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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 721 records · Page 40
All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

The Next 25 Years of Deep Space Navigation

This slide presentation reviews the missions that will be flown into deep space in the next 25 years, the navigational challenges for these missions, and the strategies that will be used to overcome these challenges. The challenges include: (1) an incresed need for autonomous navigation, (2) an increased use of in-situ and optical navigation, (3) an increased use of low-thrust propulsion, (4) an increased need for higher accuracy in guidance, navigation, and control, and an increased need for integration between flight path and attitude control. The enabling strategies that are planned for use are: (1) Advance Radio-Metric Tracking Capabilities, (2) Expand the Use of Optical Navigation, (3) Develop General-Purpose Autonomous Navigation Capabilities, (4) Improve Frequency and Timing Systems, and (5) Develop In-situ Tracking Infrastructure. Future trends that are being developed are Optical and Autonomous Navigation

optical navigation↗
NASA NTRS2010

Fiber-Optic Continuous Liquid Sensor for Cryogenic Propellant Gauging

An innovative fiber-optic sensor has been developed for low-thrust-level settled mass gauging with measurement uncertainty <0.5 percent over cryogenic propellant tank fill levels from 2 to 98 percent. The proposed sensor uses a single optical fiber to measure liquid level and liquid distribution of cryogenic propellants. Every point of the sensing fiber is a point sensor that not only distinguishes liquid and vapor, but also measures temperature. This sensor is able to determine the physical location of each point sensor with 1-mm spatial resolution. Acting as a continuous array of numerous liquid/vapor point sensors, the truly distributed optical sensing fiber can be installed in a propellant tank in the same manner as silicon diode point sensor stripes using only a single feedthrough to connect to an optical signal interrogation unit outside the tank. Either water or liquid nitrogen levels can be measured within 1-mm spatial resolution up to a distance of 70 meters from the optical interrogation unit. This liquid-level sensing technique was also compared to the pressure gauge measurement technique in water and liquid nitrogen contained in a vertical copper pipe with a reasonable degree of accuracy. It has been demonstrated that the sensor can measure liquid levels in multiple containers containing water or liquid nitrogen with one signal interrogation unit. The liquid levels measured by the multiple fiber sensors were consistent with those virtually measured by a ruler. The sensing performance of various optical fibers has been measured, and has demonstrated that they can survive after immersion at cryogenic temperatures. The fiber strength in liquid nitrogen has also been measured. Multiple water level tests were also conducted under various actual and theoretical vibration conditions, and demonstrated that the signal-to-noise ratio under these vibration conditions, insofar as it affects measurement accuracy, is manageable and robust enough for a wide variety of spacecraft applications. A simple solution has been developed to absorb optical energy at the termination of the optical sensor, thereby avoiding any feedback to the optical interrogation unit

Xu. Wei↗
NASA NTRS2010

Mars Earth Return Vehicle (MERV) Propulsion Options

The COMPASS Team was tasked with the design of a Mars Sample Return Vehicle. The current Mars sample return mission is a joint National Aeronautics and Space Administration (NASA) and European Space Agency (ESA) mission, with ESA contributing the launch vehicle for the Mars Sample Return Vehicle. The COMPASS Team ran a series of design trades for this Mars sample return vehicle. Four design options were investigated: Chemical Return /solar electric propulsion (SEP) stage outbound, all-SEP, all chemical and chemical with aerobraking. The all-SEP and Chemical with aerobraking were deemed the best choices for comparison. SEP can eliminate both the Earth flyby and the aerobraking maneuver (both considered high risk by the Mars Sample Return Project) required by the chemical propulsion option but also require long low thrust spiral times. However this is offset somewhat by the chemical/aerobrake missions use of an Earth flyby and aerobraking which also take many months. Cost and risk analyses are used to further differentiate the all-SEP and Chemical/Aerobrake options.

