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Author: Aaron J. Glaser
Publisher:
ISBN:
Category :
Languages : en
Pages : 241
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The acoustic signature of a pulse detonation engine was characterized in both the near-field and far-field regimes. Experimental measurements were performed in an anechoic test facility designed for jet noise testing. Both shock strength and speed were mapped as a function of radial distance and direction from the PDE exhaust plane. It was found that the PDE generated pressure field can be reasonably modeled by a theoretical point-source explosion. The effect of several exit nozzle configurations on the PDE acoustic signature was studies. These included various chevron nozzles, a perforated nozzle, and a set of proprietary noise attenuation mufflers.
Author: Aaron J. Glaser
Publisher:
ISBN:
Category :
Languages : en
Pages : 241
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Book Description
The acoustic signature of a pulse detonation engine was characterized in both the near-field and far-field regimes. Experimental measurements were performed in an anechoic test facility designed for jet noise testing. Both shock strength and speed were mapped as a function of radial distance and direction from the PDE exhaust plane. It was found that the PDE generated pressure field can be reasonably modeled by a theoretical point-source explosion. The effect of several exit nozzle configurations on the PDE acoustic signature was studies. These included various chevron nozzles, a perforated nozzle, and a set of proprietary noise attenuation mufflers.
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: Dean P. Petters
Publisher:
ISBN:
Category :
Languages : en
Pages :
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Author: Weipeng Hu
Publisher: Springer Nature
ISBN: 9811974357
Category : Technology & Engineering
Languages : en
Pages : 540
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Book Description
To make the content of the book more systematic, this book mainly briefs some related basic knowledge reported by other monographs and papers about geometric mechanics. The main content of this book is based on the last 20 years’ jobs of the authors. All physical processes can be formulated as the Hamiltonian form with the energy conservation law as well as the symplectic structure if all dissipative effects are ignored. On the one hand, the important status of the Hamiltonian mechanics is emphasized. On the other hand, a higher requirement is proposed for the numerical analysis on the Hamiltonian system, namely the results of the numerical analysis on the Hamiltonian system should reproduce the geometric properties of which, including the first integral, the symplectic structure as well as the energy conservation law.
Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 14
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Author: Rajendran Mohanraj
Publisher:
ISBN:
Category :
Languages : en
Pages :
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Author: James T. Peace
Publisher:
ISBN:
Category : Aerospace engineering
Languages : en
Pages : 312
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Book Description
The pulsed detonation engine (PDE) is an advanced propulsion system that makes use of intermittent detonations to provide thrust. In recent decades, the PDE has been at the center of various propulsion research efforts focused on practical implementation of a reliable detonation-based engine for aerospace propulsion applications. However, many design challenges remain to be solved due to the PDEs unsteady operating characteristics. In particular, the unsteady nature of the thrust chamber flow field inherent to the PDE operation makes the design of nozzles aimed at adequately expanding the burned detonation products especially difficult. In order to address this design challenge, a series of related analytical, numerical, and experimental studies have been conducted, which are focused on investigating the manner in which the PDE propulsive performance is governed by the various gasdynamic processes occurring within the thrust chamber and nozzle flow fields. In this study, three primary PDE configurations are considered. These configurations include fully- and partially-filled PDEs, and PDEs equipped with diverging nozzles. For each configuration, a comprehensive description of the PDE flow field is provided, whereby details concerning the evolution and interaction of various gasdynamic waves and discontinuities are discussed. Additionally, the dominant gasdynamic processes within the thrust chamber and nozzle flow fields are identified, as these processes must be appropriately modeled in order to accurately evaluate the propulsive performance.The collision of a detonation wave with a contact surface separating detonable and non-combustible mixtures is a fundamental gasdynamic interaction process that takes place every cycle in the cyclic operation of the PDE. This interaction can drastically influence the evolving thrust chamber flow field and the subsequent propulsive performance metrics. To improve its understanding, this gasdynamic interaction is investigated analytically in order to predict the resulting transmitted shock wave properties, and the necessary conditions for a shock, Mach, or rarefaction wave to reflect at the contact surface. Concurrently, this interaction is investigated experimentally with the use of a detonation-driven shock tube. The analytical and experimental results indicate that the transmitted shock can either be amplied or attenuated depending on the reflection type at the contact surface, and the ratio of the acoustic impedance across the interface. A quasi-one-dimensional method of characteristics (MOC) model is developed to evaluate the single-cycle gasdynamic flow field and associated propulsive performance of general PDE configurations. The model incorporates the current detonation-contact surface interaction results in order to accurately treat the one-dimensional collision of a detonation wave with a contact discontinuity. Additionally, the MOC model is developed using a simplified unit process approach with an explicit inverse time marching algorithm in order to readily construct the complex thrust chamber flow field along a predefined grid. A thorough validation of the model is presented over a broad range of operating conditions with existing higher-fidelity numerical and