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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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS2003

Isothermal Calorimetric Observations of the Effect of Welding on Compatibility of Stainless Steels with High-Test Hydrogen Peroxide Propellant

High-Test Hydrogen Peroxide (HTP) is receiving renewed interest as a monopropellant and as the oxidizer for bipropellant systems. HTP is hydrogen peroxide having concentrations ranging from 70 to 98%. In these applications the energy and oxygen released during decomposition of HTP is used for propulsion. In propulsion systems components must be fabricated and connected using available joining processes. Welding is a common joining method for metallic components. The goal of this study was to compare the HTP compatibility of welded vs. unwelded stainless steel.

Gostowski, Rudy↗
NASA NTRS2004

In-Space Chemical Propulsion System Model

Multiple, new technologies for chemical systems are becoming available and include high temperature rockets, very light propellant tanks and structures, new bipropellant and monopropellant options, lower mass propellant control components, and zero boil off subsystems. Such technologies offer promise of increasing the performance of in-space chemical propulsion for energetic space missions. A mass model for pressure-fed, Earth and space-storable, advanced chemical propulsion systems (ACPS) was developed in support of the NASA MSFC In-Space Propulsion Program. Data from flight systems and studies defined baseline system architectures and subsystems and analyses were formulated for parametric scaling relationships for all ACPS subsystems. The paper will first provide summary descriptions of the approaches used for the systems and the subsystems and then present selected analyses to illustrate use of the model for missions with characteristics of current interest.

Byers, David C.↗
NASA NTRS2004

Advanced Chemical Propulsion Study

A study was performed of advanced chemical propulsion technology application to space science (Code S) missions. The purpose was to begin the process of selecting chemical propulsion technology advancement activities that would provide greatest benefits to Code S missions. Several missions were selected from Code S planning data, and a range of advanced chemical propulsion options was analyzed to assess capabilities and benefits re these missions. Selected beneficial applications were found for higher-performing bipropellants, gelled propellants, and cryogenic propellants. Technology advancement recommendations included cryocoolers and small turbopump engines for cryogenic propellants; space storable propellants such as LOX-hydrazine; and advanced monopropellants. It was noted that fluorine-bearing oxidizers offer performance gains over more benign oxidizers. Potential benefits were observed for gelled propellants that could be allowed to freeze, then thawed for use.

Woodcock, Gordon↗
NASA NTRS2004

In-Space Chemical Propulsion System Model

Multiple, new technologies for chemical systems are becoming available and include high temperature rockets, very light propellant tanks and structures, new bipropellant and monopropellant options, lower mass propellant control components, and zero boil off subsystems. Such technologies offer promise of increasing the performance of in-space chemical propulsion for energetic space missions. A mass model for pressure-fed, Earth and space-storable, advanced chemical propulsion systems (ACPS) was developed in support of the NASA MSFC In-Space Propulsion Program. Data from flight systems and studies defined baseline system architectures and subsystems and analyses were formulated for parametric scaling relationships for all ACPS subsystem. The paper will first provide summary descriptions of the approaches used for the systems and the subsystems and then present selected analyses to illustrate use of the model for missions with characteristics of current interest.

Byers, David C.↗
NASA NTRS2004

Advanced Chemical Propulsion Study

A study was performed of advanced chemical propulsion technology application to space science (Code S) missions. The purpose was to begin the process of selecting chemical propulsion technology advancement activities that would provide greatest benefits to Code S missions. Several missions were selected from Code S planning data, and a range of advanced chemical propulsion options was analyzed to assess capabilities and benefits re these missions. Selected beneficial applications were found for higher-performing bipropellants, gelled propellants, and cryogenic propellants. Technology advancement recommendations included cryocoolers and small turbopump engines for cryogenic propellants; space storable propellants such as LOX-hydrazine; and advanced monopropellants. It was noted that fluorine-bearing oxidizers offer performance gains over more benign oxidizers. Potential benefits were observed for gelled propellants that could be allowed to freeze, then thawed for use.

