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Sample records for airbreathing pulse detonation

  1. Airbreathing Pulse Detonation Engine Performance

    NASA Technical Reports Server (NTRS)

    Povinelli, Louis A.; Yungster, Shaye

    2002-01-01

    This paper presents performance results for pulse detonation engines (PDE) taking into account the effects of dissociation and recombination. The amount of sensible heat recovered through recombination in the PDE chamber and exhaust process was found to be significant. These results have an impact on the specific thrust, impulse and fuel consumption of the PDE.

  2. Airbreathing Pulse Detonation Engine Performance

    NASA Technical Reports Server (NTRS)

    Povinelli, Louis A.; Yungster, Shaye

    2002-01-01

    This paper presents performance results for pulse detonation engines taking into account the effects of dissociation and recombination. The amount of sensible heat recovered through recombination in the PDE chamber and exhaust process was found to be significant. These results have an impact on the specific thrust, impulse and fuel consumption of the PDE.

  3. Analysis of Laser-Generated Impulse In An Airbreathing Pulsed Detonation Engine: Part 1

    NASA Astrophysics Data System (ADS)

    Richard, Jacques C.; Myrabo, Leik N.

    2005-04-01

    An investigation is performed on an airbreathing laser propulsion (LP) system designed to propel a 1.4 m diameter, 120-kg (dry mass) vehicle called the Lightcraft Technology Demonstrator (LTD) into low Earth orbit, along with its opto-electronics payload. The LTD concept led directly to the model ♯200 lightcraft — recently demonstrated in laboratory and flight experiments at White Sands Missile Range, NM at the High Energy Laser Systems Test Facility (HELSTF), using the 10-kW PLVTS CO2 laser. The pulsed detonation wave engine (PDE) employs repetitively ignited, laser-supported detonation (LSD) waves to develop thrust by expanding high pressure blast waves over an annular, interior shroud surface. Numerical simulation of thruster impulse is accomplished with a 1-D cylindrical model of blast waves propagating radially outward from a laser-generated `line-source' of high temperature, high pressure air. External airflow over the LTD structure is also analyzed to predict basic engine/vehicle drag characteristics, including inlet total pressure recovery, and captured air mass flow rate — all projected vs. flight Mach number and altitude.

  4. Experimental Research on Induction Systems of an Air-breathing Valveless Pulse Detonation Engine

    NASA Astrophysics Data System (ADS)

    Wang, Zhi-wu; Chen, Xinggu; Zheng, Long-xi; Peng, Changxin; Yan, Chuan-jun

    2012-06-01

    An air-breathing valveless PDE model was designed and manufactured, which was made up of subsonic inlet, mixing chamber, ignition chamber, detonation chamber. The total pressure recovery coefficient, flux coefficient and intake resistance with six different induction systems were measured by a semi free subsonic flow field. The proof-of-principle experiments of PDE model with different induction systems were all successfully carried out, by using liquid gasoline-air mixture with low-energy system (total stored energy less than 50 mJ). The measured detonation wave pressure ratio was very close to that of C-J detonation. The air-breathing PDE model was easy to initiate and worked in good condition. The deflagration to detonation transition (DDT) and operation frequency effect on pressure traces were also investigated by experiments. The results indicated the oscillation of pressure peak at P6 enhanced with the operation frequency increased. DDT accomplished before P6 and the DDT distance was about 0.9 m (from the ignitor).

  5. Thermodynamic Cycle and CFD Analyses for Hydrogen Fueled Air-breathing Pulse Detonation Engines

    NASA Technical Reports Server (NTRS)

    Povinelli, Louis A.; Yungster, Shaye

    2002-01-01

    This paper presents the results of a thermodynamic cycle analysis of a pulse detonation engine (PDE) using a hydrogen-air mixture at static conditions. The cycle performance results, namely the specific thrust, fuel consumption and impulse are compared to a single cycle CFD analysis for a detonation tube which considers finite rate chemistry. The differences in the impulse values were indicative of the additional performance potential attainable in a PDE.

  6. Role of Air-Breathing Pulse Detonation Engines in High Speed Propulsion

    NASA Technical Reports Server (NTRS)

    Povinelli, Louis A.; Lee, Jin-Ho; Anderberg, Michael O.

    2001-01-01

    In this paper, the effect of flight Mach number on the relative performance of pulse detonation engines and gas turbine engines is investigated. The effect of ram and mechanical compression on combustion inlet temperature and the subsequent sensible heat release is determined. Comparison of specific thrust, fuel consumption and impulse for the two engines show the relative benefits over the Mach number range.

  7. Optimal Area Profiles for Ideal Single Nozzle Air-Breathing Pulse Detonation Engines

    NASA Technical Reports Server (NTRS)

    Paxson, Daniel E.

    2003-01-01

    The effects of cross-sectional area variation on idealized Pulse Detonation Engine performance are examined numerically. A quasi-one-dimensional, reacting, numerical code is used as the kernel of an algorithm that iteratively determines the correct sequencing of inlet air, inlet fuel, detonation initiation, and cycle time to achieve a limit cycle with specified fuel fraction, and volumetric purge fraction. The algorithm is exercised on a tube with a cross sectional area profile containing two degrees of freedom: overall exit-to-inlet area ratio, and the distance along the tube at which continuous transition from inlet to exit area begins. These two parameters are varied over three flight conditions (defined by inlet total temperature, inlet total pressure and ambient static pressure) and the performance is compared to a straight tube. It is shown that compared to straight tubes, increases of 20 to 35 percent in specific impulse and specific thrust are obtained with tubes of relatively modest area change. The iterative algorithm is described, and its limitations are noted and discussed. Optimized results are presented showing performance measurements, wave diagrams, and area profiles. Suggestions for future investigation are also discussed.

  8. Pulse Detonation Engine Modeled

    NASA Technical Reports Server (NTRS)

    Paxson, Daniel E.

    2001-01-01

    Pulse Detonation Engine Technology is currently being investigated at Glenn for both airbreathing and rocket propulsion applications. The potential for both mechanical simplicity and high efficiency due to the inherent near-constant-volume combustion process, may make Pulse Detonation Engines (PDE's) well suited for a number of mission profiles. Assessment of PDE cycles requires a simulation capability that is both fast and accurate. It should capture the essential physics of the system, yet run at speeds that allow parametric analysis. A quasi-one-dimensional, computational-fluid-dynamics-based simulation has been developed that may meet these requirements. The Euler equations of mass, momentum, and energy have been used along with a single reactive species transport equation, and submodels to account for dominant loss mechanisms (e.g., viscous losses, heat transfer, and valving) to successfully simulate PDE cycles. A high-resolution numerical integration scheme was chosen to capture the discontinuities associated with detonation, and robust boundary condition procedures were incorporated to accommodate flow reversals that may arise during a given cycle. The accompanying graphs compare experimentally measured and computed performance over a range of operating conditions for a particular PDE. Experimental data were supplied by Fred Schauer and Jeff Stutrud from the Air Force Research Laboratory at Wright-Patterson AFB and by Royce Bradley from Innovative Scientific Solutions, Inc. The left graph shows thrust and specific impulse, Isp, as functions of equivalence ratio for a PDE cycle in which the tube is completely filled with a detonable hydrogen/air mixture. The right graph shows thrust and specific impulse as functions of the fraction of the tube that is filled with a stoichiometric mixture of hydrogen and air. For both figures, the operating frequency was 16 Hz. The agreement between measured and computed values is quite good, both in terms of trend and

  9. Pulse Detonation Engines for High Speed Flight

    NASA Technical Reports Server (NTRS)

    Povinelli, Louis A.

    2002-01-01

    Revolutionary concepts in propulsion are required in order to achieve high-speed cruise capability in the atmosphere and for low cost reliable systems for earth to orbit missions. One of the advanced concepts under study is the air-breathing pulse detonation engine. Additional work remains in order to establish the role and performance of a PDE in flight applications, either as a stand-alone device or as part of a combined cycle system. In this paper, we shall offer a few remarks on some of these remaining issues, i.e., combined cycle systems, nozzles and exhaust systems and thrust per unit frontal area limitations. Currently, an intensive experimental and numerical effort is underway in order to quantify the propulsion performance characteristics of this device. In this paper, we shall highlight our recent efforts to elucidate the propulsion potential of pulse detonation engines and their possible application to high-speed or hypersonic systems.

  10. Pulse detonation MHD power

    SciTech Connect

    Litchford, R.J.; Thompson, B.R.; Lineberry, J.T.

    1998-07-01

    A series of laboratory scale experiments were conducted to investigate the basic engineering performance characteristics of a pulse detonation driven magnetohydrodynamic electric power generator. In these experiments, stoichiometric oxy-acetylene mixtures seeded with a cesium-hydroxide/ methanol spray were detonated at atmospheric pressure in a 1 m long tube having an inside diameter of 2.54 cm. Experiments with a plasma diagnostic channel attached to the end of the tube confirmed the attainment of detonation conditions (p{sub 2}/p{sub 1} {approximately} 34 and D {approximately} 2400 m/s) and enabled the measurement of current density ({approximately} 2 A/cm{sup 2}) and electrical conductivity ({approximately} 6 mho/m) behind the detonation wave front. In a second set of experiments, a 30 cm long continuous electrode Faraday channel having a height of 2.54 cm and a width of 2.0 cm was attached to the end of the tube using an area transition duct. The Faraday channel was placed inside a permanent magnet assembly having a nominal magnetic induction of 0.6 Tesla, and the electrodes were connected to an active loading circuit in order to characterize power extraction dependence on load impedance while also simulating higher effective magnetic induction. In these single-shot experiments, the near-electrode potential drop was found to consume approximately 60% of the effective u x B induced potential. For B = 0.6 Tesla, the authors obtained a peak open circuit voltage of V{sub O}C = Bh {approximately} 10 volts implying an effective burned gas velocity relative to the tube of {approximately} 660 m/s which may be compared with the theoretical equilibrium value for the idealized case (1100 m/s). The experiments indicated peak power extraction at a load impedance between 5 and 10 Ohms. The measured peak electrical energy density ranged from 10 to 10{sup 3} J/m{sup 3} when the effective magnetic induction was varied from 0.6 to 4.2 Tesla. These results

  11. Some perspectives on pulse detonation propulsion systems

    NASA Astrophysics Data System (ADS)

    Lu, F. K.; Wilson, D. R.

    Pulse detonation engines and rockets (PDE/Rs) can potentially revolutionize air breathing and rocket propulsion [1-6]. While the PDE concept is over five decades old, it has recently enjoyed renewed interest, due mostly to theoretical and computational studies indicating high cycle efficiencies. When modeled by a constant volume, Humphrey cycle, the detonation engine is found to be superior to that of existing constant pressure, Brayton cycles, with claims of as much as 10-40% improvement in specific impulse [4,7-9]. The constant volume process is derived from the Zeldovich-von Neumann-Döring (ZND) model of the detonation wave as a high strength shock wave, followed by a region of chemical reaction and a subsequent isentropic rarefaction. Amongst other advantages of the PDE is simplicity, where the PDE is easy to manufacture and requires few moving parts, with the possibility of eliminating high-pressure pumps in rocket applications, or reducing turbomachinery stages in air-breathing propulsion systems.

  12. Pulse detonation engines and components thereof

    NASA Technical Reports Server (NTRS)

    Tangirala, Venkat Eswarlu (Inventor); Rasheed, Adam (Inventor); Vandervort, Christian Lee (Inventor); Dean, Anthony John (Inventor)

    2009-01-01

    A pulse detonation engine comprises a primary air inlet; a primary air plenum located in fluid communication with the primary air inlet; a secondary air inlet; a secondary air plenum located in fluid communication with the secondary air inlet, wherein the secondary air plenum is substantially isolated from the primary air plenum; a pulse detonation combustor comprising a pulse detonation chamber, wherein the pulse detonation chamber is located downstream of and in fluid communication with the primary air plenum; a coaxial liner surrounding the pulse detonation combustor defining a cooling plenum, wherein the cooling plenum is in fluid communication with the secondary air plenum; an axial turbine assembly located downstream of and in fluid communication with the pulse detonation combustor and the cooling plenum; and a housing encasing the primary air plenum, the secondary air plenum, the pulse detonation combustor, the coaxial liner, and the axial turbine assembly.

  13. Dissociation and Recombination Effects on the Performance of Pulse Detonation Engines

    NASA Technical Reports Server (NTRS)

    Povinelli, Louis A.

    2003-01-01

    This paper summarizes major theoretical results for pulse detonation engine performance taking into account real gas chemistry, as well as significant performance differences resulting from the presence of ram and compression heating. An unsteady CFD analysis, as well as a thermodynamic cycle analysis, was conducted in order to determine the actual and the ideal performance for an air-breathing pulse detonation engine (PDE) using either a hydrogen-air or ethylene-air mixture over a flight Mach number range from 0 to 4. The results clearly elucidate the competitive regime of PDE application relative to ramjets and gas turbines.

  14. Experimental investigation of a unique airbreathing pulsed laser propulsion concept

    NASA Technical Reports Server (NTRS)

    Myrabo, L. N.; Nagamatsu, H. T.; Manka, C.; Lyons, P. W.; Jones, R. A.

    1991-01-01

    Investigations were conducted into unique methods of converting pulsed laser energy into propulsive thrust across a flat impulse surface under atmospheric conditions. The propulsion experiments were performed with a 1-micron neodymium-glass laser at the Space Plasma Branch of the Naval Research Laboratory. Laser-induced impulse was measured dynamically by ballistic pendulums and statically using piezoelectric pressure transducers on a stationary impulse surface. The principal goal was to explore methods for increasing the impulse coupling performance of airbreathing laser-propulsion engines. A magnetohydrodynamic thrust augmentation effect was discovered when a tesla-level magnetic field was applied perpendicular to the impulse surface. The impulse coupling coefficient performance doubled and continued to improve with increasing laser-pulse energies. The resultant performance of 180 to 200 N-s/MJ was found to be comparable to that of the earliest afterburning turbojets.

  15. Pulse detonation assembly and hybrid engine

    NASA Technical Reports Server (NTRS)

    Rasheed, Adam (Inventor); Dean, Anthony John (Inventor); Vandervort, Christian Lee (Inventor)

    2010-01-01

    A pulse detonation (PD) assembly includes a number of PD chambers adapted to expel respective detonation product streams and a number of barriers disposed between respective pairs of PD chambers. The barriers define, at least in part, a number of sectors that contain at least one PD chamber. A hybrid engine includes a number of PD chambers and barriers. The hybrid engine further includes a turbine assembly having at least one turbine stage, being in flow communication with the PD chambers and being configured to be at least partially driven by the detonation product streams. A segmented hybrid engine includes a number of PD chambers and segments configured to receive and direct the detonation product streams from respective PD chambers. The segmented hybrid engine further includes a turbine assembly configured to be at least partially driven by the detonation product streams.

  16. Pulse Detonation Engine Test Bed Developed

    NASA Technical Reports Server (NTRS)

    Breisacher, Kevin J.

    2002-01-01

    A detonation is a supersonic combustion wave. A Pulse Detonation Engine (PDE) repetitively creates a series of detonation waves to take advantage of rapid burning and high peak pressures to efficiently produce thrust. NASA Glenn Research Center's Combustion Branch has developed a PDE test bed that can reproduce the operating conditions that might be encountered in an actual engine. It allows the rapid and cost-efficient evaluation of the technical issues and technologies associated with these engines. The test bed is modular in design. It consists of various length sections of both 2- and 2.6- in. internal-diameter combustor tubes. These tubes can be bolted together to create a variety of combustor configurations. A series of bosses allow instrumentation to be inserted on the tubes. Dynamic pressure sensors and heat flux gauges have been used to characterize the performance of the test bed. The PDE test bed is designed to utilize an existing calorimeter (for heat load measurement) and windowed (for optical access) combustor sections. It uses hydrogen as the fuel, and oxygen and nitrogen are mixed to simulate air. An electronic controller is used to open the hydrogen and air valves (or a continuous flow of air is used) and to fire the spark at the appropriate times. Scheduled tests on the test bed include an evaluation of the pumping ability of the train of detonation waves for use in an ejector and an evaluation of the pollutants formed in a PDE combustor. Glenn's Combustion Branch uses the National Combustor Code (NCC) to perform numerical analyses of PDE's as well as to evaluate alternative detonative combustion devices. Pulse Detonation Engine testbed.

  17. Pulse Detonation Rocket Magnetohydrodynamic Power Experiment

    NASA Technical Reports Server (NTRS)

    Litchford, R. J.; Jones, J. E.; Dobson, C. C.; Cole, J. W.; Thompson, B. R.; Plemmons, D. H.; Turner, M. W.

    2003-01-01

    The production of onboard electrical power by pulse detonation engines is problematic in that they generate no shaft power; however, pulse detonation driven magnetohydrodynamic (MHD) power generation represents one intriguing possibility for attaining self-sustained engine operation and generating large quantities of burst power for onboard electrical systems. To examine this possibility further, a simple heat-sink apparatus was developed for experimentally investigating pulse detonation driven MHD generator concepts. The hydrogen oxygen fired driver was a 90 cm long stainless steel tube having a 4.5 cm square internal cross section and a short Schelkin spiral near the head end to promote rapid formation of a detonation wave. The tube was intermittently filled to atmospheric pressure and seeded with a CsOH/methanol prior to ignition by electrical spark. The driver exhausted through an aluminum nozzle having an area contraction ratio of A*/A(sub zeta) = 1/10 and an area expansion ratio of A(sub zeta)/A* = 3.2 (as limited by available magnet bore size). The nozzle exhausted through a 24-electrode segmented Faraday channel (30.5 cm active length), which was inserted into a 0.6 T permanent magnet assembly. Initial experiments verified proper drive operation with and without the nozzle attachment, and head end pressure and time resolved thrust measurements were acquired. The exhaust jet from the nozzle was interrogated using a polychromatic microwave interferometer yielding an electron number density on the order of 10(exp 12)/cm at the generator entrance. In this case, MHD power generation experiments suffered from severe near-electrode voltage drops and low MHD interaction; i.e., low flow velocity, due to an inherent physical constraint on expansion with the available magnet. Increased scaling, improved seeding techniques, higher magnetic fields, and higher expansion ratios are expected to greatly improve performance.

  18. Plasma Sensor Measurements in Pulse Detonation Engines

    NASA Astrophysics Data System (ADS)

    Matlis, Eric; Marshall, Curtis; Corke, Thomas; Gogineni, Sivaram

    2014-11-01

    Measurements have been conducted in a pulse detonation and rotating detonation engine using a newly developed plasma sensor. This sensor relies on the novel approach of using an ac-driven, weakly-ionized electrical discharge as the main sensing element. The advantages of this approach include a native high bandwidth of 1 MHz without the need for electronic frequency compensation, a dual-mode capability that provides sensitivity to multiple flow parameters, including velocity, pressure, temperature, and gas-species, and a simple and robust design making it very cost effective. The sensor design is installation-compatible with conventional sensors commonly used in gas-turbine research such as the Kulite dynamic pressure sensor while providing much better longevity. Developmental work was performed in high temperature facilities that are relevant to the propulsion and high-speed research community. This includes tests performed in a J85 augmentor at full afterburner and pulse-detonation engines at the University of Cincinnati (UC) at temperatures approaching 2760°C (5000°F).