experimental performance data for fully- and partially-filled PDEs, and PDEs equipped with diverging nozzles. This includes PDEs operating with a variety of detonable fuels, non-combustible inert mixtures, ll fractions, blowdown pressure ratios, and nozzle expansion area ratios. Lastly, a detailed discussion of the model limitations is provided, and particular operating conditions that lead to a breakdown of the assumptions used in the development of the model are addressed. A simplified analytical model is developed based on control volume analysis for evaluating the primary performance metrics of a general fully-filled PDE. The MOC model is used to justify and establish a simplified thrust relation based solely on the ow properties at the exit plane of a fully-filled PDE. A detailed analytical description of the thrust chamber flow field is provided, from which an analytical piece wise expression for thrust is derived based on the exit plane pressure history. This expression is then used to evaluate the specific impulse, total impulse, and time-averaged thrust of a fully-filled PDE. This simplified model is validated against the current MOC model and higher-fidelity numerical and experimental performance data for a variety of detonable fuels, equivalence ratios, and blowdown pressure ratios.Using the current MOC model, the single-cycle propulsive performance of partially-filled PDEs is investigated. The results of the detonation-contact surface interaction study are used to tailor the acoustic impedance of the non-combustible mixture at a fixed fill fraction in order to demonstrate the sensitivity of the thrust chamber flow field to the non-combustible acoustic impedance. Subsequently, the detonable fill fraction and noncombustible acoustic impedance are varied simultaneously in order to characterize the general partially-lled PDE performance. The partial-filling performance benefit is also investigated by varying the initial pressure and temperature of the non-combustible mixture in order to highlight the advantage of using a cold purge gas during operation, and disadvantage of operating in sub-atmospheric environments. It is demonstrated that the partially-filled specific impulse performance results generated with the MOC model from these various parametric investigations are successfully modeled using a previously developed scaling law, whereby this scaling law is extended in the current work to partially-filled total impulse and time-averaged thrust.Similarly, the single-cycle propulsive performance of PDEs with diverging nozzles is examined. A parametric investigation is conducted to characterize the combined effects of nozzle expansion area and blowdown pressure ratios on the resulting thrust chamber and nozzle flow fields. Detailed discussion of the transient nozzle flow field is provided in order to emphasize the influence of non-combustible acoustic impedance on the partial-fill effect in diverging nozzles. Moreover, a comparative study is used to demonstrate the performance advantages of a diverging nozzle in sub-atmospheric environments compared to a straight extension nozzle. Lastly, a detailed parametric investigation is conducted by simultaneously varying the nozzle length, expansion area ratio, and blowdown pressure ratio in order to determine the optimum nozzle performance characteristics. An analytical model is formulated to predict the strength and motion of a transmitted shock wave through a general contour diverging nozzle for PDEs. The model is derived on the basis of a two-equation approximation of the generalized CCW (Chester Chisnell Whitham) theory for treating general shock dynamics in non-uniform channels. A major feature of the two-equation model is the ability to incorporate non-uniformity in the flow immediately following the shock wave, which turns out to be essential for describing the transmitted shock dynamics in PDE nozzles. This model is then used to demonstrate the effects of thrust chamber length on the magnitude of ow non-uniformity behind the transmitted shock entering the nozzle, and how drastically this can influence the nature of shock attenuation within the nozzle. Further, the shock dynamics model is used in conjunction with the MOC model to demonstrate how different nozzle wall curvature influences the PDE propulsive performance, due to the changes in transmitted shock attenuation and gasdynamic over-expansion in the nozzle flow field during the nozzle starting process.
Author: National Aeronautics and Space Administration (NASA)
Publisher: Createspace Independent Publishing Platform
ISBN: 9781721570782
Category :
Languages : en
Pages : 34
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This paper presents results for a single-pulse detonation tube wherein the effects of high temperature dissociation and the subsequent recombination influence the sensible heat release available for providing propulsive thrust. The study involved the use of ethylene and air at equivalence ratios of 0.7 and 1.0. The real gas effects on the sensible heat release were found to be significantly large so as to have an impact on the thrust, impulse and fuel consumption of a PDE. Povinelli, Louis A. and Yungster, Shaye Glenn Research Center NASA/TM-2003-212211, E-13819, NAS 1.15:212211, AIAA Paper 2003-0712, ICOMP-2003-02
Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 42
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The performance analysis of the supersonic pulse-detonation-jet engine system represents some revision of earlier results, including new ideas of scavenging flow. Curves of the drag coefficient of a typical supersonic long- range missile showed that four 36-in.-diam engines have enough reserve power to propel the missile through sonic flight velocities and up to a flight Mach number of 2.80. A unit small enough to be mounted on the blade tip of a helicopter was analyzed. The unit has a maximum diameter of 8.25 in., with combustion tubes 6 in. long and ranging in diameter from 0.60 to 0.25 in. The total weight of 1 unit would be about 35 lb. Combustion would be achieved in the unit by surface contact with the hot walls of the ceramic tubes, and it would proceed radially inward. A unit of this type would produce a thrust of 110 lb at a maximum temperature of 2000 deg F, and would have a specific fuel consumption of 1.65 lb/hr/lb of thrust.
Author: J. Kasahara
Publisher:
ISBN:
Category :
Languages : en
Pages :
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