Browse Topic: Propellants

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Graphite Coating, Thin Lubricating Film Impingement AppliedAMS2525E (Current)10/10/2019
This specification covers a coating consisting of finely-powdered graphite in a heat-resistant inorganic binder.
AMS B Finishes Processes and Fluids Committee
Organic Compound Turns Toxic Waste into Harmless ByproductsTBMG-341674/1/2019
Spinoff is NASA’s annual publication featuring successfully commercialized NASA technology. This commercialization has contributed to the development of products and services in the fields of health and medicine, consumer goods, transportation, public safety, computer technology, and environmental resources.
Ferrofluid Technology Becomes a Magnet for Pioneering ArtistsTBMG-339413/1/2019
Spinoff is NASA's annual publication featuring successfully commercialized NASA technology. This commercialization has contributed to the development of products and services in the fields of health and medicine, consumer goods, transportation, public safety, computer technology, and environmental resources.
Pulsed Microwave Plasma Instrumentation for Investigation of Plasma-Tuned Multiphase Combustion19AERP02_112/1/2019
Instrumentation developed to support the investigation of electromagnetic wave interaction with energetic materials and flames could help develop microwave-sensitive energetic materials that produce effects such as microwave ignition, acceleration of burning rate, extinguishment, and re-ignition. Air Force Research Laboratory, Arlington, Virginia Strategies to control solid rocket propellant regression rate require a robust throttling technique applicable to high performance propellant formulations. Currently, several methods to control and throttle either motors or subscale propellant strands exist, including chamber pressure control (e.g. pintle nozzles or rapid depressurization quench), infrared laser irradiation of the burning surface to increase burning rates, development of inherently unstable combustion chamber geometries (producing either local pressure or velocity perturbations), and electrically sensitive hydroxylammonium nitrate (HAN)-based formulations in which burning rate is controlled by a voltage potential. However, these techniques are limited in that they either can only be used with low flame temperature (low specific impulse) propellants, result in low propulsion system mass fraction (pintle), are only capable of producing a single perturbation, or are formulation specific. To gain control over a combustion process, combustion plasma enhancement has been demonstrated in electrothermal-chemical (ETC) launchers, in which solid gun propellant ignition flame spread, pressurization rate, and global propellant burning rate improvements were observed. With ETC enhancement, burning rate improvement of up to 35% is possible and further enhancement is speculated to be possible with higher solid loading. However, ETC launchers (e.g. capillary plasma generation) are capable only of single plasma injections or have limited volume.
Pulsed Microwave Plasma Instrumentation for Investigation of Plasma-Tuned Multiphase CombustionTBMG-337682/1/2019
Strategies to control solid rocket propellant regression rate require a robust throttling technique applicable to high performance propellant formulations. Currently, several methods to control and throttle either motors or subscale propellant strands exist, including chamber pressure control (e.g. pintle nozzles or rapid depressurization quench), infrared laser irradiation of the burning surface to increase burning rates, development of inherently unstable combustion chamber geometries (producing either local pressure or velocity perturbations), and electrically sensitive hydroxylammonium nitrate (HAN)-based formulations in which burning rate is controlled by a voltage potential. However, these techniques are limited in that they either can only be used with low flame temperature (low specific impulse) propellants, result in low propulsion system mass fraction (pintle), are only capable of producing a single perturbation, or are formulation specific.
Tungsten Disulfide Coating, Thin Lubricating Film, Binder-Less Impingement AppliedAMS2530A (Current)1/14/2019
This specification covers a coating consisting of tungsten disulfide without binders and not requiring a curing process.
AMS B Finishes Processes and Fluids Committee
Multi-Pulse Motor (MPM) for Use with Electric Solid Propellants (ESP)TBMG-275389/1/2017
The purpose of this work was to create a safe, green, controllable solid rocket motor that can be pulsed a number of times with electricity to control the ignition and extinguishment times to produce a required thrust or impulse bit. The innovation features an Electric Solid Propellant (ESP). The key problem is that the geometry of the ESP grain changes because of the evolution of the propellant to exhaust gases, but a closed electrical circuit is required to keep the electrical power applied to the grain and continue burning. The chamber pressure is utilized in the multi-pulse motor (MPM) design to ensure electrical contact is retained during the pulsing event.
