Showing papers on "Rocket published in 2000"

Restricted Shock Separation in Rocket Nozzles

Abstract: In overexpanded rocket nozzles the e ow separates from the nozzle wall at a certain pressure ratio of wall pressure to ambient pressure. Flow separation and its theoretical prediction have been the subject of several experimental and theoretical studies in the past decades. Two distinctive e ow separation phenomena, the freeshock and restricted-shock separation, were observed in experiments with nozzles. Both phenomena are discussed in detail, and the system of recompression shocks and expansion waves is described. For the free-shock case three different shock structures in theplume can occur, namely the regular shock ree ection, the Mach disk, or a cap-like shock pattern. Theappearanceofthesedifferentplumepatternsis discussed. Theseshock structuresareconserved for the full-e owing, but overexpanded, nozzle. Numerical results obtained for existing rocket nozzles, e.g., Space ShuttleMain EngineorVulcain, show a qualitativegood agreement with experimental photographs.Furthermore, an explanation for the appearance of restricted shock separation, which has been widely unknown up to now, is given, analyzing why and under what conditions it occurs. The type of nozzle contour strongly ine uences this form of e ow separation, and restricted shock separation also occursin full-scale, thrust-optimized rocket nozzles. Based on the results established for e ow separation, an outlook on the generation of side loads is given.

Injection and Mixing Processes in High-Pressure Liquid Oxygen/Gaseous Hydrogen Rocket Combustors

Abstract: The injection, mixing combustion processes in a liquid oxygen (LOX)/gaseous hydrogen (GH2) rocket engine combustor at high chamber pressures (10 MPa) are studied and modeled. An experimental LOX/GH2 rocket motor consisting of a single coaxial shear injector element and a cylindrical chamber with optical access has been used for flow visualizations and measurements. Cold-flow injection test utilizing liquid nitrogen and gaseous helium at elevated pressures have been done for flowfield characterization by different diagnostic methods such as flash-ligth photography and high-speed cinematography using a shadowgraph setup. The injection visualizations and studies under cold-flow and combusting conditions revealed a remarkable difference between subcritical spray formation and evaporation and the supercritical injection and mixing process. The study shows that aproaching supercritical chamber pressures injection can no longer be regarded as a spray formation but rather as a fluid/fluid mixing process. As the flow visualizations indicate, the effect of the coaxial atomizer gas is less effective as previously expected. The flame is attached to the LOX post and develops in the LOX post wake. The observed flame holding mechanism is discussed. An evaluation of the radiation spectrum of the flame inside the combustion chamber revealed that radiation in the visible range is mainly due to water vapor.

Thermophysics Characterization of Kerosene Combustion

Abstract: A one-formula surrogate fuel formulation and its quasi-global combustion kinetics model are developed to support the design of injectors and thrust chambers of kerosene-fueled rocket engines. This surrogate fuel model depicts a fuel blend that properly represents the general physical and chemical properties of kerosene. The accompanying gaseous-phase thermodynamics of the surrogate fuel is anchored with the heat of formation of kerosene and verified by comparing a series of one-dimensional rocket thrust chamber calculations. The quasi-global combustion kinetics model consists of several global steps for parent fuel decomposition, soot formation, and soot oxidation and a detailed wet-CO mechanism to complete the combustion process. The final thermophysics formulations are incorporated with a computational fluid dynamics model for prediction of the combustion efficiency of an unielement, tripropellant combustor and the radiation of a kerosene-fueled thruster plume. The model predictions agreed reasonably well with those of the tests.

