Showing papers on "Rocket published in 2007"

Fundamentals of Hybrid Rocket Combustion and Propulsion

Abstract: * Introduction To Hybrids * Review Of Solid-Fuel Regression Rate Behavior In Classical And Non-Classical Hybrid Rocket Motors * Solid Fuel Pyrolysis Phenomena And Regression Rate, Part I: Mechanisms * Solid Fuel Pyrolysis Phenomena And Regression Rate, Part II: Measurement Techniques * Analytical Models For Hybrid Rockets * Vortex Injection Hybrid Rockets * High Speed Flow Effects In Hybrid Rockets * Computational Fluid Dynamics Modeling Of Hybrid Rocket Flowfields * Combustion Instability And Transient Behavior In Hybrid Rocket Motors * Metals, Other Energetic Additives, And Special Binders Used In Solid Fuels For Hybrid Rocket Applications.

Propellants and explosives : thermochemical aspects of combustion

Abstract: Preface. Preface to the Second Edition. 1 Foundations of Pyrodynamics. 1.1 Heat and Pressure. 1.2 Thermodynamics in a Flow Field. 1.3 Formation of Propulsive Forces. 1.4 Formation of Destructive Forces. 2 Thermochemistry of Combustion. 2.1 Generation of Heat Energy. 2.2 Adiabatic Flame Temperature. 2.3 Chemical Reaction. 2.4 Evaluation of Chemical Energy. 3 Combustion Wave Propagation. 3.1 Combustion Reactions. 3.2 Combustion Wave of a Premixed Gas. 3.3 Structures of Combustion Waves. 3.4 Ignition Reactions. 3.5 Combustion Waves of Energetic Materials. 4 Energetics of Propellants and Explosives. 4.1 Crystalline Materials. 4.2 Polymeric Materials. 4.3 Classification of Propellants and Explosives. 4.4 Formulation of Propellants. 4.5 Nitropolymer Propellants. 4.6 Composite Propellants. 4.7 Composite-Modified Double-Base Propellants. 4.8 Black Powder. 4.9 Formulation of Explosives. 5 Combustion of Crystalline and Polymeric Materials. 5.1 Combustion of Crystalline Materials. 5.2 Combustion of Polymeric Materials. 6 Combustion of Double-Base Propellants. 6.1 Combustion of NC-NG Propellants. 6.2 Combustion of NC-TMETN Propellants. 6.3 Combustion of Nitro-Azide Propellants. 6.4 Catalyzed Double-Base Propellants. 7 Combustion of Composite Propellants. 7.1 AP Composite Propellants. 7.2 Nitramine Composite Propellants. 7.3 AP-Nitramine Composite Propellants. 7.4 TAGN-GAP Composite Propellants. 7.5 AN-Azide Polymer Composite Propellants. 7.6 AP-GAP Composite Propellants. 7.7 ADN , HNF, and HNIW Composite Propellants. 8 Combustion of CMDB Propellants. 8.1 Characteristics of CMDB Propellants. 8.2 AP-CMDB Propellants. 8.3 Nitramine-CMDB Propellants. 8.4 Plateau Burning of Catalyzed HMX-CMDB Propellants. 9 Combustion of Explosives. 9.1 Detonation Characteristics. 9.2 Density and Detonation Velocity. 9.3 Critical Diameter. 9.4 Applications of Detonation Phenomena. 10 Formation of Energetic Pyrolants. 10.1 Differentiation of Propellants, Explosives, and Pyrolants. 10.2 Energetics of Pyrolants. 10.3 Energetics of Elements. 10.4 Selection Criteria of Chemicals. 10.5 Oxidizer Components. 10.6 Fuel Components. 10.7 Metal Azides. 11 Combustion Propagation of Pyrolants. 11.1 Physicochemical Structures of Combustion Waves. 11.2 Combustion of Metal Particles. 11.3 Black Powder. 11.4 Li-SF6 Pyrolants. 11.5 Zr Pyrolants. 11.6 Mg-Tf Pyrolants. 11.7 B-KNO3 Pyrolants. 11.8 Ti-KNO3 and Zr-KNO3 Pyrolants. 11.9 Metal-GAP Pyrolants. 11.10 Ti-C Pyrolants. 11.11 NaN3 Pyrolants. 11.12 GAP-AN Pyrolants. 11.13 Nitramine Pyrolants. 11.14 B-AP Pyrolants. 11.15 Friction Sensitivity of Pyrolants. 12 Emission from Combustion Products. 12.1 Fundamentals of Light Emission. 12.2 Light Emission from Flames. 12.3 Smoke Emission. 12.4 Smokeless Pyrolants. 12.5 Smoke Characteristics of Pyrolants. 12.6 Smoke and Flame Characteristics of Rocket Motors. 12.7 HCl Reduction from AP Propellants. 12.8 Reduction of Infrared Emission from Combustion Products. 13 Transient Combustion of Propellants and Pyrolants. 13.1 Ignition Transient. 13.2 Ignition for Combustion. 13.3 Erosive Burning Phenomena. 13.4 Combustion Instability. 13.5 Combustion under Acceleration. 13.6 Wired Propellant Burning. 14 Rocket Thrust Modulation. 14.1 Combustion Phenomena in a Rocket Motor. 14.2 Dual-Thrust Motor. 14.2.3 Dual-Grain Dual-Thrust Motor. 14.4 Erosive Burning in a Rocket Motor. 14.5 Nozzleless Rocket Motor. 14.6 Gas-Hybrid Rockets. 15 Ducted Rocket Propulsion. 15.1 Fundamentals of Ducted Rocket Propulsion. 15.2 Design Parameters of Ducted Rockets. 15.3 Performance Analysis of Ducted Rockets. 15.4 Principle of the Variable Fuel-Flow Ducted Rocket. 15.5 Energetics of Gas-Generating Pyrolants. 15.6 Combustion Tests for Ducted Rockets. Appendix A: List of Abbreviations of Energetic Materials. Appendix B: Mass and Heat Transfer in a Combustion Wave. B.1 Conservation Equations at a Steady State in a One-Dimensional Flow Field. B.2 Generalized Conservation Equations at a Steady-State in a Flow Field. Appendix C: Shock Wave Propagation in a Two-Dimensional Flow Field. C.1 Oblique Shock Wave. C.2 Expansion Wave. C.3 Diamond Shock Wave. Appendix D: Supersonic Air-Intake. D.1 Compression Characteristics of Diffusers. D.2 Air-Intake System. Appendix E: Measurements of Burning Rate and Combustion Wave Structure. Index.