Combustion Characteristics of Hydrocarbon Droplets Induced by Photoignition of Aluminum NanoparticlesTBMG-273818/1/2017
In the study of combustion characteristics of liquid rocket fuels, it is customary to either study the combustion of liquid fuel droplets or the combustion of fuel sprays. However, the two are closely related to each other, because in a typical rocket combustion chamber, the burning of droplets, droplet clusters, and fuel sprays occur simultaneously.
Measuring Propellant Stress Relaxation Modulus Using Dynamic Mechanical AnalyzerTBMG-273808/1/2017
Structural analysis of solid rocket motors is challenging for several reasons, but the most important of these is the complex behavior of the propellant. The mechanical response of a solid propellant is time and temperature dependent. The complexity of the mathematical analysis of the propellant depends on the loading conditions, but for some loading situations, the linear viscoelasticity assumption is reasonable. In particular, linear viscoelasticity is perhaps the most appropriate material behavior description for use in the simulations of stresses related to storage conditions. Typically, simulations use a viscoelastic model in the form of a Prony series and a Williams–Landel–Ferry (WLF) equation. The parameters in these models are derived from stress relaxation experiments, making the stress relaxation experiment a key viscoelastic test, analogous to the tensile test for linear elastic materials.
Combustion Characteristics of Hydrocarbon Droplets Induced by Photoignition of Aluminum Nanoparticles17AERP08_076/1/2017
Test methodology allows analysis of combustion dynamics for subscale rocket injectors under super critical conditions. Air Force Research Laboratory (AFMC), Edwards Air Force Base, California In the study of combustion characteristics of liquid rocket fuels, it is customary to either study the combustion of liquid fuel droplets or the combustion of fuel sprays. However, the two are closely related to each other, because in a typical rocket combustion chamber, the burning of droplets, droplet clusters, and fuel sprays occur simultaneously. The study of droplet burning is mostly focused either on single suspended droplets or free droplets, usually in a free fall. In either case, the conditions regarding the combustion of a single fuel droplet deviates from what is expected in typical liquid rocket engines (LREs). This not only concerns the chamber pressure and the temperature that is usually much lower in a typical droplet study, but also the droplet ignition process and the related transient effects such as temperature ramp up that is much slower than what takes place in LREs. As a result, reported values of combustion parameters such as the burning rate constant and the ignition delay time obtained under typical experimental conditions may not entirely represent the combustion processes in LREs.
Measuring Propellant Stress Relaxation Modulus Using Dynamic Mechanical Analyzer17AERP08_066/1/2017
New testing technique requires less material, gives more accurate results. Air Force Research Laboratory (AFMC), Edwards Air Force Base, California Structural analysis of solid rocket motors is challenging for several reasons, but the most important of these is the complex behavior of the propellant. The mechanical response of a solid propellant is time and temperature dependent. The complexity of the mathematical analysis of the propellant depends on the loading conditions, but for some loading situations, the linear viscoelasticity assumption is reasonable. In particular, linear viscoelasticity is perhaps the most appropriate material behavior description for use in the simulations of stresses related to storage conditions. Typically, simulations use a viscoelastic model in the form of a Prony series and a Williams-Landel-Ferry (WLF) equation. The parameters in these models are derived from stress relaxation experiments, making the stress relaxation experiment a key viscoelastic test, analogous to the tensile test for linear elastic materials. A typical set of stress relaxation tests is performed at several discrete temperatures that cover a range of temperatures anticipated by the fielded motor. At each of the selected temperatures, the specimen is deformed with approximately a single step in strain, which is then held constant for the duration of the test. While held at this constant strain, the stress decays over time due to relaxation of the rubbery elastomer. During this portion of the test, the stresses are measured, and the ratio of stress to applied strain is determined. This ratio is termed the stress relaxation modulus ER. Using time-temperature superposition, the set of curves at the various temperatures can be shifted horizontally relative to each other to form a master curve. The translation of the curves takes a specific mathematical form, viz., the WLF equation. From this master curve, the Prony series at any given temperature can be calculated, and the calculation can be incorporated into finite element analyses along with the WLF equation, making linear viscoelastic analysis of rocket motors possible.
