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Author: Mathias Ross
Publisher:
ISBN:
Category :
Languages : en
Pages : 0
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Book Description
With efficiencies in conventional rocket designs reaching the limit of theoretical possibility,there has been renewed interest in technologies which may be able to shift the boundaries of efficiency. One such technology is the rotating detonation rocket engine, which has the potential to create highly efficient engines in a small form factor. However, the detonation dynamics and complex flowfields inside the combustion chamber are greatly dependent on geometry; in particular, the downstream nozzle design affects dynamics inside the combustion chamber. In this work, high fidelity large eddy simulations of gaseous methane-oxygen rotating detonation rocket engines are presented for five engine configurations. The first simulation discussed is a validation case from the AIAA model validation in propulsion workshop. A laser model based on the Beer-Lambert law was developed for com- paring simulations with experimental laser absorbance measurements, and used to directly relate the simulation with experimental measurements of temperature, pressure, and CO column density in the exhaust of the engine. The analysis found that the simulation over- predicted pressure and thrust in the engine, as has been the case in other simulations of the engine, but that features in the exhaust flowfield closely matched experimental measurements. Close agreement between simulation and experiment was also seen in the measured CO mole fraction of the exhaust. The effect of adding a converging-diverging nozzle to a rotating detonation rocket engine was explored in the other four simulations, which consider an engine of two different lengths, with and without a constriction. The geometries matched experimental tests previously conducted at the Air Force Research Laboratory, and the operational modes attained in the simulations were found in all cases to directly relate to experimental observations. In the unconstricted geometries, flow in the chamber exceeded Mach 1 in pockets up- stream of the chamber exit. However, geometries with a diverging-converging nozzle directly followed the Mach-area relationship, with supersonic flow existing only in the diverging regions of the nozzle. This suggests a fundamental difference between the flowfield present in RDRE geometries with and without an area constriction, even though the constriction studied was gradual enough that no reflected shocks were observed traveling upstream. The formation enthalpy of the flow was measured inside the chamber for all configurations, and demonstrated that the difference in pressures and detonation structures associated with the chamber area constriction did not result in a significant change in the amount of energy released through combustion. Adding a constriction increased the average pressure of the combustion chamber, which would typically result in increased combustive energy release, but no associated release through combustion was observed. As such, although the use of a converging-diverging nozzle increased overall performance, the induced change in operating mode was detrimental to the extraction of energy from the flow. Changing chamber length was found to have little impact on the operation of an unconstricted rotating detonation rocket engine. However, changing the length of a chamber with a constriction resulted in a change in operating mode, and decrease in the strength of the counter-propagating waves. This suggests that, although unconstricted chamber geometries are likely optimized at short lengths, the length of the chamber is an important parameter to be considered when the engine utilizes a chamber area constriction.