Oleson, Steven R.↗
NASA NTRS2003

Human Exploration of Earth's Neighborhood and Mars

The presentation examines Mars landing scenarios, Earth to Moon transfers comparing direct vs. via libration points. Lunar transfer/orbit diagrams, comparison of opposition class and conjunction class missions, and artificial gravity for human exploration missions. Slides related to Mars landing scenarios include: mission scenario; direct entry landing locations; 2005 opportunity - Type 1; Earth-mars superior conjunction; Lander latitude accessibility; Low thrust - Earth return phase; SEP Earth return sequence; Missions - 200, 2007, 2009; and Mission map. Slides related to Earth to Moon transfers (direct vs. via libration points (L1, L2) include libration point missions, expeditionary vs. evolutionary, Earth-Moon L1 - gateway for lunar surface operations, and Lunar mission libration point vs. lunar orbit rendezvous (LOR). Slides related to lunar transfer/orbit diagrams include: trans-lunar trajectory from ISS parking orbit, trans-Earth trajectories, parking orbit considerations, and landing latitude restrictions. Slides related to comparison of opposition class (short-stay) and conjunction class (long-stay) missions for human exploration of Mars include: Mars mission planning, Earth-Mars orbital characteristics, delta-V variations, and Mars mission duration comparison. Slides related to artificial gravity for human exploration missions include: current configuration, NEP thruster location trades, minor axis rotation, and example load paths.

Condon, Gerald↗
NASA NTRS2010

Application of the NEXT Ion Thruster Lifetime Assessment to Thruster Throttling

Ion thrusters are low thrust, high specific impulse devices with typical operational lifetimes of 10,000 to 30,000 hr over a range of throttling conditions. The NEXT ion thruster is the latest generation of ion thrusters under development. The NEXT ion thruster currently has a qualification level propellant throughput requirement of 450 kg of xenon, which corresponds to roughly 22,000 hr of operation at the highest input power throttling point. This paper will provide a brief review the previous life assessment predictions for various throttling conditions. A further assessment will be presented examining the anticipated accelerator grid hole wall erosion and related electron backstreaming limit. The continued assessment of the NEXT ion thruster indicates that the first failure mode across the throttling range is expected to be in excess of 36,000 hr of operation from charge exchange induced groove erosion. It is at this duration that the groove is predicted to penetrate the accelerator grid possibly resulting in structural failure. Based on these lifetime and mission assessments, a throttling approach is presented for the Long Duration Test to demonstrate NEXT thruster lifetime and validate modeling.

VanNoord, Jonathan L.↗
NASA NTRS2011

Energy Systems Divisions

This slide presentation reviews the JSC Energy Systems Divisions work in propulsion. Specific work in LO2/CH4 propulsion, cryogenic propulsion, low thrust propulsion for Free Flyer, robotic and Extra Vehicular Activities, and work on the Morpheus terrestrial free flyer test bed is reviewed. The back-up slides contain a chart with comparisons of LO2/LCH4 with other propellants, and reviewing the advantages especially for spacecraft propulsion.

Applewhite, John↗
NASA NTRS2011

Application of Solar-Electric Propulsion to Robotic and Human Missions in Near-Earth Space

Interest in applications of solar electric propulsion (SEP) is increasing. Application of SEP technology is favored when: (1) the mission is compatible with low-thrust propulsion, (2) the mission needs high total delta V such that chemical propulsion is disadvantaged; and (3) performance enhancement is needed. If all such opportunities for future missions are considered, many uses of SEP are likely. Representative missions are surveyed and several SEP applications selected for analysis, including orbit raising, lunar science, lunar exploration, lunar exploitation, planetary science, and planetary exploration. These missions span SEP power range from 10s of kWe to several MWe. Modes of use and benefits are described, and potential SEP evolution is discussed.

Woodcock, Gordon R.↗
NASA NTRS2011

Eagleworks Laboratories: Advanced Propulsion Physics Research

NASA/JSC is implementing an advanced propulsion physics laboratory, informally known as "Eagleworks", to pursue propulsion technologies necessary to enable human exploration of the solar system over the next 50 years, and enabling interstellar spaceflight by the end of the century. This work directly supports the "Breakthrough Propulsion" objectives detailed in the NASA OCT TA02 In-space Propulsion Roadmap, and aligns with the #10 Top Technical Challenge identified in the report. Since the work being pursued by this laboratory is applied scientific research in the areas of the quantum vacuum, gravitation, nature of space-time, and other fundamental physical phenomenon, high fidelity testing facilities are needed. The lab will first implement a low-thrust torsion pendulum (<1 uN), and commission the facility with an existing Quantum Vacuum Plasma Thruster. To date, the QVPT line of research has produced data suggesting very high specific impulse coupled with high specific force. If the physics and engineering models can be explored and understood in the lab to allow scaling to power levels pertinent for human spaceflight, 400kW SEP human missions to Mars may become a possibility, and at power levels of 2MW, 1-year transit to Neptune may also be possible. Additionally, the lab is implementing a warp field interferometer that will be able to measure spacetime disturbances down to 150nm. Recent work published by White [1] [2] [3] suggests that it may be possible to engineer spacetime creating conditions similar to what drives the expansion of the cosmos. Although the expected magnitude of the effect would be tiny, it may be a "Chicago pile" moment for this area of physics.