  19. Future Modeling Needs in Pulse Detonation Rocket Engine Design

    NASA Technical Reports Server (NTRS)

    Meade, Brian; Talley, Doug; Mueller, Donn; Tew, Dave; Guidos, Mike; Seymour, Dave

    2001-01-01

    This paper presents a performance model rocket engine design that takes advantage of pulse detonation to generate thrust. The contents include: 1) Introduction to the Pulse Detonation Rocket Engine (PDRE); 2) PDRE modeling issues and options; 3) Discussion of the PDRE Performance Workshop held at Marshall Space Flight Center; and 4) Identify needs involving an open performance model for Pulse Detonation Rocket Engines. This paper is in viewgraph form.

  20. Integrated Pulse Detonation Propulsion and Magnetohydrodynamic Power

    NASA Technical Reports Server (NTRS)

    Litchford, R. J.; Lyles, Garry M. (Technical Monitor)

    2001-01-01

    The prospects for realizing an integrated pulse detonation propulsion and magnetohydrodynamic (MHD) power system are examined. First, energy requirements for direct detonation initiation of various fuel-oxygen and fuel-air mixtures are deduced from available experimental data and theoretical models. Second, the pumping power requirements for effective chamber scavenging are examined through the introduction of a scavenging ratio parameter and a scavenging efficiency parameter. A series of laboratory experiments were carried out to investigate the basic engineering performance characteristics of a pulse detonation-driven MHD electric power generator. In these experiments, stoichiometric oxy-acetylene mixtures seeded with a cesium hydroxide/methanol spray were detonated at atmospheric pressure in a 1-m-long tube having an i.d. of 2.54 cm. Experiments with a plasma diagnostic channel attached to the end of the tube confirmed the attainment of detonation conditions (p(sub 2)/p(sub 1) approx. 34 and D approx. 2,400 m/sec) and enabled the direct measurement of current density and electrical conductivity (=6 S/m) behind the detonation wave front. In a second set of experiments, a 30-cm-long continuous electrode Faraday channel, having a height of 2.54 cm and a width of 2 cm, was attached to the end of the tube using an area transition duct. The Faraday channel was inserted in applied magnetic fields of 0.6 and 0.95 T. and the electrodes were connected to an active loading circuit to characterize power extraction dependence on load impedance while also simulating higher effective magnetic induction. The experiments indicated peak power extraction at a load impedance between 5 and 10 Ohm. The measured power density was in reasonable agreement with a simple electrodynamic model incorporating a correction for near-electrode potential losses. The time-resolved thrust characteristics of the system were also measured, and it was found that the MHD interaction exerted a

  1. Integrated Pulse Detonation Propulsion and Magnetohydrodynamic Power

    NASA Technical Reports Server (NTRS)

    Litchford, Ron J.

    2001-01-01

    The prospects for realizing an integrated pulse detonation propulsion and magnetohydrodynamic (MHD) power system are examined. First, energy requirements for direct detonation initiation of various fuel-oxygen and fuel-air mixtures are deduced from available experimental data and theoretical models. Second, the pumping power requirements for effective chamber scavenging are examined through the introduction of a scavenging ratio parameter and a scavenging efficiency parameter. A series of laboratory experiments were carried out to investigate the basic engineering performance characteristics of a pulse detonation-driven MHD electric power generator. In these experiments, stoichiometric oxy-acetylene mixtures seeded with a cesium hydroxide/methanol spray were detonated at atmospheric pressure in a 1-m-long tube having an i.d. of 2.54 cm. Experiments with a plasma diagnostic channel attached to the end of the tube confirmed the attainment of detonation conditions (p2/p1 approximately 34 and D approximately 2,400 m/sec) and enabled the direct measurement of current density and electrical conductivity (approximately = 6 S/m) behind the detonation wave front, In a second set of experiments, a 30-cm-long continuous electrode Faraday channel, having a height of 2.54 cm and a width of 2 cm, was attached to the end of the tube using an area transition duct. The Faraday channel was inserted in applied magnetic fields of 0.6 and 0.95 T, and the electrodes were connected to an active loading circuit to characterize power extraction dependence on load impedance while also simulating higher effective magnetic induction. The experiments indicated peak power extraction at a load impedance between 5 and 10 Omega. The measured power density was in reasonable agreement with a simple electrodynamic model incorporating a correction for near-electrode potential losses. The time-resolved thrust characteristics of the system were also measured, and it was found that the NM interaction

  2. Frequency content of current pulses in slapper detonator bridges

    SciTech Connect

    Carpenter, K H

    2006-12-18

    DFT amplitudes are obtained for digital current pulse files. The frequency content of slapper detonator bridge current pulses is obtained. The frequencies are confined well within the passband of the CVR used to sample them.

  3. Modeling and laser-based sensing of pulsed detonation engines

    NASA Astrophysics Data System (ADS)

    Barbour, Ethan A.

    This work is concerned with two major aspects of pulse detonation engines (PDE) research: modeling and laser-based sensing. The modeling addresses both ideal and real considerations relevant to PDE design. First, an ideal nozzle model is developed which provides a tool for choosing area ratios for fixed-geometry converging, diverging, or converging-diverging nozzles. Next, losses associated with finite-rate chemistry are investigated. It was found that PDEs can experience up to 10% reduction in specific impulse from this effect if 02 is used as the oxidizer, whereas the losses are negligible for air-breathing applications. Next, heat transfer and friction losses were investigated and found to be greater than the losses from simple straight-tube PDEs. These losses are most pronounced (˜15%) when converging nozzles are used. The second portion of this work focuses on laser-based absorption sensing for PDEs. The mid-infrared was chosen as the best way to address the challenges of signal-to-noise ratio, sensitivity, robustness, and sensor bandwidth. A water vapor sensor was developed and applied to the PDE at the Naval Postgraduate School. This sensor provided improvements in temperature accuracy, and it revealed that water (generated by the vitiator) inhibited performance of the engine. Next, a JP-10 absorption sensor was developed and applied to the same engine. This sensor provided thermometry data at a higher temporal resolution than the water sensor. The sensor also provided crucial information on equivalence ratio and fuel arrival time which enabled the engine to be successfully operated on JP-10 and air for the first time.

  4. Pulse Detonation Rocket MHD Power Experiment

    NASA Technical Reports Server (NTRS)

    Litchford, Ron J.; Cook, Stephen (Technical Monitor)

    2002-01-01

    A pulse detonation research engine (MSFC (Marshall Space Flight Center) Model PDRE (Pulse Detonation Rocket Engine) G-2) has been developed for the purpose of examining integrated propulsion and magnetohydrodynamic power generation applications. The engine is based on a rectangular cross-section tube coupled to a converging-diverging nozzle, which is in turn attached to a segmented Faraday channel. As part of the shakedown testing activity, the pressure wave was interrogated along the length of the engine while running on hydrogen/oxygen propellants. Rapid transition to detonation wave propagation was insured through the use of a short Schelkin spiral near the head of the engine. The measured detonation wave velocities were in excess of 2500 m/s in agreement with the theoretical C-J velocity. The engine was first tested in a straight tube configuration without a nozzle, and the time resolved thrust was measured simultaneously with the head-end pressure. Similar measurements were made with the converging-diverging nozzle attached. The time correlation of the thrust and head-end pressure data was found to be excellent. The major purpose of the converging-diverging nozzle was to configure the engine for driving an MHD generator for the direct production of electrical power. Additional tests were therefore necessary in which seed (cesium-hydroxide dissolved in methanol) was directly injected into the engine as a spray. The exhaust plume was then interrogated with a microwave interferometer in an attempt to characterize the plasma conditions, and emission spectroscopy measurements were also acquired. Data reduction efforts indicate that the plasma exhaust is very highly ionized, although there is some uncertainty at this time as to the relative abundance of negative OH ions. The emission spectroscopy data provided some indication of the species in the exhaust as well as a measurement of temperature. A 24-electrode-pair segmented Faraday channel and 0.6 Tesla permanent

  5. Analysis of pulse detonation turbojet engines

    NASA Astrophysics Data System (ADS)

    Vutthivithayarak, Ronnachai

    2011-12-01

    Research over the last two decades has shown the potential advantages of pulse detonation engines (PDEs) over existing aero-engines in terms of improved thermodynamics efficiency, improved thrust performance, simplicity of design, and flexibility to operate over a wide speed range. The inherently unsteady characteristic of PDEs makes it difficulty to analyze and evaluate their performance. The conventional method that relies on steady-state assumptions cannot be directly applied. PDE studies have to employ unsteady gasdynamics behavior. In this study, the thermodynamic cycle of a PDE, which can be called the ZND cycle, is theoretically analyzed. A parametric analysis of turbojet PDEs is considered for both ideal and non-ideal cases. The conventional turbojet with a Brayton cycle is brought in the comparison to verify that PDEs can provide better performance.

  6. Quasi-One-Dimensional Modeling of Pulse Detonation Rocket Engines

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2003-01-01

    This viewgraph presentation provides information on the engine cycle of a pulse detonation rocket engine (PDRE), models for optimizing the performance of a PDRE, and the performance of PDREs in comparison to Solid State Rocket Engines (SSREs).

  7. Numerical Modeling of Pulse Detonation Rocket Engine Gasdynamics and Performance

    NASA Technical Reports Server (NTRS)

    Morris, C. I.

    2003-01-01

    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.

  8. Development of a chemical microthruster based on pulsed detonation

    NASA Astrophysics Data System (ADS)

    Wu, Ming-Hsun; Lu, Tsung-Hsun

    2012-10-01

    The development of a microthruster based on gaseous pulsed detonation is presented in this study. The feasibility of cyclic valveless pulsed detonation at frequencies over 100 Hz is first experimentally investigated in a microchannel with 1 mm × 0.6 mm rectangular cross-section. Highly reactive ethylene/oxygen mixtures are utilized to reduce the time and distance required for the reaction wave to run up to detonation in a smooth channel. High-speed visualizations have shown that the reaction waves reach detonative state through highly repeatable flame acceleration and deflagration-to-detonation transition processes in the channel. The validated concepts are implemented for the development of an integrated pulsed detonation microthruster. The microthruster was fabricated using low temperature co-fired ceramic tape technology. The volume of the reaction channel in the microthruster was 58 mm3. Spark electrodes and ion probes were embedded in the ceramic microthruster. The channel and via holes were fabricated using laser cutting techniques. Ion probe measurements showed that the reaction wave propagated at velocities larger than 2000 m s-1 before reaching the channel exit. The pulsed detonation microthruster has been successfully operated at frequencies as high as 200 Hz.

  9. Efficiency of Pulsed Detonation Thermal Spraying

    NASA Astrophysics Data System (ADS)

    Cannon, Jacob E.; Alkam, Mohammad; Butler, P. Barry

    2008-12-01

    Pulsed detonation thermal spray coating is used to enhance the material properties at the surface of an object. The present research implements computational fluid dynamic modeling to identify the efficiency of energy and mass delivered to potential target locations. Six cases of a hydrogen-air mixture are used to investigate the gas flow from the instant of ignition to the instant of flow reversal at the tube exit. Flow monitors are included in the model to represent potential target locations. These monitors are placed at different axial locations in order to record mass flow rate and the flow rate of enthalpy over time. The results indicate that there exists a quasi-steady jet that is efficient and predictable in delivery of energy and mass from the tube exit to potential target locations positioned on the centerline. The duration of the quasi-steady jet is dependent on the fraction of combustible gas (i.e., % fill). Much of the initial energy and mass delivered from the jet avoids the flow monitors. This is found to be related to the evolution of the jet behind the blast wave where energy is lost in expansion and vorticity production. It is also found that nearly 11-18% of the available energy and 20-23% of the available mass remains in the tube after flow reversal.

  10. Pulse Detonation Engine Air Induction System Analysis

    NASA Technical Reports Server (NTRS)

    Pegg, R. J.; Hunter, L. G.; Couch, B. D.

    1996-01-01

    A preliminary mixed-compression inlet design concept for potential pulse-detonation engine (PDE) powered supersonic aircraft was defined and analyzed. The objectives of this research were to conceptually design and integrate an inlet/PDE propulsion system into a supersonic aircraft, perform time-dependent CFD analysis of the inlet flowfield, and to estimate the installed PDE cycle performance. The study was baselined to a NASA Mach 5 Waverider study vehicle in which the baseline over/under turboramjet engines were replaced with a single flowpath PDE propulsion system. As much commonality as possible was maintained with the baseline configuration, including the engine location and forebody lines. Modifications were made to the inlet system's external ramp angles and a rotating cowl lip was incorporated to improve off-design inlet operability and performance. Engines were sized to match the baseline vehicle study's ascent trajectory thrust requirement at Mach 1.2. The majority of this study was focused on a flight Mach number of 3.0. The time-dependent Navier Stokes CFD analyses of a two-dimensional approximation of the inlet was conducted for the Mach 3.0 condition. The Lockheed Martin Tactical Aircraft Systems-developed FALCON CFD code with a two equation 'k-1' turbulence model was used. The downstream PDE was simulated by an array of four sonic nozzles in which the flow areas were rapidly varied in various opening/closing combinations. Results of the CFD study indicated that the inlet design concept operated successfully at the Mach 3.0 condition, satisfying mass capture, total pressure recovery, and operability requirements. Time-dependent analysis indicated that pressure and expansion waves from the simulated valve perturbations did not effect the inlet's operability or performance.

  11. Toward a High-Frequency Pulsed-Detonation Actuator

    NASA Technical Reports Server (NTRS)

    Cutler, Andrew D.; Drummond, J. Philip

    2006-01-01

    This paper describes the continued development of an actuator, energized by pulsed detonations, that provides a pulsed jet suitable for flow control in high-speed applications. A high-speed valve, capable of delivering a pulsed stream of reactants a mixture of H2 and air at rates of up to 1500 pulses per second, has been constructed. The reactants burn in a resonant tube and the products exit the tube as a pulsed jet. High frequency pressure transducers have been used to monitor the pressure fluctuations in the device at various reactant injection frequencies, including both resonant and off-resonant conditions. Pulsed detonations have been demonstrated in the lambda/4 mode of an 8 inch long tube at approximately 600 Hz. The pulsed jet at the exit of the device has been observed using shadowgraph and an infrared camera.

  12. Toward a High-Frequency Pulsed-Detonation Actuator

    NASA Technical Reports Server (NTRS)

    Cutler, Andrew D.; Drummond, J. Philip

    2006-01-01

    This paper describes the continued development of an actuator, energized by pulsed detonations, that provides a pulsed jet suitable for flow control in high-speed applications. A high-speed valve, capable of delivering a pulsed stream of reactants a mixture of H2 and air at rates of up to 1500 pulses per second, has been constructed. The reactants burn in a resonant tube and the products exit the tube as a pulsed jet. High frequency pressure transducers have been used to monitor the pressure fluctuations in the device at various reactant injection frequencies, including both resonant and off-resonant conditions. Pulsed detonations have been demonstrated in the lambda/4 mode of an 8 inch long tube at approx. 600 Hz. The pulsed jet at the exit of the device has been observed using shadowgraph and an infrared camera.

  13. A Performance Map for Ideal Air Breathing Pulse Detonation Engines

    NASA Technical Reports Server (NTRS)

    Paxson, Daniel E.

    2001-01-01

    The performance of an ideal, air breathing Pulse Detonation Engine is described in a manner that is useful for application studies (e.g., as a stand-alone, propulsion system, in combined cycles, or in hybrid turbomachinery cycles). It is shown that the Pulse Detonation Engine may be characterized by an averaged total pressure ratio, which is a unique function of the inlet temperature, the fraction of the inlet flow containing a reacting mixture, and the stoichiometry of the mixture. The inlet temperature and stoichiometry (equivalence ratio) may in turn be combined to form a nondimensional heat addition parameter. For each value of this parameter, the average total enthalpy ratio and total pressure ratio across the device are functions of only the reactant fill fraction. Performance over the entire operating envelope can thus be presented on a single plot of total pressure ratio versus total enthalpy ratio for families of the heat addition parameter. Total pressure ratios are derived from thrust calculations obtained from an experimentally validated, reactive Euler code capable of computing complete Pulse Detonation Engine limit cycles. Results are presented which demonstrate the utility of the described method for assessing performance of the Pulse Detonation Engine in several potential applications. Limitations and assumptions of the analysis are discussed. Details of the particular detonative cycle used for the computations are described.

  14. Parametric Study of High Frequency Pulse Detonation Tubes

    NASA Technical Reports Server (NTRS)

    Cutler, Anderw D.

    2008-01-01

    This paper describes development of high frequency pulse detonation tubes similar to a small pulse detonation engine (PDE). A high-speed valve injects a charge of a mixture of fuel and air at rates of up to 1000 Hz into a constant area tube closed at one end. The reactants detonate in the tube and the products exit as a pulsed jet. High frequency pressure transducers are used to monitor the pressure fluctuations in the device and thrust is measured with a balance. The effects of injection frequency, fuel and air flow rates, tube length, and injection location are considered. Both H2 and C2H4 fuels are considered. Optimum (maximum specific thrust) fuel-air compositions and resonant frequencies are identified. Results are compared to PDE calculations. Design rules are postulated and applications to aerodynamic flow control and propulsion are discussed.

  15. Multi-Level Analysis of Pulsed Detonation Engines

    NASA Technical Reports Server (NTRS)

    Ebrahimi, Houshang B.; Mohanraj, Rajendran; Merkle, Charles L.

    2001-01-01

    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 multi-dimensional effects and provides an initial multi-dimensional 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 PDE out-performs 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 result in a modest improvement in predicted specific impulse over the constant pressure boundary condition.

  16. Metallized Gelled Propellants Combustion Experiments in a Pulse Detonation Engine

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan; Jurns, John; Breisacher, Kevin; Kearns, Kim

    2006-01-01

    A series of combustion tests were performed with metallized gelled JP 8/aluminum fuels in a Pulse Detonation Engine (PDE). Nanoparticles of aluminum were used in the 60 to 100 nanometer diameter. Gellants were also of a nanoparticulate type composed of hydrocarbon alkoxide materials. Using simulated air (a nitrogen-oxygen mixture), the ignition potential of metallized gelled fuels with nanoparticle aluminum was investigated. Ignition of the JP 8/aluminum was possible with less than or equal to a 23-wt% oxygen loading in the simulated air. JP 8 fuel alone was unable to ignite with less than 30 percent oxygen loaded simulated air. The tests were single shot tests of the metallized gelled fuel to demonstrate the capability of the fuel to improve fuel detonability. The tests were conducted at ambient temperatures and with maximal detonation pressures of 1340 psia.

  17. Direct thrust force measurement of pulse detonation engine

    NASA Astrophysics Data System (ADS)

    Wahid, Mazlan Abdul; Faiz, M. Z. Ahmad; Saqr, Khalid M.