Woodcock, Gordon↗
NASA NTRS1951

Combustion Instability in an Acid-Heptane Rocket with a Pressurized-Gas Propellant Pumping System

Results of experimental measurements of low-frequency combustion instability of a 300-pound thrust acid-heptane rocket engine were compared to the trends predicted by an analysis of combustion instability in a rocket engine with a pressurized-gas propellant pumping system. The simplified analysis, which assumes a monopropellant model, was based on the concept of a combustion the delay occurring from the moment of propellant injection to the moment of propellant combustion. This combustion time delay was experimentally measured; the experimental values were of approximately half the magnitude predicted by the analysis. The pressure-fluctuation frequency for a rocket engine with a characteristic length of 100 inches and operated at a combustion-chamber pressure of 280 pounds per square inch absolute was 38 cycles per second; the analysis indicated. a frequency of 37 cycles per second. Increasing combustion-chamber characteristic length decreased the pressure-fluctuation frequency, in conformity to the analysis. Increasing the chamber operating pressure or increasing the injector pressure drop increased the frequency. These latter two effects are contrary to the analysis; the discrepancies are attributed to the conflict between the assumptions made to simplify the analysis and the experimental conditions. Oxidant-fuel ratio had no apparent effect on the experimentally measured pressure-fluctuation frequency for acid-heptane ratios from 3.0 to 7.0. The frequencies decreased with increased amplitude of the combustion-chamber pressure variations. The analysis indicated that if the combustion time delay were sufficiently short, low-frequency combustion instability would be eliminated.

Tischler, Adelbert O.↗
NASA NTRS2002

Initial Cassini propulsion system in-flight characterization

The Cassini Propulsion Module Subsystem has performed excellently throughout the first four years of mission operations. The PMS is the most complex interplanetary propulsion subsystem ever flown, with separate monopropellant and bipropellant propulsion modules, each replete with many redundant components.

Cassini propulsion performance mission operations ↗
NASA NTRS2003

Minimum impulse thruster valve design and development

The design and development of a minimum impulse thruster valve was conducted, by Moog, under contract by NASA's Jet Propulsion Laboratory, California Institute of Technology, for deep space propulsion systems. The effort was focused on applying known solenoid design techniques scaled to provide a 1 -millisecond response capability for monopropellant, hydrazine ACS thruster applications. The valve has an extended operating temperature range of 20(deg)F to +350(deg)F with a total mass of less than 25 grams and nominal power draw of 7 watts. The design solution resulted in providing a solenoid valve that is one-tenth the scale of the standard product line. The valve has the capability of providing a mass flow rate of 0.0009 pounds per second hydrazine. The design life of 1,000,000 cycles was demonstrated both dry and wet. Not all design factors scaled as expected and proved to be the focus of the final development effort. These included the surface interactions, hydrodynamics and driver electronics. The resulting solution applied matured design approaches to minimize the program risk with innovative methods to address the impacts of scale.

impulse thruster valve↗
NASA NTRS2005

The Minimum Impulse Thruster

The Minimum Impulse Thruster (MIT) was developed to improve the state-of-the-art minimum impulse capability of hydrazine monopropellant thrusters. Specifically, a new fast response solenoid valve was developed, capable of responding to a much shorter electrical pulse width, thereby reducing the propellant flow time and the minimum impulse bit. The new valve was combined with the Aerojet MR-103, 0.2 lbf (0.9 N) thruster and put through an extensive Delta-qualification test program, resulting in a factor of 5 reduction in the minimum impulse bit, from roughly 1.1 milli-lbf-seconds (5 milliNewton seconds) to - 0.22 milli-lbf-seconds (1 mN-s). To maintain it's extensive heritage, the thruster itself was left unchanged. The Minimum Impulse Thruster provides mission and spacecraft designers new design options for precision pointing and precision translation of spacecraft.

thruster↗
NASA NTRS2006

Determination of Silicon in Hydrazine

Inductively coupled plasma-mass spectrometry (ICP-MS) is a highly sensitive technique sometimes used for the trace determination of silicon at a mass-to-charge (m/z) ratio of 28, the most abundant natural isotope of silicon. Unfortunately, ICP-MS is unable to differentiate between other sources of m/z 28 and false positive results for silicon will result when other sources of m/z 28 are present. Nitrogen was a major source of m/z 28 and contributes to the m/z 28 signal when hydrazine sample or nitric acid preservative is introduced into the plasma. Accordingly, this work was performed to develop a sample preparation step coupled with an ICP-MS analysis that minimized non-silicon sources of m/z 28. In the preparatory step of this method, the hydrazine sample was first decomposed predominately to nitrogen gas and water with copper-catalyzed hydrogen peroxide. In the analysis step, ICP-MS was used without nitric acid preservative in samples or standards. Glass, a potential source of silicon contamination, was also avoided where possible. The method was sensitive, accurate, and reliable for the determination of silicon in monopropellant grade hydrazine (MPH) in AF-E-332 elastomer leaching tests. Results for silicon in MPH were comparable to those reported in the literature for other studies.

