Browse Topic: Solid 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
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.
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.
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).
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.
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).
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.
2013 Create the Future Design: Aerospace & Defense Category WinnerTBMG-1773711/1/2013
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
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
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.
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
Prototype Rocket Engine for a Nanosat Launch VehicleTBMG-493912/1/2008
A paper discusses a prototype rocket engine for a suborbital Reusable Nanosat Launch Vehicle (RNLV). The engine must operate with thrust levels from 3,000 to 5,000 lbf. The engine is comprised of three major subassemblies: the injector, igniter, and the combustion chamber nozzle. Propellants are introduced and mixed in the combustion chamber utilizing a combination of triplet and unlike doublet injector elements. In addition, film cooling is provided in order to extend the life of the ablative chamber. Ignition is accomplished with solid propellant ports mounted on the side of the chamber.
Perspective on Research in Laser PropulsionTBMG-491910/1/2008
A document summarizes the history of research on laser propulsion, starting with the first publication of the laser-propulsion concept in the year 1969, proceeding through studies of laser propulsion within the broader context of advanced propulsion systems in general, and culminating in proof-of-concept experiments and associated studies under the auspices of the Air Force Research Laboratory during the years 1996 through 2007. The document briefly discusses the basic physical principles and engineering concepts of laser propulsion as demonstrated in the proof-of-concept experiments, in which a model rocket denoted a laser lightcraft was lofted to altitudes exceeding 200 ft. (about 61 m). In these experiments, a CO2-laser beam of pulse energy of 400 J at a repetition rate of 25 Hz was used to vaporize a solid propellant (Delrin® polyoxymethylene) to generate thrust. There is a brief discussion of energy-conversion efficiency in laser propulsion. This is followed by an analysis of the concept of supplementing laser ablation energy, and thereby effectively increasing the energy-conversion efficiency, by incorporating a chemically energetic ingredient (ammonium nitrate) into the Delrin propellant. The document includes an extensive bibliography of research on laser propulsion.
Graphite Coating, Thin Lubricating Film Impingement AppliedAMS2525C (Historical)11/21/2007
This specification covers a coating consisting of finely-powdered graphite in a heat-resistant inorganic binder.
AMS B Finishes Processes and Fluids Committee
Molybdenum Disulfide Coating, Thin Lubricating Film Impingement AppliedAMS2526C (Historical)10/5/2007
This specification covers a coating consisting of finely-powdered molybdenum disulfide in a heat-resistant inorganic binder applied to parts.
AMS B Finishes Processes and Fluids Committee
Simulating Flights of Future Launch Vehicles and SpacecraftTBMG-11822/1/2007
Marshall Aerospace Vehicle Representation in C (MAVERIC) is a computer program for generic, low-to-high-fidelity simulation of the flight(s) of one or more launch vehicle(s) or spacecraft. MAVERIC is designed to accommodate multi-staged vehicles, powered serially or in parallel, with multiple engines, tanks, and cargo elements. Engines can be of jet or conventional rocket types, using either liquid or solid propellants.
Embedded Sensors for Measuring Surface RegressionTBMG-777/1/2006
The development and evaluation of new hybrid and solid rocket motors requires accurate characterization of the propellant surface regression as a function of key operational parameters. These characteristics establish the propellant flow rate and are prime design drivers affecting the propulsion system geometry, size, and overall performance. There is a similar need for the development of advanced ablative materials, and the use of conventional ablatives exposed to new operational environments. The Miniature Surface Regression Sensor (MSRS) was developed to serve these applications. It is designed to be cast or embedded in the material of interest and regresses along with it. During this process, the resistance of the sensor is related to its instantaneous length, allowing the real-time thickness of the host material to be established. The time derivative of this data reveals the instantaneous surface regression rate.