Rocket Grade Hydrogen Peroxide (RGHP) for use in Propulsion and Power Devices - Historical Discussion of Hazards

Abstract: Common historical anecdotal negative characteristics for propellant grade hydrogen peroxide are described and reviewed. Supporting evidence, analysis, historical technical data, recent test data, prior experience, current experience, modern and literature test data are provided to address the issues and perceived concerns and to provide referenced and established scientific data and evidence which can be used to make informed decisions.

Numerical Parametric Analysis of Dual-Bell Nozzle Flows

Abstract: Dual-bell nozzles permit two different operating modes, which provide higher performance than conventional bell nozzles when applied to rocket engines operating from sea level During the low-altitude mode, the flow is separated and the separation line is located near the inflection of the nozzle wall, in a region characterized by a negative value of the wall pressure gradient, as in conventional nozzles, and in which side loads may occur The characteristics of this region in hot full-scale applications are addressed by means of numerical simulations, and scaling laws are sought for cold subscale side-load experiments Moreover, the flow behavior during the transition between the two operating modes is analyzed by time-accurate simulations

On the efficiency of rocket-borne particle detection in the mesosphere

Abstract: Meteoric smoke particles have been proposed as a key player in the formation and evolution of mesospheric phenomena. Despite their apparent importance still very little is known about these particl ...