McClure, Mark B.↗
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 NTRS2010

Predicting Ares I Reaction Control System Performance by Utilizing Analysis Anchored with Development Test Data

The Ares I launch vehicle is an integral part of NASA s Constellation Program, providing a foundation for a new era of space access. The Ares I is designed to lift the Orion Crew Module and will enable humans to return to the Moon as well as explore Mars.1 The Ares I is comprised of two inline stages: a Space Shuttle-derived five-segment Solid Rocket Booster (SRB) First Stage (FS) and an Upper Stage (US) powered by a Saturn V-derived J-2X engine. A dedicated Roll Control System (RoCS) located on the connecting interstage provides roll control prior to FS separation. Induced yaw and pitch moments are handled by the SRB nozzle vectoring. The FS SRB operates for approximately two minutes after which the US separates from the vehicle and the US Reaction Control System (ReCS) continues to provide reaction control for the remainder of the mission. A representation of the Ares I launch vehicle in the stacked configuration and including the Orion Crew Exploration Vehicle (CEV) is shown in Figure 1. Each Reaction Control System (RCS) design incorporates a Gaseous Helium (GHe) pressurization system combined with a monopropellant Hydrazine (N2H4) propulsion system. Both systems have two diametrically opposed thruster modules. This architecture provides one failure tolerance for function and prevention of catastrophic hazards such as inadvertent thruster firing, bulk propellant leakage, and over-pressurization. The pressurization system on the RoCS includes two ambient pressure-referenced regulators on parallel strings in order to attain the required system level single Fault Tolerant (FT) design for function while the ReCS utilizes a blow-down approach. A single burst disk and relief valve assembly is also included on the RoCS to ensure single failure tolerance for must-not-occur catastrophic hazards. The Reaction Control Systems are designed to support simultaneously firing multiple thrusters as required

Stein, William B.↗
NASA NTRS2009

Advanced Development Program for a 625 lbf thrust engine for Ares First Stage Roll Control System

NASA's new Ares Launch Vehicle will require twelve thrusters to provide roll control of the vehicle during the first stage firing. All twelve roll control thrusters will be located at the inter-stage segment that separates the solid rocket booster first stage from the second stage. NASA selected a mono propellant hydrazine solution and as a result awarded Aerojet-General a contract in 2007 for an advanced development program for an MR-80- series 625 Ibf vacuum thrust monopropellant hydrazine thruster. This thruster has heritage dating back to the 1976 Viking Landers and most recently for the 2011 Mars Science Laboratory. Prior to the Ares application, the MR-80-series thrusters had been equipped with throttle valves and not typically operated in pulse mode. The primary objective of the advanced development program was to increase the technology readiness level and retire major technical risks for the future flight qualification test program. Aerojet built on their heritage MR-80 rocket engine designs to achieve the design and performance requirements. Significant improvements to cost and lead-time were achieved by applying Design for Manufacturing and Assembly (DFMA) principles. AerojetGeneral has completed Preliminary and Critical Design Reviews, followed by two successful rocket engine development test programs. The test programs included qualification random vibration and firing lite that significantly exceed the flight qualification requirements. This paper discusses the advanced development program and the demonstrated capability of the MR-80C engine. Y;

Dawson, Matt↗
NASA NTRS2010

NASA Ares I Launch Vehicle Upper Stage Reaction Control System (ReCS) Cold Flow Development Test Overview

NASA s Ares I launch vehicle, consisting of a five segment solid rocket booster first stage and a liquid bi-propellant J2-X engine Upper Stage, is the vehicle that s been chosen to launch the Orion Crew Module, which will return humans to the Moon, Mars, and beyond. After First Stage booster separation, the Reaction Control System (ReCS), a monopropellant hydrazine system, will provide the Upper Stage element with three degrees of freedom control as needed. This paper provides an overview of the system level development testing that has taken place on the Ares I launch vehicle Upper Stage ReCS. The ReCS System Development Test Article (SDTA) was built as a flight representative water flow test article whose primary test objective was to obtain fluid system performance data to evaluate the integrate system performance characteristics and verify analytical models. Water is the industry standard for cold flow testing of hydrazine systems, because the densities are very close and the speeds of sound are well characterized. The completion of this development level test program was considered necessary to support the ReCS Critical Design Review. This paper will address the design approach taken in building the test article, the objectives of the test program, types of testing completed, general results, the ability of the program to meet the test objectives, and lessons learned

Dervan, Melanie↗
NASA NTRS2010

Propellant Readiness Level: A Methodological Approach to Propellant Characterization

