sleep per 24 h across the 8-day duty patterns. After night duty, subjects awoke around 1400 local time, even when they had slept 2-3 h less than a normal nocturnal sleep episode. Consequently, the duration of morning sleep episodes was correlated with the off-duty time (multiple r(sup 2)=0.44, F=37.23, p less than 0.0001). Anecdotally, crew members complained of being unable to sleep longer and not feeling well-rested. These wakeups were clustered 6 h after the temperature minimum, which suggests that they may have been a response to the circadian wakeup signal. Daytime layovers in which crew members were able to sleep again in the evening ended later (0200-0300) and were longer (average 19.2 h versus 14.8 h) than those in which they slept only once in the morning. Overnight cargo crew members are working around the time of the circadian nadir with an accumulating sleep debt. Two scheduling factors affect sleep loss during these operations: how long before the circadian wakeup signal crew members come off duty, and whether the layover lasts long enough to permit a second sleep episode in the early evening.

Gander, Philippa H.↗