Thrust transient prediction and control of solid rocket engines.
Analytical model for thrust-time curve during ignition transient of solid propellant rocket engines, discussing flame spreading
Macdonald, B. W.↗
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Low-Cost Gas Generator and Ignitor
A gas generator and ignitor assembly for powering a turbine of a turbopump assembly for a rocket engine comprises an injector and a combustion chamber. the injector having a body member including a fuel inlet and an oxidizer inlet spaced one from the other and communicating with respective radially spaced apart annular members in the body member. Three annuli communicate with the fuel inlet and two annuli communicate with the oxidizer inlet. the annuli which communicates with the oxidizer being positioned between pairs of the other annuli, The body member is enclosed by a plate having an array of bores arranged in two series with three radially spaced apart groups of circular rows in each series. The outer series has 28 groups of triplet bores while the inner series has 14 groups of triplet bores. The annuli which communicate with oxidizer feed bores of each series that are between the other bores of a triplet. the latter bores communicating with annuli that communicate with fuel. The inner and outer bores of the triplets of each series are inclined relatively to each other and to the third bore of the triplet so that fuel and oxidizer atomizes as it is sprayed into the inlet of the combustion chamber where the propellants are mixed. and burned. The burning of the propellants is effected by ignition of a plug of solid propellant fuel mounted to communicate with the interior of the combustion chamber.
Henry J Dennis, Jr↗
JANNAF 36th Combustion Subcommittee Meeting
Volume 1, the first of three volumes is a compilation of 47 unclassified/unlimited-distribution technical papers presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 36th Combustion Subcommittee held jointly with the 24th Airbreathing Propulsion Subcommittee and 18th Propulsion Systems Hazards Subcommittee. The meeting was held on 18-21 October 1999 at NASA Kennedy Space Center and The DoubleTree Oceanfront Hotel, Cocoa Beach, Florida. Solid phase propellant combustion topics covered in this volume include cookoff phenomena in the pre- and post-ignition phases, solid rocket motor and gun propellant combustion, aluminized composite propellant combustion, combustion modeling and combustion instability and instability measurement techniques.
Fry, Ronald S.↗
Solid propellant engineering
Surface temperature relations for ignition material deflagration onset, and static tests of Applications Technology Satellite motor
Source record↗
Recommendations for Safe Separation Distances from the Kennedy Space Center (KSC) Vehicle Assembly Building (VAB) Using a Heat-Flux-Based Analytical Approach (Abridged)
The NASA Engineering and Safety Center (NESC) was requested to provide computational modeling to support the establishment of a safe separation distance surrounding the Kennedy Space Center (KSC) Vehicle Assembly Building (VAB). The two major objectives of the study were 1) establish a methodology based on thermal flux to determine safe separation distances from the Kennedy Space Center's (KSC's) Vehicle Assembly Building (VAB) with large numbers of solid propellant boosters containing hazard division 1.3 classification propellants, in case of inadvertent ignition; and 2) apply this methodology to the consideration of housing eight 5-segment solid propellant boosters in the VAB. The results of the study are contained in this report.
Cragg, Clinton H.↗
Utilization of solid-propellant upper stages in STS payload orbital operations
The main purpose of this report is to discuss techniques of trajectory design, maneuver execution, and stage loading that are compatible with the use of SRM's (solid rocket motors) which, once ignited, must burn to propellant depletion. It is anticipated that some shuttle payloads will use non-IUS (interim upper stage) solid propellant kick stages; therefore this subject is also pertinent to shuttle flights other than those involving the use of the IUS. The SRM utilization techniques can be divided into two major categories: (1) those in which the stage performance is adjusted to match the velocity increment magnitude requirements of a preselected trajectory, and (2) those in which the trajectory is designed to match the velocity increment magnitude capability of the stage(s).
Wilson, S. W.↗
The Alamo multiphysics solver for phase field simulations with strong-form mechanics and block structured adaptive mesh refinement
Alamo is a high-performance scientific code that uses block-structured adaptive mesh refinement to solve such problems as: the ignition and burn of solid rocket propellant, plasticity, damage and fracture in materials undergoing loading, and the interaction of compressible flow with eroding solid materials. Alamo is powered by AMReX, and provides a set of unique methods, models, and algorithms that enable it to solve solid-mechanics problems (coupled to other physical behavior such as fluid flow or thermal diffusion) using the power of block-structured adaptive mesh refinement.
36 MATERIALS SCIENCE↗
A Preliminary Investigation on the Destruction of Solid-Propellant Rocket Motors by Impact from Small Particles
An investigation was conducted to determine whether solid-propellant rocket motors could be ignited and destroyed by small-particle impacts at particle velocities up to a approximately 10,940 feet per second. Spheres ranging from 1/16 to 7/32 inch in diameter were fired into simulated rocket motors containing T-22 propellant over a range of ambient pressures from sea level to 0.12 inch of mercury absolute. Simulated cases of stainless steel, aluminum alloy, and laminated Fiberglas varied in thickness from 1/50 to 1/8 inch. Within the scope of this investigation, it was found that ignition and explosive destruction of simulated steel-case rocket motors could result from impacts by steel spheres at the lowest attainable pressure.