Ultralow-Temperature-Operable Solid Propellant BinderTBMG-249657/1/2016
A unique binder was developed that exhibits a glass transition temperature of –100 °C, which is more than 50 °C lower than that of traditional HTPB (hydroxyl-terminated polybutadiene) and CTPB (carboxyl-terminated polybutadiene) binders. This innovation would be a solid propellant that would ameliorate low-temperature operability problems for the two-stage Mars Ascent Vehicle (MAV).
Spray Characterization and Ignition Delay Measurements of JP-8 and IPK in a Constant-Pressure Flow Chamber2016-01-07364/5/2016
This research compares the spray development and combustion characteristics of jet propellant 8 (JP-8) and iso-paraffinic kerosene (IPK) through a range of diesel engine in-cylinder operating conditions. Non-reacting spray experiments were performed in a constant-pressure flow chamber with 99% nitrogen gas composition at constant temperature (900 K) and densities ranging from 11-56 kg/m3. Near-simultaneous, high-speed Mie and schlieren images of the spray were acquired to measure the liquid and vapor penetration lengths of the non-reacting jet. Reacting experiments, consisting of photodiode measurements and intensified high-speed movies of OH* chemiluminescence, were performed at the same thermodynamic conditions as the non-reacting experiments, except with a 21%/79% oxygen/nitrogen ambient gas composition. Measurements of the rate of injection, issued from a single-hole axial common-rail fuel injector, showed negligible differences between the fuels. The non-reacting liquid length of IPK was approximately 20% shorter than JP-8 for the range of tested conditions, which was consistent with the average difference in volatility between the fuels. The ignition delay, determined from the photodiode record, was up to 80% longer for IPK fuel at low density conditions. Additionally, the lift-off length of IPK was measured to be approximately 50% more sensitive to ambient pressure than JP-8. These results provide fundamental information for the calibration and optimization of military diesel engines operating on conventional and alternatively sourced jet fuels.
Tess, Michael, Kurman, Matthew, Kweon, Chol-Bum
Cyclops: the Space Station Integrated Kinetic Launcher for Orbital Payload Systems (SSIKLOPS)TBMG-242584/1/2016
The Space Station Integrated Kinetic Launcher for Orbital Payload Systems (SSIKLOPS), also known as “Cyclops,” deployed the largest satellite ever from the International Space Station (ISS) on November 28, 2014. The satellite, SpinSat, a Naval Research Laboratory (NRL)/Department of Defense Space Test Program (DoD STP) satellite, is pioneering the utilization of electronically controlled solid propellant thrusters as well as acquiring vital atmospheric density data. It is a spherical satellite 22 inches in diameter, weighing 115 pounds, and will remain in orbit for over two years.
Multi-Pulse Motor (MPM) Designed for Use with Electric Solid PropellantsTBMG-241043/1/2016
The multi-pulse motor is a solid-propellant rocket motor that is able to produce a number of pulses for various thrust levels (5 to 30 pulses and thrusts between 0.25 and 1.5 N, depending on electric power delivery system) and can be turned on and off through the application of electrical power.
Six-Degree-of-Freedom Control With Only Eight ThrustersTBMG-2346212/1/2015
Typical spacecraft thruster configurations are often unable to provide full six-degree-of-freedom control and may have unwanted interaction between their attitude control and trajectory control functions, have undesirably high instantaneous electrical power demands, and use more thrusters than desirable. These last two potential problems gain increased significance if a spacecraft is required to have especially small size and mass, and have very low cost.
Optical Hotspot Conductive Fluid Flow SensorTBMG-228479/1/2015
The Optical Hotspot Flow Sensor (OHFS) enables optical measurement of electrically conductive liquid metal flow rate by exploiting the electrically conductive nature of liquid metals or any other conductive fluid to enable thermal tagging of liquid flow in tubes (US Patent 7,409,875). It is well suited for measuring very low flow rates, on the order of mg/s.
