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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 775 records · Page 43

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NASA NTRS2007

Propellant Gauging for Exploration

This paper presents a brief overview of propellant gauging needs and requirements in the context of lunar exploration missions defined by the Exploration Systems Architecture Study (ESAS) report. A timeline for the development and testing of gauging technologies, and a few key design review dates are presented. A lunar exploration mission scenario is discussed which aids in defining the propellant gauging needs. The fleet of new exploration vehicles includes the Ares I and Ares V launch vehicles, Earth Departure Stage (EDS), Lunar Surface Access Module (LSAM) ascent and descent stages, and the Orion Crew Exploration Vehicle (CEV). The liquid propellant choices are currently oxygen - hydrogen for the launch vehicles, the EDS, and LSAM descent module; oxygen - methane for LSAM ascent module; and monomethylhydrazine nitrogen tetroxide (MMH-NTO) for the CEV. Estimated tank sizes, temperatures, pressures, and storage durations are presented. A baseline propellant gauging system is proposed that is based on high Technology Readiness Level (TRL) gauging technologies. In order to be considered for use on the new exploration vehicles, any new gauging technologies will have to show a clear benefit over the baseline methods in terms of performance and/or cost.

Zimmerli, Gregory A.↗

NASA NTRS1992

Propeller speed and phase sensor

A speed and phase sensor counterrotates aircraft propellers. A toothed wheel is attached to each propeller, and the teeth trigger a sensor as they pass, producing a sequence of signals. From the sequence of signals, rotational speed of each propeller is computer based on time intervals between successive signals. The speed can be computed several times during one revolution, thus giving speed information which is highly up-to-date. Given that spacing between teeth may not be uniform, the signals produced may be nonuniform in time. Error coefficients are derived to correct for nonuniformities in the resulting signals, thus allowing accurate speed to be computed despite the spacing nonuniformities. Phase can be viewed as the relative rotational position of one propeller with respect to the other, but measured at a fixed time. Phase is computed from the signals.

Collopy, Paul D.↗

NASA NTRS2004

Reusable module for the storage, transportation, and supply of multiple propellants in a space environment

A space module has an outer structure designed for traveling in space, a docking mechanism for facilitating a docking operation therewith in space, a first storage system storing a first propellant that burns as a result of a chemical reaction therein, a second storage system storing a second propellant that burns as a result of electrical energy being added thereto, and a bi-directional transfer interface coupled to each of the first and second storage systems to transfer the first and second propellants into and out thereof. The space module can be part of a propellant supply architecture that includes at least two of the space modules placed in an orbit in space.

Mazanek, Daniel D.↗

NASA NTRS1936

Relative Efficiencies and Design Charts for Various Engine-Propeller Combinations, Special Report

The relative efficiencies of various engine-propeller combinations were the subject of a study that covered the important flight conditions, particularly the take-off. Design charts that graphically correlate the various propeller parameters were prepared to facilitate the solution of problems and also to c1arify the conception of the relationships of the various engine-propeller design factors. It is shown that, among the many methods for improving the take-off thrust, the use of high-pitch, large-diameter controllable propellers turning at low rotational speeds is probably the most generally promising. With such a combination the take-off thrust may be further increased, at the expense of a small loss in cruising efficiency, by compromise designs wherein the pitch setting is slightly reduced and the diameter is further increased. The degree of compromise necessary to accomplish the maximum possible take-off improvement depends on such design factors as overspeeding and overboosting at take-off as well as depending on the design altitude. Both overspeeding and designing for altitude operation have the same effect on the take-off thrust as compromising in that the propulsive efficiency is increased thereby; boosting the engine, however, has the reverse effect on the propulsive efficiency, although the brake horsepower is increased.

Biermann, David↗

NASA NTRS2009

Parametric Optimization of Ares I Propellant Slosh Characteristics Using Frequency Response Criteria

A novel technique for developing propellant slosh damping requirements with respect to the stability characteristics of large flexible launch vehicles is presented. A numerical algorithm is devised which allows an automated software program to rapidly converge to pseudo-optimal solutions that minimize required propellant slosh damping for multiple tanks while maintaining constraints on the frequency response characteristics of a particular open-loop plant transfer function. An implementation of the algorithm using a high-order linear model of the Ares I plant dynamics considers all relevant dynamic interactions of flexible body modes, propellant slosh, and nozzle inertia effects. A high-resolution propellant damping requirements table is produced that can be used for baffle design. The method is demonstrated to provide exceptional speed and accuracy when compared with the alternative human-in-the-loop approach.