TRW Pintle Engine Heritage and Performance Characteristics

Abstract: The pintle injector rocket engine is fundamentally different from other rocket engines, which nearly universally employ a series of separate propellant injection orifices distributed across the diameter of the headend of the combustion chamber. The pintle’s central, singular injection geometry results in a combustion chamber flowfield that varies greatly from that of conventional rocket engines. These differences result in certain operational characteristics of great benefit to rocket engine design, performance, stability, and test flexibility. The mid-1950’s origin of the pintle injector concept and the subsequent early development work and applications in rocket engines are reviewed. The pintle engine’s key design and operational features are compared to conventional rocket engines. Pintle injector design refinements and associated recent applications are discussed. The presentation includes photographs and summaries of many different rocket engines that TRW has developed and successfully flown, each of which used the pintle injector.

Rocket and Spacecraft Propulsion: Principles, Practice and New Developments

Abstract: Foreward Authors Preface Acknowledgements Principles of Rocket Propulsion The Thermal Rocket Engine Liquid Propellant Rocket Engines Solid Propellant Rocket Motors Launch Vehicle Dynamics Electric Propulsion Advanced Thermal Rockets Appendix 1: Current Launcher portfolio Appendix 2: Ariane 5 and the Space Shuttle Appendix 3: Orbital motion Bibliography Index.

The physics and engineering of the vasimr engine

Abstract: The Variable Specific Impulse Magnetoplasma Rocket (VASIMR) is a high power, radio frequency-driven magnetoplasma rocket, capable of I s]/thrust modulation at constant power. The physics and engineering of this device have been under study since 1980. The plasma is produced in an integrated plasma injector by a helicon discharge. However, the bulk of the plasma energy is added downstream by ion cyclotron resonance. The system features a magnetic nozzle, which accelerates the plasma particles by converting their azimuthal energy into directed momentum. A NASA-led, research effort, involving several teams in the United States, continues to explore the physics and engineering of the VASIMR, and its extrapolation as a high power, in-space propulsion system. These studies have produced attractive results in a number of areas, involving plasma theory and experiments, systems engineering and mission analysis. A conceptual point design for a 10 kW space demonstrator experiment has been completed.

Application of a Genetic Algorithm to the Optimization of Hybrid Rockets

Abstract: A Genetic Algorithm (GA) optimization technique has been successfully applied to the design of a large hybrid rocket booster. Optimizations on gross liftoff weight and total inert weight have been carried out using the Hybrid ROcket Sizing (HYROCS) code developed at Purdue University. The GA was able to optimize designs which contained both continuous and discrete variables. Discrete variables which were optimized included the propellant combination and the number of fuel ports, while continuous variables such as tank and chamber pressure, and oxidizer massflux level were simultaneously optimized using the GA procedure. Results with optimal or very-nearly optimal solutions have been verified on a design space which happened to contain a very broad, shallow minimum in weight.

Convergence of Two Analytical Flowfield Models Predicting a Destabilizing Agent in Rocket Combustion

Abstract: In the combustion stability assessment of solid propellant rocket motors, several new destabilizing terms are introduced when rotational-flow effects are properly accounted for. Such effects must be included when the wave motion is parallel to the burning surface. A normal fluctuating velocity component then appears in a careful resolution of intrinsic fluid dynamics, including acoustico‐vortical interactions that must satisfy mass and momentum conservation principles while accommodating the no-slip condition at the propellant surface. The source of this destabilizing term appears explicitly in two separate, independently derived, analytical formulations of the internal flowfield. Predictions generated by these analytical models are shown to agree with reliable computational data produced recently by a numerical code that solves the unsteady nonlinear Navier‐Stokes equations. Verification of the analytical formulations by means of theoretical considerations, numerical comparisons, and global error assessments are also undertaken before examining the impact of the new time-dependent radial-velocity correction on rocket stability. The new radial-velocity fluctuations introduce a correction comparable in importance to the classical pressure coupling at the propellant surface. This effect along with several companion terms must be accounted for properly in the assessment of motor stability characteristics.