Non-Thermal, Plasma-Assisted Catalytic Reactor for CO2 MethanationTBMG-224137/1/2015
Converting in-situ resources such as CO2, which is the main component of the Mars atmosphere, into methane for rocket propellants can significantly reduce the cost and risk of human exploration while at the same time enabling new mission concepts and long-term exploration sustainability. Methanation of CO2, also called a Sabatier reaction, is hence a key enabling technology required for sustainable and affordable human exploration of Mars.
Improved Hall Thrusters Fed by Solid Phase PropellantTBMG-224157/1/2015
Hall thrusters normally use Xe propellant, which is expensive and scarce in the solar system. The weight of Xe is such that typical Hall thrusters are limited in specific impulse to approximately 3,000 s. The objective of this program was to improve and demonstrate Mg Hall thruster systems. Mg is abundant in the solar system and has an atomic mass approximately one-fifth that of Xe, which means much higher specific impulse is achieved than with Xe at typical thruster operating conditions (power, voltage).
Micropulse Detonation Rocket Engine for Nano-Satellite PropulsionTBMG-214132/1/2015
An efficient propulsion system would use a micropulse detonation rocket engine (–PDRE) for nano-satellite maneuverability in space. Technical objectives are to design, build, and conduct a small detonation tube experiment in order to explore the feasibility of using –PDRE for propelling a nano-satellite. The plan is to study the requirement and predict the performance of –PDRE using various candidate propellants, as well as to conduct ground experiments, demonstrate useful thrust, and measure the specific impulse in a two-year time frame, so that a follow-on project can be proposed in a future NRI Center Innovation Fund.
Nitrous Oxide Ethane-Ethylene EngineTBMG-2071210/1/2014
The Nitrous Oxide Ethylene-Ethane (NEE) engine uses nitrous oxide as an autogenously pressurizing oxidizer, and a mixture of ethane and ethylene is used in the same manner as fuel. Initially, the ethane and ethylene mixture has the same vapor pressure as the nitrous oxide. By using the autogenous pressurization capabilities of these propellants, instead of an additional pressurization system, greater system simplicity and reliability can be attained. The NEE can obtain a specific impulse of 320 s, making it the highest-performing, non-toxic, storable bipropellant rocket propulsion system in existence at the time of this reporting.
Production of Satellite with First All-Electric Propulsion System AdvancesTBMG-2076710/1/2014
Boeing is “running on schedule” as it continues to achieve production milestones for the first of its all-electric-propulsion 702SP (small platform) satellites.
Pulsed Plasma Lubricator (PPL) Technology for the In Situ Replenishment of Dry Lubricants in Extreme EnvironmentsTBMG-193744/1/2014
NASA missions employing mobility systems and other moving mechanical assemblies for application on Mars, the Moon, and in deep space depend on the reliable operation of these assemblies and their tribological components. Wet lubricants are sometimes used in space applications, but in order to avoid solidification, they often require active heating due to the extreme cold temperatures that are encountered. Dry lubricants, such as molybdenum disulfide (MoS2), are more commonly chosen for space mechanisms because they are not subject to the low-temperature limitations of wet lubricants while also providing superior lubricating properties. A major drawback of dry lubricants is low wear resistance that eventually leads to failure of the assembly as the lubricant is removed.
Thrusters Support Communications Satellite LaunchTBMG-193584/1/2014
Monopropellant (hydrazine) thrusters Aerojet Rocketdyne Sacramento, CA 916-355-4000
2013 Create the Future Design: Aerospace & Defense Category WinnerTBMG-1773711/1/2013
High-Pressure Lightweight ThrustersTBMG-1744510/1/2013
Returning samples of Martian soil and rock to Earth is of great interest to scientists. There were numerous studies to evaluate Mars Sample Return (MSR) mission architectures, technology needs, development plans, and requirements. The largest propulsion risk element of the MSR mission is the Mars Ascent Vehicle (MAV). Along with the baseline solid-propellant vehicle, liquid propellants have been considered. Similar requirements apply to other lander ascent engines and reaction control systems.