White, Harold↗
NASA NTRS2012

Mars Sample Return Using Solar Sail Propulsion

Many Mars Sample Return (MSR) architecture studies have been conducted over the years. A key element of them is the Earth Return Stage (ERS) whose objective is to obtain the sample from the Mars Ascent Vehicle (MAV) and return it safely to the surface of the Earth. ERS designs predominantly use chemical propulsion [1], incurring a significant launch mass penalty due to the low specific impulse of such systems coupled with the launch mass sensitivity to returned mass. It is proposed to use solar sail propulsion for the ERS, providing a high (effective) specific impulse propulsion system in the final stage of the multi-stage system. By doing so to the launch mass of the orbiter mission can be significantly reduced and hence potentially decreasing mission cost. Further, solar sailing offers a unique set of non-Keplerian low thrust trajectories that may enable modifications to the current approach to designing the Earth Entry Vehicle by potentially reducing the Earth arrival velocity. This modification will further decrease the mass of the orbiter system. Solar sail propulsion uses sunlight to propel vehicles through space by reflecting solar photons from a large, mirror-like surface made of a lightweight, reflective material. The continuous photonic pressure provides propellantless thrust to conduct orbital maneuvering and plane changes more efficiently than conventional chemical propulsion. Because the Sun supplies the necessary propulsive energy, solar sails require no onboard propellant, thus reducing system mass. This technology is currently at TRL 7/8 as demonstrated by the 2010 flight of the Japanese Aerospace Exploration Agency, JAXA, IKAROS mission. [2]

Johnson, Les↗
NASA NTRS2012

Using Gravity Assists in the Earth-moon System as a Gateway to the Solar System

For spacecraft departing the Earth - Moon system, lunar flybys can significantly increase the hype rbolic escape energy (C3, in km 2 /sec 2 ) for a modest increase in flight time. Within ~2 months, lunar flybys can produce a C3 of ~2. Over 4 - 6 months, lunar flybys alone can increase the C3 to ~4.5, or they can provide for additional periapsis burns to increase the C3 from ~2 -3 to 10 or more, suitable for planetary missions. A lunar flyby departure can be followed by additional ∆ -V (such as that efficiently provided by a low thrust system, eg. Solar Electric Propulsion (SEP)) to raise the Earth - relative velocity (at a ratio of more than 2:1) before a subsequent Earth flyby, which redirects that velocity to a more di stant target, all within not much more than a year. This paper describes the applicability of lunar flybys for different flight times and propulsi on systems, and illustrates this with instances of past usage and future possibilities. Examples discussed i nclude ISEE - 3, Nozomi, STEREO, 2018 Mars studies (which showed an 8% payload increase), and missions to Near Earth Objects (NEOs). In addition, the options for the achieving the initial lunar flyby are systematically discussed, with a view towards their p ractical use with in a compact launch period. In particular, we show that launches to geosynchronous transfer orbit (GTO) as a secondary payload provide a feasible means of obtaining a lunar flyby for an acceptable cost, even for SEP systems that cannot ea sily deliver large ∆ - Vs at periapsis. Taken together, these results comprise a myriad of options for increasing the mission performance, by the efficient use of lunar flybys within an acceptable extension of the flight time.

lunar fly-bys↗
NASA NTRS2012

Using Gravity Assists in the Earth-moon System as a Gateway to the Solar System

For spacecraft departing the Earth - Moon system, lunar flybys can significantly increase the hyperbolic escape energy (C3, in km (exp 2) /sec (exp 2) ) for a modest increase in flight time. Within approx 2 months, lunar flybys can produce a C3 of approx 2. Over 4 - 6 months, lunar flybys alone can increase the C3 to approx 4.5, or they can provide for additional periapsis burns to increase the C3 from approx 2 -3 to 10 or more, suitable for planetary missions. A lunar flyby departure can be followed by additional delta -V (such as that efficiently provided by a low thrust system, eg. Solar Electric Propulsion (SEP)) to raise the Earth - relative velocity (at a ratio of more than 2:1) before a subsequent Earth flyby, which redirects that velocity to a more distant target, all within not more than a year. This paper describes the applicability of lunar flybys for different flight times and propulsion systems, and illustrates this with instances of past usage and future possibilities. Examples discussed include ISEE-3, Nozomi, STEREO, 2018 Mars studies (which showed an 8% payload increase), and missions to Near Earth Objects (NEOs). In addition, the options for the achieving the initial lunar flyby are systematically discussed, with a view towards their practical use within a compact launch period. In particular, we show that launches to geosynchronous transfer orbit (GTO) as a secondary payload provide a feasible means of obtaining a lunar flyby for an acceptable cost, even for SEP systems that cannot easily deliver large delta-Vs at periapsis. Taken together, these results comprise a myriad of options for increasing the mission performance, by the efficient use of lunar flybys within an acceptable extension of the flight time.