Author: Mathias Ross
Publisher:
ISBN:
Category :
Languages : en
Pages : 0
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Book Description
With efficiencies in conventional rocket designs reaching the limit of theoretical possibility,there has been renewed interest in technologies which may be able to shift the boundaries of efficiency. One such technology is the rotating detonation rocket engine, which has the potential to create highly efficient engines in a small form factor. However, the detonation dynamics and complex flowfields inside the combustion chamber are greatly dependent on geometry; in particular, the downstream nozzle design affects dynamics inside the combustion chamber. In this work, high fidelity large eddy simulations of gaseous methane-oxygen rotating detonation rocket engines are presented for five engine configurations. The first simulation discussed is a validation case from the AIAA model validation in propulsion workshop. A laser model based on the Beer-Lambert law was developed for com- paring simulations with experimental laser absorbance measurements, and used to directly relate the simulation with experimental measurements of temperature, pressure, and CO column density in the exhaust of the engine. The analysis found that the simulation over- predicted pressure and thrust in the engine, as has been the case in other simulations of the engine, but that features in the exhaust flowfield closely matched experimental measurements. Close agreement between simulation and experiment was also seen in the measured CO mole fraction of the exhaust. The effect of adding a converging-diverging nozzle to a rotating detonation rocket engine was explored in the other four simulations, which consider an engine of two different lengths, with and without a constriction. The geometries matched experimental tests previously conducted at the Air Force Research Laboratory, and the operational modes attained in the simulations were found in all cases to directly relate to experimental observations. In the unconstricted geometries, flow in the chamber exceeded Mach 1 in pockets up- stream of the chamber exit. However, geometries with a diverging-converging nozzle directly followed the Mach-area relationship, with supersonic flow existing only in the diverging regions of the nozzle. This suggests a fundamental difference between the flowfield present in RDRE geometries with and without an area constriction, even though the constriction studied was gradual enough that no reflected shocks were observed traveling upstream. The formation enthalpy of the flow was measured inside the chamber for all configurations, and demonstrated that the difference in pressures and detonation structures associated with the chamber area constriction did not result in a significant change in the amount of energy released through combustion. Adding a constriction increased the average pressure of the combustion chamber, which would typically result in increased combustive energy release, but no associated release through combustion was observed. As such, although the use of a converging-diverging nozzle increased overall performance, the induced change in operating mode was detrimental to the extraction of energy from the flow. Changing chamber length was found to have little impact on the operation of an unconstricted rotating detonation rocket engine. However, changing the length of a chamber with a constriction resulted in a change in operating mode, and decrease in the strength of the counter-propagating waves. This suggests that, although unconstricted chamber geometries are likely optimized at short lengths, the length of the chamber is an important parameter to be considered when the engine utilizes a chamber area constriction.
Author: Jiun-Ming Li
Publisher: Springer
ISBN: 3319689061
Category : Technology & Engineering
Languages : en
Pages : 246
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Book Description
This book focuses on the latest developments in detonation engines for aerospace propulsion, with a focus on the rotating detonation engine (RDE). State-of-the-art research contributions are collected from international leading researchers devoted to the pursuit of controllable detonations for practical detonation propulsion. A system-level design of novel detonation engines, performance analysis, and advanced experimental and numerical methods are covered. In addition, the world’s first successful sled demonstration of a rocket rotating detonation engine system and innovations in the development of a kilohertz pulse detonation engine (PDE) system are reported. Readers will obtain, in a straightforward manner, an understanding of the RDE & PDE design, operation and testing approaches, and further specific integration schemes for diverse applications such as rockets for space propulsion and turbojet/ramjet engines for air-breathing propulsion. Detonation Control for Propulsion: Pulse Detonation and Rotating Detonation Engines provides, with its comprehensive coverage from fundamental detonation science to practical research engineering techniques, a wealth of information for scientists in the field of combustion and propulsion. The volume can also serve as a reference text for faculty and graduate students and interested in shock waves, combustion and propulsion.