Conceptual Study of a Rocket-Ramjet Combined-Cycle Engine for an Aerospace Plane

Abstract: Operating conditions of a rocket-ramjet combined-cycle engine for a single-stage-to-orbit aerospace plane were studied. The engine was composed of an ejector-jet mode, a ramjet mode, a scramjet mode, and a rocket mode. Characteristics of the engine operating conditions were studied analytically. The thrust augmentation effect of the ejector-jet mode was found to be small at low subsonic speed and to increase with an increase of the flight Mach number. Study of the effective impulse function clarified that higher specific impulse was preferable in supersonic flight, whereas greater thrust coefficient was preferable in hypersonic flight. The mentioned characteristics were examined by simulation of engine operating in an aerospace plane flight. Transportation of a mass into orbit was compared among several engines with different combinations of thrust and specific impulse. The mass which could be carried into orbit was larger with a ramjet mode of higher specific impulse and with a scramjet mode of greater thrust.

Metals, Energetic Additives, and Special Binders Used in Solid Fuels for Hybrid Rockets

Abstract: R low mass and linear regression rates of solid fuels have been among the major drawbacks of classical hybrid rocket engine technology due to low density, inertness of conventional solid fuels, and the diffusion-controlled combustion process. The thermal degradation in the pyrolysis process of inert polymeric fuels has been considered to be one of the key processes occurring in hybrid rocket engines and solid-fuel ramjet engines [1]. The fuel-surface regression rate generated by this process is a very important design and performance parameter and is strongly affected by the operating conditions and the composition and thermophysical properties of the solid fuel. In addition, fluid dynamic, heat-transfer, and combustion processes in these solid-fuel systems are characterized by complex interactions involving numerous physical phenomena, simultaneously taking place in the combustion chamber and the fuel grain. These complex interactions include solid-fuel pyrolysis; metal vaporization for metallized solid fuels; oxidizer atomization and vaporization; gas-phase species mass diffusion; mixing and combustion between the fuel-rich and oxidizer-rich species; turbulent flow with mass addition; conductive, convective, and radiative energy transfer; and time-varying

Acoustic Instability of the Slab Rocket Motor

Abstract: This article provides rigorous mathematical details leading to an improved formulation of acoustic instability in solid rocket motors. The evaluation of stability growth rate factors is carried out both numerically and asymptotically. Analytical expressions for the stability factors are obtained over a broad spectrum of operating parameters. For all representative rocket motors under investigation, the analytical estimates exhibit an error of 5% or less. Both numerics and asymptotics converge in predicting markedly less stable systems than projected by classic stability theory. The dramatic differences can be ascribed to the dismissal of time-dependent rotational coupling in the previous formulation. The current study unravels the details of six additional growth rate corrections not accounted for previously. These include the rotational flow, mean vorticity, viscosity, pseudo acoustic, pseudo vorticity and unsteady nozzle growth rate factors. The fourth and fifth terms are due to acoustical and vortical interactions with the often neglected pseudopressure. The sixth is due to the energy associated with the unsteady rotational flow exiting the nozzle. This study enables us to isolate the impact of various flow attributes on stability. These involve the motor aspect ratio, surface Mach number, viscous parameter, oscillation mode shape number, and surface admittance. Based on the slab motor geometry, we find that the flow turning correction is cancelled identically by another rotational term not accounted for previously. We also find that the unsteady nozzle damping effect is offset by another source of instability due to the pseudopressure.