Orr, Jeb S.↗

NASA NTRS2006

Low-Cost Propellant Launch From a Tethered Balloon

A document presents a concept for relatively inexpensive delivery of propellant to a large fuel depot in low orbit around the Earth, for use in rockets destined for higher orbits, the Moon, and for remote planets. The propellant is expected to be at least 85 percent of the mass needed in low Earth orbit to support the NASA Exploration Vision. The concept calls for the use of many small ( 10 ton) spin-stabilized, multistage, solid-fuel rockets to each deliver 250 kg of propellant. Each rocket would be winched up to a balloon tethered above most of the atmospheric mass (optimal altitude 26 2 km). There, the rocket would be aimed slightly above the horizon, spun, dropped, and fired at a time chosen so that the rocket would arrive in orbit near the depot. Small thrusters on the payload (powered, for example, by boil-off gases from cryogenic propellants that make up the payload) would precess the spinning rocket, using data from a low-cost inertial sensor to correct for small aerodynamic and solid rocket nozzle misalignment torques on the spinning rocket; would manage the angle of attack and the final orbit insertion burn; and would be fired on command from the depot in response to observations of the trajectory of the payload so as to make small corrections to bring the payload into a rendezvous orbit and despin it for capture by the depot. The system is low-cost because the small rockets can be mass-produced using the same techniques as those to produce automobiles and low-cost munitions, and one or more can be launched from a U.S. territory on the equator (Baker or Jarvis Islands in the mid-Pacific) to the fuel depot on each orbit (every 90 minutes, e.g., any multiple of 6,000 per year).

Wilcox, Brian↗

NASA NTRS2010

Combustion Stability Analyses of Coaxial Element Injectors with Liquid Oxygen/Liquid Methane Propellants

Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for in-space vehicles. This propellant combination has not been previously used in a flight-qualified engine system, so limited test data and analysis results are available at this stage of early development. NASA has funded several hardware-oriented activities with oxygen and methane propellants over the past several years with the Propulsion and Cryogenic Advanced Development (PCAD) project, under the Exploration Technology Development Program. As part of this effort, the NASA Marshall Space Flight Center has conducted combustion stability analyses of several of the configurations. This paper presents test data and analyses of combustion stability from the recent PCAD-funded test programs at the NASA MSFC. These test programs used swirl coaxial element injectors with liquid oxygen and liquid methane propellants. Oxygen was injected conventionally in the center of the coaxial element, and swirl was provided by tangential entry slots. Injectors with 28-element and 40-element patterns were tested with several configurations of combustion chambers, including ablative and calorimeter spool sections, and several configurations of fuel injection design. Low frequency combustion instability (chug) occurred with both injectors, and high-frequency combustion instability occurred at the first tangential (1T) transverse mode with the 40-element injector. In most tests, a transition between high-amplitude chug with gaseous methane flow and low-amplitude chug with liquid methane flow was readily observed. Chug analyses of both conditions were conducted using techniques from Wenzel and Szuch and from the Rocket Combustor Interactive Design and Analysis (ROCCID) code. The 1T mode instability occurred in several tests and was apparent by high-frequency pressure measurements as well as dramatic increases in calorimeter-measured heat flux throughout the chamber. Analyses of the transverse mode were conducted with ROCCID and empirical methods such as Hewitt d/V. This paper describes the test hardware configurations, test data, analysis methods, and presents results of the various analyses.

Hulka, J. R.↗

NASA NTRS2010

Fuel-Cell Power Source Based on Onboard Rocket Propellants

The use of onboard rocket propellants (dense liquids at room temperature) in place of conventional cryogenic fuel-cell reactants (hydrogen and oxygen) eliminates the mass penalties associated with cryocooling and boil-off. The high energy content and density of the rocket propellants will also require no additional chemical processing. For a 30-day mission on the Moon that requires a continuous 100 watts of power, the reactant mass and volume would be reduced by 15 and 50 percent, respectively, even without accounting for boiloff losses. The savings increase further with increasing transit times. A high-temperature, solid oxide, electrolyte-based fuel-cell configuration, that can rapidly combine rocket propellants - both monopropellant system with hydrazine and bi-propellant systems such as monomethyl hydrazine/ unsymmetrical dimethyl hydrazine (MMH/UDMH) and nitrogen tetroxide (NTO) to produce electrical energy - overcomes the severe drawbacks of earlier attempts in 1963-1967 of using fuel reforming and aqueous media. The electrical energy available from such a fuel cell operating at 60-percent efficiency is estimated to be 1,500 Wh/kg of reactants. The proposed use of zirconia-based oxide electrolyte at 800-1,000 C will permit continuous operation, very high power densities, and substantially increased efficiency of conversion over any of the earlier attempts. The solid oxide fuel cell is also tolerant to a wide range of environmental temperatures. Such a system is built for easy refueling for exploration missions and for the ability to turn on after several years of transit. Specific examples of future missions are in-situ landers on Europa and Titan that will face extreme radiation and temperature environments, flyby missions to Saturn, and landed missions on the Moon with 14 day/night cycles.