Dual-Laser-Pulse IgnitionTBMG-64571/1/2006
A dual-pulse laser (DPL) technique has been demonstrated for generating laser-induced sparks (LIS) to ignite fuels. The technique was originally intended to be applied to the ignition of rocket propellants, but may also be applicable to ignition in terrestrial settings in which electric igniters may not be suitable. Laser igniters have been sought as alternatives to such conventional devices as electrical spark plugs and torch igniters for the following main reasons:
Dual-Laser-Pulse IgnitionTBMG-119871/1/2006
A dual-pulse laser (DPL) technique has been demonstrated for generating laser-induced sparks (LIS) to ignite fuels. The technique was originally intended to be applied to the ignition of rocket propellants, but may also be applicable to ignition in terrestrial settings in which electric igniters may not be suitable. Laser igniters have been sought as alternatives to such conventional devices as electrical spark plugs and torch igniters for the following main reasons:
Self Contained Atmospheric Protective Ensemble (SCAPE) Suits Redesign and Implementation at Kennedy Space Center2005-01-29597/11/2005
The Self Contained Atmospheric Protective Ensemble (SCAPE) suits, worn at the Kennedy Space Center (KSC) have been updated from the original 1970's design. The suits were renamed Propellant Handlers Ensemble (PHE) but are still commonly referred to as SCAPE. Several modifications to the suit were done over the last 20 years to improve the design for operational use. However, anthropometric changes in the user population over time have not been addressed. The following study addressed anthropometric concerns in the current SCAPE population. It was found that all suits had at least one area in which the recommended upper limit was exceeded by technicians. The most common areas to exceed the upper limit were: waist circumference, chest circumference and upper thigh circumference. Forearm circumference posed the least concern unless using long gauntlet glove which cause the twist lock ring to be located at the forearm rather than the wrist.
Mock, Jessica, Miller, Darcy
Membranes on Mars for In-Situ Resource Utilization Processes2004-01-23167/19/2004
The goal of this project is the development and characterization of synthetic membranes for the separation and purification of CO2 from the Martian atmosphere for in-situ resource utilization (ISRU) applications such as in-situ propellant production. Candidate materials should have high selectivity for carbon dioxide over nitrogen and argon, and a glass transition temperature of -40 °C or less to remain in rubbery state at low temperature for high permeance (flux/driving force). Membrane materials we identified include the rubbery polymers poly(dimethyl siloxane) (PDMS) and the copolymer poly(dimethyl, methylphenyl siloxane) (PMPS). Pure and mixed gas permeation experiments with CO2, N2 and Ar were performed with these membrane materials in the temperature range -25 to 21 °C. In experiments with the commercially available PDMS membranes, the pure gas CO2 permeability increases from 1932 Barrers to 2755 Barrers as the temperature decreases from 22 to -30 °C. The ideal CO2/N2 separation factor (ratio of pure gas permeabilities/permeances) increases from 7.5 to 17.5 over the same temperature range. However, in mixed gas experiments, the CO2/N2 separation factor for PDMS was much lower, increasing from 4.5 to 6 as the temperature decreased from 22 to -30 °C. Pure gas permeation results with PMPS membranes also show an increase in CO2 permeability from 1450 Barrers to 1650 Barrers as the temperature decreases from 21 °C to -10 °C for differential feed pressure of 20 psi. The CO2/N2 ideal separation factor increased from 12 to 27 over the same range of temperature. Unlike PDMS, the mixed gas CO2/N2 selectivities are nearly the same as the pure gas values. We will also present reasons for these differences and discuss the design of membrane modules.
Jha, Praveen, Way, J. Douglas, Srinivasan, Ananth, Mason, Larry W.
Novel Acoustical Polyolefin Foams2001-01-15564/30/2001
Novel acoustical materials have been developed. The materials are thermoplastic foams extruded from blends of a polypropylene (PP) resin with an ethylenic polymer resin. One material is an open-cell sheet product made from a blend of a PP resin and a polyolefin elastomer (POE). Another is a large-celled plank foam of substantially closed-cell structure made from a blend of a polypropylene resin and a low density polyethylene (PE) resin. The foam materials are of lightweight, hydrophobic, dust free, recyclable and withstand the temperatures prevailing in automotive uses.