    2012-06-01

    In this paper we present the result of High-Speed Reacting Flow Laboratory (HiREF) pulse detonation engine (PDE) experimental study on direct thrust measurement. The thrust force generated by the repetitive detonation from a 50 mm inner diameter and 600 mm length tube was directly measured using load cell. Shchelkin spiral was used as an accelerator for the Deflagration to Detonation Transition (DDT) phenomenon. Propane-oxygen at stoichiometric condition was used as the combustible fuel-air mixture for the PDE. The PDE was operated at the operation frequency of 3Hz during the test. The amount of thrust force that was measured during the test reaching up to 70N. These values of thrust force were found to be fluctuating and its combustion phenomenon has been analyzed and discussed.

  18. A flash vaporization system for detonation of hydrocarbon fuels in a pulse detonation engine

    NASA Astrophysics Data System (ADS)

    Tucker, Kelly Colin

    Current research by both the US Air Force and Navy is concentrating on obtaining detonations in a pulse detonation engine (PDE) with low vapor pressure, kerosene based jet fuels. These fuels, however, have a low vapor pressure and the performance of a liquid hydrocarbon fueled PDE is significantly hindered by the presence of fuel droplets. A high pressure, fuel flash vaporization system (FVS) has been designed and built to reduce and eliminate the time required to evaporate the fuel droplets. Four fuels are tested: n-heptane, isooctane, aviation gasoline, and JP-8. The fuels vary in volatility and octane number and present a clear picture on the benefits of flash vaporization. Results show the FVS quickly provided a detonable mixture for all of the fuels tested without coking or clogging the fuel lines. Combustion results validated the model used to predict the fuel and air temperatures required to achieve gaseous mixtures with each fuel. The most significant achievement of the research was the detonation of flash vaporized JP-8 and air. The results show that the flash vaporized JP-8 used 20 percent less fuel to ignite the fuel air mixture twice as fast (8 ms from 16 ms) when compared to the unheated JP-8 combustion data. Likewise, the FVS has been validated as a reliable method to create the droplet free mixtures required for liquid hydrocarbon fueled PDEs.

  19. Thrust Augmentation Measurements Using a Pulse Detonation Engine Ejector

    NASA Technical Reports Server (NTRS)

    Santoro, Robert J.; Pal, Sibtosh

    2005-01-01

    Results of an experimental effort on pulse detonation driven ejectors are presented and discussed. The experiments were conducted using a pulse detonation engine (PDE)/ejector setup that was specifically designed for the study and operated at frequencies up to 50 Hz. The results of various experiments designed to probe different aspects of the PDE/ejector setup are reported. The baseline PDE was operated using ethylene (C2H4) as the fuel and an oxygen/nitrogen O2 + N2) mixture at an equivalence ratio of one. The PDE only experiments included propellant mixture characterization using a laser absorption technique, high fidelity thrust measurements using an integrated spring-damper system, and shadowgraph imaging of the detonation/shock wave structure emanating from the tube. The baseline PDE thrust measurement results at each desired frequency agree with experimental and modeling results reported in the literature. These PDE setup results were then used as a basis for quantifying thrust augmentation for various PDE/ejector setups with constant diameter ejector tubes and various ejector lengths, the radius of curvature for the ejector inlets and various detonation tube/ejector tube overlap distances. For the studied experimental matrix, the results showed a maximum thrust augmentation of 106% at an operational frequency of 30 Hz. The thrust augmentation results are complemented by shadowgraph imaging of the flowfield in the ejector tube inlet area and high frequency pressure transducer measurements along the length of the ejector tube.

  20. Experimental Study of a Pulse Detonation Engine Driven Ejector

    NASA Technical Reports Server (NTRS)

    Santoro, Robert J.; Pal, Sibtosh; Shehadeh, R.; Saretto, S.; Lee, S.-Y.

    2005-01-01

    Results of an experimental effort on pulse detonation driven ejectors are presented and discussed. The experiments were conducted using a pulse detonation engine (PDE)/ejector setup that was specifically designed for the study. The results of various experiments designed to probe different aspects of the PDE/ejector setup are reported. The baseline PDE was operated using ethylene (C2H4) as the fuel and an oxygen/nitrogen (O2 + N2) mixture at an equivalence ratio of one. The PDE only experiments included propellant mixture characterization using a laser absorption technique, high fidelity thrust measurements using an integrated spring-damper system, and shadowgraph imaging of the detonation/shock wave structure emanating from the tube. The baseline PDE thrust measurement results are in excellent agreement with experimental and modeling results reported in the literature. These PDE setup results were then used as a basis for quantifying thrust augmentation for various PDE/ejector setups with constant diameter ejector tubes and various detonation tube/ejector tube overlap distances. The results show that for the geometries studied here, a maximum thrust augmentation of 24% is achieved. Further increases are possible by tailoring the ejector geometry based on CFD predictions conducted elsewhere. The thrust augmentation results are complemented by shadowgraph imaging of the flowfield in the ejector tube inlet area and high frequency pressure transducer measurements along the length of the ejector tube.

  1. Analytical Investigation of an Airbreathing, Repetitively Pulsed LSC-Wave Thruster: Part 1

    NASA Astrophysics Data System (ADS)

    Myrabo, L. N.; Borkowski, C. A.; Kaminski, D. A.

    2006-05-01

    The prospects for an airbreathing repetitively pulsed, LSC-wave thruster are assessed from the perspective of both thrust production and radiation heat transfer rates into the vehicle undersurface. Pulsed LSC wave thrusters are characterized by high plasma temperatures (e.g., 11,000 K), moderate overpressures (a function of laser intensity), and subsonic LSC wave propagation velocities. The axisymmetric Lightcraft engine configuration examined in this study has a flat underbody impulsive surface of radius 1.25 m. Energy is supplied to the LSC wave by an annular laser beam converging uniformly from the edges of the disk (i.e., reflected inward towards the axis of symmetry from secondary optics positioned around the entire disc circumference). The LSC wave propagates radially outward toward the edge of the disc. At some point the beam is extinguished, and a new LSC wave is initiated at the disk center. The process is repeated at a frequency that produces the desired thrust, and continued until maximum temperatures are exceeded. Results indicate that thrust levels in the range of 104 to 4×105 N are indeed feasible — i.e., sufficient to sustain a 1 tonne Lightcraft in hover, or accelerate it vertically at 40 g. Depending on the desired thrust and LSC wave geometry, viable PRFs can range from ˜100 Hz up to 750 Hz (or more), pulse energies from 0.01 to 4 MJ, and pulse durations from 10 to 350 ms. Momentum coupling coefficients range from 100 to 1000 N/MW, varying as a function of laser intensity incident upon the LSC wave.

  2. Towards Integrated Pulse Detonation Propulsion and MHD Power

    NASA Technical Reports Server (NTRS)

    Litchford, Ron J.; Thompson, Bryan R.; Lineberry, John T.

    1999-01-01

    The interest in pulse detonation engines (PDE) arises primarily from the advantages that accrue from the significant combustion pressure rise that is developed in the detonation process. Conventional rocket engines, for example, must obtain all of their compression from the turbopumps, while the PDE provides additional compression in the combustor. Thus PDE's are expected to achieve higher I(sub sp) than conventional rocket engines and to require smaller turbopumps. The increase in I(sub sp) and the decrease in turbopump capacity must be traded off against each other. Additional advantages include the ability to vary thrust level by adjusting the firing rate rather than throttling the flow through injector elements. The common conclusion derived from these aggregated performance attributes is that PDEs should result in engines which are smaller, lower in cost, and lighter in weight than conventional engines. Unfortunately, the analysis of PDEs is highly complex due to their unsteady operation and non-ideal processes. Although the feasibility of the basic PDE concept has been proven in several experimental and theoretical efforts, the implied performance improvements have yet to be convincingly demonstrated. Also, there are certain developmental issues affecting the practical application of pulse detonation propulsion systems which are yet to be fully resolved. Practical detonation combustion engines, for example, require a repetitive cycle of charge induction, mixing, initiation/propagation of the detonation wave, and expulsion/scavenging of the combustion product gases. Clearly, the performance and power density of such a device depends upon the maximum rate at which this cycle can be successfully implemented. In addition, the electrical energy required for direct detonation initiation can be significant, and a means for direct electrical power production is needed to achieve self-sustained engine operation. This work addresses the technological issues associated

  3. Modeling of Multi-Tube Pulse Detonation Engine Operation

    NASA Technical Reports Server (NTRS)

    Ebrahimi, Houshang B.; Mohanraj, Rajendran; Merkle, Charles L.

    2001-01-01

    The present paper explores some preliminary issues concerning the operational characteristics of multiple-tube pulsed detonation engines (PDEs). The study is based on a two-dimensional analysis of the first-pulse operation of two detonation tubes exhausting through a common nozzle. Computations are first performed to assess isolated tube behavior followed by results for multi-tube flow phenomena. The computations are based on an eight-species, finite-rate transient flow-field model. The results serve as an important precursor to understanding appropriate propellant fill procedures and shock wave propagation in multi-tube, multi-dimensional simulations. Differences in behavior between single and multi-tube PDE models are discussed, The influence of multi-tube geometry and the preferred times for injecting the fresh propellant mixture during multi-tube PDE operation are studied.

  4. Pulse Detonation Rocket Engine Research at NASA Marshall

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2003-01-01

    Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modeling task. A quasi 1-D, finite-rate chemistry CFD model for a PDRE is described and implemented. A parametric study of the effect of blowdown pressure ratio on the performance of an optimized, fixed PDRE nozzle configuration is reported. The results are compared to a steady-state rocket system using similar modeling assumptions.

  5. Quasi-One-Dimensional Modeling of Pulse Detonation Rocket Engines

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2003-01-01

    Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modeling task. A quasi 1-D, finite-rate chemistry CFD model for a PDRE & described and implemented. A parametric study of the effect of blowdown pressure ratio on the performance of an optimized, fixed PDRE nozzle configuration is reported. The results are compared to a steady-state rocket system using similar modeling assumptions.

  6. Thrust Augmentation Measurements Using a Pulse Detonation Engine Ejector

    NASA Technical Reports Server (NTRS)

    Santoro, Robert J.; Pal, Sibtosh

    2003-01-01

    The present NASA GRC-funded three-year research project is focused on studying PDE driven ejectors applicable to a hybrid Pulse Detonation/Turbofan Engine. The objective of the study is to characterize the PDE-ejector thrust augmentation. A PDE-ejector system has been designed to provide critical experimental data for assessing the performance enhancements possible with this technology. Completed tasks include demonstration of a thrust stand for measuring average thrust for detonation tube multi-cycle operation, and design of a 72-in.-long, 2.25-in.-diameter (ID) detonation tube and modular ejector assembly. This assembly will allow testing of both straight and contoured ejector geometries. Initial ejectors that have been fabricated are 72-in.-long-constant-diameter tubes (4-, 5-, and 6-in.-diameter) instrumented with high-frequency pressure transducers. The assembly has been designed such that the detonation tube exit can be positioned at various locations within the ejector tube. PDE-ejector system experiments with gaseous ethylene/ nitrogen/oxygen propellants will commence in the very near future. The program benefits from collaborations with Prof. Merkle of University of Tennessee whose PDE-ejector analysis helps guide the experiments. The present research effort will increase the TRL of PDE-ejectors from its current level of 2 to a level of 3.

  7. Numerical Modeling of Pulse Detonation Rocket Engine Gasdynamics And Performance

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2004-01-01

    Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modeling task. A quasi-1-D, finite-rate chemistry computational fluid dynamics model for PDREs is described and implemented. Four different PDRE geometries are evaluated in this work: a baseline detonation tube, a detonation tube with a straight extension, and a detonation tube with two types of converging-diverging (C-D) nozzles. The effect of extension length and C-D nozzle area ratio on the single-shot gasdynamics and performance of a PDRE is studied over a wide range of blowdown pressure ratios (1-1000). The results indicate that a C-D nozzle is generally more effective than a straight extension in improving PDRE performance, particularly at higher pressure ratios. Additionally, the results show that the blowdown process of the C-D nozzle systems could be beneficially cut off well before the pressure at the end-wall reaches the ambient value. The performance results are also compared to a steady-state rocket system using similar modeling assumptions.

  8. Experimental Study of the Momentum Coupling Coefficient with the Pulse Frequency and Ambient Pressure for Air-Breathing Laser Propulsion

    NASA Astrophysics Data System (ADS)

    Tang, Zhiping; Cai, Jian; Gong, Ping; Hu, Xiaojun; Tan, Rongqin; Zheng, Zhijun; Wu, Jin; Lu, Yan

    2006-05-01

    The air-breathing laser propulsion tests are conducted for parabolic models by using a high power TEA-CO2 pulsed laser. It is found the momentum coupling coefficient Cm varies with the pulse repeatable frequency and reaches the maximum near 50Hz. With a multi-use pendulum chamber, the change of Cm at different ambient pressure is measured. The experimental results show that the propulsion efficiency Cm does not decrease below the altitude of 10km, even increases a little bit. The calculated Cm fits the experimental result up to altitude 3km, then, they are separated. One possible reason is the temperature which is constant in the experiments.

  9. Thrust Measurements for a Pulse Detonation Engine Driven Ejector

    NASA Technical Reports Server (NTRS)

    Santoro, Robert J.; Pak, Sibtosh; Shehadeh, R.; Saretto, S. R.; Lee, S.-Y.

    2005-01-01

    Results of an experimental effort on pulse detonation driven ejectors aimed at probing different aspects of PDE ejector processes, are presented and discussed. The PDE was operated using ethylene as the fuel and an equimolar oxygen/nitrogen mixture as the oxidizer at an equivalence ratio of one. The thrust measurements for the PDE alone are in excellent agreement with experimental and modeling results reported in the literature and serve as a Baseline for the ejector studies. These thrust measurements were then used as a basis for quantifying thrust augmentation for various PDE/ejector setups using constant diameter ejector tubes and various detonation tube/ejector tube overlap distances. The results show that for the geometries studied here, a maximum thrust augmentation of 24% is achieved. The thrust augmentation results are complemented by shadowgraph imaging of the flowfield in the ejector tube inlet area and high frequency pressure transducer measurements along the length of the ejector tube.

  10. Development of a Gas-Fed Pulse Detonation Research Engine

    NASA Technical Reports Server (NTRS)

    Litchford, Ron J.; Hutt, John (Technical Monitor)

    2001-01-01

    In response to the growing need for empirical data on pulse detonation engine performance and operation, NASA Marshall Space Flight Center has developed and placed into operation a low-cost gas-fed pulse detonation research engine. The guiding design strategy was to achieve a simple and flexible research apparatus, which was inexpensive to build and operate. As such, the engine was designed to operate as a heat sink device, and testing was limited to burst-mode operation with run durations of a few seconds. Wherever possible, maximum use was made of standard off-the-shelf industrial or automotive components. The 5-cm diameter primary tube is about 90-cm long and has been outfitted with a multitude of sensor and optical ports. The primary tube is fed by a coaxial injector through an initiator tube, which is inserted directly into the injector head face. Four auxiliary coaxial injectors are also integrated into the injector head assembly. All propellant flow is controlled with industrial solenoid valves. An automotive electronic ignition system was adapted for use, and spark plugs are mounted in both tubes so that a variety of ignition schemes can be examined. A microprocessor-based fiber-optic engine control system was developed to provide precise control over valve and ignition timing. Initial shakedown testing with hydrogen/oxygen mixtures verified the need for Schelkin spirals in both the initiator and primary tubes to ensure rapid development of the detonation wave. Measured pressure wave time-of-flight indicated detonation velocities of 2.4 km/sec and 2.2 km/sec in the initiator and primary tubes, respectively. These values implied a fuel-lean mixture corresponding to an H2 volume fraction near 0.5. The axial distribution for the detonation velocity was found to be essentially constant along the primary tube. Time-resolved thrust profiles were also acquired for both underfilled and overfilled tube conditions. These profiles are consistent with previous time

  11. Dynamics of galloping detonations: inert hydrodynamics with pulsed energy release

    NASA Astrophysics Data System (ADS)

    Radulescu, Matei I.; Shepherd, Joseph E.

    2015-11-01

    Previous models for galloping and cellular detonations of Ulyanitski, Vasil'ev and Higgins assume that the unit shock decay or cell can be modeled by Taylor-Sedov blast waves. We revisit this concept for galloping detonations, which we model as purely inert hydrodynamics with periodically pulsed energy deposition. At periodic time intervals, the chemical energy of the non-reacted gas accumulating between the lead shock and the contact surface separating reacted and non reacted gas is released nearly instantaneously. In between these pulses, the gas evolves as an inert medium. The resulting response of the gas to the periodic forcing is a sudden gain in pressure followed by mechanical relaxation accompanied by strong shock waves driven both forward and backwards. It is shown that the decay of the lead shock in-between pulses follows an exponential decay, whose time constant is controlled by the frequency of the energy deposition. More-over, the average speed of the lead shock is found to agree within 2 percent to the ideal Chapman-Jouguet value, while the large scale dynamics of the wave follows closely the ideal wave form of a CJ wave trailed by a Taylor expansion. When friction and heat losses are accounted for, velocity deficits are predicted, consistent with experiment. Work performed while MIR was on sabbatical at Caltech.

  12. Thrust Augmentation Measurements for a Pulse Detonation Engine Driven Ejector

    NASA Technical Reports Server (NTRS)

    Pal, S.; Santoro, Robert J.; Shehadeh, R.; Saretto, S.; Lee, S.-Y.

    2005-01-01

    Thrust augmentation results of an ongoing study of pulse detonation engine driven ejectors are presented and discussed. The experiments were conducted using a pulse detonation engine (PDE) setup with various ejector configurations. The PDE used in these experiments utilizes ethylene (C2H4) as the fuel, and an equi-molar mixture of oxygen and nitrogen as the oxidizer at an equivalence ratio of one. High fidelity thrust measurements were made using an integrated spring damper system. The baseline thrust of the PDE engine was first measured and agrees with experimental and modeling results found in the literature. Thrust augmentation measurements were then made for constant diameter ejectors. The parameter space for the study included ejector length, PDE tube exit to ejector tube inlet overlap distance, and straight versus rounded ejector inlets. The relationship between the thrust augmentation results and various physical phenomena is described. To further understand the flow dynamics, shadow graph images of the exiting shock wave front from the PDE were also made. For the studied parameter space, the results showed a maximum augmentation of 40%. Further increase in augmentation is possible if the geometry of the ejector is tailored, a topic currently studied by numerous groups in the field.

  13. Impact of Dissociation and Sensible Heat Release on Pulse Detonation and Gas Turbine Engine Performance

    NASA Technical Reports Server (NTRS)

    Povinelli, Louis A.

    2001-01-01

    A thermodynamic cycle analysis of the effect of sensible heat release on the relative performance of pulse detonation and gas turbine engines is presented. Dissociation losses in the PDE (Pulse Detonation Engine) are found to cause a substantial decrease in engine performance parameters.