Solar Electric Propulsion (SEP)↗
NASA NTRS2014

The Near Earth Object (NEO) Scout Spacecraft: A Low-cost Approach to In-situ Characterization of the NEO Population

This paper describes a microsatellite spacecraft with supporting mission profile and architecture, designed to enable preliminary in-situ characterization of a significant number of Near Earth Objects (NEOs) at reasonably low cost. The spacecraft will be referred to as the NEO-Scout. NEO-Scout spacecraft are to be placed in Geosynchronous Equatorial Orbit (GEO), cis-lunar space, or on earth escape trajectories as secondary payloads on launch vehicles headed for GEO or beyond, and will begin their mission after deployment from the launcher. A distinguishing key feature of the NEO-Scout system is to design the spacecraft and mission timeline so as to enable rendezvous with and landing on the target NEO during NEO close approach (<0.3 AU) to the Earth-Moon system using low-thrust/high-impulse propulsion systems. Mission durations are on the order 100 to 400 days. Mission feasibility and preliminary design analysis are presented, along with detailed trajectory calculations.

Woeppel, Eric A.↗
NASA NTRS2014

A High Power Solar Electric Propulsion - Chemical Mission for Human Exploration of Mars

Recently Solar Electric Propulsion (SEP) as a main propulsion system has been investigated as an option to support manned space missions to near-Earth destinations for the NASA Gateway spacecraft. High efficiency SEP systems are able to reduce the amount of propellant long duration chemical missions require, ultimately reducing the required mass delivered to Low Earth Orbit (LEO) by a launch vehicle. However, for long duration interplanetary Mars missions, using SEP as the sole propulsion source alone may not be feasible due to the long trip times to reach and insert into the destination orbit. By combining an SEP propulsion system with a chemical propulsion system the mission is able to utilize the high-efficiency SEP for sustained vehicle acceleration and deceleration in heliocentric space and the chemical system for orbit insertion maneuvers and trans-earth injection, eliminating the need for long duration spirals. By capturing chemically instead of with low-thrust SEP, Mars stay time increases by nearly 200 days. Additionally, the size the of chemical propulsion system can be significantly reduced from that of a standard Mars mission because the SEP system greatly decreases the Mars arrival and departure hyperbolic excess velocities (V(sub infinity)).

Solar Electric Propulsion↗
NASA NTRS2014

Interplanetary Trajectory Design for the Asteroid Robotic Redirect Mission Alternate Approach Trade Study

This paper presents mission performance analysis methods and results for the Asteroid Robotic Redirect Mission (ARRM) option to capture a free standing boulder on the surface of a 100 m or larger NEA. It details the optimization and design of heliocentric low-thrust trajectories to asteroid targets for the ARRM solar electric propulsion spacecraft. Extensive searches were conducted to determine asteroid targets with large pick-up mass potential and potential observation opportunities. Interplanetary trajectory approximations were developed in method based tools for Itokawa, Bennu, 1999 JU3, and 2008 EV5 and were validated by end-to-end integrated trajectories.

Merrill, Raymond Gabriel↗
NASA NTRS2013

Dawn Orbit Determination Team : Trajectory Modeling and Reconstruction Processes at Vesta

The NASA Dawn spacecraft was launched on September 27, 2007 on a mission to study the asteroid belt's two largest objects, Vesta and Ceres. It is the first deep space orbiting mission to demonstrate solar-electric ion propulsion, providing the necessary delta-V to enable capture and escape from two extraterrestrial bodies. At this time, Dawn has completed its science campaign at Vesta and is currently on its journey to Ceres, where it will arrive in mid-2015. The spacecraft spent over a year in orbit around Vesta from July 2011 through August 2012, capturing science data during four dedicated orbit phases. In order to maintain the reference orbits necessary for science and enable the transfers between those orbits, precise and timely orbit determination was required. The constraints associated with low-thrust ion propulsion coupled with the relatively unknown a priori gravity and rotation models for Vesta presented unique challenges for the Dawn orbit determination team. While [1] discusses the prediction performance of the orbit determination products, this paper discusses the dynamics models, filter configuration, and data processing implemented to deliver a rapid orbit determination capability to the Dawn project.