Development and Testing of ADN-Based Monopropellants in Small Rocket Engines

Abstract: This paper presents the initial research, development and testing of a novel rocket monopropellant which has potentially higher performance, significantly less toxic and environmentally benign compared with hydrazine, which is the current state-of-the-art. The new monopropellant is based on the oxidizer Ammonium Dinitramide (ADN). Swedish Space Corporation (SSC) has since 1995 been working on new chemical propulsion systems for small spacecraft. In 1997, SSC in cooperation with the Swedish Defence Research Establishment (FOA) and Chalmers University of Technology began research and development of ADN-based liquid monopropellants for small rocket engines. The objectives of the work presented here, have been research of the basic principals, design and experimental tests of critical functions and characteristic. The outcome of this work was formulation of a new ADNbased monopropellant candidate, LMP-101, consisting of 61% ADN, 26 % water and 13% glycerol. LMP-101 has a theoretical vacuum specific impulse of 2420 Ns/kg (exp. ratio 50) and an adiabatic combustion temperature of 1970 K. LMP-101 has been tested in different experimental rocket engines and operated both pulsed in steady state. "Proof of concept" has been demonstrated, i.e. the propellant ignites rapidly and is capable of a sustained, complete combustion with clean exhaust gases (i.e. approximately 50% water, 30% nitrogen and 20% carbon dioxide). Restarts have been performed without observed degradation in ignition characteristics. A two-year development phase of a flight-like rocket engine is planned to start during 2000. The work was supported by the Swedish National Space Board and the European Space Agency (ESA). INTRODUCTION Hydrazine and hydrogen peroxide have been recognized as rocket propellants for more than 50 years. During the 1950's and 1960's the use of hydrogen peroxide decreased due to its inferior storage stability and hazards in production and handling. Simultaneously, the hydrazine propulsion system technology developed. With the development of a reliable and long lived catalyst, hydrazine-based monopropellant propulsion system became commonly used for space propulsion applications. Hydrazine emerged as the standard liquid monopropellant. Since then, hydrazine-based systems have performed a vast number of space flights demonstrating the capability of millions of pulses and mission duration of more than 20 years. Moreover, there are today a wide range of qualified commercial off the shelf components suitable for hydrazine based propulsion systems. Personnel safety and recent increased environmental awareness are, and will continue to be, important issues in the context of propulsion system handling and operation. There is also an increased awareness on how safety and handling contributes to the vehicles overall costs. This is of particular importance for small spacecraft handled and operated by organizations not having the infrastructure to handle a propellant such as hydrazine. In the past decades, the payload cost has decreased while the cost of fuelling a spacecraft has increased along with the repeated lowering of the limit value of hydrazine exposure. Over the past few years, hydroxylammonium nitrate (HAN)based monopropellants have emerged as "Green Propellant" candidates for space propulsion '''. Although other candidates such as hydrogenperoxide also could be envisaged as a "Green Propellant" candidate, SSC decided at an early stage to focus on ADN-based monopropellants as the "Green Propellant" candidate for the work described in this paper. For the successful development of a new propellant, it must lead to a more cost effective overall propulsion system. * Head of Dept. Propulsion R&D t Research Engineer, Dept. Propulsion R&D I Research Engineer, FOA Defense Research Establishment Copyright 2)2, is a solid oxidizer salt, mainly intended for high performance solid rocket propellants. It is synthesized from mixed acid (nitric and sulfuric acid), ammonia and salts based on sulfamicacid. All components are standard industrial chemicals. No solvents, except water, are needed to produce ADN. All waste chemicals are recycled. ADN is highly hygroscopic which makes it possible to formulate a liquid monopropellant by dissolving ADN in water and adding a suitable fuel. The amount of ADN that can be dissolved in a solvent depends on the temperature at which the blend is saturated. E.g. at room temperature measurements have shown that it is possible to dissolve 80% ADN in water, while at 0 °C, 70% ADN can be dissolved. Hence, low temperature requirements on the propellant will decrease the specific impulse and the density of the propellant. Figure 4 shows a Differential Scanning Calorimeter (DSC) curve of ADN. The DSC used was a Mettler DSC 30 with a ceramic sensor. The curve shows the endotherm melting peak at 91 °C, followed by an exothermal decomposition at 154 °C. Table 1 gives of the basic properties of ADN. Melting point Heat of formation Density Molecular weight Enthalpy of melting Oxygen balance 91 °C -148kJ/mol 1.81 g/cm 1 24.056 g/mol 140 J/g +25.79 % Table 1: Basic properties of ADN