  14. Preliminary Studies on a Small-Scale Single-Tube Pulse Detonation Rocket Prototype

    NASA Astrophysics Data System (ADS)

    Wang, Ke; Fan, Wei; Yan, Yu; Jin, Le

    2013-06-01

    As a new concept propulsion system, the pulse detonation engine has received extensive concerns from all over the world in the past few years. With oxidizer on board, it operates as a rocket engine which is known as pulse detonation rocket engine. In this study, a rocket model powered by a single-tube pulse detonation rocket engine was fabricated to demonstrate and validate whether or not it could operate stably and reliably independently. The single-tube pulse detonation rocket prototype consisted of a wireless control unit, three tanks for oxidizer, fuel and purge gas, various valves and a detonation tube. With compact design, the pulse detonation rocket prototype had an outer diameter of 260 mm and a length of 2200 mm. Oxygen, liquid aviation kerosene and nitrogen were utilized as oxidizer, fuel and purge gas, respectively. Operation tests were carried out to obtain proper operating conditions for the pulse detonation rocket prototype first, and then sliding test was conducted. It was concluded that the pulse detonation rocket prototype could operate stably and reliably. The generated thrust was estimated and compared with theoretical value.

  15. Intake flow analysis of a pulsed detonation engine

    NASA Astrophysics Data System (ADS)

    Strafaccia, Joshua A.

    A CFD program is converted and modified to explore unsteady flow within the intake system of a pulse detonation engine (PDE). Using a quasi-one-dimensional approach the program provides insight into the unsteady nature of localized equivalence ratios to include their effects on PDE performance. The original FORTRAN program is converted into the MATLAB architecture, taking full advantage of user availability and post processing convenience. The converted program was validated against the original program and modified to include a primitive intake manifold system with a single fuel injector located approximately 10 feet upstream of the primary intake valve. Constant fuel mass flow rate at the injector end creates local variations in equivalence ratio throughout the PDE that may have significant impact on overall engine performance. The results of the current thesis research suggest that performance effects of up to 21% can be attributed to non-uniform fuel distribution throughout the detonation process and are most prevalent at lower frequencies and fill ratios.

  16. Pulse Detonation Rocket Engine Research at NASA Marshall

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2003-01-01

    This viewgraph representation provides an overview of research being conducted on Pulse Detonation Rocket Engines (PDRE) by the Propulsion Research Center (PRC) at the Marshall Space Flight Center. PDREs have a theoretical thermodynamic advantage over Steady-State Rocket Engines (SSREs) although unsteady blowdown processes complicate effective use of this advantage in practice; PRE is engaged in a fundamental study of PDRE gas dynamics to improve understanding of performance issues. Topics covered include: simplified PDRE cycle, comparison of PDRE and SSRE performance, numerical modeling of quasi 1-D rocket flows, time-accurate thrust calculations, finite-rate chemistry effects in nozzles, effect of F-R chemistry on specific impulse, effect of F-R chemistry on exit species mole fractions and PDRE performance optimization studies.

  17. Numerical Modeling of Pulse Detonation Rocket Engine Gasdynamics and Performance

    NASA Technical Reports Server (NTRS)

    2003-01-01

    This paper presents viewgraphs on the numerical modeling of pulse detonation rocket engines (PDRE), with an emphasis on the Gasdynamics and performance analysis of these engines. The topics include: 1) Performance Analysis of PDREs; 2) Simplified PDRE Cycle; 3) Comparison of PDRE and Steady-State Rocket Engines (SSRE) Performance; 4) Numerical Modeling of Quasi 1-D Rocket Flows; 5) Specific PDRE Geometries Studied; 6) Time-Accurate Thrust Calculations; 7) PDRE Performance (Geometries A B C and D); 8) PDRE Blowdown Gasdynamics (Geom. A B C and D); 9) PDRE Geometry Performance Comparison; 10) PDRE Blowdown Time (Geom. A B C and D); 11) Specific SSRE Geometry Studied; 12) Effect of F-R Chemistry on SSRE Performance; 13) PDRE/SSRE Performance Comparison; 14) PDRE Performance Study; 15) Grid Resolution Study; and 16) Effect of F-R Chemistry on SSRE Exit Species Mole Fractions.

  18. Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2005-01-01

    Pulse detonation rocket engines (PDREs) have generated research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional rocket engines. The detonative mode of combustion employed by these devices offers a thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional rocket engines and gas turbines. However, while this theoretical advantage has spurred considerable interest in building PDRE devices, the unsteady blowdown process intrinsic to the PDRE has made realistic estimates of the actual propulsive performance 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. In recent work by the author, a quasi-one-dimensional, finite rate chemistry CFD model was utilized to study the gasdynamics and performance characteristics of PDREs over a range of blowdown pressure ratios from 1-1000. Models of this type are computationally inexpensive, and enable first-order parametric studies of the effect of several nozzle and extension geometries on PDRE performance over a wide range of conditions. However, the quasi-one-dimensional approach is limited in that it cannot properly capture the multidimensional blast wave and flow expansion downstream of the PDRE, nor can it resolve nozzle flow separation if present. Moreover, the previous work was limited to single-pulse calculations. In this paper, an axisymmetric finite rate chemistry model is described and utilized to study these issues in greater detail. Example Mach number contour plots showing the multidimensional blast wave and nozzle exhaust plume are shown. The performance results are compared with the quasi-one-dimensional results from the previous paper. Both Euler and Navier-Stokes solutions are calculated in order to determine the effect of viscous

  19. Quasi-One-Dimensional Modeling of Pulse Detonation Rocket Engines

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2002-01-01

    Pulsed detonation rocket engines (PDREs) have generated considerable research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional rocket engines. The detonative mode of combustion employed by these devices offers a thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional rocket engines and gas turbines. However, while this theoretical advantage has spurred a great deal of interest in building PDRE devices, the unsteady blowdown process intrinsic to the PDRE has made realistic estimates of the actual propulsive performance problematic. The recent review article by Kailasanath highlights some of the difficulties in comparing the available experimental measurements with numerical models. In a previous paper by the author, parametric studies of the performance of a single, straight-tube PDRE were reported. A 1-D, unsteady method of characteristics code, employing a constant-gamma assumption behind the detonation front, was developed for that study. Models of this type are computationally inexpensive, and are particularly useful for parametric performance comparisons. For example, a plot showing the specific impulse of various PDRE and steady-state rocket engine (SSRE) configurations as a function of blowdown pressure ratio. The performance curves clearly indicate that a straight-tube PDRE is superior in specific impulse to a SSRE with a sonic nozzle over the entire range of pressure ratios. Note, however, that a straight-tube PDRE in general does not compare favorably to a SSRE fitted with an optimized de Laval supersonic nozzle, particularly at the high pressure ratios typical for boost or in-space rocket applications. However, the calculations also show that if a dynamically optimized, supersonic de Laval nozzle could be could be fitted to a PDRE, then the specific impulse of the device would exceed that of a comparable SSRE

  20. Unsteady thrust measurement techniques for pulse detonation engines

    NASA Astrophysics Data System (ADS)

    Joshi, Dibesh Dhoj

    Thrust is a critical performance parameter and its correct determination is necessary to characterize an engine. Many conventional thrust measurement techniques prevail. However, further developments are required for correct measurement of thrust in the case of a pulse detonation engine (PDE), since the entire thrust generation process is intermittent. The significant effect of system dynamics in the form of inertial forces, stress wave propagation and reflections initiated in the structure due to detonations and pulse-to-pulse interaction in a fast operating PDE further complicate the thrust measurement process. These complications call for a further, detailed study of the unsteady thrust characteristics. A general approach was first developed to recover actual thrust from the measured thrust generated by the PDE. The developed approach consisted of two steps. The first step incorporated a deconvolution procedure using a pre-established system transfer function and measured input to reconstruct the output yielding the deconvolved thrust. The second step accounted for inertial forces through an acceleration compensation procedure. These two steps allowed the actual thrust to be determined. A small scale PDE operating at 10 and 20 Hz with varied filling fractions and mixture equivalence ratios was used for the experimental application of the general approach. The analytical study of gas dynamics in the PDE while in operation and the measured pressure histories at the exit of the engine allowed the generated thrust during a cycle to be determined semi-empirically. The thrust values determined semi-empirically were compared against the experimental results. A dynamical model of the PDE was created for the study of the unsteady thrust characteristics using finite element analysis. The results from finite element analysis were compared against semi-empirical and experimental results. In addition, finite element analysis also facilitated to numerically determine the

  1. Laser-supported detonation waves and pulsed laser propulsion

    SciTech Connect

    Kare, J.T.

    1989-01-01

    A laser thermal rocket uses the energy of a large remote laser, possibly ground-based, to heat an inert propellant and generate thrust. Use of a pulsed laser allows the design of extremely simple thrusters with very high performance compared to chemical rockets. The temperatures, pressures, and fluxes involved in such thrusters (10{sup 4} K, 10{sup 2} atmospheres, 10{sup 7} w/cm{sup 2}) typically result in the creation of laser-supported detonation (LSD) waves. The thrust cycle thus involves a complex set of transient shock phenomena, including laser-surface interactions in the ignition if the LSD wave, laser-plasma interactions in the LSD wave itself, and high-temperature nonequilibrium chemistry behind the LSD wave. The SDIO Laser Propulsion Program is investigating these phenomena as part of an overall effort to develop the technology for a low-cost Earth-to-orbit laser launch system. We will summarize the program's approach to developing a high performance thruster, the double-pulse planar thruster, and present an overview of some results obtained to date, along with a discussion of the many research questions still outstanding in this area. 16 refs., 7 figs.

  2. Laser-supported detonation waves and pulsed laser propulsion

    SciTech Connect

    Kare, J. )

    1990-07-30

    A laser thermal rocket uses the energy of a large remote laser, possibly ground-based, to heat an inert propellant and generate thrust. Use of a pulsed laser allows the design of extremely simple thrusters with very high performance compared to chemical rockets. The temperatures, pressures, and fluxes involved in such thrusters (10{sup 4} K, 10{sup 2} atmospheres, 10{sup 7} w/cm{sup 2}) typically result in the creation of laser-supported detonation (LSD) waves. The thrust cycle thus involves a complex set of transient shock phenomena, including laser-surface interactions in the ignition of the LSD wave, laser-plasma interactions in the LSD wave itself, and high-temperature nonequilibrium chemistry behind the LSD wave. The SDIO Laser Propulsion Program is investigating these phenomena as part of an overall effort to develop the technology for a low-cost Earth-to-orbit laser launch system. We will summarize the Program's approach to developing a high performance thruster, the double-pulse planar thruster, and present an overview of some results obtained to date, along with a discussion of the many research question still outstanding in this area.

  3. Airbreathing Hypersonic Vision-Operational-Vehicles Design Matrix

    NASA Technical Reports Server (NTRS)

    Hunt, James L.; Pegg, Robert J.; Petley, Dennis H.

    1999-01-01

    This paper presents the status of the airbreathing hypersonic airplane and space-access vision-operational-vehicle design matrix, with emphasis on horizontal takeoff and landing systems being, studied at Langley, it reflects the synergies and issues, and indicates the thrust of the effort to resolve the design matrix including Mach 5 to 10 airplanes with global-reach potential, pop-up and dual-role transatmospheric vehicles and airbreathing launch systems. The convergence of several critical systems/technologies across the vehicle matrix is indicated. This is particularly true for the low speed propulsion system for large unassisted horizontal takeoff vehicles which favor turbines and/or perhaps pulse detonation engines that do not require LOX which imposes loading concerns and mission Flexibility restraints.

  4. Airbreathing Hypersonic Vision-Operational-Vehicles Design Matrix

    NASA Technical Reports Server (NTRS)

    Hunt, James L.; Pegg, Robert J.; Petley, Dennis H.

    1999-01-01

    This paper presents the status of the airbreathing hypersonic airplane and space-access vision-operational-vehicle design matrix, with emphasis on horizontal takeoff and landing systems being studied at Langley; it reflects the synergies and issues, and indicates the thrust of the effort to resolve the design matrix including Mach 5 to 10 airplanes with global-reach potential, pop-up and dual-role transatmospheric vehicles and airbreathing launch systems. The convergence of several critical systems/technologies across the vehicle matrix is indicated. This is particularly true for the low speed propulsion system for large unassisted horizontal takeoff vehicles which favor turbines and/or perhaps pulse detonation engines that do not require LOX which imposes loading concerns and mission flexibility restraints.

  5. Studies of DDT enhancement approaches for kerosene-fueled small-scale pulse detonation engines applications

    NASA Astrophysics Data System (ADS)

    Huang, Y.; Tang, H.; Li, J.; Zhang, C.

    2012-11-01

    Two-phase small-scale pulse detonation engine (SPDE) offers a competitive alternative for small-scale propulsion systems from a high cycle efficiency and structural simplicity standpoint. SPDE models are designed with the aero-valve, and three different cases of obstacle combinations are used as deflagration-to-detonation transition (DDT) devices. The inner diameters of detonation tubes are 29 mm, and the lengths of three SPDEs are 995, 1,100, and 1,175 mm. Using kerosene-air as the fuel-oxidizer, a series of high-frequency detonation tests is conducted to seek efficient DDT enhancement approaches that reduce DDT distance and time and increase the frequency of kerosene-fueled SPDE. The results show that the fully developed detonation wave can be achieved at a distance of 3.4 times the minimum characteristic distance for gaseous detonation formation from the igniter and that the SPDE can steadily operate at a maximal frequency of 62.5 Hz. By adopting these DDT enhancement approaches, the detonability of kerosene is significantly improved. In addition, experiments are performed to study the effects of firing frequencies on detonation transitions. The results clearly indicate that the values of detonation wave pressures and velocities, the degree of overdriven wave, the ignition delay times, and detonation initiation times vary with frequencies. In terms of the performance, the optimal frequencies of three SPDE models are 20, 42.5, and 50 Hz, respectively.

  6. The Fluidic Obstacle Technique: An Approach for Enhancing Deflagration-to-Detonation Transition in Pulsed Detonation Engines

    NASA Astrophysics Data System (ADS)

    Knox, Benjamin W.

    The current research explored the fluidic obstacle technique and obtained relative performance estimates of this new approach for enhancement of de agration-to-detonation transition. Optimization of conventional physical obstacles has comprised the majority of de agration-to-detonation enhancement research but these devices ultimately degrade the performance of a pulsed detonation engine. Therefore, a new approach has been investigated that demonstrates a fluidic obstacle has the potential to maximize turbulence production and enhance the flame acceleration process, leading to successful DDT. A fluidic obstacle is also able to reduce total pressure losses, "heat soaking", and ignition times. A reduction in these variables serves to maximize available thrust. In addition, the fluidic obstacle technique is an active combustion control method capable of adapting to off-design conditions. Steady non-reacting and unsteady reacting flow have been utilized in two facilities, namely the UB Combustion Laboratory and AFRL Detonation Engine Research facility, to provide experimental measurements and observations into the feasibility of this new approach.

  7. Multiple-cycle Simulation of a Pulse Detonation Engine Ejector

    NASA Technical Reports Server (NTRS)

    Yungster, S.; Perkins, H. D.

    2002-01-01

    This paper presents the results of a study involving single and multiple-cycle numerical simulations of various PDE-ejector configurations utilizing hydrogen-oxygen mixtures. The objective was to investigate the thrust, impulse and mass flow rate characteristics of these devices. The results indicate that ejector systems can utilize the energy stored in the strong shock wave exiting the detonation tube to augment the impulse obtained from the detonation tube alone. Impulse augmentation ratios of up to 1.9 were achieved. The axial location of the converging-diverging ejectors relative to the end of the detonation tube were shown to affect the performance of the system.

  8. Parametric Investigation of Thrust Augmentation by Ejectors on a Pulsed Detonation Tube

    NASA Technical Reports Server (NTRS)

    Wilson, Jack; Sgondea, Alexandru; Paxson, Daniel E.; Rosenthal, Bruce N.

    2005-01-01

    A parametric investigation has been made of thrust augmentation of a 1 inch diameter pulsed detonation tube by ejectors. A set of ejectors was used which permitted variation of the ejector length, diameter, and nose radius, according to a statistical design of experiment scheme. The maximum augmentations for each ejector were fitted using a polynomial response surface, from which the optimum ejector diameters, and nose radius, were found. Thrust augmentations above a factor of 2 were measured. In these tests, the pulsed detonation device was run on approximately stoichiometric air-hydrogen mixtures, at a frequency of 20 Hz. Later measurements at a frequency of 40 Hz gave lower values of thrust augmentation. Measurements of thrust augmentation as a function of ejector entrance to detonation tube exit distance showed two maxima, one with the ejector entrance upstream, and one downstream, of the detonation tube exit. A thrust augmentation of 2.5 was observed using a tapered ejector.

  9. The development and testing of pulsed detonation engine ground demonstrators

    NASA Astrophysics Data System (ADS)

    Panicker, Philip Koshy

    2008-10-01

    The successful implementation of a PDE running on fuel and air mixtures will require fast-acting fuel-air injection and mixing techniques, detonation initiation techniques such as DDT enhancing devices or a pre-detonator, an effective ignition system that can sustain repeated firing at high rates and a fast and capable, closed-loop control system. The control system requires high-speed transducers for real-time monitoring of the PDE and the detection of the detonation wave speed. It is widely accepted that the detonation properties predicted by C-J detonation relations are fairly accurate in comparison to experimental values. The post-detonation flow properties can also be expressed as a function of wave speed or Mach number. Therefore, the PDE control system can use C-J relations to predict the post-detonation flow properties based on measured initial conditions and compare the values with those obtained from using the wave speed. The controller can then vary the initial conditions within the combustor for the subsequent cycle, by modulating the frequency and duty cycle of the valves, to obtain optimum air and fuel flow rates, as well as modulate the energy and timing of the ignition to achieve the required detonation properties. Five different PDE ground demonstrators were designed, built and tested to study a number of the required sub-systems. This work presents a review of all the systems that were tested, along with suggestions for their improvement. The PDE setups, ranged from a compact PDE with a 19 mm (3/4 in.) i.d., to two 25 mm (1 in.) i.d. setups, to a 101 mm (4 in.) i.d. dual-stage PDE setup with a pre-detonator. Propane-oxygen mixtures were used in the smaller PDEs. In the dual-stage PDE, propane-oxygen was used in the pre-detonator, while propane-air mixtures were used in the main combustor. Both rotary valves and solenoid valve injectors were studied. The rotary valves setups were tested at 10 Hz, while the solenoid valves were tested at up to 30 Hz

  10. The Use of Steady and Pulsed Detonations for Propulsion Systems

    SciTech Connect

    Adelman, H.G.; Menees, G.P.; Cambier, J.L.; Bowles, J.V.