Ganapathi, Gani↗

NASA NTRS2011

Cryogenic Propellant Feed System Analytical Tool Development

The Propulsion Systems Branch at NASA s Lyndon B. Johnson Space Center (JSC) has developed a parametric analytical tool to address the need to rapidly predict heat leak into propellant distribution lines based on insulation type, installation technique, line supports, penetrations, and instrumentation. The Propellant Feed System Analytical Tool (PFSAT) will also determine the optimum orifice diameter for an optional thermodynamic vent system (TVS) to counteract heat leak into the feed line and ensure temperature constraints at the end of the feed line are met. PFSAT was developed primarily using Fortran 90 code because of its number crunching power and the capability to directly access real fluid property subroutines in the Reference Fluid Thermodynamic and Transport Properties (REFPROP) Database developed by NIST. A Microsoft Excel front end user interface was implemented to provide convenient portability of PFSAT among a wide variety of potential users and its ability to utilize a user-friendly graphical user interface (GUI) developed in Visual Basic for Applications (VBA). The focus of PFSAT is on-orbit reaction control systems and orbital maneuvering systems, but it may be used to predict heat leak into ground-based transfer lines as well. PFSAT is expected to be used for rapid initial design of cryogenic propellant distribution lines and thermodynamic vent systems. Once validated, PFSAT will support concept trades for a variety of cryogenic fluid transfer systems on spacecraft, including planetary landers, transfer vehicles, and propellant depots, as well as surface-based transfer systems. The details of the development of PFSAT, its user interface, and the program structure will be presented.

Lusby, Brian S.↗

NASA NTRS2012

Thermal Optimization and Assessment of a Long Duration Cryogenic Propellant Depot

A Cryogenic Propellant Depot (CPD) operating in Low Earth Orbit (LEO) could provide many near term benefits to NASA space exploration efforts. These benefits include elongation/extension of spacecraft missions and reduction of launch vehicle up-mass requirements. Some of the challenges include controlling cryogenic propellant evaporation and managing the high costs and long schedules associated with new spacecraft hardware development. This paper describes a conceptual CPD design that is thermally optimized to achieve extremely low propellant boil-off rates. The CPD design is based on existing launch vehicle architecture, and its thermal optimization is achieved using current passive thermal control technology. Results from an integrated thermal model are presented showing that this conceptual CPD design can achieve propellant boil-off rates well under 0.05% per day, even when subjected to the LEO thermal environment.

Honour, Ryan↗

NASA NTRS2011

Sloshing in the Liquid Hydrogen and Liquid Oxygen Propellant Tanks After Main Engine Cut Off

NASA Marshall Space Flight Center is designing and developing the Main Propulsion System (MPS) for Ares launch vehicles. Propellant sloshing in the liquid hydrogen (LH2) and liquid oxygen (LO2) propellant tanks after Main Engine Cut Off (MECO) was modeled using the Volume of Fluid (VOF) module of the computational fluid dynamics code, CFD-ACE+. The present simulation shows that there is substantial sloshing side forces acting on the LH2 tank during the deceleration of the vehicle after MECO. The LH2 tank features a side wall drain pipe. The side loads result from the residual propellant mass motion in the LH2 tank which is initiated by the stop of flow into the drain pipe at MECO. The simulations show that radial force on the LH2 tank wall is less than 50 lbf and the radial moment calculated based up through the center of gravity of the vehicle is predicted to be as high as 300 lbf-ft. The LO2 tank features a bottom dome drain system and is equipped with sloshing baffles. The remaining LO2 in the tank slowly forms a liquid column along the centerline of tank under the zero gravity environments. The radial force on the LO2 tank wall is predicted to be less than 100 lbf. The radial moment calculated based on the center of gravity of the vehicle is predicted as high as 4500 lbf-ft just before MECO and dropped down to near zero after propellant draining stopped completely.