Development of design methods for short cylindrical supersonic exhaust diffuser

Abstract: Short cylindrical supersonic exhaust diffuser (SED) is needed for use in vertical firing rocket test stands. Design methods are developed and presented in this paper for short SEDs. Incorporation of shock generators further helps in reducing its starting pressure.

The use of GPS arrays in detecting the ionospheric response during rocket launchings

Abstract: In this paper we investigate the form and dynamics of shock acoustic waves (SAW) generated during the rocket Proton launching from the Baikonur cosmodrome in 1998 and 1999. In spite of the difference of geophysical conditions, the ionospheric response for all launchings has a period of about 300 s and the amplitude exceeding background fluctuations under quiet and moderate geomagnetic conditions by factors of 2 to 5 as a minimum. The angle of elevation of the SAW wave vector varies from 45° to 60°, and the SAW phase velocity (900–1200 m/s) approaches the sound velocity at heights of the ionospheric F region maximum. The position of the SAW source, inferred by neglecting refraction corrections, corresponds to the segment of the rockets path at a distance no less than 700–900 km from the launch pad, which is consistent with the estimated delay time of SAW source triggering (250–300 s).

Experimental investigation of the rbcc rocket-ejector mode

Abstract: The experimental results from the study of the rocket-ejector mode of a Rocket Based Combined Cycle (RBCC) engine are presented and discussed. The experiments involved systematic flowfield measurements in a two-dimensional variable geometry rocket-ejector system. The rocket-ejector system utilizes a single two-dimensional gaseous oxygen/gaseous hydrogen rocket as the ejector. To gain a thorough understanding of the rocket-ejector's internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static configurations for a range of rocket operating conditions. Overall system performance was obtained through global measurements of wall static pressure profiles, heat flux profiles and engine thrust, whereas detailed mixing and combustion information was obtained through Raman spectroscopy measurements of major species (oxygen, hydrogen, nitrogen and water vapor). Rocket-ejector experimental results for both the Diffusion and Afterburning (DAB) and Simultaneous Mixing and Combustion (SMC) geometries are presented here. For the DAB configuration, experiments were conducted for both direct-connect and sea-level static configurations, whereas SMC

Virtual prototyping of solid propellant rockets

Abstract: Researchers seek a detailed, whole-system simulation of solid propellant rockets under normal and abnormal operating conditions. A virtual prototyping tool for solid propellant rocket motors based on first principles models of rocket components and their dynamic interactions meets this goal. Given a design specification (geometry, materials, and so on), we hope to predict the entire system's resulting collective behavior with sufficient fidelity to determine both nominal performance characteristics and potential weaknesses or failures. Such a response tool could explore the space of design parameters much more quickly, cheaply, and safely than traditional build-and-test methods. The article is a progress report on a project to design such a virtual prototyping tool for SRMs (solid rocket motors).

Driving dayside convection with northward IMF: Observations by a sounding rocket launched from Svalbard