An Approach to Improved Credibility of CFD Simulations for Rocket Injector Design

Abstract: Computational fluid dynamics (CFD) has the potential to improve the historical rocket injector design process by simulating the sensitivity of performance and injector-driven thermal environments to. the details of the injector geometry and key operational parameters. Methodical verification and validation efforts on a range of coaxial injector elements have shown the current production CFD capability must be improved in order to quantitatively impact the injector design process.. This paper documents the status of an effort to understand and compare the predictive capabilities and resource requirements of a range of CFD methodologies on a set of model problem injectors. Preliminary results from a steady Reynolds-Average Navier-Stokes (RANS), an unsteady Reynolds-Average Navier Stokes (URANS) and three different Large Eddy Simulation (LES) techniques used to model a single element coaxial injector using gaseous oxygen and gaseous hydrogen propellants are presented. Initial observations are made comparing instantaneous results, corresponding time-averaged and steady-state solutions in the near -injector flow field. Significant differences in the flow fields exist, as expected, and are discussed. An important preliminary result is the identification of a fundamental mixing mechanism, accounted for by URANS and LES, but missing in the steady BANS methodology. Since propellant mixing is the core injector function, this mixing process may prove to have a profound effect on the ability to more correctly simulate injector performance and resulting thermal environments. Issues important to unifying the basis for future comparison such as solution initialization, required run time and grid resolution are addressed.

High-Altitude Plume Simulations for a Solid Propellant Rocket

Abstract: A simulation scheme is proposed for flowfield and radiation analysis of solid rocket exhaust plumes at high altitude. Several recently developed numerical procedures are used to determine properties of the gas and condensed phase Al2O3 particles, and spectrally resolved plume radiation calculations are performed using a Monte Carlo ray trace model. Simulations are run for a representative plume flow at 114 km, and a comparison is made with experimental measurements of UV radiance. A series of parametric studies involving simulations of this same flow are used to evaluate the influence of physical processes and input parameters related to gas-particle interaction, particle radiation, and the presence of soot. I. Introduction n of m the flowfield simulation and radiation analysis of solid rocket exhaust plumes at very high altitudes, a number approximations and simplifying assumptions are typically made due to computational cost, a lack of existing odels, or uncertainty over the influence of various physical phenomena. These flows tend to include a large mass fraction of Al2O3 particles, which can significantly influence bulk flow properties and dominate plume radiative emission through much of the IR, visible, and UV range. Some determination of particle phase characteristics is typically required for useful and accurate simulation results, and may be necessary to assess base heating rates, radiation signatures, surface contamination effects, or other flow properties of interest. As a result, important physical processes and phenomena associated with gas-particle interaction must be recognized and incorporated into simulation procedures. Several potentially important effects have received little attention in the literature, and the significance of coupling between many of these effects still remains an open question. In this paper, we attempt to address the uncertainty in the significance of various effects, and describe a general procedure for the simulation of rarefied plume flows from solid propellant rockets. I

Advanced Rocket Engines

Abstract: SUMMARY Starting with some basics about space transportation systems such as the thrust equation and some mission requirements, the paper explains the underlying physical and technical challenges every rocket engine design engineer faces at the beginning of a project. A brief overview about the main subsystems of a liquid rocket engine such as turbopumps and gas generators is followed by a more detailed description of the thrust chamber assembly, the injector head, the ignition system and the combustion chamber liner which includes the first part of the nozzle and finally the nozzle extension. The technological challenges of these components are presented which result from the severe operating conditions and their current design principles and production technologies. Finally, a series of new concepts and techniques for propellant injection, ignition, combustion chamber liners and nozzles are proposed. The major challenges of new materials and production technologies lay in the still missing detailed knowledge about the behaviour of these materials under operating conditions, material-related crack initiation and propagation laws and reliable life prediction tools.

Investigation of Feed System Coupled Low Frequency Combustion Instabilities in Hybrid Rockets

Abstract: In this paper, the transient combustion theory has been extended to hybrid rockets using liquid oxidizers with feed systems characterized by finite response times. Models for generic hybrid feed systems have been developed and these have been coupled to the combustion chamber dynamics with lag times introduced to model the delays associated with oxidizer vaporization and fuel gasification processes. This study has been limited to a simplified behavior for the combustion chamber that only includes the filling/emptying dynamics. The set of Ordinary Differential Equations that represent the system behavior have been linearized and nondimensionalized. Using the technique of Laplace Transformation, transfer functions for the feed coupled system with and with out flow isolation elements have been developed. The model has been used to investigate the stability behavior of the feed coupled system and to develop stability criteria in terms of the practical operational parameters for liquid fed hybrid motors. The oxidizer vaporization delay is treated as an input parameter which can be adjusted to match the observed oscillation frequency to the model prediction. As a practical application of the model, the estimated vaporization delay can be used to evaluate the atomization characteristics of injectors and pre-combustion chamber designs for hybrid rocket motors.