Park, Chung P., Burgun, Sandrine, Brucker, Michel, Subramonian, Suresh
Graphite Coating, Thin Lubricating Film Impingement AppliedAMS2525B (Historical)7/1/2000
This specification covers a coating consisting of finely-powdered graphite in a heat-resistant inorganic binder.
AMS B Finishes Processes and Fluids Committee
Molybdenum Disulfide Coating, Thin Lubricating Film Impingement AppliedAMS2526B (Historical)7/1/2000
This specification covers a coating consisting of finely-powdered molybdenum disulfide in a heat-resistant inorganic binder.
AMS B Finishes Processes and Fluids Committee
Engine; Rocket, Solid Propellant 5 Second Duration, 4500 Pounds ThrustAS8581 (Current)3/1/2000
This specification established the requirements for a 5 second duration, 4500 pound thrust (when determined at 60°F) solid propellant rocket engine hereinafter specified as 5 S-4500.
Aerospace Council
GFSSP — Program for Analyzing Flows in a Complex NetworkTBMG-67241/1/2000
The Generalized Fluid System Simulation Program (GFSSP) version 2.01 is a general-purpose computer program for analyzing steady-state and time-dependent flowrates, pressures, temperatures, and concentrations in a complex flow network. The program is capable of modeling phase changes, compressibility, mixture thermodynamics, and external body forces, such as gravity and centrifugal.
Hybrisol Rocket EnginesTBMG-21129/1/1999
"Hybrisol" denotes a proposed rocket engine that would contain hybrid and solid-propellant parts within a single combustion chamber. ["Hybrid" as used here denotes a type of rocket engine in which a solid fuel is burned by use of a liquid or gaseous oxidizer and the flow of the oxidizer can be throttled to control the engine. Unlike conventional solid rocket propellants, a solid fuel for a hybrid rocket engine can be made relatively inert in the absence of the oxidizer and therefore presents little hazard of explosion or inadvertent ignition.] Inside a hybrisol rocket engine, there would be two concentric tubular energy-storage layers: an outer layer of conventional solid rocket propellant and an inner layer of solid hybrid fuel (see figure).
Safety Criteria for Pneumatic Starting SystemsAIR1639A (Current)3/1/1999
This SAE Aerospace Information Report (AIR) presents safety criteria for pneumatic type engine starting system design and component hardware. Included are safety criteria in design of both starter control valves and starters as well as in design of airframe control systems. Safety topics concern starter valve operation and material application, airframe controls and instrumentation installations and starter rotor integrity and containment.
AE-6 Starting Systems and Auxiliary Power Committee
H-ISP - A Hybrid Inflator Simulation Program1999-01-10693/1/1999
A hybrid-inflator simulation program, H-ISP, has been developed to facilitate the evaluation of hybrid inflators for airbag systems. Based on fundamental principles of species mass and energy conservation and gas dynamics, this program can simulate the complex thermo-chemical processes of a hybrid inflator, including solid propellant combustion, single-phase or two-phase flow, multiple chemical species with temperature- and composition-dependent thermal properties, and heat transfer phenomena. In this paper, we present the theory, computer program structure and usage of H-ISP with an example.
Kang, Jian, Wang, J. T.
Experimental Analysis of Aspirating Airbag Units1999-01-04363/1/1999
Aspirating airbag modules are unique from other designs in that the gas entering the airbag is a mixture of inflator-delivered gas and ambient-temperature air entrained from the atmosphere surrounding the module. Today's sophisticated computer simulations of an airbag deployment typically require as input the mass-flow rate, chemical composition and thermal history of the gas exiting the canister and entering the airbag. While the mass-flow rate and temperature of the inflator-delivered gas can be obtained from a standard tank test, information on air entrainment into an aspirated canister is limited. The purpose of this study is to provide quantitative information about the aspirated mass-flow rate during airbag deployment. Pressure and velocity measurements are combined with high-speed photography in order to gain further insight into the relationship between the canister pressure, the rate of cabin-air entrainment and the airbag deployment. Fully instrumented, production level aspirating airbag modules serve as the principal test device.