    1996-02-01

    Objectives of the ODWE concept studies are: demonstrate the feasibility of the oblique detonation wave engine (ODWE) for hypersonic propulsion; demonstrate the existance and stability of an oblique detonation wave in hypersonic wind tunnels; develop engineering codes which predict the performance characteristics of the ODWE including specific impulse and thrust coefficients for various operating conditions; develop multi-dimensional computer codes which can model all aspects of the ODWE including fuel injection, mixing, ignition, combustion and expansion with fully detailed chemical kinetics and turbulence models; and validate the codes with experimental data use the simulations to predict the ODWE performance for conditions not easily obtained in wind tunnels.

  11. Formation of detonation in a pulse combustion chamber with a porous obstacle

    NASA Astrophysics Data System (ADS)

    Alhussan, Kh.; Assad, M. S.; Penyazkov, O. G.; Sevruk, K. L.

    2012-09-01

    A study has been made of the influence of a porous obstacle on deflagration-to-detonation transition in a pulse combustion chamber of small length. Dependences of the detonation-wave velocity on the distance have been obtained for two samples of a porous material (steel spheres and a ceramic porous body). It has been shown that the use of an insert from a porous material leads to a reduction of 40% in the predetonation distance without changing substantially the structure of the pulse combustion chamber.

  12. Laser High-Cycle Thermal Fatigue of Pulse Detonation Engine Combustor Materials Tested

    NASA Technical Reports Server (NTRS)

    Zhu, Dong-Ming; Fox, Dennis S.; Miller, Robert A.

    2001-01-01

    Pulse detonation engines (PDE's) have received increasing attention for future aerospace propulsion applications. Because the PDE is designed for a high-frequency, intermittent detonation combustion process, extremely high gas temperatures and pressures can be realized under the nearly constant-volume combustion environment. The PDE's can potentially achieve higher thermodynamic cycle efficiency and thrust density in comparison to traditional constant-pressure combustion gas turbine engines (ref. 1). However, the development of these engines requires robust design of the engine components that must endure harsh detonation environments. In particular, the detonation combustor chamber, which is designed to sustain and confine the detonation combustion process, will experience high pressure and temperature pulses with very short durations (refs. 2 and 3). Therefore, it is of great importance to evaluate PDE combustor materials and components under simulated engine temperatures and stress conditions in the laboratory. In this study, a high-cycle thermal fatigue test rig was established at the NASA Glenn Research Center using a 1.5-kW CO2 laser. The high-power laser, operating in the pulsed mode, can be controlled at various pulse energy levels and waveform distributions. The enhanced laser pulses can be used to mimic the time-dependent temperature and pressure waves encountered in a pulsed detonation engine. Under the enhanced laser pulse condition, a maximum 7.5-kW peak power with a duration of approximately 0.1 to 0.2 msec (a spike) can be achieved, followed by a plateau region that has about one-fifth of the maximum power level with several milliseconds duration. The laser thermal fatigue rig has also been developed to adopt flat and rotating tubular specimen configurations for the simulated engine tests. More sophisticated laser optic systems can be used to simulate the spatial distributions of the temperature and shock waves in the engine. Pulse laser high

  13. Experimental Investigation on Heating Kerosene Using Thrust Tube Waste Heat of Pulse Detonation Engine

    NASA Astrophysics Data System (ADS)

    Huang, Xiqiao; Yan, Yu; Mu, Yang; Zheng, Longxi; Chen, Le

    2013-12-01

    Performance of liquid hydrocarbon fueled pulse detonation engines are hindered by the time required to evaporate liquid fuel droplets within the mixture. Efforts such as heating the liquid fuel before introduced into the engine were made to overcome this problem. The engine wall suffered high temperature of gaseous detonative products, which is harmful to the engine. Therefore the waste heat of PDE engine wall was used here to heat the liquid kerosene of the engine by using two concentric heat-exchangers with helical slot inside mounted at the tail of the pulse detonation engine. The pulse detonation engine (PDE) used here are 110 mm in inner diameter and 1700 mm in length and liquid kerosene and gaseous air were used as fuel and oxidizer respectively. Heating liquid kerosene was found effective in enhancing the engine performance, which was helpful for deflagration to detonation transition process. It was found that the waste heat could elevate the kerosene temperature effectively and the wall temperature could also be decreased obviously.

  14. Pressure and Thrust Measurements of a High-Frequency Pulsed Detonation Tube

    NASA Technical Reports Server (NTRS)

    Nguyen, N.; Cutler, A. D.

    2008-01-01

    This paper describes measurements of a small-scale, high-frequency pulsed detonation tube. The device utilized a mixture of H2 fuel and air, which was injected into the device at frequencies of up to 1200 Hz. Pulsed detonations were demonstrated in an 8-inch long combustion volume, at about 600 Hz, for the quarter wave mode of resonance. The primary objective of this experiment was to measure the generated thrust. A mean value of thrust was measured up to 6.0 lb, corresponding to H2 flow based specific impulse of 2970 s. This value is comparable to measurements in H2-fueled pulsed detonation engines (PDEs). The injection and detonation frequency for this new experimental case was much higher than typical PDEs, where frequencies are usually less than 100 Hz. The compact size of the device and high frequency of detonation yields a thrust-per-unit-volume of approximately 2.0 pounds per cubic inch, and compares favorably with other experiments, which typically have thrust-per-unit-volume of order 0.01 pound per cubic inch. This much higher volumetric efficiency results in a potentially much more practical device than the typical PDE, for a wide range of potential applications, including high-speed boundary layer separation control, for example in hypersonic engine inlets, and propulsion for small aircraft and missiles.

  15. Pressure and Thrust Measurements of a High-Frequency Pulsed-Detonation Actuator

    NASA Technical Reports Server (NTRS)

    Nguyen, Namtran C.; Cutler, Andrew D.

    2008-01-01

    This paper describes the development of a small-scale, high-frequency pulsed detonation actuator. The device utilized a fuel mixture of H2 and air, which was injected into the device at frequencies of up to 1200 Hz. Pulsed detonations were demonstrated in an 8-inch long combustion volume, at approx.600 Hz, for the lambda/4 mode. The primary objective of this experiment was to measure the generated thrust. A mean value of thrust was measured up to 6.0 lb, corresponding to specific impulse of 2611 s. This value is comparable to other H2-fueled pulsed detonation engines (PDEs) experiments. The injection and detonation frequency for this new experimental case was approx.600 Hz, and was much higher than typical PDEs, where frequencies are usually less than 100 Hz. The compact size of the model and high frequency of detonation yields a thrust-per-unit-volume of approximately 2.0 lb/cu in, and compares favorably with other experiments, which typically have thrust-per-unit-volume values of approximately 0.01 lb/cu in.

  16. Slag Characterization and Removal Using Pulse Detonation for Coal Gasification.

    SciTech Connect

    Hugue, Z; Mei, D.; Biney, P.O.; Zhou, J.; Ali, M.R.

    1997-09-25

    The research activities performed in this quarter (reporting period: 07/01/97- 09/30/97) are summarized as follows: The research activities concentrated on (1) Design, development, and fabrication of a 9 positions (3x3 matrix form) fixture (2) Preparation of the test parameters (3) Multi-cycle detonation wave slag removal test (4) Partial analysis of the test results and (5) Interpretation and discussion of the test results.

  17. Quasi 1-D Study of Pulse Detonation Rocket Engine Blowdown Gasdynamics and Performance

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2002-01-01

    Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modeling task. A quasi 1-D, finite-rate chemistry CFD model for a PDRE is described and implemented. A parametric study of the effect of blowdown pressure ratio on the performance of several different PDRE nozzle configurations is reported.

  18. Parametric Investigation of Thrust Augmentation by Ejectors on a Pulsed Detonation Tube

    NASA Technical Reports Server (NTRS)

    Wilson, Jack; Sgondea, Alexandru; Paxson, Daniel E.; Rosenthal, Bruce N.

    2006-01-01

    A parametric investigation has been made of thrust augmentation of a 1 in. diameter pulsed detonation tube by ejectors. A set of ejectors was used which permitted variation of the ejector length, diameter, and nose radius, according to a statistical design of experiment scheme. The maximum augmentation ratios for each ejector were fitted using a polynomial response surface, from which the optimum ratios of ejector diameter to detonation tube diameter, and ejector length and nose radius to ejector diameter, were found. Thrust augmentation ratios above a factor of 2 were measured. In these tests, the pulsed detonation device was run on approximately stoichiometric air-hydrogen mixtures, at a frequency of 20 Hz. Later measurements at a frequency of 40 Hz gave lower values of thrust augmentation. Measurements of thrust augmentation as a function of ejector entrance to detonation tube exit distance showed two maxima, one with the ejector entrance upstream, and one downstream, of the detonation tube exit. A thrust augmentation of 2.5 was observed using a tapered ejector.

  19. Slang characterization and removal using pulse detonation technology during coal gasification

    SciTech Connect

    Huque, Z.; Mei, D.; Biney, P.O.; Zhou, J.

    1997-03-25

    Boiler slagging and fouling as a result of inorganic impurities in combustion gases being deposited on heat transfer tubes have caused severe problems in coal-fired power plant operation. These problems are fuel, system design, and operating condition dependent. Pulse detonation technology for the purpose of removing slag and fouling deposits in coal-fired utility power plant boilers offers great potential. The detonation wave technique based on high impact velocity with sufficient energy and thermal shock on the slag deposited on gas contact surfaces offers a convenient, inexpensive, yet efficient and effective way to supplement existing slag removal methods. These detonation waves have been demonstrated experimentally to have exceptionally high shearing capability important to the task of removing slag and fouling deposits. The experimental results show that the single shot detonation wave is capable of removing the entire slag (types of slag deposited on economizer) even at a distance of 8 in. from the exit of a detonation engine tube. Wave strength and slag orientation also have different effects on the chipping off of the slag. This paper discusses about the results obtained in effectively removing the economizer slag.

  20. Investigation of the effect of the ejector on the performance of the pulse detonation engine nozzle extension

    NASA Astrophysics Data System (ADS)

    Korobov, A. E.; Golovastov, S. V.

    2015-11-01

    Influence of an ejector nozzle extension on gas flow at a pulse detonation engine was investigated numerically and experimentally. Detonation formation was organized in stoichiometric hydrogen-oxygen mixture in cylindrical detonation tube. Cylindrical ejector was constructed and mounted at the open end of the tube. Thrust, air consumption and parameters of the detonation were measured in single and multiple regimes of operation. Axisymmetric model was used in numerical investigation. Equations of Navies-Stokes were solved using a finite-difference scheme Roe of second order of accuracy. Initial conditions were estimated on a base of experimental data. Numerical results were validated with experiments data.

  1. Experimental investigations on pulse detonation rocket engine with various injectors and nozzles

    NASA Astrophysics Data System (ADS)

    Yan, Yu; Fan, Wei; Wang, Ke; Zhu, Xu-dong; Mu, Yang

    2011-07-01

    Pulse detonation engines (PDEs) may represent a revolutionary approach to propulsion. The engine of simple construction can be easily manufactured. The pulse detonation rocket engine (PDRE) used here are 30 mm in inner diameter and 860 mm in length. Liquid kerosene, gaseous oxygen and nitrogen were used as fuel, oxidizer and purge gas, respectively. Two-phase detonation generating is harder than gaseous detonation due to liquid fuel atomization and mixing of two-phase reactants. It is a difficult task for liquid fuel and gaseous oxidizer to mix and form uniformly distributed mixture in the entire long engine during filling process in a short time. Therefore the velocities of fuel and oxidizer must be well designed to achieve not only the requirement of filling the entire engine but also the requirement of liquid fuel atomization and reactants mixing. Four injectors were tested to improve the atomization of liquid fuel and mixing process of reactants for performance enhancement of PDRE. Injector with small fuel exit area and large gas exit area was found to be effective for liquid fuel atomization and reactants mixing process. The PDRE with injector B performed the best among all the injectors tested. Nozzles are critical components in improving the performance of PDRE. Four kinds of bell-shaped converging-diverging nozzles were also tested here in order to enhance the performance of PDRE. It was found that a nozzle with high contraction ratio and high expansion ratio generated the highest thrust augmentation of 27.3%.

  2. JANNAF Airbreathing Propulsion Subcommittee and 35th Combustion Subcommittee Meeting. Volume 1

    NASA Technical Reports Server (NTRS)

    Fry, Ronald S. (Editor); Gannaway, Mary T. (Editor); Rognan, Melanie (Editor)

    1998-01-01

    This document, CPIA Publication 682, Volume 1, is a compilation of 5 unclassified/unlimited technical papers (approved for public release) which were presented at the 1 998 meeting of the Joint Army-Navy-NASA-Air Force (JANNAF) Airbreathing Propulsion Subcommittee (APS) and Combustion Subcommittee (CS) held jointly with the Propulsion Systems Hazards Subcommittee (PSHS). The meeting was held on 7-11 December 1998 at Raytheon Systems Company and the Marriott Hotel, Tucson, AZ. Topics covered include HyTech technology development, hydrocarbon fuel development for hypersonic applications, pulse detonation propulsion system development and arc heaters for direct-connect scramjet testing.

  3. Numerical study of sidewall filling for gas-fed pulse detonation engines

    NASA Astrophysics Data System (ADS)

    Rongrat, Wunnarat

    Pulse detonation engines for aerospace propulsion are required to operate at 50-100 Hz meaning that each pulse is 10-20 ms long. Filling of the engine and the related purging process become dominant due to their long duration compared to ignition and detonation wave propagation. This study uses ANSYS FLUENT to investigate the filling of a 1 m long tube with an internal diameter of 100 mm. Six different configurations were investigated with an endwall port and various sidewall arrangements, including stagger and inclination. A stoichiometric mixture of gaseous octane and air at STP was used to fill the tube at injection rates of 40, 150 and 250 m/s. Phase injection was also investigated and it showed performance improvements such as reduced lling time and reduced propellant escape from the exit.

  4. Real Gas Effects on the Performance of Hydrocarbon-fueled Pulse Detonation Engines

    NASA Technical Reports Server (NTRS)

    Povinelli, Louis A.; Yungster, Shaye

    2003-01-01

    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.

  5. Quasi 1-D Study of Pulse Detonation Rocket Engine Blowdown Gasdynamics and Performance

    NASA Technical Reports Server (NTRS)

    Morris, Christopher

    2002-01-01

    This viewgraph representation provides an overview of research which develops a quasi one dimensional chemistry computational fluid dynamics code to study the effect of nozzle design on the performance of pulse detonation rocket engines (PDREs). Topics considered include: PDREs vs. steady-state rocket engines (SSREs), PDRE cycles, numerical models of idealized PDRE performance, thrust determination of PDRE, specific geometries, and nozzle design and geometry.

  6. Performance Evaluation of Hybrid Gas Turbine Engine Embedded with Pulse Detonation Combustor

    NASA Astrophysics Data System (ADS)

    Deng, Jun-Xiang; Yan, Chuan-Jun; Zheng, Long-Xi; Huang, Xi-Qiao

    2011-09-01

    The numerical investigations of performance evaluation of a hybrid gas turbine engine embedded with a pulse detonation combustor (PDC) were performed to examine the improvement of the performance of the hybrid propulsion system. The calculation model and method were described. The architecture, configuration and size of detonation tubes were investigated in the calculation. Two models of detonation tube exit temperature were utilized. Eight configuration choices for the PDC based on the calculation model were designed. Specific fuel consumption of a hybrid gas turbine engine was compared with that of the baseline engine at the condition of the same engine net thrust. The experimental research of a PDC interacted with a radial flow turbine of a turbocharger was conducted. The numerical results show that if the net thrust of hybrid PDC engine is matched to that of baseline engine, specific fuel consumption of hybrid PDC engine is 20-25% less than that of baseline engine. The total volume of the hybrid engine combustor is reduced. The incorporation of PDC into gas turbine engine can improve the performance of hybrid PDC engine, decrease the combustor weight, and increase the thrust-weight ratio. The experimental results show that the fully developed detonation waves are achieved in the experimental apparatus.

  7. On the Exit Boundary Condition for One-Dimensional Calculations of Pulsed Detonation Engine Performance

    NASA Technical Reports Server (NTRS)

    Wilson, Jack; Paxson, Daniel E.

    2002-01-01

    In one-dimensional calculations of pulsed detonation engine (PDE) performance, the exit boundary condition is frequently taken to be a constant static pressure. In reality, for an isolated detonation tube, after the detonation wave arrives at the exit plane, there will be a region of high pressure, which will gradually return to ambient pressure as an almost spherical shock wave expands away from the exit, and weakens. Initially, the flow is supersonic, unaffected by external pressure, but later becomes subsonic. Previous authors have accounted for this situation either by assuming the subsonic pressure decay to be a relaxation phenomenon, or by running a two-dimensional calculation first, including a domain external to the detonation tube, and using the resulting exit pressure temporal distribution as the boundary condition for one-dimensional calculations. These calculations show that the increased pressure does affect the PDE performance. In the present work, a simple model of the exit process is used to estimate the pressure decay time. The planar shock wave emerging from the tube is assumed to transform into a spherical shock wave. The initial strength of the spherical shock wave is determined from comparison with experimental results. Its subsequent propagation, and resulting pressure at the tube exit, is given by a numerical blast wave calculation. The model agrees reasonably well with other, limited, results. Finally, the model was used as the exit boundary condition for a one-dimensional calculation of PDE performance to obtain the thrust wall pressure for a hydrogen-air detonation in tubes of length to diameter ratio (L/D) of 4, and 10, as well as for the original, constant pressure boundary condition. The modified boundary condition had no performance impact for values of L/D > 10, and moderate impact for L/D = 4.

  8. A Resonant Pulse Detonation Actuator for High-Speed Boundary Layer Separation Control

    NASA Technical Reports Server (NTRS)

    Beck, B. T.; Cutler, A. D.; Drummond, J. P.; Jones, S. B.

    2004-01-01

    A variety of different types of actuators have been previously investigated as flow control devices. Potential applications include the control of boundary layer separation in external flows, as well as jet engine inlet and diffuser flow control. The operating principles for such devices are typically based on either mechanical deflection of control surfaces (which include MEMS flap devices), mass injection (which includes combustion driven jet actuators), or through the use of synthetic jets (diaphragm devices which produce a pulsating jet with no net mass flow). This paper introduces some of the initial flow visualization work related to the development of a relatively new type of combustion-driven jet actuator that has been proposed based on a pulse detonation principle. The device is designed to utilize localized detonation of a premixed fuel (Hydrogen)-air mixture to periodically inject a jet of gas transversely into the primary flow. Initial testing with airflow successfully demonstrated resonant conditions within the range of acoustic frequencies expected for the design. Schlieren visualization of the pulsating air jet structure revealed axially symmetric vortex flow, along with the formation of shocks. Flow visualization of the first successful sustained oscillation condition is also demonstrated for one configuration of the current test section. Future testing will explore in more detail the onset of resonant combustion and the approach to conditions of sustained resonant detonation.