Application of a Generalized Instability Model to a Longitudinal Mode Combustion Instability

Abstract: An analytic model is used to study combustion insta bilities via the solution of an inhomogeneous pressure wave equation by a modified Galerkin method. This generalized instability model (GIM) is part of a testbed develo ped to study combustion instabilities in liquid rocket engines and augmentors. This model represents the lower fidelity spectrum of the testbed, and is compared with its higher order counterparts to asses its accuracy as well as its limitations. The model is used to study spec ial problems in acoustics, such as ducts with discontinuous changes in flow properties, continuous and discontinuous changes in mean flow Mach numbers, as well as their coupling with unsteady heat addition. The model is then used to determine the longitudinal stability charac teristics of a model rocket combustor and compared with relevant experimental results.

Hypersonic Missile Performance and Sensitivity Analysis

Abstract: The performance of a conceptual, missile-class, hypersonic vehicle was modeled using a 2-degree-of-freedom dynamics model. The vehicle was assumed to be air launched, accelerated to a Mach number of 3 using a solid propellant rocket, and subsequently propelled using a dual-mode scramjet engine to cruise at Mach numbers between 4 and 9. Modeling results show that fuel storage capacity and dynamic pressure have a significant effect on vehicle range and average speed. A perturbation analysis was also performed that ranked the sensitivity of missile rangetosmallchangesin14designparameters.Thekineticenergyefficiency(totalpressureloss)ofthescramjethas the highest effect on performance at high cruise Mach numbers (6.7–9), followed by structural mass, combustion efficiency, and aerodynamics parameters. At low cruise Mach numbers (4–6.7), structural concerns dominate performance, followed by the aerodynamics and scramjet operating parameters.

Compact, high performance swirl combustion rocket engine

Abstract: A rocket engine includes a combustor assembly for carrying out a combustion process of fuel and oxidizer rocket propellants to produce thrust. A swirl generator is positioned within the combustor assembly to produce a turbulent flowfield of the fuel and oxidizer rocket propellants within the combustor assembly.

Technical Background and Challenges of the SpaceLiner Concept

Abstract: At the Space Launcher System Analysis (SART) department of DLR-Cologne, a hypersonic spaceplane for passenger transportation is being investigated. The spaceplane is called the “SpaceLiner”. The vehicle performs its rocket powered, intercontinental flight via a suborbital trajectory. The paper describes the latest developments and improvements on the design of the SpaceLiner. The aerodynamic heating of the vehicle is discussed, and a possible solution for handling the extreme heatloads will be presented. The solution involves an innovative new way of transpiration cooling, using liquid water.

Coning motion stability of wrap around fin rockets

Abstract: Both the asymptotical stability criterion and the bounded stability criterion of the coning motion for wrap around fin (WAF) rockets are proposed through the analysis of coning motion equations, which can be easily used to determine the existence of the coning motion during the rocket design. The correctness of the criterions is verified by mathematical simulation examples of a WAF rocket with different setting angles. It is also found that the setting angle of WAF has great effects on the rolling moment and side moment of the rocket.

Experimental Program to Assess Erosive Burning in Segmented Solid Rocket Motors

Abstract: A 2 -D windowed combustor has been developed for the study of erosive burnin g in composite solid propellants. The modular design permits study of planar and slotted propellant configurations (as are present in segmented rocket motors) with adjustable port to -throat area ratios. The current focus is on a simple planar configurati on without optical access. Head, aft, and mid -motor pressure instrumentation is included in addition to thermocouples designed to sense local flame front approach time at head and aft ends of the grain. To date, the combustor has been fired 12 times and useful data are now being obtained. A ballistic element ballistic model has also been developed and has been used to analyze some of the tests using the classical Lenoir -Robillard erosive burning model. While more analysis is required, reasonable fits of the results are being obtained by suitable adjustment of the Lenoir -Robillard erosive burning parameters.