Lee, Y. G., Wang, J. T., Green, P. W., Yu, C. S., Butler, P. B., Chen, L. D.
Subliming Solid MicrothrustersTBMG-320428/1/1998
Devices are proposed that would sublime solid propellants to generate small thrusts for maneuvering spacecraft with masses of no more than 15 kg. With solids rather than gases in propellant tanks, there would be no leakage and thus no waste of limited quantities of propellants. There would also be none of the bulk, weight, and cost of plumbing like that needed for handling liquid or gaseous propellants if the propellant tank would be integrated with the subliming solid thruster.
Optimization of the ARIANE 5 EPS Thermal Control Concept9816457/13/1998
The European ARIANE 5 launcher is designed to put a payload mass of about 6.0 t into geostationary transfer orbit. It consists of a lower central stage with two solid propellant strap-on boosters and a single-engined storable propellant upper stage. The ARIANE 5 was successfully launched on October 30th, 1997. This paper deals with the thermal protection of the ARIANE 5 upper stage. The tanks, the actuators, the instrumentation and electronic equipment have to be protected from the heat input of the hot engine nozzle during the 1100 second thrust phase and maintained operational for a further time period for passivation or re-ignition of the engine. The paper provides a short history of the upper stage development and the thermal design for the configuration of the qualification flights. For the production phase the optimization potential with regard to mass saving and simplified integration has been investigated. Thermal design alternatives are revealed and it is shown where the insulation hardware could be further improved.
Schwarz, B., Menn, F.
Hybrid Propulsion System for Returning a Sample From MarsTBMG-322176/1/1998
A paper suggests the development of a hybrid rocket engine and associated equipment for returning a sample of material from Mars at relatively low cost. In a hybrid rocket engine, a solid fuel is burned by use of a liquid or gaseous oxidizer, the flow of which can be throttled to control the engine. Unlike conventional solid rocket propellants, a solid rocket fuel can be made relatively inert in the absence of the oxidizer and therefore presents little hazard of explosion or inadvertent ignition. Unlike conventional (and relatively expensive) liquid rocket propellants, a solid rocket fuel is not corrosive or susceptible to leakage. The solid fuel in the proposed system would be in granular form, packed into the rocket motor. Oxygen or another suitable oxidizer could be transported from Earth together with this solid fuel. Alternatively, oxygen could be generated from CO2 in the Martian atmosphere by use of in-situ resource utilization (ISRU) equipment. Inasmuch as ISRU is not yet a mature technological discipline, some research on ISRU would be necessary to estimate the reduction in cost achieved by not having to carry the oxidizer to Mars.
Martian Atmospheric Utilization by Temperature-Swing Adsorption9615977/1/1996
Technologies that can be used to extract oxygen and other useful products from the Martian atmosphere for exploration missions will require compression of the low-pressure Martian gas. One technique that appears ideally suited for this application is temperature-swing adsorption, which can produce purified and compressed CO2 in a virtually solid-state process whose energy requirements can be met mainly through the diurnal temperature cycle. This paper focuses on material selection and sensitivity of this adsorption process to variations in Mars surface conditions. Experimental results indicate that, of the zeolite and carbon materials studied, a NaX zeolite is a superior adsorbent in terms of the amount of pressurized gas it can produce per unit mass of sorbent.
Finn, John E., McKay, Christopher P.
ROCKET IGNITION AND RELATED NOMENCLATUREAIR913 (Current)6/1/1994
This information report provides a short glossary of rocket ignition and related terms.