Kim, Sura↗

NASA NTRS2012

Thermal Optimization of an On-Orbit Long Duration Cryogenic Propellant Depot

A Cryogenic Propellant Depot (CPD) operating in Low Earth Orbit (LEO) could provide many near term benefits to NASA's space exploration efforts. These benefits include elongation/extension of spacecraft missions and requirement reduction of launch vehicle up-mass. Some of the challenges include controlling cryogenic propellant evaporation and managing the high costs and long schedules associated with the new development of spacecraft hardware. This paper describes a conceptual CPD design that is thermally optimized to achieve extremely low propellant boil-off rates. The CPD design is based on existing launch vehicle architecture, and its thermal optimization is achieved using current passive thermal control technology. Results from an integrated thermal model are presented showing that this conceptual CPD design can achieve propellant boil-off rates well under 0.05% per day, even when subjected to the LEO thermal environment.

Honour, Ryan↗

NASA NTRS2012

The PROPEL Electrodynamic Tether Mission and Connecting to the Ionosphere

The exponential increase of launch system size.and cost.with delta-V makes missions that require large total impulse cost prohibitive. Led by NASA's Marshall Space Flight Center, a team from government, industry, and academia has developed a flight demonstration mission concept of an integrated electrodynamic (ED) tethered satellite system called PROPEL: "Propulsion using Electrodynamics". The PROPEL Mission is focused on demonstrating a versatile configuration of an ED tether to overcome the limitations of the rocket equation, enable new classes of missions currently unaffordable or infeasible, and significantly advance the Technology Readiness Level (TRL) to an operational level. We are also focused on establishing a far deeper understanding of critical processes and technologies to be able to scale and improve tether systems in the future. Here, we provide an overview of the proposed PROPEL mission. One of the critical processes for efficient ED tether operation is the ability to inject current to and collect current from the ionosphere. Because the PROPEL mission is planned to have both boost and deboost capability using a single tether, the tether current must be capable of flowing in both directions and at levels well over 1 A. Given the greater mobility of electrons over that of ions, this generally requires that both ends of the ED tether system can both collect and emit electrons. For example, hollow cathode plasma contactors (HCPCs) generally are viewed as state-of-the-art and high TRL devices; however, for ED tether applications important questions remain of how efficiently they can operate as both electron collectors and emitters. Other technologies will be highlighted that are being investigated as possible alternatives to the HCPC such as Solex that generates a plasma cloud from a solid material (Teflon) and electron emission (only) technologies such as cold-cathode electron field emission or photo-electron beam generation (PEBG) techniques.

Gilchrist, Brian↗

NASA NTRS2012

The PROPEL Electrodynamic Tether Demonstration Mission

The PROPEL ("Propulsion using Electrodynamics") mission will demonstrate the operation of an electrodynamic tether propulsion system in low Earth orbit and advance its technology readiness level for multiple applications. The PROPEL mission has two primary objectives: first, to demonstrate the capability of electrodynamic tether technology to provide robust and safe, near-propellantless propulsion for orbit-raising, de-orbit, plane change, and station keeping, as well as to perform orbital power harvesting and formation flight; and, second, to fully characterize and validate the performance of an integrated electrodynamic tether propulsion system, qualifying it for infusion into future multiple satellite platforms and missions with minimal modification. This paper provides an overview of the PROPEL system and design reference missions; mission goals and required measurements; and ongoing PROPEL mission design efforts.

Bilen, Sven G.↗

NASA NTRS2011

High Power Flex-Propellant Arcjet Performance

A MW-class electrothermal arcjet based on a water-cooled, wall-stabilized, constricted arc discharge configuration was subjected to extensive performance testing using hydrogen and simulated ammonia propellants with the deliberate aim of advancing technology readiness level for potential space propulsion applications. The breadboard design incorporates alternating conductor/insulator wafers to form a discharge barrel enclosure with a 2.5-cm internal bore diameter and an overall length of approximately 1 meter. Swirling propellant flow is introduced into the barrel, and a DC arc discharge mode is established between a backplate tungsten cathode button and a downstream ringanode/ spin-coil assembly. The arc-heated propellant then enters a short mixing plenum and is accelerated through a converging-diverging graphite nozzle. This innovative design configuration differs substantially from conventional arcjet thrusters, in which the throat functions as constrictor and the expansion nozzle serves as the anode, and permits the attainment of an equilibrium sonic throat (EST) condition. During the test program, applied electrical input power was varied between 0.5-1 MW with hydrogen and simulated ammonia flow rates in the range of 4-12 g/s and 15-35 g/s, respectively. The ranges of investigated specific input energy therefore fell between 50-250 MJ/kg for hydrogen and 10-60 MJ/kg for ammonia. In both cases, observed arc efficiencies were between 40-60 percent as determined via a simple heat balance method based on electrical input power and coolant water calorimeter measurements. These experimental results were found to be in excellent agreement with theoretical chemical equilibrium predictions, thereby validating the EST assumption and enabling the utilization of standard TDK nozzle expansion analyses to reliably infer baseline thruster performance characteristics. Inferred specific impulse performance accounting for recombination kinetics during the expansion process implied nearly frozen flow in the nozzle and yielded performance ranges of 800-1100 sec for hydrogen and 400-600 sec for ammonia. Inferred thrust-to-power ratios were in the range of 30-10 lbf/MWe for hydrogen and 60-20 lbf/MWe for ammonia. Successful completion of this test series represents a fundamental milestone in the progression of high power arcjet technology, and it is hoped that the results may serve as a reliable touchstone for the future development of MW-class regeneratively-cooled flex-propellant plasma rockets.