  9. SLAG CHARACTERIZATION AND REMOVAL USING PULSE DETONATION TECHNOLOGY DURING COAL GASIFICATION

    SciTech Connect

    DR. DANIEL MEI; DR. JIANREN ZHOU; DR. PAUL O. BINEY; DR. ZIAUL HUQUE

    1998-07-30

    Pulse detonation technology for the purpose of removing slag and fouling deposits in coal-fired utility power plant boilers offers great potential. Conventional slag removal methods including soot blowers and water lances have great difficulties in removing slags especially from the down stream areas of utility power plant boilers. The detonation wave technique, based on high impact velocity with sufficient energy and thermal shock on the slag deposited on gas contact surfaces offers a convenient, inexpensive, yet efficient and effective way to supplement existing slag removal methods. A slight increase in the boiler efficiency, due to more effective ash/deposit removal and corresponding reduction in plant maintenance downtime and increased heat transfer efficiency, will save millions of dollars in operational costs. Reductions in toxic emissions will also be accomplished due to reduction in coal usage. Detonation waves have been demonstrated experimentally to have exceptionally high shearing capability, important to the task of removing slag and fouling deposits. The experimental results describe the parametric study of the input parameters in removing the different types of slag and operating condition. The experimental results show that both the single and multi shot detonation waves have high potential in effectively removing slag deposit from boiler heat transfer surfaces. The results obtained are encouraging and satisfactory. A good indication has also been obtained from the agreement with the preliminary computational fluid dynamics analysis that the wave impacts are more effective in removing slag deposits from tube bundles rather than single tube. This report presents results obtained in effectively removing three different types of slag (economizer, reheater, and air-heater) t a distance of up to 20 cm from the exit of the detonation tube. The experimental results show that the softer slags can be removed more easily. Also closer the slag to the exit of

  10. Simplified Analysis of Pulse Detonation Rocket Engine B1owdown Gasdynamics and Performance

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2001-01-01

    Pulsed detonation rocket engines (PDREs) have generated considerable research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional rocket engines. The detonative mode of combustion employed by these devices offers a thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional rocket engines and gas turbines. However, while this theoretical advantage has spurred a great deal of interest in building PDRE devices, the unsteady blowdown process intrinsic to the PDRE has made realistic estimates of the actual propulsive performance problematic. The recent review article by Kailasanath highlights some of the difficulties in comparing the available experimental measurements with numerical models. The goal of this paper is to improve understanding of PDRE blowdown gasdynamics and performance issues through use of a simplified model that captures the essential features of the unsteady blowdown process, and yet remains computationally inexpensive. The PDRE system studied here is highly idealized, consisting of a constant-area detonation tube with one end closed and the other end open to the environment. The tube is prefilled with a gaseous propellant mixture with no initial velocity or outflow to the environment. The detonation is initiated instantaneously at the closed end of the device. Chapman-Jouguet (C-J) post-detonation gas conditions are calculated using the CET89 version of the NASA thermochemical code. The I-D, unsteady method of characteristics is used to calculate the flowfield following the detonation front. See the compressible flow texts by Thompson and Zucrow and Hoffman for details of this method. Parametric studies of the effect of mixture stoichiometry, fill temperature, and blowdown pressure ratio on performance are reported. A comparison of the performance of an idealized straight-tube PDRE with a conventional steady

  11. Simplified Analysis of Pulse Detonation Rocket Engine Blowdown Gasdynamics and Performance

    NASA Technical Reports Server (NTRS)

    Morris, C. I.; Rodgers, Stephen L. (Technical Monitor)

    2002-01-01

    Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modellng task. A simplified model for an idealized, straight-tube, single-shot PDRE blowdown process and thrust determination is described and implemented. In order to form an assessment of the accuracy of the model, the flowfield time history is compared to experimental data from Stanford University. Parametric Studies of the effect of mixture stoichiometry, initial fill temperature, and blowdown pressure ratio on the performance of a PDRE are performed using the model. PDRE performance is also compared with a conventional steady-state rocket engine over a range of pressure ratios using similar gasdynamic assumptions.

  12. Experimental study of a valveless pulse detonation rocket engine using nontoxic hypergolic propellants

    NASA Astrophysics Data System (ADS)

    Kan, Brandon K.

    A pulsed detonation rocket engine concept was explored through the use of hypergolic propellants in a fuel-centered pintle injector combustor. The combustor design yielded a simple open ended chamber with a pintle type injection element and pressure instrumentation. High-frequency pressure measurements from the first test series showed the presence of large pressure oscillations in excess of 2000 psia at frequencies between 400-600 hz during operation. High-speed video confirmed the high-frequency pulsed behavior and large amounts of after burning. Damaged hardware and instrumentation failure limited the amount of data gathered in the first test series, but the experiments met original test objectives of producing large over-pressures in an open chamber. A second test series proceeded by replacing hardware and instrumentation, and new data showed that pulsed events produced under expanded exhaust prior to pulsing, peak pressures around 8000 psi, and operating frequencies between 400-800 hz. Later hot-fires produced no pulsed behavior despite undamaged hardware. The research succeeded in producing pulsed combustion behavior using hypergolic fuels in a pintle injector setup and provided insights into design concepts that would assist future injector designs and experimental test setups.

  13. Thrust Estimates of a Partially Filled Multi-Cycle Pulse Detonation Rocket Engine

    NASA Astrophysics Data System (ADS)

    Wang, Ke; Fan, Wei

    2013-09-01

    Partially filled multi-cycle operation usually happens in pulse detonation rocket engines (PDREs). Thrust estimates of PDREs under such operating conditions will provide significant reference for relevant studies. An analytical model for thrust estimates is proposed in the present study. This model is based on Wintenberger' model for impulse calculation and empirical formula for specific impulse prediction of partially filled tubes; moreover, it takes into account performance penalty created by obstacles preliminarily, which are introduced to accelerate deflagration to detonation transition. Four specific analytical models are developed according to three previously proposed empirical formulas and a fitting one proposed for partial filling effect. Comparisons between model predictions and experimental measurements are carried out to validate the reliability of the models. It is found that the models make reasonable estimates and one of the models performs very well over a wide range of conditions. Discussion and analysis on the proposed model and partial filling effect are also performed in this study. Although there are many other factors that should be considered in evaluating thrust of partially filled multi-cycle PDREs, the present study supplies a rapid and effective means for thrust estimation and provides some modeling ideas meanwhile.

  14. Slag characterization and removal using pulse detonation for coal gasification. Quarterly report, October 1--December 31, 1996

    SciTech Connect

    Huque, Z.; Mei, D.; Biney, P.O.; Zhou, J.

    1997-12-31

    This report summarizes the costs associated with ash fouling and its removal from the heat exchanger of the Sherburne Co. plant`s 800 MW unit. The lead heat exchanger in the convective pass takes the brunt of the ash accumulation. This study concluded that for the amount of money used one cleaning cycle, 33 pulse detonation units could be purchased.

  15. Thermal Spray Using a High-Frequency Pulse Detonation Combustor Operated in the Liquid-Purge Mode

    NASA Astrophysics Data System (ADS)

    Endo, T.; Obayashi, R.; Tajiri, T.; Kimura, K.; Morohashi, Y.; Johzaki, T.; Matsuoka, K.; Hanafusa, T.; Mizunari, S.

    2016-02-01

    Experiments on thermal spray by pulsed detonations at 150 Hz were conducted. Two types of pulse detonation combustors were used, one operated in the inert gas purge (GAP) mode and the other in the liquid-purge (LIP) mode. In both modes, all gases were supplied in the valveless mode. The GAP mode is free of moving components, although the explosive mixture is unavoidably diluted with the inert gas used for the purge of the hot burned gas. In the LIP mode, pure fuel-oxygen combustion can be realized, although a liquid-droplet injector must be actuated cyclically. The objective of this work was to demonstrate a higher spraying temperature in the LIP mode. First, the temperature of CoNiCrAlY particles heated by pulsed detonations was measured. As a result, the spraying temperature in the LIP mode was higher than that in the GAP mode by about 1000 K. Second, the temperature of yttria-stabilized zirconia (YSZ) particles, whose melting point was almost 2800 °C, heated by pulsed detonations in the LIP mode was measured. As a result, the YSZ particles were heated up to about 2500 °C. Finally, a thermal spray experiment using YSZ particles was conducted, and a coating with low porosity was successfully deposited.

  16. Millimeter-wave Driven Shock Wave for a Pulsed Detonation Microwave Rocket

    NASA Astrophysics Data System (ADS)

    Yamaguchi, Toshikazu; Komatsu, Reiji; Fukunari, Masafumi; Komurasaki, Kimiya; Oda, Yasuhisa; Kajiwara, Ken; Takahashi, Koji; Sakamoto, Keishi

    2011-11-01

    A shock wave driven by millimeter wave ionization can be applied into a pulsed detonation engine as a Microwave Rocket. A high pressure induced inside the thruster generates the thrust, thus the shock wave propagation driven by the plasma is important. In this study, to obtain a different propagating structure, the beam profile was transformed from a Gaussian into a Ring and a Flat-top profile by using a pair of phase correcting mirrors. As a result, the shape of the propagating plasma was changed into a no-center shape in case of the Ring beam, and it was changed to a wider shape in case of the Flat-top beam. The propagating velocity of the ionization front of the Flat-top beam was much lower than that of the Gaussian due to the lower peak power density, and a higher plateau pressure and higher thrust impulse were generated by the Flat-top beam.

  17. Numerical Analysis of a Pulse Detonation Cross Flow Heat Load Experiment

    NASA Technical Reports Server (NTRS)

    Paxson, Daniel E.; Naples, Andrew .; Hoke, John L.; Schauer, Fred

    2011-01-01

    A comparison between experimentally measured and numerically simulated, time-averaged, point heat transfer rates in a pulse detonation (PDE) engine is presented. The comparison includes measurements and calculations for heat transfer to a cylinder in crossflow and to the tube wall itself using a novel spool design. Measurements are obtained at several locations and under several operating conditions. The measured and computed results are shown to be in substantial agreement, thereby validating the modeling approach. The model, which is based in computational fluid dynamics (CFD) is then used to interpret the results. A preheating of the incoming fuel charge is predicted, which results in increased volumetric flow and subsequent overfilling. The effect is validated with additional measurements.

  18. Slag characterization and removal using pulse detonation for coal gasification. Quarterly research report, July 1--September 31, 1996

    SciTech Connect

    Huque, Z.; Mei, D.; Biney, P.O.; Zhou, J.; Ali, M.R.

    1996-10-25

    Boiler slagging and fouling as a result of inorganic impurities in combustion gases being deposited on heat transfer tubes have caused severe problems in coal-fired power plant operation. These problems are fuel, system design, and operating condition dependent. Conventional slag and ash removal methods include the use of in situ blowing or jet-type devices such as air or steam soot blowers and water lances. Pulse detonation technology for the purpose of removing slag and fouling deposits in coal-fired utility power plant boilers offers great potential. The detonation wave technique based on high impact velocity with sufficient energy and thermal shock on the slag deposited on gas contact surfaces offers a convenient, inexpensive, yet efficient and effective way to supplement existing slag removal methods. These detonation waves have been demonstrated experimentally to have exceptionally high shearing capability important to the task of removing slag and fouling deposits. Several tests have been performed with single shot detonation wave at University of Texas at Arlington to remove the slag deposit. To hold the slag deposit samples at the exit of detonation tube, two types of fixture was designed and fabricated. They are axial arrangement and triangular arrangement. The slag deposits from the utility boilers have been used to prepare the slag samples for the test. The experimental results show that the single shot detonation wave is capable of removing the entire slag (types of slag deposited on economizer, and air-heater, i.e., relatively softer slags) and 30% of the reheater slag (which is harder) even at a distance of 6 in. from the exit of a detonation engine tube. Wave strength and slag orientation also have different effects on the chipping off of the slag. The annual report discusses about the results obtained in effectively removing the slag.

  19. Studies of Operating Frequency Effects On Ejector-based Thrust Augmentation in a Pulse Detonation Engine

    NASA Technical Reports Server (NTRS)

    Landry, K.

    2005-01-01

    Studies were performed in order to characterize the thrust augmentation potential of an ejector in a Pulse Detonation Engine application. A 49-mm diameter tube of 0.914-m length was constructed with one open end and one closed end. Ethylene, oxygen, and nitrogen were introduced into the tube at the closed end through the implementation of a fast mixing injector. The tube was completely filled with a stoichiometric mixture containing a one to one molar ratio of nitrogen to oxygen. Ethylene was selected as the fuel due to its detonation sensitivity and the molar ratio of the oxidizer was chosen for heat transfer purposes. Detonations were initiated in the tube through the use of a spark ignition system. The PDE was operated in a multi-cycle mode at frequencies ranging from 20-Hz to 50-Hz. Baseline thrust measurements with no ejector present were performed while operating the engine at various frequencies and compared to theoretical estimates. The baseline values were observed to agree with the theoretical model at low operating frequencies and proved to be increasingly lower than the predicted values as the operating frequency was increased. The baseline thrust measurements were observed to agree within 15 percent of the model for all operating frequencies. A straight 152-mm diameter ejector was installed and thrust augmentation percentages were measured. The length of the ejector was varied while the overlap percentage (percent of the ejector length which overlapped the tube) was maintained at 25 percent for all tests. In addition, the effect of ejector inlet geometry was investigated by comparing results with a straight inlet to those of a 38-mm inlet diameter. The thrust augmentation of the straight inlet ejector proved to be independent of engine operating frequency, augmenting thrust by 40 percent for the 0.914-m length ejector. In contrast, the rounded lip ejector of the same length seemed to be highly dependent on the engine operating frequency. An optimum

  20. Unsteady specific work and isentropic efficiency of a radial turbine driven by pulsed detonations

    NASA Astrophysics Data System (ADS)

    Rouser, Kurt P.

    There has been longstanding government and industry interest in pressure-gain combustion for use in Brayton cycle based engines. Theoretically, pressure-gain combustion allows heat addition with reduced entropy loss. The pulsed detonation combustor (PDC) is a device that can provide such pressure-gain combustion and possibly replace typical steady deflagration combustors. The PDC is inherently unsteady, however, and comparisons with conventional steady deflagration combustors must be based upon time-integrated performance variables. In this study, the radial turbine of a Garrett automotive turbocharger was coupled directly to and driven, full admission, by a PDC in experiments fueled by hydrogen or ethylene. Data included pulsed cycle time histories of turbine inlet and exit temperature, pressure, velocity, mass flow, and enthalpy. The unsteady inlet flowfield showed momentary reverse flow, and thus unsteady accumulation and expulsion of mass and enthalpy within the device. The coupled turbine-driven compressor provided a time-resolved measure of turbine power. Peak power increased with PDC fill fraction, and duty cycle increased with PDC frequency. Cycle-averaged unsteady specific work increased with fill fraction and frequency. An unsteady turbine efficiency formulation is proposed, including heat transfer effects, enthalpy flux-weighted total pressure ratio, and ensemble averaging over multiple cycles. Turbine efficiency increased with frequency but was lower than the manufacturer reported conventional steady turbine efficiency.

  1. Ideal cycle analysis of a regenerative pulse detonation engine for power production

    NASA Astrophysics Data System (ADS)

    Bellini, Rafaela

    Over the last few decades, considerable research has been focused on pulse detonation engines (PDEs) as a promising replacement for existing propulsion systems with potential applications in aircraft ranging from the subsonic to the lower hypersonic regimes. On the other hand, very little attention has been given to applying detonation for electric power production. One method for assessing the performance of a PDE is through thermodynamic cycle analysis. Earlier works have adopted a thermodynamic cycle for the PDE that was based on the assumption that the detonation process could be approximated by a constant volume process, called the Humphrey cycle. The Fickett-Jacob cycle, which uses the one--dimensional Chapman--Jouguet (CJ) theory of detonation, has also been used to model the PDE cycle. However, an ideal PDE cycle must include a detonation based compression and heat release processes with a finite chemical reaction rate that is accounted for in the Zeldovich -- von Neumann -- Doring model of detonation where the shock is considered a discontinuous jump and is followed by a finite exothermic reaction zone. This work presents a thermodynamic cycle analysis for an ideal PDE cycle for power production. A code has been written that takes only one input value, namely the heat of reaction of a fuel-oxidizer mixture, based on which the program computes all the points on the ZND cycle (both p--v and T--s plots), including the von Neumann spike and the CJ point along with all the non-dimensionalized state properties at each point. In addition, the program computes the points on the Humphrey and Brayton cycles for the same input value. Thus, the thermal efficiencies of the various cycles can be calculated and compared. The heat release of combustion is presented in a generic form to make the program usable with a wide variety of fuels and oxidizers and also allows for its use in a system for the real time monitoring and control of a PDE in which the heat of reaction

  2. Experimental Study on DDT Characteristics in Spiral Configuration Pulse Detonation Engines

    NASA Astrophysics Data System (ADS)

    Wang, Wei; Qiu, Hua; Fan, Wei; Xiong, Cha

    2013-09-01

    This work investigated features of the deflagration to detonation transition in a curved tube. A number of experiments were performed to acquire the transition rule of DDT, which would provide the design data and theoretical basis for the curved detonation chamber. The content of research is as follows: (1) Flow resistance experiments of nine detonation chambers have been explored. The results show that the spiral configuration can reduce the axial length of DC, and the total pressure recovery coefficient increases with the spiral pitch. (2) Single-cycle detonation experiments have been conducted using the 9 tubes in the resistance experiments. Liquid-gasoline/air is used as the detonative mixture in all the experiments. The detonation experimental results indicate that there is no detonation wave formed in the straight tube, but in all the selected spiral tubes fully-developed detonation waves have been obtained; compared to the straight tube case, the DDT time decrease with the decreasing of the radius of curvature (RC) by 6.2%˜19.8% in the spiral detonation tubes.

  3. Study of ejector geometry on thrust augmentation for pulse detonation engine ejector systems

    NASA Astrophysics Data System (ADS)

    Shehadeh, Ra'fat

    Pulse detonation engine (PDE) technology is a novel form of propulsion that offers the potential of high efficiency combustion with reduced hardware complexity. Although the primary interest of the research in the pulse detonation engine field is directed towards overcoming the problems associated with operating a pure PDE system, there are other worthy options to be considered for these engines. The PDE driven ejector concept is one such option where the system would be part of a hybrid PD/Turbofan engine. This system offers the promise of replacing the high-pressure turbine sections of the core of a high bypass turbofan engine. The purpose of the current research is to investigate the thrust augmentation capabilities of a PDE driven ejector and provide experimental data that would assist in understanding the behavior of such a system. The major potential advantages of the PDE-ejector include reduced costs due to the reduced engine weight, along with improved specific fuel consumption and specific power inherent in the incorporation of a PDE component. To achieve the goal of this research, the thrust augmentation of a PDE driven ejector was characterized for a set of configurations. Two separate PDE's were utilized in this study. The first PDE was capable of operating at a constant frequency of 10 Hz de to flow rate limitations, and another PDE built to have an operational frequency range of 10 Hz-70 Hz to test the effect of operational frequency on PDE-ejector systems. Optical diagnostics were employed at specific positions of interest to understand the physical behavior of the flow. Baseline experimental results helped define and understand the operational characteristics of the PDE's utilized in this study. Thrust measurements were then made for PDE driven ejector configurations. The parameters that were independently changed were the inlet geometry of a constant diameter ejector, as well as the overlap distance between the PDE tube exit and ejector tube inlet

  4. Oxidation- and Creep-Enhanced Fatigue of Haynes 188 Alloy-Oxide Scale System Under Simulated Pulse Detonation Engine Conditions

    NASA Technical Reports Server (NTRS)

    Zhu, Dongming; Fox, Dennis S.; Miller, Robert A.