Vesta↗
NASA NTRS2011

Power and Propulsion System Design for Near-Earth Object Robotic Exploration

Near-Earth Objects (NEOs) are exciting targets for exploration; they are relatively easy to reach but relatively little is known about them. With solar electric propulsion, a vast number of interesting NEOs can be reached within a few years and with extensive flexibility in launch date. An additional advantage of electric propulsion for these missions is that a spacecraft can be small, enabling a fleet of explorers launched on a single vehicle or as secondary payloads. Commercial, flight-proven Hall thruster systems have great appeal based on their performance and low cost risk, but one issue with these systems is that the power processing units (PPUs) are designed for regulated spacecraft power architectures which are not attractive for small NEO missions. In this study we consider the integrated design of power and propulsion systems that utilize the capabilities of existing PPUs in an unregulated power architecture. Models for solar array and engine performance are combined with low-thrust trajectory analyses to bound spacecraft design parameters for a large class of NEO missions, then detailed array performance models are used to examine the array output voltage and current over a bounded mission set. Operational relationships between the power and electric propulsion systems are discussed, and it is shown that both the SPT-100 and BPT-4000 PPUs can perform missions over a solar range of 0.7 AU to 1.5 AU - encompassing NEOs, Venus, and Mars - within their operable input voltage ranges. A number of design trades to control the array voltage are available, including cell string layout, array offpointing during mission operations, and power draw by the Hall thruster system.

Near Earth Object (NEO)↗
NASA NTRS2009

Water Contaminant Mitigation in Ionic Liquid Propellant

Appropriate system and operational requirements are needed in order to ensure mission success without unnecessary cost. Purity requirements applied to thruster propellants may flow down to materials and operations as well as the propellant preparation itself. Colloid electrospray thrusters function by applying a large potential to a room temperature liquid propellant (such as an ionic liquid), inducing formation of a Taylor cone. Ions and droplets are ejected from the Taylor cone and accelerated through a strong electric field. Electrospray thrusters are highly efficient, precise, scaleable, and demonstrate low thrust noise. Ionic liquid propellants have excellent properties for use as electrospray propellants, but can be hampered by impurities, owing to their solvent capabilities. Of foremost concern is the water content, which can result from exposure to atmosphere. Even hydrophobic ionic liquids have been shown to absorb water from the air. In order to mitigate the risks of bubble formation in feed systems caused by water content of the ionic liquid propellant, physical properties of the ionic liquid EMI-Im are analyzed. The effects of surface tension, material wetting, physisorption, and geometric details of the flow manifold and electrospray emitters are explored. Results are compared to laboratory test data.

Conroy, David↗
NASA NTRS2015

Enhanced Gravity Tractor Technique for Planetary Defense

Given sufficient warning time, Earth-impacting asteroids and comets can be deflected with a variety of different "slow push/pull" techniques. The gravity tractor is one technique that uses the gravitational attraction of a rendezvous spacecraft to the impactor and a low-thrust, high-efficiency propulsion system to provide a gradual velocity change and alter its trajectory. An innovation to this technique, known as the Enhanced Gravity Tractor (EGT), uses mass collected in-situ to augment the mass of the spacecraft, thereby greatly increasing the gravitational force between the objects. The collected material can be a single boulder, multiple boulders, regolith or a combination of different sources. The collected mass would likely range from tens to hundreds of metric tons depending on the size of the impactor and warning time available. Depending on the propulsion system's capability and the mass collected, the EGT approach can reduce the deflection times by a factor of 10 to 50 or more, thus reducing the deflection times of several decades to years or less and overcoming the main criticism of the traditional gravity tractor approach. Additionally, multiple spacecraft can orbit the target in formation to provide the necessary velocity change and further reduce the time needed by the EGT technique to divert hazardous asteroids and comets. The robotic segment of NASA's Asteroid Redirect Mission (ARM) will collect a multi-ton boulder from the surface of a large Near-Earth Asteroid (NEA) and will provide the first ever demonstration of the EGT technique and validate one method of collecting in-situ mass on an asteroid of hazardous size.

Mazanek, Daniel D.↗