Litchford, Ron J.↗

NASA NTRS2012

NASA's Evolutionary Xenon Thruster (NEXT) Long-Duration Test as of 736 kg of Propellant Throughput

The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation solar-electric ion propulsion system with significant enhancements beyond the state-of-the-art NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) ion propulsion system to provide future NASA science missions with enhanced mission capabilities. A Long-Duration Test (LDT) was initiated in June 2005 to validate the thruster service life modeling and to qualify the thruster propellant throughput capability. The thruster has set electric propulsion records for the longest operating duration, highest propellant throughput, and most total impulse demonstrated. At the time of this publication, the NEXT LDT has surpassed 42,100 h of operation, processed more than 736 kg of xenon propellant, and demonstrated greater than 28.1 MN s total impulse. Thruster performance has been steady with negligible degradation. The NEXT thruster design has mitigated several lifetime limiting mechanisms encountered in the NSTAR design, including the NSTAR first failure mode, thereby drastically improving thruster capabilities. Component erosion rates and the progression of the predicted life-limiting erosion mechanism for the thruster compare favorably to pretest predictions based upon semi-empirical ion thruster models used in the thruster service life assessment. Service life model validation has been accomplished by the NEXT LDT. Assuming full-power operation until test article failure, the models and extrapolated erosion data predict penetration of the accelerator grid grooves after more than 45,000 hours of operation while processing over 800 kg of xenon propellant. Thruster failure due to degradation of the accelerator grid structural integrity is expected after

Shastry, Rohit↗

NASA NTRS2012

NASA's Evolutionary Xenon Thruster (NEXT) Long-Duration Test as of 736 kg of Propellant Throughput

Shastry, Rohit↗

NASA NTRS2013

Optimal Propellant Maneuver Flight Demonstrations on ISS

In this paper, first ever flight demonstrations of Optimal Propellant Maneuver (OPM), a method of propulsive rotational state transition for spacecraft controlled using thrusters, is presented for the International Space Station (ISS). On August 1, 2012, two ISS reorientations of about 180deg each were performed using OPMs. These maneuvers were in preparation for the same-day launch and rendezvous of a Progress vehicle, also a first for ISS visiting vehicles. The first maneuver used 9.7 kg of propellant, whereas the second used 10.2 kg. Identical maneuvers performed without using OPMs would have used approximately 151.1kg and 150.9kg respectively. The OPM method is to use a pre-planned attitude command trajectory to accomplish a rotational state transition. The trajectory is designed to take advantage of the complete nonlinear system dynamics. The trajectory choice directly influences the cost of the maneuver, in this case, propellant. For example, while an eigenaxis maneuver is kinematically the shortest path between two orientations, following that path requires overcoming the nonlinear system dynamics, thereby increasing the cost of the maneuver. The eigenaxis path is used for ISS maneuvers using thrusters. By considering a longer angular path, the path dependence of the system dynamics can be exploited to reduce the cost. The benefits of OPM for the ISS include not only reduced lifetime propellant use, but also reduced loads, erosion, and contamination from thrusters due to fewer firings. Another advantage of the OPM is that it does not require ISS flight software modifications since it is a set of commands tailored to the specific attitude control architecture. The OPM takes advantage of the existing ISS control system architecture for propulsive rotation called USTO control mode1. USTO was originally developed to provide ISS Orbiter stack attitude control capability for a contingency tile-repair scenario, where the Orbiter is maneuvered using its robotic manipulator relative to the ISS. Since 2005 USTO has been used for nominal ISS operations.

Bhatt, Sagar↗