Impact and Manufacturing Defect Visualization of Space Launcher Aluminum Liner/Filament Wound Composite Fuel Tank using Ultrasonic Propagation Imaging System2013-01-22569/17/2013
We applied ultrasonic propagation imaging (UPI) system for rapid and reliable quality control of fuel tanks for a space launcher. The fuel tank is an aluminum-lined CFRP propellant tank. The UPI system uses Q-switched laser (QL) to generate ultrasonic wave on the test specimen, and laser mirror scanner (LMS) to control the laser impinging point that scans the area of interest with high speed. Each ultrasonic wave generated by laser impinging was received by a piezoelectric sensor with coordinate information of the scanned area. After ultrasonic propagation image processing, results with impact damage and manufacturing defect information of the fuel tank were presented.
Jeong, Hyomi
Graphite Coating, Thin Lubricating Film Impingement AppliedAMS2525D (Historical)7/10/2013
This specification covers a coating consisting of finely-powdered graphite in a heat-resistant inorganic binder.
AMS B Finishes Processes and Fluids Committee
Pneumatic Proboscis Heat-Flow ProbeTBMG-163235/1/2013
Heat flow is a fundamental property of a planet, and provides significant constraints on the abundance of radiogenic isotopes, the thermal evolution and differentiation history, and the mechanical properties of the lithosphere. Heat-flow measurements are also essential in achieving at least four of the goals set out by the National Research Council for future lunar exploration. The heat-flow probe therefore directly addresses the goal of the Lunar Geophysical Network, which is to understand the interior structure and composition of the Moon.
Molybdenum Disulfide Coating, Thin Lubricating Film Impingement AppliedAMS2526D (Current)1/12/2013
This specification covers requirements for a coating consisting of finely-powdered molybdenum disulfide in a heat-resistant inorganic binder applied to parts.
AMS B Finishes Processes and Fluids Committee
Standard Electrical and Logical Interface for Airborne Fuzing SystemsAS5716A (Current)12/3/2012
This interface standard applies to fuzes/fuzing systems (referred to as fuzing system hereafter) in airborne weapons that use a MIL-STD-1760 type interface. It defines the powers, the discrete signals and the serial data interface for the communications at the interface between the fuzing system and the remainder of the weapon, including the weapon control unit. The Class 1 interface is an electrical only interface that facilitates use of MIL-STD-1760 type platform store interfaces for the fuze to monitor intentional release and defines the fuze interface bus communications protocol to allow sending and receiving data from fuzing systems. Class 2 interfaces add a defined connector and additional interfaces to facilitate the exchange of compatible fuzing systems. Class 3 interfaces add further interface definitions to facilitate the exchange of AS5680A compatible fuzing systems components. The bus communications protocol provides a means by which the weapon may set mission parameters within the fuzing system main housing and in other devices external to the fuzing system main housing. The standard also defines the target detection signal to be provided by a target detection device external to the fuzing system. This standard does not address the mechanical interfaces to target detection devices or to other sensors that are external to the fuzing system main housing.