    2002-01-01

    The development of the pulse detonation engine (PDE) requires robust design of the engine components that are capable of enduring harsh detonation environments. In this study, a high cycle thermal fatigue test rig was developed for evaluating candidate PDE combustor materials using a CO2 laser. The high cycle thermal fatigue behavior of Haynes 188 alloy was investigated under an enhanced pulsed laser test condition of 30 Hz cycle frequency (33 ms pulse period, and 10 ms pulse width including 0.2 ms pulse spike). The temperature swings generated by the laser pulses near the specimen surface were characterized by using one-dimensional finite difference modeling combined with experimental measurements. The temperature swings resulted in significant thermal cyclic stresses in the oxide scale/alloy system, and induced extensive surface cracking. Striations of various sizes were observed at the cracked surfaces and oxide/alloy interfaces under the cyclic stresses. The test results indicated that oxidation and creep-enhanced fatigue at the oxide scale/alloy interface was an important mechanism for the surface crack initiation and propagation under the simulated PDE condition.

  5. Numerical Study and Performance Evaluation for Pulse Detonation Engine with Exhaust Nozzle

    NASA Astrophysics Data System (ADS)

    Kimura, Yuichiro; Tsuboi, Nobuyuki; Hayashi, A. Koichi; Yamada, Eisuke

    This paper presents the propulsive performance evaluation for the H2/Air Pulse Detonation Engine (PDE) with a converging-diverging exhaust nozzle by the system-level modeling and multi-cycle numerical simulations. This study deals with the two-dimensional and axisymmetric compressible Euler equations with a detail chemical reaction model. First, single-shot propulsive performance of simplified-PDE, which is without exhaust nozzle, is evaluated to show the validity of the numerical and performance evaluation method. The influences of the initial conditions, ignition energy, grid resolution, and scale effects on the propulsive performance are studied with the multi-cycle simulations. The present results are compared with the results calculated by Ma et al. and Harris et al. and the difference between their results and the present simulations are approximately 2-3% because their chemical reactions use one-step model with one-γ model. The effects of the specific heat ratio should be estimated for various nozzle configurations and flight conditions.

  6. Progress in the Development of a Nozzle Design Methodology for Pulsed Detonation Engines

    NASA Technical Reports Server (NTRS)

    Leary, B. A.; Waltrup, P. J.; Rice, T.; Cybyk, B. Z.

    2002-01-01

    The Johns Hopkins University Applied Physics Laboratory (JHU/APL), in support of the NASA Glenn Research Center (NASA GRC), is investigating performance methodologies and system integration issues related to Pulsed Detonation Engine (PDE) nozzles. The primary goal of this ongoing effort is to develop design and performance assessment methodologies applicable to PDE exit nozzle(s). APL is currently focusing its efforts on a common plenum chamber design that collects the exhaust products from multiple PDE tubes prior to expansion in a single converging-diverging exit nozzle. To accomplish this goal, a time-dependent, quasi-one-dimensional analysis for determining the flow properties in and through a single plenum and exhaust nozzle is underway. In support of these design activities, parallel modeling efforts using commercial Computational Fluid Dynamics (CFD) software are on-going. These efforts include both two and three-dimensional as well as steady and time-dependent computations to assess the flow in and through these devices. This paper discusses the progress in developing this nozzle design methodology.

  7. JANNAF 24th Airbreathing Propulsion Subcommittee and 36th Combustion Subcommittee Joint Meeting. Volume 1

    NASA Technical Reports Server (NTRS)

    Fry, Ronald S. (Editor); Gannaway, Mary T. (Editor)

    1999-01-01

    Volume 1, the first of three volumes is a compilation of 16 unclassified/unlimited-technical papers presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 24th Airbreathing Propulsion Subcommittee and 36th Combustion Subcommittee held jointly with the 181 Propulsion Systems Hazards Subcommittee. The meeting was held on 18-21 October 1999 at NASA Kennedy Space Center and The DoubleTree Oceanfront Hotel, Cocoa Beach, Florida. Topics covered include overviews of RBCC and PDE hypersonic technology, Hyper-X propulsion ground testing, development of JP-8 for hypersonic vehicle applications, numerical simulation of dual-mode SJ combustion, V&V of M&S computer codes, MHD SJ and Rocket Based Combined Cycle (RBCC) launch vehicle concepts, and Pulse Detonation Engine (PDE) propulsion technology development including fundamental investigations, modeling, aerodynamics, operation and performance.

  8. Thermal Barrier and Protective Coatings to Improve the Durability of a Combustor Under a Pulse Detonation Engine Environment

    NASA Technical Reports Server (NTRS)

    Ghosn, Louis J.; Zhu, Dongming

    2008-01-01

    Pulse detonation engine (PDE) concepts are receiving increasing attention for future aeronautic propulsion applications, due to their potential thermodynamic cycle efficiency and higher thrust to density ratio that lead to the decrease in fuel consumption. But the resulting high gas temperature and pressure fluctuation distributions at high frequency generated with every detonation are viewed to be detrimental to the combustor liner material. Experimental studies on a typical metal combustion material exposed to a laser simulated pulse heating showed extensive surface cracking. Coating of the combustor materials with low thermal conductivity ceramics is shown to protect the metal substrate, reduce the thermal stresses, and hence increase the durability of the PDE combustor liner material. Furthermore, the temperature fluctuation and depth of penetration is observed to decrease with increasing the detonation frequency. A crack propagation rate in the coating is deduced by monitoring the variation of the coating apparent thermal conductivity with time that can be utilized as a health monitoring technique for the coating system under a rapid fluctuating heat flux.

  9. Development and Characterization Testing of an Air Pulsation Valve for a Pulse Detonation Engine Supersonic Parametric Inlet Test Section

    NASA Technical Reports Server (NTRS)

    Tornabene, Robert

    2005-01-01

    In pulse detonation engines, the potential exists for gas pulses from the combustor to travel upstream and adversely affect the inlet performance of the engine. In order to determine the effect of these high frequency pulses on the inlet performance, an air pulsation valve was developed to provide air pulses downstream of a supersonic parametric inlet test section. The purpose of this report is to document the design and characterization tests that were performed on a pulsation valve that was tested at the NASA Glenn Research Center 1x1 Supersonic Wind Tunnel (SWT) test facility. The high air flow pulsation valve design philosophy and analyses performed are discussed and characterization test results are presented. The pulsation valve model was devised based on the concept of using a free spinning ball valve driven from a variable speed electric motor to generate air flow pulses at preset frequencies. In order to deliver the proper flow rate, the flow port was contoured to maximize flow rate and minimize pressure drop. To obtain sharp pressure spikes the valve flow port was designed to be as narrow as possible to minimize port dwell time.

  10. Optically detonated explosive device

    NASA Technical Reports Server (NTRS)

    Yang, L. C.; Menichelli, V. J. (Inventor)

    1974-01-01

    A technique and apparatus for optically detonating insensitive high explosives, is disclosed. An explosive device is formed by containing high explosive material in a house having a transparent window. A thin metallic film is provided on the interior surface of the window and maintained in contact with the high explosive. A laser pulse provided by a Q-switched laser is focussed on the window to vaporize the metallic film and thereby create a shock wave which detonates the high explosive. Explosive devices may be concurrently or sequentially detonated by employing a fiber optic bundle to transmit the laser pulse to each of the several individual explosive devices.

  11. High-voltage pulse testing of DSSL (detonator strong safety link) actuator drive cables

    SciTech Connect

    Neau, E.L.; Laderach, G.E.

    1990-05-01

    A series of test has evaluated the current-carrying characteristics of several proposed detonator strong safety link (DSSL) drive cables to conduct current into an electrically simulated radiation case enclosure. The drive cables tested included a dielectric cable made up of either one or two braided Kevlar threads in a polysulfone sleeve, an existing DSSL conductive drive cable similar to a bicycle shift cable, with and without an external ground-strap diversion feature, and the flex circuit hook assembly used for DSSL instrumentation purposes. Each of the test samples was connected to a 200-kV, 30-kA Marx generator and terminated inside an electrically simulated radiation case enclosure. Two lengths of drive cables (4 and 13 in.) within the simulated radiation case were tested for each type. The results indicate that the conductive drive cables without an external short-circuit diversion feature conducted about 5 to 9 times as much current into the simulated radiation case as either the dielectric drive cables or the conductive drive cables with an external short-circuit feature. Surprisingly, the flex circuit hook assemblies, both the short and long cables, conducted nearly 100% of the available current into the simulated radiation case enclosure. The next series of experiments will repeat the tests at the SNL Lightning Facility in order to scale the results up to the maximum lightning threat levels of 200 kA. 18 figs., 1 tab.

  12. New detonation concepts for propulsion and power generation

    NASA Astrophysics Data System (ADS)

    Braun, Eric M.

    A series of related analytical and experimental studies are focused on utilizing detonations for emerging propulsion and power generation devices. An understanding of the physical and thermodynamic processes for this unsteady thermodynamic cycle has taken over 100 years to develop. An overview of the thermodynamic processes and development history is provided. Thermodynamic cycle analysis of detonation-based systems has often been studied using surrogate models. A real gas model is used for a thermal efficiency prediction of a detonation wave based on the work and heat specified by process path diagrams and a control volume analysis. A combined first and second law analysis aids in understanding performance trends for different initial conditions. A cycle analysis model for an airbreathing, rotating detonation wave engine (RDE) is presented. The engine consists of a steady inlet system with an isolator which delivers air into an annular combustor. A detonation wave continuously rotates around the combustor with side relief as the flow expands towards the nozzle. Air and fuel enter the combustor when the rarefaction wave pressure behind the detonation front drops to the inlet supply pressure. To create a stable RDE, the inlet pressure is matched in a convergence process with the average combustor pressure by increasing the annulus channel width with respect to the isolator channel. Performance of this engine is considered using several parametric studies. RDEs require a fuel injection system that can cycle beyond the limits of mechanical valves. Fuel injectors composed of an orifice connected to a small plenum cavity were mounted on a detonation tube. These fuel injectors, termed fluidic valves, utilize their geometry and a supply pressure to deliver fuel and contain no moving parts. Their behavior is characterized in order to determine their feasibility for integration with high-frequency RDEs. Parametric studies have been conducted with the type of fuel injected

  13. Innovative Airbreathing Propulsion Concepts for High-speed Applications

    NASA Technical Reports Server (NTRS)

    Whitlow, Woodrow, Jr.

    2002-01-01

    The current cost to launch payloads to low earth orbit (LEO) is approximately loo00 U.S. dollars ($) per pound ($22000 per kilogram). This high cost limits our ability to pursue space science and hinders the development of new markets and a productive space enterprise. This enterprise includes NASA's space launch needs and those of industry, universities, the military, and other U.S. government agencies. NASA's Advanced Space Transportation Program (ASTP) proposes a vision of the future where space travel is as routine as in today's commercial air transportation systems. Dramatically lower launch costs will be required to make this vision a reality. In order to provide more affordable access to space, NASA has established new goals in its Aeronautics and Space Transportation plan. These goals target a reduction in the cost of launching payloads to LEO to $lo00 per pound ($2200 per kilogram) by 2007 and to $100' per pound by 2025 while increasing safety by orders of magnitude. Several programs within NASA are addressing innovative propulsion systems that offer potential for reducing launch costs. Various air-breathing propulsion systems currently are being investigated under these programs. The NASA Aerospace Propulsion and Power Base Research and Technology Program supports long-term fundamental research and is managed at GLenn Research Center. Currently funded areas relevant to space transportation include hybrid hyperspeed propulsion (HHP) and pulse detonation engine (PDE) research. The HHP Program currently is addressing rocket-based combined cycle and turbine-based combined cycle systems. The PDE research program has the goal of demonstrating the feasibility of PDE-based hybrid-cycle and combined cycle propulsion systems that meet NASA's aviation and access-to-space goals. The ASTP also is part of the Base Research and Technology Program and is managed at the Marshall Space Flight Center. As technologies developed under the Aerospace Propulsion and Power Base

  14. Integrated System Test of an Airbreathing Rocket

    NASA Technical Reports Server (NTRS)

    Mack, Gregory; Beaudry, Charles; Ketchum, Andrew; McArthur, J. Craig (Technical Monitor)

    2002-01-01

    This viewgraph presentation provides information on NASA's attempts to develop an air-breathing propulsion in an effort to make future space transportation safer, more reliable and significantly less expensive than today's missions. Spacecraft powered by air-breathing rocket engines would be completely reusable, able to take off and land at airport runways and ready to fly again within days. A radical new engine project is called the Integrated System Tests of an Air-breathing Rocket, or ISTAR.

  15. Reverse slapper detonator

    DOEpatents

    Weingart, Richard C.

    1990-01-01

    A reverse slapper detonator (70), and methodology related thereto, are provided. The detonator (70) is adapted to be driven by a pulse of electric power from an external source (80). A conductor (20) is disposed along the top (14), side (18), and bottom (16) surfaces of a sheetlike insulator (12). Part of the conductor (20) comprises a bridge (28), and an aperture (30) is positioned within the conductor (20), with the bridge (28) and the aperture (30) located on opposite sides of the insulator (12). A barrel (40) and related explosive charge (50) are positioned adjacent to and in alignment with the aperture (30), and the bridge (28) is buttressed with a backing layer (60). When the electric power pulse vaporizes the bridge (28), a portion of the insulator (12) is propelled through the aperture (30) and barrel (40), and against the explosive charge (50), thereby detonating it.

  16. Airbreathing Acceleration Toward Earth Orbit

    SciTech Connect

    Whitehead, J C

    2007-05-09

    As flight speed increases, aerodynamic drag rises more sharply than the availability of atmospheric oxygen. The ratio of oxygen mass flux to dynamic pressure cannot be improved by changing altitude. The maximum possible speed for airbreathing propulsion is limited by the ratio of air capture area to vehicle drag area, approximately Mach 6 at equal areas. Simulation of vehicle acceleration shows that the use of atmospheric oxygen offers a significant potential for minimizing onboard consumables at low speeds. These fundamental calculations indicate that a practical airbreathing launch vehicle would accelerate to near steady-state speed while consuming only onboard fuel, then transition to rocket propulsion. It is suggested that an aircraft carrying a rocket-propelled vehicle to approximately Mach 5 could be a realistic technical goal toward improving access to orbit.

  17. Air-Breathing Rocket Engines

    NASA Technical Reports Server (NTRS)

    1998-01-01

    This photograph depicts an air-breathing rocket engine prototype in the test bay at the General Applied Science Lab facility in Ronkonkoma, New York. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced Space Transportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  18. Hypersonic Airbreathing Vehicles/Technologies

    NASA Technical Reports Server (NTRS)

    Hunt, James L.

    1996-01-01

    Hypersonic airbreathing horizontal takeoff and landing (HTOL) vehicles are highly integrated systems involving many advanced technologies. The design environment is variable rich, intricately networked, and sensitivity intensive; as such, it represents a tremendous challenge. Creating a viable design requires addressing three main elements: (1) an understanding of the 'figures of merit' and their relationship, (2) the development of sophisticated configuration discipline prediction methods and a synthesis procedure, and (3) the synergistic integration of advanced technologies across the discipline spectrum. This paper will focus on the vision for hypersonic airbreathing vehicles and the advanced technologies that forge the designs. Airbreathing hypersonics encompass endoatmospheric (airplanes...missiles are a part of the matrix but will not be included in this paper since they are an air force focus) and space access vehicles with speed from Mach 4 up to Mach 25 (orbital). These vehicles can be divided into two classes...cruisers and accelerators. The cruiser designs reflect high lift-to-drag whereas the accelerators reflect low drag per unit inlet capture; thus, the cross section of the accelerator attributes a much larger percentage to propulsion. One of the more design influencing items is fuel. The hydrogen fueled vehicles must be very volumetric efficient to contain the low density fuel and thus tend to be a bit bulgy (more conducive to lifting bodies or wing bodies) whereas with hydrocarbon fueled vehicles, the concern is loading because of the high density fuel; thus, they may tend to be more towards waveriders which are not usually very volumetric efficient. Hydrocarbon fuels (endothermic) are limited in their engine cooling capacity to below Mach 8.

  19. Airbreathing engines for space shuttle

    NASA Technical Reports Server (NTRS)

    Glassman, A. J.; Stewart, W. L.; Nosek, S. M.

    1972-01-01

    The requirements imposed on the airbreathing engines by the shuttle mission and some results from engine design studies are discussed. In particular, some of the engine system weight study results are presented, potential problem areas and required engine modifications are identified, and testing requirements for a development and qualification program are discussed. The engines of interest for the shuttle are engines that are currently being developed for other applicatons. The potential problems, engine modifications, and testing requirements result primarily from the new environments associated with launch, space residence, and reentry.

  20. Detonator comprising a nonlinear transmission line

    SciTech Connect

    Elizondo-Decanini, Juan M

    2014-12-30

    Detonators are described herein. In a general embodiment, the detonator includes a nonlinear transmission line that has a variable capacitance. Capacitance of the nonlinear transmission line is a function of voltage on the nonlinear transmission line. The nonlinear transmission line receives a voltage pulse from a voltage source and compresses the voltage pulse to generate a trigger signal. Compressing the voltage pulse includes increasing amplitude of the voltage pulse and decreasing length of the voltage pulse in time. An igniter receives the trigger signal and detonates an explosive responsive to receipt of the trigger signal.

  1. Development of a numerical tool to study the mixing phenomenon occurring during mode one operation of a multi-mode ejector-augmented pulsed detonation rocket engine

    NASA Astrophysics Data System (ADS)

    Dawson, Joshua

    A novel multi-mode implementation of a pulsed detonation engine, put forth by Wilson et al., 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 Ejector-Augmented Pulsed Detonation Rocket Engine operates in mode one during take-off 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 research paper. Several advantages can be obtained by use of this technology. Geometrically the propulsion system is fairly

  2. Measurements of multiple gas parameters in a pulsed-detonation combustor using time-division-multiplexed Fourier-domain mode-locked lasers.