A methodological approach to defining propellant characterization is presented. The method is based on the well-established Technology Readiness Level nomenclature. This approach establishes the Propellant Readiness Level as a metric for ascertaining the readiness of a propellant or a propellant combination by evaluating the following set of propellant characteristics: thermodynamic data, toxicity, applications, combustion data, heat transfer data, material compatibility, analytical prediction modeling, injector/chamber geometry, pressurization, ignition, combustion stability, system storability, qualification testing, and flight capability. The methodology is meant to be applicable to all propellants or propellant combinations; liquid, solid, and gaseous propellants as well as monopropellants and propellant combinations are equally served. The functionality of the proposed approach is tested through the evaluation and comparison of an example set of hydrocarbon fuels.

Bossard, John A.↗
NASA NTRS2010

Survey of Propulsion Technologies Applicable to Cubesats

At present, no Cubesat has flown in space featuring propulsion. This was acceptable as long as CubeSats were flown mostly as university experiments. As CubeSats become of interest to other users in the government and industry communities as well, a larger range of capabilities may be required than exhibited so far, while maintaining the uniqueness of the Cubesat platform. Propulsion capability is crucial in increasing mission capabilities of future CubeSats, such as orbit change and raising, formation flying, proximity operations, fine attitude control, or drag-make-up and de-orbit. While some of these tasks may be accomplished with propellantless devices, their applications are limited, applicable mostly to a single task, and bear their own risks. In this study, a survey was conducted of propulsion technologies applicable to CubeSats. Only few off-the-shelf design solutions exist today. The survey was thus expanded to such devices as well that are under significant development, and are approaching the required design envelope for CubeSats with respect to mass, volume, and power. In some cases, such as electric propulsion devices, CubeSat architectures themselves may need to be adapted, required to feature deployable solar arrays to increase power capabilities. Given the vast scope of this survey, only thruster technologies could be surveyed. However, valves and other feed system components, as well as their integration, are equally important, but have to be left to a future survey. Three major propulsion technology areas applicable to CubeSats emerged when conducting this review: (1) Existing technologies, such as butane systems, pulsed plasma thrusters, and vacuum arc thrusters are applicable to CubeSats today with no or only minor changes, (2) New thruster technologies under significant development, such as hydrazine monopropellant systems, ion engines, or colloid thrusters could be adapted to CubeSats with some further development, especially also in other subsystem areas such as feed systems and power processing units. They will also require increased power capabilities, and (3) emerging technologies, such as micro electrospray arrays and micro cavity discharge arrays that offer even higher flexibility due to scalability for CubeSats, and enable compact integration.

micropropulsion↗
NASA NTRS2010

Hydrogen peroxide catalytic decomposition

Nitric oxide in a gaseous stream is converted to nitrogen dioxide using oxidizing species generated through the use of concentrated hydrogen peroxide fed as a monopropellant into a catalyzed thruster assembly. The hydrogen peroxide is preferably stored at stable concentration levels, i.e., approximately 50%-70% by volume, and may be increased in concentration in a continuous process preceding decomposition in the thruster assembly. The exhaust of the thruster assembly, rich in hydroxyl and/or hydroperoxy radicals, may be fed into a stream containing oxidizable components, such as nitric oxide, to facilitate their oxidation.

Parrish, Clyde F.↗
NASA NTRS2003

Oxide Protective Coats for Ir/Re Rocket Combustion Chambers

An improved material system has been developed for rocket engine combustion chambers for burning oxygen/ hydrogen mixtures or novel monopropellants, which are highly oxidizing at operating temperatures. The baseline for developing the improved material system is a prior iridium/rhenium system for chambers burning nitrogen tetroxide/monomethyl hydrazine mixtures, which are less oxidizing. The baseline combustion chamber comprises an outer layer of rhenium that provides structural support, plus an inner layer of iridium that acts as a barrier to oxidation of the rhenium. In the improved material system, the layer of iridium is thin and is coated with a thermal fatigue-resistant refractory oxide (specifically, hafnium oxide) that serves partly as a thermal barrier to decrease the temperature and thus the rate of oxidation of the rhenium. The oxide layer also acts as a barrier against the transport of oxidizing species to the surface of the iridium. Tests in which various oxygen/hydrogen mixtures were burned in iridium/rhenium combustion chambers lined with hafnium oxide showed that the operational lifetimes of combustion chambers of the improved material system are an order of magnitude greater than those of the baseline combustion chambers.

Fortini, Arthur↗