AS-1B Aircraft Store Integration Committee
Activation and Control of Autoignition in HCCI Engines Using Volumetrically-Distributed Ignition of As-Produced Single-Walled Carbon Nanotubes2012-01-16919/10/2012
The discovery that nanostructured materials exhibit properties different than their bulk materials provided many exciting opportunities with technological applications. One such opportunity is the observed ignition of the single-walled carbon nanotubes (SWCNTs) with an ordinary camera flash. In this paper, light-activated ignition characteristics of the as-produced SWCNTs (50 wt% iron nanoparticle content) with a camera flash are presented. The primary objective of this work is to use nanostructured materials as means for distributed (or volumetric) ignition and improved combustion in propulsion systems. Important examples are homogeneous-charged compression ignition (HCCI) engines, liquid rocket fuel sprays, and enhanced flame stabilization in gas turbine engines. The idea was originally proposed by the author in April 2003 and the first patent filed in July 2004 following a series of initial investigations. Based on these and additional tests, this new ignition method is now considered as a potential enabling technology for volumetric and distributed ignition of liquid fuel sprays or gaseous fuel-air mixtures with the lowest incident power intensity possible. This means remote and spatial ignition within any desired and adjustable region defined by the shape of the light from a pulsed light source. Average intensities in between 10 to 150 W/cm₂ are required for ignition of SWCNTs. This is a factor of 80 less than cases where lasers (pulsed and continuous wave (cw)) are used in coal particles. Results acquired in a premixed gaseous fuel-air mixture in a cylindrical combustion chamber, comparing a spark plug with the light-activated distributed ignition of SWCNTs, confirmed the patented concept and showed a truly on-demand activation of the autoignition process for HCCI engine applications. Faster fuel-air mixture burn rate reaching up to a factor of 3 has been demonstrated for distributed ignition under lean mixture as compared with a conventional spark ignition system.
Chehroudi, Bruce
Modular, Rapid Propellant Loading System/Cryogenic TestbedTBMG-146139/1/2012
The Cryogenic Test Laboratory (CTL) at Kennedy Space Center (KSC) has designed, fabricated, and installed a modular, rapid propellant-loading system to simulate rapid loading of a launch-vehicle composite or standard cryogenic tank. The system will also function as a cryogenic testbed for testing and validating cryogenic innovations and ground support equipment (GSE) components. The modular skid-mounted system is capable of flow rates of liquid nitrogen from 1 to 900 gpm (≈3.8 to 3,400 L/min), of pressures from ambient to 225 psig (≈1.5 MPa), and of temperatures to –320 °F (≈–195 °C). The system can be easily validated to flow liquid oxygen at a different location, and could be easily scaled to any particular vehicle interface requirements.
Navigation/Prop Software SuiteTBMG-138706/1/2012
Navigation (Nav)/Prop software is used to support shuttle mission analysis, production, and some operations tasks. The Nav/Prop suite containing configuration items (CIs) resides on IPS/Linux workstations. It features lifecycle documents, and data files used for shuttle navigation and propellant analysis for all flight segments. This suite also includes trajectory server, archive server, and RAT software residing on MCC/Linux workstations.
Thermodynamic Vent System for an On-Orbit Cryogenic Reaction Control EngineTBMG-135845/1/2012
A report discusses a cryogenic reaction control system (RCS) that integrates a Joule-Thompson (JT) device (expansion valve) and thermodynamic vent system (TVS) with a cryogenic distribution system to allow fine control of the propellant quality (subcooled liquid) during operation of the device. It enables zero-venting when coupled with an RCS engine. The proper attachment locations and sizing of the orifice are required with the propellant distribution line to facilitate line conditioning. During operations, system instrumentation was strategically installed along the distribution/TVS line assembly, and temperature control bands were identified.
Oxygen-Methane ThrusterTBMG-132714/1/2012
An oxygen-methane thruster was conceived with integrated igniter/injector capable of nominal operation on either gaseous or liquid propellants. The thruster was designed to develop 100 lbf (≈445 N) thrust at vacuum conditions and use oxygen and methane as propellants. This continued development included refining the design of the thruster to minimize part count and manufacturing difficulties/cost, refining the modeling tools and capabilities that support system design and analysis, demonstrating the performance of the igniter and full thruster assembly with both gaseous and liquid propellants, and acquiring data from this testing in order to verify the design and operational parameters of the thruster.
Technique for Configuring an Actively Cooled Thermal Shield in a Flight SystemTBMG-1136310/1/2011
Broad area cooling shields are a massefficient alternative to conductively cooled thermal radiation shielding. The shield would actively intercept a large portion of incident thermal radiation and transport the heat away using cryogenic helium gas. The design concept consists of a conductive and conformable surface that maximizes heat transfer and formability.