Abstract: The first sounding rocket flight into the dayside cusp with dark ground and northward interplanetary magnetic field (IMF) conditions was launched from the new SvalRak range at Ny-Alesund in the Svalbard archipelago in early December 1997. Extensive ground observations of auroral emissions and radar backscatter provided contexts for in situ rocket measurements. Real-time interplanetary measurements from the Wind satellite aided launch selection with foreknowledge of impending conditions. NASA rocket flight 36.153 was launched near local magnetic noon while the IMF was dominated by positive B X and had lesser northward B Z and negative B Y components. The rocket's westward trajectory carried it toward auroral forms associated with morningside boundary layers. The rich set of vector dc electric and magnetic fields, energetic particles, thermal plasma, plasma waves, and optical emissions gathered by the rocket reveal a complex electrodynamic picture of the cusp/boundary-layer region. Four factors were important in separating temporal and spatial effects: (1) Near the winter solstice the Earth's north magnetic pole tilts away from the Sun, (2) at the UT of the flight the dipole axis was rotated toward dawn, (3) the variability of solar wind driving was low, and (4) B X was the dominant IMF component. We conclude that no signatures of dayside merging in the Northern Hemisphere were detected in either the rocket or ground sensors. Electric field variations in the interplanetary medium directly correlate with those observed by the sounding rocket, with significantly shorter lag times than estimated for simple propagation between Wind and the ionosphere. The correlation requires that the observed Northern Hemisphere convection structures were stirred in part by merging of the IMF with closed field lines in the Southern Hemisphere, thereby adding open flux to the northern polar cap. Subsequent motions of adiaroic polar cap boundaries were detected in the rocket electric field measurements. The observations indicate that IMF B X significantly affected the location and timing of merging interactions.

System and method for reducing dispersion of small rockets

Abstract: An active damping method and a self-contained active damping system that can be retrofitted to existing rockets are provided which for reducing the dispersion of rockets by using lateral thrusters to oppose any initial yawing motion. The self-contained system of the present invention can be installed in a cylindrical section of a rocket body by insertion between other flight body parts.

Hydrodynamics of a Water Rocket

Abstract: Applications of mass, momentum, and energy balances are central to the teaching of fluid dynamics. In this study, a model to determine the trajectory of a water rocket is given that is far simpler than the system of coupled partial differential equations that typically results in modern hydrodynamic problems of interest. This makes the problem an excellent choice for a student project---it can reasonably be completed with a day or two of effort. In addition to the fundamental mathematics, this problem offers opportunities in scale analysis, numerical methods for IVPs, balance principles in accelerated frames of reference, and the collection and assessment of flight test data.

Oxidizer package for propellant system for rockets

Abstract: An oxidizer package for a propellant system for a motor in which the oxidizer is separated from fuel grain, the oxidizer package comprising oxidizer material and an ignition system therefor in a wrapping or sealing material. A hybrid rocket comprising oxidizer material and fuel grain, the oxidizer material being separated from the fuel grain and being in the form of a single package or plurality of packages of oxidizer material and an ignition system therefor, said packages generally conforming to the shape of the rocket. A grid of a pyrotechnic material. A propulsion system for a hybrid rocket comprising oxidizer material in a matrix, mesh, wool, foamed metal or wires of structural or pyrotechnic material.

A Labscale Hybrid Rocket Motor for Instrumentation Studies

Abstract: An interest in plume spectroscopy led to the development of a labscale Hybrid Rocket Facility at the University of Arkansas at Little Rock (UALR). The goal of this project was to develop a reliable, consistent rocket motor testbed for the development of plume spectroscopy instrumentation. Hybrid motor technology was selected because it has proven to be safe and inexpensive to operate. The project included the design and construction of the labscale hybrid rocket motor, the supporting facility, the istrumentation and computer control of the motor, and the characterization of this particular thruster, including the regression rate of h ydroxyl-terminated polybutadiene (HTPB) fuel grains. For plume spectroscopy experiments, the fuel is doped with metal salts, to simulate either solid motors or liquid engines. It was determined the labscale hybrid motor produces a reliable and consistent plume, resulting in an excellent tool for the development of plume spectroscopy and other instrumentation.