Characterization and Fate of Gun and Rocket Propellant Residues on Testing and Training Ranges: Interim Report 1

Abstract: : The objectives of the research described in this report are to characterize the deposition and accumulation of propellant residues at the various types of firing points at military firing ranges, develop process descriptors to allow estimation of environmental transport rates of individual energetic chemicals from these residues, and collect lysimeter and groundwater monitoring well samples to experimentally assess off-site transport of residues. Estimates of residue deposition are presented for the firing of 60- and 81-mm mortars and 105-mm howitzers. Experimental results are provided for propellant residue accumulation at antitank rocket, mortar, artillery, and small arms ranges at several installations. Results from soil column experiments on the transport of nitroglycerin, nitroguanidine, and diphenylamine also are presented with resulting transport property estimates. Also, an experiment to assess the deposition of ammonium perchlorate from Mk58 rocket motors is described.

Numerical Analysis of Flow Separation Structures in Rocket Nozzles

Abstract: Numerical simulations of overexpanded rocket nozzle flows show three kinds of flow separation structure: the free shock separation (FSS), the restricted shock separation (RSS), and an intermediate configuration which is a FSS with “inviscid separation” in the nozzle core. The paper presents a number of numerical simulations carried out with dierent nozzle geometries and chamber-to-ambient pressure ratios, with the goal to understand which are the geometries and the range of pressure ratio which lead to each dierent flow separation structure. The results show a good agreement with available experimental data and that RSS takes place only in case of strong radial gradients in the nozzle divergent section. Moreover, it is found that “inviscid separation” is much more likely to occur than RSS.

Combustion Stability Boundaries of the Subscale Rocket Chamber with Impinging Jet Injectors

Abstract: Combustion stability boundaries are investigated experimentally using a subscale chamber for screening the split-triplet impinging jet injectors of a liquid propellant rocket engine. Geometrical dimensions and operating conditions of the subscale chamber are determined using the scaling methods proposed in the previous works. From the experimental tests, two major instability regions are identified by the parameters of combustion-chamber pressure and mixture (oxidizer/fuel) ratio. A key instability mechanism can be explained by the correlation between the characteristic burning or mixing time and the characteristic acoustic time. In each instability region, the dynamic behavior of the flame is. investigated to verify the characteristic lengths of the jet injectors derived from hydrodynamic theory. It is found that pressure oscillations with large amplitude are generated by lifted-off flames. Stability margins are evaluated as a function of the impingement angle and thereby, the optimum angle is identified.

Investigation of Non-Conventional Bio-Derived Fuels for Hybrid Rocket Motors

Abstract: : Non-conventional bio-derived fuels have been evaluated for use in hybrid rocket motors. Tests were conducted at combustion pressures in the range of 100 - 220 psig and thrust levels of 40 - 170 newtons. Beeswax was tested with oxygen as the oxidizer and showed a regression rate at least three times as high as traditional hybrid propellant combinations such as hydroxyl-terminated polybutadiene (HTPB) and liquid oxygen (LOX). This provides the promise of a high thrust hybrid rocket motor using a simple, single port geometry and overcomes the main weakness of traditional hybrid rocket motor propellants, which are low regression rates. Beeswax was also tested with nitrous oxide as an oxidizer, but further testing is needed to attain high enough combustion chamber pressures to achieve stable combustion. Experimental evaluation of the specific impulse for beeswax and oxygen was moderately successful for lab scale testing, but needs further refinement. Analytical studies were performed to evaluate the theoretical performance of non-conventional hybrid rocket motors This analysis indicates beeswax, lard, a mixture of paraffin and lard, and combinations of beeswax and aluminum should all perform better than traditional hybrid rocket propellants considered when burned with oxygen.