Author: James Koch
Publisher:
ISBN:
Category : Detonation waves
Languages : en
Pages : 147
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Book Description
The Rotating Detonation Engine (RDE) is a novel rocket combustor configuration that features a periodic, high aspect ratio (length of flowpath versus transverse thickness) combustion chamber designed to promote tangential high-frequency combustion instabilities typical of conventional rockets. Most RDEs are comprised of concentric cylinders whereby the annular gap between the cylinders constitutes the flow domain. The annular gap acts as geometric confinement suitable for robustly promoting the self-steepening of pressure and density gradients caused by combustion. The RDE's steady operation is the saturation of this highly nonlinear self-steepening process: a number of circumferentially traveling detonation waves. The benefits of the RDE include a larger stable operating envelope and potentially higher thermodynamic cycle efficiency over conventional rockets. However, the collective behavior of the detonation waves present in the RDE combustion chamber, while readily observable, is not well understood nor sufficiently characterized, especially with respect to engineering metrics such as thrust or engine stability. This dissertation is a comprehensive experimental, theoretical, and numerical study that aims to link observed the gasdynamic engine behavior to the nonlinear dynamics of the detonation waves. The experimental test campaign features engines of two sizes: a 154-mm flowpath outer diameter (OD) engine and a 76mm OD engine. A sweep of boundary conditions (inlet and outlet pressures) was conducted using the 154-mm RDE at reduced mass flux conditions to establish the engine0́9s gasdynamic operating regimes, namely the attainment of a thermal choke at the exit of the annular duct. Similarly, using the 76-mm RDE at elevated mass flux conditions, the engine0́9s response to the attainment of an axial thermal choke is investigated with respect to changes in total injection area. From both sets of testing, found is that the choked annular duct acts as a boundary condition that fixes the upstream pressure required to steadily deliver a given mass flow rate of propellant of a specified chemical energy potential. By recording the space-time history of the detonation waves with a high speed camera, a diverse set of behavior was recorded and collected during the experimental test campaign. Such behavior includes wave nucleation, destruction, mode-locking of multiple waves, persistent wave modulation, and pulsating plane waves. By drawing upon the well-established fields of nonlinear waves and detonation analog modeling, a Rotating Detonation Engine analog system is proposed. This model system is an adaptation of the Majda detonation analog to a periodic domain with imposed dissipation and propellant regeneration. The dissipative process is constrained to enforce the same global behavior seen in experiments, namely the self-similar pressure operating profiles attained with a thermal choke point at the exit of the engine. Within the reduced-scope of co-rotating detonation waves, the RDE analog system is found to qualitatively reproduce all transients and modes of operation seen in experiments. The propagating waves are classified as autosolitons, or localized structures with offsetting dominant balance physics. Within this context, the dominant balance physics are identified and found to be strongly influenced by input-output energy dynamics and act across several orders of spatial and temporal scales. In this manner, the global multi-scale balance physics give rise to the traveling detonation waves and their associated dynamics - not exclusively the frontal dynamics prescribed by classical detonation theory. Furthermore, the underlying fundamental Hopf bifurcation to time-periodic modulation of the collection of waves is confirmed to exist in the RDE analog. Comparisons between computed Hopf orbits of the model and experimentally-extracted kinematic traces are made showing good qualitative agreement.
Author: National Aeronautics and Space Administration (NASA)
Publisher: Createspace Independent Publishing Platform
ISBN: 9781721604586
Category :
Languages : en
Pages : 28
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Book Description
Pulse detonation engines (PDB) have generated considerable research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional gas turbines and rocket engines. The detonative mode of combustion employed by these devices offers a theoretical thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional engines. However, the unsteady blowdown process intrinsic to all pulse detonation devices has made realistic estimates of the actual propulsive performance of PDES problematic. The recent review article by Kailasanath highlights some of the progress that has been made in comparing the available experimental measurements with analytical and numerical models. Morris, C. I. Marshall Space Flight Center AIAA Paper 2004-0463
Author: MICHIGAN UNIV ANN ARBOR.
Publisher:
ISBN:
Category :
Languages : en
Pages : 1
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Theoretical study of an analytical model for the internal gas dynamics associated with a rotating detonation wave engine are presented. The dif ferential equations resulting from this analysis have been put in a form convenient for solution. Two cases of physical significance have been identified; the case of instantaneous, complete mixing between the burned and unburned propel lants and the case of no mixing. An analytical solution to the equations has been found for the case where the local chamber pressure is equal or greater than the injector pressure and the injector mass flow is zero (blocked injector case). For the complete mixing case a digital computer program has been developed f the IBM 7090 computer which integrates numerically the differential equations. (Author).
Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 32
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Author:
Publisher: AIAA
ISBN: 9781600863875
Category : Detonation waves
Languages : en
Pages : 422
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Author: Joshua Dawson
Publisher:
ISBN:
Category :
Languages : en
Pages : 105
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A novel multi-mode implementation of a pulsed detonation engine, put forth by Wilson et al. [2], consists of four modes; each specifically designed to capitalize on flow features unique to the various flow regimes. This design enables the propulsion system to generate thrust through the entire flow regime. The multi-mode ejectoraugmented pulsed detonation rocket engine operates in mode one during take-o_ conditions through the acceleration to supersonic speeds. Once the mixing chamber internal flow exceeds supersonic speed, the propulsion system transitions to mode two. While operating in mode two, supersonic air is compressed in the mixing chamber by an upstream propagating detonation wave and then exhausted through the convergent-divergent nozzle. Once the velocity of the air flow within the mixing chamber exceeds the Chapman-Jouguet Mach number, the upstream propagating detonation wave no longer has sufficient energy to propagate upstream and consequently the propulsive system shifts to mode three. As a result of the inability of the detonation wave to propagate upstream, a steady oblique shock system is established just upstream of the convergent-divergent nozzle to initiate combustion. And finally, the propulsion system progresses on to mode four operation, consisting purely of a pulsed detonation rocket for high Mach number flight and use in the upper atmosphere as is needed for orbital insertion. Modes three and four appear to be a fairly significant challenge to implement, while the challenge of implementing modes one and two may prove to be a more practical goal in the near future. A vast number of potential applications exist for a propulsion system that would utilize modes one and two, namely a high Mach number hypersonic cruise vehicle. There is particular interest in the dynamics of mode one operation, which is the subject of this study. Several advantages can be obtained by use of this technology. Geometrically, the propulsion system is fairly simple and the rapid combustion process results in an engine cycle which is more efficient compared to its combined-cycle counterparts. The flow path geometry consists of an inlet system, followed just downstream by a mixing chamber where an ejector structure is placed within the flow path. Downstream of the ejector structure is a duct leading to a convergent-divergent nozzle. During mode one operation and within the ejector, products from the detonation of a stoichiometric hydrogen/air mixture are exhausted directly into the surrounding secondary air stream. Mixing then occurs between both the primary and secondary flow streams, at which point the air mass containing the high pressure, high temperature reaction products is convected downstream towards the nozzle. The engine cycle is engineered to a specific number of detonations per second, creating the pulsating characteristic of the primary flow. The pulsing nature of the primary flow serves as a momentum augmentation, enhancing the thrust and specific impulse at low speeds. Consequently, it is necessary to understand the transient mixing process between the primary and secondary flow streams occurring during mode one operation. Using OPENFOAM®, a numerical tool is developed to simulate the dynamics of the turbulent detonation process along with detailed chemistry in order to understand the physics involved with the stream interactions. The computational code has been developed within the framework of OPENFOAM®, an open-source alternative to commercial CFD software. A conservative formulation of the Farve averaged Navier-Stokes equations are used to facilitate programming and numerical stability. Time discretization is accomplished by using the Crank-Nicolson method, achieving second-order convergence in time. Species mass fraction transport equations are implemented and a Seulex ODE solver was used to resolve the system of ordinary differential equations describing the hydrogen-air reaction mechanism detailed in Appendix A. The Seulex ODE solution algorithm is an extrapolation method based on the linearly implicit Euler method with step size control. A second-order total variation diminishing method with a modified Sweby ux limiter was used for space discretization. And finally the use of operator splitting (PISO algorithm, and chemical kinetics) is essential due to the significant differences in characteristic time scales evolving simultaneously in turbulent reactive flow. Capturing the turbulent nature of the combustion process was done using the k-w-SST turbulence model, as formulated by, [1]. Mentor's formulation is well suited to resolve the boundary layer while remaining relatively insensitive to freestream conditions, blending the merits of both the k-w and k-E models. Further developement of the tool is possible, most notably with the Numerical Propulsion System Simulation application. NPSS allows the user to take advantage of a zooming functionality in which high-fidelity models of engine components can be integrated into NPSS models, allowing for a more robust propulsion system simulation. A more comprehensive understanding of the multi-mode ejector-augmented pulsed detonation rocket engine can be achieved with a systematic study of the impact pulsed flow has on thrust production. Although a significant increase in computational requirements, adding nozzle geometry to this study would illuminate any problems associated with pulsed flow through a nozzle. Additionally, a study including nozzle geometry would bring more clarity in regards to the efficiency of the propulsion design.