Dynamics of a Variable-Mass, Flexible-Body System

Abstract: A new formulation of the dynamics of a variable-mass, e exible-body system is presented by extending Kane’ s equations for variable mass particles to e exible bodies characterized by load-dependent stiffness and assumed modes. The method captures the effects of thrust, mass center change, and changes in transverse vibration frequencies due to mass loss and thrust. An order- n formulation with prescribed motion of a gimballed nozzle is given for a e exible rocket. Equations for a planar e exible rocket are illustrated and solved numerically. Open-loop simulations with prescribed gimbal motion show that the difference in large motion e ight behavior between a rigid-body model and a e exible-body model of a rocket increases as the e exibility increases. Changes in bending frequencies of a e exible rocket due to mass loss and thrust-induced softening are shown.

Airship shaped space craft

Abstract: The present airship-shaped space craft has a middle fuselage extending in a fore-and-aft direction, and a pair of two outer fuselages extending in the fore-and-aft direction located symmetrically on both sides of the middle fuselage. In the above fuselages, gas of a specific gravity of which is smaller than that of air is filled, and the middle fuselage is connected with the outer fuselages by a horizontal wing. The horizontal wing is provided with propelling devices supported in a gimbal fashion for generating thrust in any optional direction, jet engines with backwards directed nozzles to be controlled within a range from a slantwise upward direction to a slantwise downward direction, and rocket engines with ejection nozzles to be controlled in right-and-left and up-and-down directions. During ascent of the space craft, at first the propelling devices, then the jet engines and at last the rocket engines are actuated so as to make the space craft reach and fly along a satellite orbit. Upon return of the space craft, the rocket engines, the jet engines and the propelling devices are actuated in reverse order, and aerodynamic heating at reentry into the atmospheric space can be reduced by making the space craft descend slowly. Also, control of the flight in the atmospheric space becomes easy, so that the space craft can safely and easily land on a predetermined narrow area of the ground.

Analysis of a water-propelled rocket: A problem in honors physics

Abstract: The air-pumped, water-propelled rocket is a common child’s toy, yet forms a reasonably complicated system when carefully analyzed. A lab based on this system was included as the final laboratory project in the honors version of General Physics I at the USAF Academy. The numerical solution for the height of the rocket is presented, as well as several analytic approximations. Five out of six lab groups predicted the maximum height of the rocket within experimental error.

TRW 40 KLBf LOX/RP-1 Low Cost Pintle Engine Test Results

Abstract: One of NASA's goals is to develop and demonstrate next-generation technology that will enable industry to provide truly affordable and reliable access to space. Their mission is to provide the necessary technology to reduce dramatically both the cost of placing payloads in space and the cost of in-space transportation. This includes the next generation manned and unmanned shuttle and the heavy lift launch vehicles for future Mars and space exploration. History shows that about 40% of the vehicle cost is the cost of the rocket engines. A low-cost rocket engine could dramatically reduce the overall vehicle cost For the last 30 years, TRW has focused on low cost, simplicity and reliability in the design of the TRW single element coaxial pintle injector rocket engine. Early work at TRW involving the LOX/Kerosene propellants started with a 2K engine, tested in the late 1960s, achieving 96% combustion efficiency. In the early 1970s, the TRW Holloman sled engine was retrofitted to burn LOX/RP-1 at 50 kfbf thrust. This engine achieved 93% combustion efficiency. Also, stable combustion was demonstrated in bomb tests. In the early 1990s, TRW tested an LOX/LH2 engine at both the 16.5 and 40K thrust levels, using the same basic injector hardware. In 1995, this engine was retrofitted with injection rings sized for LOX/RP-1 and tested at 13K thrust. This engine demonstrated 98% combustion efficiency in runs with an ablative chamber and was also shown to be stable after bomb testing. Recently, TRW tested a pressure fed 40K LOX/RP-1 low cost pintle engine (LCPE) and the test results are reported herein. *Project Engineer, Member of AlAA **Chief Engineer, Member of AiAA INTRODUCTION SUMMARY In a test program co-sponsored by TRW and the Air Force, TRW tested a pressure-fed 40K LOX/RP-1 engine at the Energetic Materials Research and Testing Center (EMRTC) Rocket Test Site in Socorro New Mexico in November and December of 1999. Testing at 25 klbf thrust and 40 klbf thrust were accomplished, using one injector and chamber assembly with only changes to the injection orifice elements, fn all, 15 hot fire tests were conducted, six at 25 klbf thrust level, and nine at 40 kibf thrust level, demonstrating combustion efficiencies of up to 98.4% of theoretical C* (ODE). Fuel film cooling of the chamber was performed at the 40K thrust level in 4 separate tests, with fuel film cooling flow rates of 4%, 6% and 9% of the total fuel flow demonstrated. Test durations were one to five seconds using a copper heat-sink chamber on loan from NASA Glen Research Center. The 25 klbf tests were operated at 250 psia chamber pressure and a target mixture ratio of 2.4, while the 40 klbf tests were operated at 390 psia chamber pressure and a mixture ratio of 2.25. Data was gathered on seven different oxidizer slot geometries, and three different fuel injection velocities, providing basic information needed for optimization of engine performance. HARDWARE DESCRIPTION The injector and chamber for this test series (Figure 1) were originally built in the early 1990's for testing of the 16 kibf and 40 klbf LOX/LH2, and 13K LOX/RP-1 at NASA Glen (then NASA LeRC), described in References 1 through 4. Modifications for this program involved replacing the original 4-inch pintle diameter with a new 5.25 "Copyright © 2000 by TRW Inc. Published by the American institute of Aeronautics and Astronautics, Inc., with permission." American Institute of Aeronautics and Astronautics (c)2000 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