Carter, David J., Jr.↗
Solid State Digital Propulsion "Cluster Thrusters" For Small Satellites Using High Performance Electrically Controlled Extinguishable Solid Propellants (ECESP)
Electrically controlled extinguishable solid propellants (ESCSP) are capable of multiple ignitions, extinguishments and throttle control by the application of electrical power. Both core and end burning no moving parts ECESP grains/motors to three inches in diameter have now been tested. Ongoing research has led to a newer family of even higher performance ECESP providing up to 10% higher performance, manufacturing ease, and significantly higher electrical conduction. The high conductivity was not found to be desirable for larger motors; however it is ideal for downward scaling to micro and pico- propulsion applications with a web thickness of less than 0.125 inch/ diameter. As a solid solution propellant, this ECESP is molecularly uniform, having no granular structure. Because of this homogeneity and workable viscosity it can be directly cast into thin layers or vacuum cast into complex geometries. Both coaxial and grain stacks have been demonstrated. Combining individual propellant coaxial grains and/or grain stacks together form three-dimensional arrays yield modular cluster thrusters. Adoption of fabless manufacturing methods and standards from the electronics industry will provide custom, highly reproducible micro-propulsion arrays and clusters at low costs. These stack and cluster thruster designs provide a small footprint saving spacecraft surface area for solar panels and/or experiments. The simplicity of these thrusters will enable their broad use on micro-pico satellites for primary propulsion, ACS and formation flying applications. Larger spacecraft may find uses for ECESP thrusters on extended booms, on-orbit refueling, pneumatic actuators, and gas generators.
Sawka, Wayne N.↗
Impulse Measurements of Electric Solid Propellant in an Electrothermal Ablation-Fed Pulsed Plasma Thruster
Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled and extinguished) through the application and removal of an electric current. These propellants are also being considered for use in the ablative pulsed plasma thruster. In this paper, the performance of an electric solid propellant operating in an electrothermal ablation-fed pulsed plasma thruster was investigated using an inverted pendulum micro-Newton thrust stand. The impulse bit and specific impulse of the device using the electric solid propellant were measured for short-duration test runs of 100 pulses and longer-duration runs to end-of-life, at energy levels of 5, 10, 15 and 20 J. Also, the device was operated using the current state-of-the-art ablation-fed pulsed plasma thruster propellant, polytetrafluoroethylene or PTFE. Impulse bit measurements for PTFE indicate 100±20 μN-s at an initial energy level of 5 J, which increases linearly by ~30 μN-s/J with increased initial energy. Measurements of the impulse bit for the electric solid propellant are on average lower than PTFE by 10% or less. Specific impulse for when operating on PTFE is calculated to be about 450 s compared to 225 s for the electric solid propellant. The 50% reduction in specific impulse is due to increased mass ablated during operation with the electric solid propellant relative to PTFE.
Glascock, Matthew S.↗
Ignition of sounding rocket motors with hand-pumped air
Method demonstrates inexpensive, safe, and foolproof concept for solid propellant rocket motors, using simple handpump to deliver air. Flueric ignition was accomplished using system without stored energy and with complete absence of electrical energy and wiring.
Rakowsky, E. L.↗
Performance Measurements of Electric Solid Propellant in an Ablative Pulsed Electric Thruster
Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled and extinguished) through the application and removal of an electric current. These propellants may also be used for electric in-space propulsion, specifically in the ablative pulsed plasma thruster. In this paper, we will investigate the performance of an electric solid propellant operating in an ablation-fed pulsed plasma device by use of an inverted pendulum micro-Newton thrust stand. Namely, the impulse-per-pulse and the specific impulse of the device using the electric solid propellant will be reported for test runs of 100 pulses and energy levels of 5, 10, 15 and 20 J. Further, the device will also be tested using the current state-of-the-art pulsed plasma thruster propellant, polytetrafluoroethylene. The performance of each propellant will be compared for each energy level using an identical setup and apparatus. This comparison of performance between propellants in a controlled setting will allow for better understanding of previous experimental observations.
Glascock, Matthew S.↗
Specific Impulse of Electric Solid Propellant in an Electrothermal Ablation-Fed Pulsed Plasma Thruster
Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled, and extinguished) through the application and removal of an electric current. Recent work has focused on application of this propellant in an electrothermal ablation-fed pulsed plasma thruster. In this paper, impulse bit measurements in such devices fed by either the electric solid propellant or a traditional state-of-the-art propellant, polytetrafluoroethylene, are expanded upon. It is demonstrated that a surface layer in the hygroscopic electric solid propellant is rapidly ablated over the first few discharges of the device, which correspondingly decreases specific impulse relative to the traditional polytetrafluoroethylene propellant. Correcting these data by subtracting the early discharge ablation mass loss measurements yields a corrected electric solid propellant specific impulse of approximately 300 s. As the test duration increases to a large number of discharges, and the initial mass loss is a reduced fraction of the total, the effect of absorbed water in the propellant is decreased and the specific impulse without any corrections approaches the corrected 300 s value.