Ionic Liquids to Replace HydrazineTBMG-94443/1/2011
A method for developing safe, easy-to-handle propellants has been developed based upon ionic liquids (ILs) or their eutectic mixtures. An IL is a binary combination of a typically organic cation and anion, which generally produces an ionic salt with a melting point below 100 °C. Many ILs have melting points near, or even below, room temperature (room temperature ionic liquids, RTILs). More importantly, a number of ILs have a positive enthalpy of formation. This means the thermal energy released during decomposition reactions makes energetic ILs ideal for use as propellants. Unlike traditional, storable propellants like hydrazine, ILs also exhibit near-zero vapor pressure. This makes them safer to handle because it eliminates hazardous inhalation — the primary pathway — as a route for toxicity in humans. Thus, ILs are ideal candidates for replacing hydrazine, which is expensive and dangerous, and poses significant handling difficulties.
Crane Brake Release ToolTBMG-91892/1/2011
At the Rotation, Processing and Surge Facility (RPSF), the Solid Rocket Booster (SRB) segments are processed after they arrive by rail prior to a Shuttle launch. Using overhead bridge cranes capable of lifting up to 200 tons, the engineers remove the rail cover of the SRB segments, inspect the propellant grain and metal surfaces and rotate the SRB segments to vertical and secure them on a pallet until launch.
Plasma Igniter for Reliable Ignition of Combustion in Rocket EnginesTBMG-90101/1/2011
A plasma igniter has been developed for initiating combustion in liquid-propellant rocket engines. The device propels a hot, dense plasma jet, consisting of elemental fluorine and fluorine compounds, into the combustion chamber to ignite the cold propellant mixture. The igniter consists of two coaxial, cylindrical electrodes with a cylindrical bar of solid Teflon plastic in the region between them. The outer electrode is a metal (stainless steel) tube; the inner electrode is a metal pin (mild steel, stainless steel, tungsten, or thoriated-tungsten). The Teflon bar fits snugly between the two electrodes and provides electrical insulation between them. The Teflon bar may have either a flat surface, or a concave, conical surface at the open, down-stream end of the igniter (the igniter face). The igniter would be mounted on the combustion chamber of the rocket engine, either on the injector-plate at the upstream side of the engine, or on the sidewalls of the chamber. It also might sit behind a valve that would be opened just prior to ignition, and closed just after, in order to prevent the Teflon from melting due to heating from the combustion chamber.
Self-Adjusting Liquid Injectors for CombustorsTBMG-863310/1/2010
A class of self-adjusting injectors for spraying liquid oxidizers and/or fuels into combustion chambers has been proposed. The proposed injectors were originally intended for use in rocket-engine combustion chambers, but could also be used to improve control over flows of liquid propellants in other combustion chambers.
Maglev Facility for Simulating Variable GravityTBMG-84979/1/2010
An improved magnetic levitation apparatus (“Maglev Facility”) has been built for use in experiments in which there are requirements to impose variable gravity (including zero gravity) in order to assess the effects of gravity or the absence thereof on physical and physiological processes. The apparatus is expected to be especially useful for experiments on the effects of gravity on convection, boiling, and heat transfer in fluids and for experiments on mice to gain understanding of bone loss induced in human astronauts by prolonged exposure to reduced gravity in space flight.
Microwave Plasma Hydrogen Recovery SystemTBMG-84848/31/2010
A microwave plasma reactor was developed for the recovery of hydrogen contained within waste methane produced by Carbon Dioxide Reduction Assembly (CRA), which reclaims oxygen from CO2. Since half of the H2 reductant used by the CRA is lost as CH4, the ability to reclaim this valuable resource will simplify supply logistics for long-term manned missions. Microwave plasmas provide an extreme thermal environment within a very small and precisely controlled region of space, resulting in very high energy densities at low overall power, and thus can drive high-temperature reactions using equipment that is smaller, lighter, and less power-consuming than traditional fixed-bed and fluidized-bed catalytic reactors. The high energy density provides an economical means to conduct endothermic reactions that become thermodynamically favorable only at very high temperatures.