Determination of Triggering Condition of Vortex-Driven Acoustic Combustion Instability in Rocket Motors

Abstract: A method to determine the sufe cient condition for the occurrence of acoustic combustion instability in solid rocket motors with slotted-tube grain is proposed. The condition is obtained by comparing the frequency of the shedding vortices at the entrance of the slots and the acoustic oscillation frequency in the upstream cylindrical port. To obtain the vortex shedding frequency, a transient e ow analysis is conducted with the consideration of grain-surfaceregression. The method is assessed by analyzing e ve practical solid rocket motors employing slottedtube grain in which acousticcombustion instability occurs in threecases, whereastheothertwo showno signie cant pressureoscillations.Theresultsprovetobesuccessfulforalle vemotors.Furthermore,goodagreementisobtained between the measurements and predictions, not only for the oscillatory frequencies but also for the occurrence time.

Spectral Study of Metallic Molecular Bands in Hybrid Rocket Plumes

Abstract: A labscale hybrid rocket was used to study spectral bands produced by metal combustion. Bands in the ultraviolet visible region (300-750) are of interest. The rubber-like fuel, hydroxyl-terminated polybutadiene (HTPB), was doped with a metallic salt for introduction into the plume during combustion. When introduced, the metals produce atomic line emissions as well as molecular bands due to excited forms of metallic molecules formed in combustion. The most likely molecular band emissions are from the excited states of metal oxides or metal hydroxides formed by these metals in the presence of the oxygen flow of the hybrid rocket. As the concentration of metallic dopants increases in the flame, the molecular band emissions also increase. The fashion by which they increase is observed here. The high concentrations observed for these metals result in intensity versus concentration curves that alter from the expected linear progression for manganese, magnesium and strontium. The molecular band emissions observed for calcium, barium and copper in this study followed linear progression, as does the atomic line emission for barium. The line emissions for manganese, strontium and calcium lean toward the concentration axis. The curves are attributed to self-absorption or increased interactions among mixing species as metal concentration increases in the flame. A pattern-like combustion routine for each metal can be characterized with further study. Introduction Atomic spectral techniques have been used in the past to provide diagnostics for engine health monitoring. The National Aeronautics and Space Administration (NASA) and Stennis Space Center in particular have taken interest in these studies as health monitoring techniques for the Space Shuttle Main Engine (SSME).--' These techniques depend on the relationship of excited atomic species in the motor plume to the amounts introduced by failures in the engine system. It is important that a linear or otherwise describable and reproducible relationship exist, in order to be able to quantify wear or other elemental introduction factors in the motor or engine system. Molecular emissions as seen in the normal realm of atomic spectroscopy are viewed as interference. A classic example is that encountered with analysis of barium in the presence of calcium. The analytical atomic line of barium is swamped by the presence of an overwhelming molecular emission from calcium, such as CaOH. Steps are usually taken to minimize the presence of these molecular bands in such work. However, these type precautions are not applicable to the field of engine health monitoring or in combustion diagnostics when applied to exhaust plumes. Molecular emissions are present in rocket * Graduate Student, Student Member AIAA t Professor, Member of AIAA Copyright © 2000 by Keith Hudson. Published by the American Institute of Aeronautics and Astronautcs, Inc. with permission. 