E-30 Propulsion Ignition Systems Committee
SAFETY CRITERIA FOR PNEUMATIC STARTING SYSTEMSAIR1639 (Historical)2/1/1994
AE-6 Starting Systems and Auxiliary Power Committee
Rocket Propulsion Experimental Facilities at the Phillips Laboratory, Edwards Air Force Base, California9314464/1/1993
The Phillips Laboratory has extensive rocket propulsion facilities at Edwards Air Force Base, California. These facilities range from small scale cells capable of firing liquid rocket engines as small as 0.01 lbf (0.04 Newtons (N)) thrust up to very large liquid/solid test stands designed for firings up to 10 million lbf (44.5 million N) thrust. This paper gives a review of the variety and capabilities of the numerous facilities at Edwards Air Force Base, as well as an overview of current projects. Emphasis is placed on the general capabilities of these facilities; specifics are available upon request to the author. Trends in the evolution of rocket test facilities are discussed.
Beckmann, Joel W., Kalliomaa, Wayne M.
ROCKET ENGINE TEST FACILITIES AT THE U.S. ARMY TEST & EVALUATION COMMAND9314484/1/1993
The Static Test Branch of the Redstone Technical Test Center (RTTC) has developed a unique test capability, in the Army, for static firing hypergolic liquid rocket engines. To accomplish this test requirement, the Army built a dual position test stand for vertical testing of liquid and solid rocket engines. This test stand is supplied with test support facilities both mechanical and electrical in nature. RTTC's facilities and procedures may be used for future testing of Army, DoD, NASA, and private industry's liquid rocket engines.
LILES, WILLIAM M.
Relationships Between Thrust Vectoring/Reversing Fighter Effectiveness and Air-to-Air Missile Technology9018459/1/1990
Air combat simulator studies consistently show large fighter effectiveness payoffs due to thrust vectoring and/or thrust reversing exhaust nozzles. However, the air-to-air missile related kill ratios achieved during these evaluations vary considerably as a function of maximum missile launch angle-of-attack (AOA). This paper will review the ways thrust vectoring/reversing is used during air combat and then discuss some missile related multi-function nozzle survivability benefits. Typical air combat simulator fighter agility advantages and disadvantages attributable to vectoring/reversing nozzles will then be examined. These include relative antenna train (look) angles, capture times, and capture frequencies during close-in combat. The high AOA launch desirability will be defined along with the payoffs of attaining that capability. Nozzle max vector angles and utilization frequencies along with high AOA missile launch problems and potential solutions will also be addressed. In addition, the relationship between pilot training and missile launch AOA will be shown as it relates to kill ratio trends.
Herrick, Paul W.
The Titan IV launch Vehicle8923069/1/1989
Titan launch vehicles have been contributing to our national space accomplishments for more than 20 years. As the U.S. space program has grown, the Titan family has expanded to meet the changing requirements. The dependability and versatility of Titan vehicles have been demonstrated by their selection for various missions, including strategic intercontinental ballistic missile weapon systems, manned Gemini space flights, NASA interplanetary missions, and critical national security programs. This article summarizes the Titan legacy and is an overview of the newest Titan family member-the Titan IV.
Morrissey, Arthur C., O'Neill, Stephen T.
Engineering Test Satellite-V(ETS-V) and Experiments of Mobile Satellite Communication87242111/13/1987
ENGINEERING TEST SATELLITE V (ETS-V) is the first Japanese three-axis stabilized geostationary satellite of 550kg class. The development objectives of ETS-V are the verification of the H-I launch vehicle, the establishment of key technologies for three-axis-stabilized satellite bus, test of new technologies for the future application satellite and mobile communication experiments. ETS-V was launched on Aug 27, 1987 and was positioned on geostationary orbit at 150°E on Sept. 17. This article reviews the outline of the ETS-V system, bus components and the plan of mobile communication experiments using ETS-V.
Tabata, Joji, Yamada, Shigeo
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