Glascock, Matthew S.↗
Initiation system for low thrust motor igniter
A test program was carried out to demonstrate an igniter motor initiation system utilizing the bimetallic material Pyrofuze for a solid propellant rocket with controlled low rate of thrust buildup. The program consisted of a series of vacuum ignition tests using a slab burning window motor that simulated the principal initial ballistic parameters of the full scale igniter motor. A Pyrofuze/pyrotechnic igniter system was demonstrated that uses a relatively low electrical current level for initiation and that eliminates the necessity of a pyrotechnic squib, with its accompanying accidental firing hazards and the typical basket of pyrotechnic pellets. The Pyrofuze ignition system does require an initial constraining of the igniter motor nozzle flow, and at the low initiating electrical current level the ignition delay time of this system was found to be quite sensitive to factors affecting local heat generation or loss rates.
Strand, L. D.↗
Presentation hydrogen-oxygen reaction studies - experimental investigation.
Explosion hazards associated with hydrogen-oxygen system ignition at various pressures and fuel- oxidant ratios, solid and liquid state
PROPELLANT COMBUSTION↗
JANNAF Combustion Subcommittee Meeting
This volume, the first of four volumes, is a collection of 33 unclassified/unlimited papers which were presented at the 33rd Joint Army-Navy-NASA-Air Force (JANNAF) Combustion Subcommittee Meeting in conjunction with the Propulsion Systems Hazards at the Naval Postgraduate School, Monterey, CA. The JANNAF papers contained in this volume review airbreathing combustion measurement, performance and control; electrothermal-chemical gun experiment and modeling; liquid propellant gun experiment and modeling; solid propellant gun interior ballistics; formulation modeling and diagnostics, ignition, and gun barrel wear modeling, blast and flash; and ram, fastcore and gas generator gun propulsion concepts.
Fry, Ronald S.↗
A study of ignition phenomena of bulk metals by radiant heating
Early research on combustion of metals was motivated by the knowledge of the large heat release and corresponding high temperatures associated with metal-oxygen reactions. The advent of space flight brought about an increased interest in the ignition and combustion of metallic particles as additives in solid rocket propellants. More recently, attention has been given to the flammability properties of bulk, structural metals due to the number of accidental explosions of metal components in high-pressure oxygen systems. The following work represents a preliminary study that is part of a broader research effort aimed at providing further insight into the phenomena of bulk metal combustion by looking at the effects of gravity on the ignition behavior of metals. The scope of this preliminary experimental study includes the use of a non-coherent, continuous radiation ignition source, the measurement of temperature profiles of a variety of metals and a qualitative observation of the ignition phenomena at normal gravity. The specific objectives of the investigation include: (1) a feasibility study of the use of a continuous radiation source for metal ignition; (2) testing and characterization of the ignition behavior of a variety of metals; and (3) building a preliminary experimental database on ignition of metals under normal gravity conditions.
Branch, Melvin C.↗
Ignition propagation and heat effects of propellant chips embedded in castable inhibitor using a laser flux test bomb
A castable inhibitor is applied to the aft face of the Space Shuttle Redesigned Solid Rocket Motor (RSRM) forward segment propellant grain to control propellant surface burn area. During fabrication, the propellant surface is trimmed prior to the inhibitor application. This produces a potential for small propellant chips to remain undetected on the propellant surface and contaminate the inhibitor during application. The concern was that undetected propellant chips in the inhibitor might provide a fuse path for premature propellant ignition underneath the inhibitor. To evaluate the fuse path potential, testing was performed on inhibitor samples with embedded propellant. The internal motor environment was simulated with a calibrated CO2 laser beam directed onto a sample which was placed in a 4100 kPa (600 psi) nitrogen pressurized bomb (laser bomb). The testing showed definitive results pertaining to fuse path formation. Embedded propellant chips did not autoignite until the receding heat affected inhibitor surface reached, or passed, the propellant chip. Samples with embedded propellant chips in alignment did not propagate ignition from one chip to another with separation distances as small as 0.010 cm(0.004 inc) and some as little as 0.0051 cm (0.002 in). Propellant chips with volumes approximately less than 0.025 cu cm (0.0015 cu in) (which did not propagate ignition) did not increase the inhibitor material decomposition depth more than the resulting void cavity of the burned out propellant chip. In addition, the depth of this void cavity did not increase until it was overtaken by the surrounding material decomposition depth. This was due, in part, to the retention of the protective inhibitor char layer. Samples with embedded propellant strings, whose thicknesses were below 0.023 cm (0.009 in), did not propagate ignition. Propellant string thicknesses above 0.038 cm (0.015 in) did propagate ignition. Test sample char and heat affected layer measurements and observations compared well with those from the Space Shuttle Solid Rocket Motor (SRM) Technical Evaluation Motor no. 9(TEM-9).
Bolton, Douglas E., Jr.↗