1 American Institute of Aeronautics and Astronautics (c)2000 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. combustion, and should be factored in when quantitative data is required. A thorough study of the effect of molecular emissions in exhaust plumes is necessary to determine interference, fraction of species present as molecular versus atomic and other parameters. The Hybrid Rocket Facility at the University of Arkansas at Little Rock (UALR) was constructed to provide combustion diagnostic testing, and uses a 2 X 10 inch labscale hybrid thruster.' Previous studies have revealed the usefulness of the labscale hybrid rocket system as a plume simulator for other propulsion systems, and characterized it for both atomic and molecular emissions.'' ' The presence of molecular bands was noted in these studies, both from the combustion of HTBP fuel and as formed by metallic dopants, such as manganese. In order to study tike molecular bands in rocket plumes, the labscale hybrid rocket fuel was doped with various levels of metallic salts. Combustion of these salts results in band emissions attributed to metal oxides or metal hydroxides. Some of the metals were chosen due to their presence in alloys used in certain engine components, and because they appear to have produced molecular bands in previous combustion studies.8'10'11'12'13 Other metals were added to the study based on their tendency to oxidize easily, thus are likely to produce refractory particles in combustion.

Multi-level analysis of pulsed detonation engines

Abstract: The present study explores some issues concerning the operational performance of pulsed detonation engines. Zero-, one-, and two-dimensional transient models are employed in a synergistic manner to elucidate the various characteristics that can be expected from each level of analysis. The zero-dimensional model provides rapid parametric trends that help to identify the global characteristics of pulsed detonation engines. The one-dimensional model adds key wave propagation issues that are omitted in the zero-dimensional model and helps to assess its limitations. Finally, the two-dimensional model allows estimates of the first-order multidimensional effects and provides an initial multidimensional end correction for the one-dimensional model. The zero-dimensional results indicate that the pulsed detonation engine is competitive with a rocket engine when exhausting to vacuum conditions. At finite back pressures, the pulsed detonation engine outperforms the rocket if the combustion pressure rise from the detonation is added to the chamber pressure in the rocket. If the two peak pressures are the same, the rocket performance is higher. Two-dimensional corrections added to the one-dimensional model appear to result in a modest improvement in predicted specific impulse over the constant pressure boundary condition.

Bottle rocket launcher

Abstract: A bottle rocket launcher is provided for rockets constructed for educational and entertainment purposes out of synthetic resin bottles used primarily for soft drinks. The rocket launcher includes a bottle plug for filling the bottle rocket with compressed gas such as air and guiding it as it exits the launcher, a release mechanism for initially retaining and then selectively releasing the rocket adjacent its nozzle, and a base. The release mechanism provides multiple hooks which grab a rim adjacent the nozzle, and selectively and simultaneously releases each of the hooks, whereby the upward force applied by the compressed gas against the liquid causes the hooks to slide off of the rim and permits the rocket to lift off of the launcher. The bottle plug is releasably connected to the release mechanism, while a gas delivery conduit remains connected to the bottle plug for inhibiting spillage of liquid from the bottle rocket until the bottle rocket is secured to the release mechanism and filled with compressed gas prior to launch.