Non-Pyrotechnic Zero-Leak Normally Closed ValveTBMG-82988/1/2010
This valve is designed to create a zero-leak seal in a liquid propulsion system that is a functional replacement for the normally closed pyrovalve. Unlike pyrovalves, Nitinol is actuated by simply heating the material to a certain temperature, called the transition temperature. Like a pyrovalve, before actuation, the upstream and downstream sections are separated from one another and from the external environment by closed welded seals. Also like pyrovalves, after actuation, the propellant or pressurant gas can flow without a significant pressure drop but are still separated from the external environment by a closed welded seal.
Fuel-Cell Power Source Based on Onboard Rocket PropellantsTBMG-81717/1/2010
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.
Radio-Frequency Tank Eigenmode Sensor for Propellant Quantity GaugingTBMG-81517/1/2010
Although there are several methods for determining liquid level in a tank, there are no proven methods to quickly gauge the amount of propellant in a tank while it is in low gravity or under low-settling thrust conditions where propellant sloshing is an issue. Having the ability to quickly and accurately gauge propellant tanks in low-gravity is an enabling technology that would allow a spacecraft crew or mission control to always know the amount of propellant onboard, thus increasing the chances for a successful mission.
Enriched Storable Oxidizers for Rocket EnginesTBMG-79635/1/2010
The name “enriched storable oxidizers” (ESOs) has been coined for a family of optimized mixtures of between two and four oxidizer fluids. For most applications, the constituents of these mixtures would be nitrogen tetroxide (N2O4), nitrous oxide (N2O), and nitrogen dioxide (NO2); in some applications, the mixtures might include inhibited red fuming nitric acid [IRFN (which consists of red fuming nitric acid to which some hydrogen fluoride is added to reduce its corrosive effect]. The optimum proportions of these constituents would be different for different applications. ESOs were originally proposed for use in spacecraft and launch-rocket propulsion systems: ESOs could be especially useful in advanced spacecraft propulsion systems that could operate in multiple modes. ESOs might also be useful in special terrestrial applications that could include ramjet and scramjet aircraft engines.
Compact Rare Earth Emitter Hollow CathodeTBMG-76133/1/2010
A compact, high-current, hollow cathode utilizing a lanthanum hexaboride (LaB6) thermionic electron emitter has been developed for use with high-power Hall thrusters and ion thrusters. LaB6cathodes are being investigated due to their long life, high current capabilities, and less stringent xenon purity and handling requirements compared to conventional barium oxide (BaO) dispenser cathodes. The new cathode features a much smaller diameter than previously developed versions that permit it to be mounted on axis of a Hall thruster (“internally mounted”), as opposed to the conventional side-mount position external to the outer magnetic circuit (“externally mounted”). The cathode has also been reconfigured to be capable of surviving vibrational loads during launch and is designed to solve the significant heater and materials compatibility problems associated with the use of this emitter material. This has been accomplished in a compact design with the capability of high-emission current (10 to 60 A). The compact, high-current design has a keeper diameter that allows the cathode to be mounted on the centerline of a 6-kW Hall thruster, inside the iron core of the inner electromagnetic coil.
Standard Electrical and Logical Interface for Airborne Fuzing SystemsAS5716 (Historical)2/11/2010
This interface standard applies to fuzes/fuzing systems (referred to as fuzing system hereafter) in airborne weapons that use a MIL-STD-1760 interface. It defines the powers, the discrete signals and the serial data interface for the communications at the interface between the fuzing system and the remainder of the weapon, including the weapon control unit, for Class 1 interfaces. Future issues of the standard will provide for additional fuzing system related functionality defined as Class 2 and Class 3 interfaces. For future issues of this standard, the connector definition is contained in AS5680. This standard does not impose any safety requirements and does not supersede or replace any existing applicable safety standards.
AS-1B Aircraft Store Integration Committee
Tungsten Disulfide Coating, Thin Lubricating Film, Binder-Less Impingement AppliedAMS2530 (Historical)1/5/2010
This specification covers a coating consisting of tungsten disulfide without binders and not requiring a curing process.
AMS B Finishes Processes and Fluids Committee
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