Author: Lord Kahil Cole
Publisher:
ISBN:
Category :
Languages : en
Pages : 227
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Book Description
A number of promising alternative rocket propulsion concepts have been developed over the past two decades that take advantage of unsteady combustion waves in order to produce thrust. These concepts include the Pulse Detonation Rocket Engine (PDRE), in which repetitive ignition, propagation, and reflection of detonations and shocks can create a high pressure chamber from which gases may be exhausted in a controlled manner. The Pulse Detonation Rocket Induced Magnetohydrodynamic Ejector (PDRIME) is a modification of the basic PDRE concept, developed by Cambier (1998), which has the potential for performance improvements based on magnetohydrodynamic (MHD) thrust augmentation. The PDRIME has the advantage of both low combustion chamber seeding pressure, per the PDRE concept, and efficient energy distribution in the system, per the rocket-induced MHD ejector (RIME) concept of Cole, et al. (1995). In the initial part of this thesis, we explore flow and performance characteristics of different configurations of the PDRIME, assuming quasi-one-dimensional transient flow and global representations of the effects of MHD phenomena on the gas dynamics. By utilizing high-order accurate solvers, we thus are able o investigate the fundamental physical processes associated with the PDRIME and PDRE concepts and identify potentially promising operating regimes. In the second part of this investigation, the detailed coupling of detonations and electric and magnetic fields are explored. First, a one-dimensional spark-ignited detonation with complex reaction kinetics is fully evaluated and the mechanisms for the different instabilities are analyzed. It is found that complex kinetics in addition to sufficient spatial resolution are required to be able to quantify high frequency as well as low frequency detonation instability modes. Armed with this quantitative understanding, we then examine the interaction of a propagating detonation and the applied MHD, both in one-dimensional and two-dimensional transient simulations. The dynamics of the detonation are found to be affected by the application of magnetic and electric fields. We find that the regularity of one-dimensional cesium-seeded detonations can be ignificantly altered by reasonable applied magnetic fields (Bz & le 8T), but that it takes a stronger applied field (Bz> 16T) to significantly alter the cellular structure and detonation velocity of a two-dimensional detonation in the time in which these phenomena were observed. This observation is likely attributed to the additional coupling of the two-dimensional detonation with the transverse waves, which are not captured in the one-dimensional simulations. Future studies involving full ionization kinetics including collisional-radiative processes, will be used to examine these processes in further detail.
Author: Andrew M. Brown
Publisher: Springer Nature
ISBN: 3031182073
Category : Technology & Engineering
Languages : en
Pages : 177
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Book Description
This is the first Structural Dynamics book focused on this indispensable aspect of liquid rocket engine design. This book begins by reviewing basic concepts in Structural Dynamics, including the free and forced response of SDOF and MDOF systems, along with some discussion of how numerical solutions are generated. The book then moves to a discussion of specific applications of these techniques in LREs, progressing from component level (turbomachinery and combustion devices), up through engine system models, and finally to integration with a launch vehicle. Clarifies specific topics including the Campbell and SAFE Diagrams for resonance identification in turbomachinery, the complications of component analysis in the pump side due to a host of complication factors such as acoustic/structure interaction, the "side-loads" fluid/structure interaction problem in overexpanded rocket nozzles, and competing methods for generation overall engine system interface loads. Includes specific examples for illustration while closing with rotordynamic analysis, dynamic data analysis, and vibroacoustics.