    PubMed

    Caswell, Andrew W; Roy, Sukesh; An, Xinliang; Sanders, Scott T; Schauer, Frederick R; Gord, James R

    2013-04-20

    Hyperspectral absorption spectroscopy is being used to monitor gas temperature, velocity, pressure, and H(2)O mole fraction in a research-grade pulsed-detonation combustor (PDC) at the Air Force Research Laboratory. The hyperspectral source employed is termed the TDM 3-FDML because it consists of three time-division-multiplexed (TDM) Fourier-domain mode-locked (FDML) lasers. This optical-fiber-based source monitors sufficient spectral information in the H(2)O absorption spectrum near 1350 nm to permit measurements over the wide range of conditions encountered throughout the PDC cycle. Doppler velocimetry based on absorption features is accomplished using a counterpropagating beam approach that is designed to minimize common-mode flow noise. The PDC in this study is operated in two configurations: one in which the combustion tube exhausts directly to the ambient environment and another in which it feeds an automotive-style turbocharger to assess the performance of a detonation-driven turbine. Because the enthalpy flow [kilojoule/second] is important in assessing the performance of the PDC in various configurations, it is calculated from the measured gas properties. PMID:23669701

  3. Wavelength-agile diode-laser sensing strategies for monitoring gas properties in optically harsh flows: application in cesium-seeded pulse detonation.

    PubMed

    Sanders, Scott; Mattison, Daniel; Ma, Lin; Jeffries, Jay; Hanson, Ronald

    2002-06-17

    The rapid, broad wavelength scanning capabilities of advanced diode lasers allow extension of traditional diode-laser absorption techniques to high pressure, transient, and generally hostile environments. Here, we demonstrate this extension by applying a vertical cavity surface-emitting laser (VCSEL) to monitor gas temperature and pressure in a pulse detonation engine (PDE). Using aggressive injection current modulation, the VCSEL is scanned through a 10 cm-1 spectral window at megahertz rates - roughly 10 times the scanning range and 1000 times the scanning rate of a conventional diode laser. The VCSEL probes absorption lineshapes of the ~ 852 nm D2 transition of atomic Cs, seeded at ~ 5 ppm into the feedstock gases of a PDE. Using these lineshapes, detonated-gas temperature and pressure histories, spanning 2000 - 4000 K and 0.5 - 30 atm, respectively, are recorded with microsecond time response. The increasing availability of wavelength-agile diode lasers should support the development of similar sensors for other harsh flows, using other absorbers such as native H2O. PMID:19436388

  4. Exhaust Nozzle for a Multitube Detonative Combustion Engine

    NASA Technical Reports Server (NTRS)

    Bratkovich, Thomas E.; Williams, Kevin E.; Bussing, Thomas R. A.; Lidstone, Gary L.; Hinkey, John B.

    2004-01-01

    An improved type of exhaust nozzle has been invented to help optimize the performances of multitube detonative combustion engines. The invention is applicable to both air-breathing and rocket engines used to propel some aircraft and spacecraft, respectively. In a detonative combustion engine, thrust is generated through the expulsion of combustion products from a detonation process in which combustion takes place in a reaction zone coupled to a shock wave. The combustion releases energy to sustain the shock wave, while the shock wave enhances the combustion in the reaction zone. The coupled shockwave/reaction zone, commonly referred to as a detonation, propagates through the reactants at very high speed . typically of the order of several thousands of feet per second (of the order of 1 km/s). The very high speed of the detonation forces combustion to occur very rapidly, thereby contributing to high thermodynamic efficiency. A detonative combustion engine of the type to which the present invention applies includes multiple parallel cylindrical combustion tubes, each closed at the front end and open at the rear end. Each tube is filled with a fuel/oxidizer mixture, and then a detonation wave is initiated at the closed end. The wave propagates rapidly through the fuel/oxidizer mixture, producing very high pressure due to the rapid combustion. The high pressure acting on the closed end of the tube contributes to forward thrust. When the detonation wave reaches the open end of the tube, it produces a blast wave, behind which the high-pressure combustion products are expelled from the tube. The process of filling each combustion tube with a detonable fuel/oxidizer mixture and then producing a detonation repeated rapidly to obtain repeated pulses of thrust. Moreover, the multiple combustion tubes are filled and fired in a repeating sequence. Hence, the pressure at the outlet of each combustion tube varies cyclically. A nozzle of the present invention channels the

  5. Optically triggered fire set/detonator system

    DOEpatents

    Chase, Jay B.; Pincosy, Philip A.; Chato, Donna M.; Kirbie, Hugh; James, Glen F.

    2007-03-20

    The present invention is directed to a system having a plurality of capacitor discharge units (CDUs) that includes electrical bridge type detonators operatively coupled to respective explosives. A pulse charging circuit is adapted to provide a voltage for each respective capacitor in each CDU. Such capacitors are discharged through the electrical bridge type detonators upon receiving an optical signal to detonate respective operatively coupled explosives at substantially the same time.

  6. The microspace launcher: first step to the fully air-breathing space launcher

    NASA Astrophysics Data System (ADS)

    Falempin, F.; Bouchez, M.; Calabro, M.

    2009-09-01

    booster, the main stage ensuring the initial acceleration in liquid rocket mode and a complementary acceleration phase in rocket mode beyond the air-breathing propulsion system operation. Finally, the liquid rocket engine of this third variant can be replaced by a continuous detonation wave rocket engine. The paper describes the main guidelines for the design of these variants and provides their main characteristics. On this basis, the achievable performance, estimated by trajectory simulation, are detailed.

  7. Environmentally Benign Stab Detonators

    SciTech Connect

    Gash, A

    2005-12-21

    Many energetic systems can be activated via mechanical means. Percussion primers in small caliber ammunition and stab detonators used in medium caliber ammunition are just two examples. Current medium caliber (20-60mm) munitions are detonated through the use of impact sensitive stab detonators. Stab detonators are very sensitive and must be small, as to meet weight and size limitations. A mix of energetic powders, sensitive to mechanical stimulus, is typically used to ignite such devices. Stab detonators are mechanically activated by forcing a firing pin through the closure disc of the device and into the stab initiating mix. Rapid heating caused by mechanically driven compression and friction of the mixture results in its ignition. The rapid decomposition of these materials generates a pressure/temperature pulse that is sufficient to initiate a transfer charge, which has enough output energy to detonate the main charge. This general type of ignition mix is used in a large variety of primers, igniters, and detonators.[1] Common primer mixes, such as NOL-130, are made up of lead styphnate (basic) 40%, lead azide (dextrinated) 20%, barium nitrate 20%, antimony sulfide 15%, and tetrazene 5%.[1] These materials pose acute and chronic toxicity hazards during mixing of the composition and later in the item life cycle after the item has been field functioned. There is an established need to replace these mixes on toxicity, health, and environmental hazard grounds. This effort attempts to demonstrate that environmentally acceptable energetic solgel coated flash metal multilayer nanocomposites can be used to replace current impact initiated devices (IIDs), which have hazardous and toxic components. Successful completion of this project will result in IIDs that include innocuous compounds, have sufficient output energy for initiation, meet current military specifications, are small, cost competitive, and perform as well as or better than current devices. We expect flash

  8. Detonating apparatus

    DOEpatents

    Johnston, Lawrence H.

    1976-01-01

    1. Apparatus for detonation of high explosive in uniform timing comprising in combination, an outer case, spark gap electrodes insulatedly supported in spaced relationship within said case to form a spark gap, high explosive of the class consisting of pentaerythritol tetranitrate and trimethylene trinitramine substantially free from material sensitive to detonation by impact compressed in surrounding relation to said electrodes including said spark gap under a pressure from about 100 psi to about 500 psi, said spark gap with said compressed explosive therein requiring at least 1000 volts for sparking, and means for impressing at least 1000 volts on said spark gap.

  9. Guidance and control of an earth to orbit vehicle with optimum transition from airbreathing to concurrent airbreathing/rocket propulsion

    NASA Astrophysics Data System (ADS)

    Kauffman, H. G.; Grandhi, R. V.; Hankey, W. L.; Belcher, P. J.

    1990-07-01

    An efficient performance analysis method (suitable for PC operation) is developed to evaluate potential airbreathing/rocket propulsion systems for advanced technology single-stage-to-orbit (SSTO) launch vehicles. Evaluated are tradeoffs between airbreathing (AB), rocket, and concurrent airbreathing/rocket propulsion in minimizing fuel consumption for a given ascent flight trajectory. Many mission, flight, and vehicle related requirements and constraints are satisfied in the process. With the analysis method, several modes of airbreathing/rocket propulsion are compared to a baseline 'airbreather alone' propulsion system in terms of fuel required to attain orbital velocity. The optimal switch point (staging) is identified for the transition from airbreathing to rocket propulsion.

  10. Experimental Investigation of a Multi-Cycle Single-Tube Pulse Detonation Rocket Engine with a Coaxial Rotary Valve

    NASA Astrophysics Data System (ADS)

    Matsuoka, Ken; Esumi, Motoki; Ikeguchi, Ken Bryan; Kasahara, Jiro; Matsuo, Akiko; Funaki, Ikkoh

    We developed a novel coaxial rotary valve for a multi-tube PDE. Since this single valve can supply three different gases (fuel, oxidizer and purge gas) into a combustor, the unification of the valve systems for three different gases is possible by using our newly designed valve. A PDRE system can be simple and lightweight by using this valve, and thus its thrust-weight ratio can be increased. We proposed the design of a multi-tube rotary-valved PDRE system by this rotary valve. Moreover, in preparation for a multi-tube rotary-valved PDRE, we carried out the multi-cycle operation experiment by the single-tube rotary-valved PDRE system. The combustion wave velocity was measured to confirm the operation of the PDRE system. Deflagration-to-detonation transition (DDT) was confirmed and DDT distance decreased under the condition of high operation frequency. In addition, a maximum operation frequency was 159 Hz.

  11. Effect of Detonation through a Turbine Stage

    NASA Technical Reports Server (NTRS)

    Ellis, Matthew T.

    2004-01-01

    Pulse detonation engines (PDE) have been investigated as a more efficient means of propulsion due to its constant volume combustion rather than the more often used constant pressure combustion of other propulsion systems. It has been proposed that a hybrid PDE-gas turbine engine would be a feasible means of improving the efficiency of the typical constant pressure combustion gas turbine cycle. In this proposed system, multiple pulse detonation tubes would replace the conventional combustor. Also, some of the compressor stages may be removed due to the pressure rise gained across the detonation wave. The benefits of higher thermal efficiency and reduced compressor size may come at a cost. The first question that arises is the unsteadiness in the flow created by the pulse detonation tubes. A constant pressure combustor has the advantage of supplying a steady and large mass flow rate. The use of the pulse detonation tubes will create an unsteady mass flow which will have currently unknown effects on the turbine located downstream of the combustor. Using multiple pulse detonation tubes will hopefully improve the unsteadiness. The interaction between the turbine and the shock waves exiting the tubes will also have an unknown effect. Noise levels are also a concern with this hybrid system. These unknown effects are being investigated using TURBO, an unsteady turbomachinery flow simulation code developed at Mississippi State University. A baseline case corresponding to a system using a constant pressure combustor with the same mass flow rate achieved with the pulse detonation hybrid system will be investigated first.

  12. Prompt detonation of secondary explosives by laser

    SciTech Connect

    Paisley, D.L.

    1989-01-01

    Secondary high explosives have been promptly detonated by directing a laser beam of various wavelengths from 266 nanometers to 1.06 micron on the surface of the explosives. For this paper ''prompt'' means the excess transit time through an explosive charge is /approximately/250 nanoseconds (or less) less than the accepted full detonation velocity time. Timing between laser pulse, explosive initiation and detonation velocity and function time have been recorded. The laser parameters studied include: wavelength, pulse length, energy and power density, and beam diameter (spot size). Explosives evaluated include: PETN, HNS, HMX, and graphited PETN, HNS, and HMX. Explosive parameters that have been correlated with optical parameters include: density, surface area, critical diameter (spot size), spectral characteristics and enhance absorption. Some explosives have been promptly detonated over the entire range of wavelengths, possibly by two competing initiating mechanisms. Other explosives could not be detonated at any of the wavelengths or power densities tested. 8 refs., 12 figs., 1 tab.

  13. Flight testing of airbreathing hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Hicks, John W.

    1993-01-01

    Using the scramjet engine as the prime example of a hypersonic airbreathing concept, this paper reviews the history of and addresses the need for hypersonic flight tests. It also describes how such tests can contribute to the development of airbreathing technology. Aspects of captive-carry and free-flight concepts are compared. An incremental flight envelope expansion technique for manned flight vehicles is also described. Such critical issues as required instrumentation technology and proper scaling of experimental devices are addressed. Lastly, examples of international flight test approaches, existing programs, or concepts currently under study, development, or both, are given.

  14. Effects of Fuel Distribution on Detonation Tube Performance

    NASA Technical Reports Server (NTRS)

    Perkins, H. Douglas; Sung, Chih-Jen

    2003-01-01

    A pulse detonation engine uses a series of high frequency intermittent detonation tubes to generate thrust. The process of filling the detonation tube with fuel and air for each cycle may yield non-uniform mixtures. Uniform mixing is commonly assumed when calculating detonation tube thrust performance. In this study, detonation cycles featuring idealized non-uniform Hz/air mixtures were analyzed using a two-dimensional Navier-Stokes computational fluid dynamics code with detailed chemistry. Mixture non-uniformities examined included axial equivalence ratio gradients, transverse equivalence ratio gradients, and partially fueled tubes. Three different average test section equivalence ratios were studied; one stoichiometric, one fuel lean, and one fuel rich. All mixtures were detonable throughout the detonation tube. Various mixtures representing the same average test section equivalence ratio were shown to have specific impulses within 1% of each other, indicating that good fuel/air mixing is not a prerequisite for optimal detonation tube performance under conditions investigated.

  15. Detonation wave compression in gas turbines

    NASA Technical Reports Server (NTRS)

    Wortman, A.

    1986-01-01

    A study was made of the concept of augmenting the performance of low pressure ratio gas turbines by detonation wave compression of part of the flow. The concept exploits the constant volume heat release of detonation waves to increase the efficiency of the Brayton cycle. In the models studied, a fraction of the compressor output was channeled into detonation ducts where it was processed by transient transverse detonation waves. Gas dynamic studies determined the maximum cycling frequency of detonation ducts, proved that upstream propagation of pressure pulses represented no problems and determined the variations of detonation duct output with time. Mixing and wave compression were used to recombine the combustor and detonation duct flows and a concept for a spiral collector to further smooth the pressure and temperature pulses was presented as an optional component. The best performance was obtained with a single firing of the ducts so that the flow could be re-established before the next detonation was initiated. At the optimum conditions of maximum frequency of the detonation ducts, the gas turbine efficiency was found to be 45 percent while that of a corresponding pressure ratio 5 conventional gas turbine was only 26%. Comparable improvements in specific fuel consumption data were found for gas turbines operating as jet engines, turbofans, and shaft output machines. Direct use of the detonation duct output for jet propulsion proved unsatisfactory. Careful analysis of the models of the fluid flow phenomena led to the conclusion that even more elaborate calculations would not diminish the uncertainties in the analysis of the system. Feasibility of the concept to work as an engine now requires validation in an engineering laboratory experiment.

  16. Airbreathing hypersonic vehicle design and analysis methods

    NASA Technical Reports Server (NTRS)

    Lockwood, Mary Kae; Petley, Dennis H.; Hunt, James L.; Martin, John G.

    1996-01-01

    The design, analysis, and optimization of airbreathing hypersonic vehicles requires analyses involving many highly coupled disciplines at levels of accuracy exceeding those traditionally considered in a conceptual or preliminary-level design. Discipline analysis methods including propulsion, structures, thermal management, geometry, aerodynamics, performance, synthesis, sizing, closure, and cost are discussed. Also, the on-going integration of these methods into a working environment, known as HOLIST, is described.

  17. Supersonic Air-Breathing Stage For Commercial Launch Rocket

    NASA Technical Reports Server (NTRS)

    Martin, James A.

    1993-01-01

    Concept proposed to expand use of air-breathing, reusable stages to put more payload into orbit at less cost. Stage with supersonic air-breathing engines added to carry expendable stages from subsonic airplane to supersonic velocity. Carry payload to orbit. Expendable stages and payload placed in front of supersonic air-breathing stage. After releasing expendable stages, remotely piloted supersonic air-breathing stage returns to takeoff site and land for reuse. New concept extends use of low-cost reusable hardware and increases payload delivered from B-52.

  18. Optimal air-breathing launch vehicle design

    NASA Technical Reports Server (NTRS)

    Hattis, P. D.

    1981-01-01

    A generalized two-point boundary problem methodology, similar to techniques used in deterministic optimal control studies, is applied to the design and flight analysis of a two-stage air-breathing launch vehicle. Simultaneous consideration is given to configuration and trajectory by treating geometry, dynamic discontinuities, and time-dependent flight variables all as controls to be optimized with respect to a single mathematical performance measure. While minimizing fuel consumption, inequality constraints are applied to dynamic pressure and specific force. The optimal system fuel consumption and staging Mach number are found to vary little with changes in the inequality constraints due to substantial geometry and trajectory adjustments. Staging, from an air-breathing first stage to a rocket-powered second stage, consistently occurs near Mach 3.5. The dynamic pressure bound has its most pronounced effects on vehicle geometry, particularly the air-breathing propulsion inlet area, and on the first-stage altitude profile. The specific force has its greatest influence on the second-stage thrust history.

  19. Air-breathing Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1999-01-01

    This Quick Time movie depicts the Rocketdyne static test of an air-breathing rocket. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's advanced Transportation Program at the Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  20. Cellular structure of detonation utilized in propulsion system

    NASA Astrophysics Data System (ADS)

    Zhang, XuDong; Fan, BaoChun; Gui, MingYue; Pan, ZhenHua

    2012-10-01

    How to confine a detonation in a combustor is a key issue of detonation applications in propulsion systems. Based on achieving schemes, detonations applied in the combustor, including pulse detonation wave (PDW), oblique detonation wave (ODW) and rotating detonation wave (RDW), are different from that described by the classic CJ theory in fine structures and its self-sustaining mechanisms. In this work, the cellular structures and flow fields of ODW and RDW were obtained numerically, and the fundamental characteristics and self-sustaining mechanisms of the detonations were analyzed and discussed. ODW front consists of three parts: the ZND-like front, the single-headed triple point front and the dual-headed triple point front. Cellular structures of RDW are heterogeneous, and the cell size near the outer wall is smaller than that near the inner wall.

  1. The flow field in a rotating detonation-wave engine

    NASA Astrophysics Data System (ADS)

    Kailasanath, Kazhikathra; Schwer, Douglas

    2011-11-01

    Rotating detonation-wave engines (RDE) are a form of continuous detonation-wave engine. They potentially provide further gains than an intermittent or pulsed detonation-wave engine (PDE). However, significantly less work has been on this concept when compared to the PDE. In this talk, we present the detailed flow field in an idealized RDE, primarily consisting of two concentric cylinders. A premixed detonable mixture is injected into the annulus between the two concentric cylinders. Once a detonation is initiated, it keeps travelling around in the annulus as long as there is fresh detonable mixture ahead of it. Hence, the injection process is critically important to the stability and performance of the RDE. Furthermore, we show that the flow field is quite complex consisting of multiple shock waves and the outflow is primarily axial, although the detonation-wave is travelling around circumferentially. Sponsored by the NRL 6.1 Computational Physics Task Area.

  2. Continuous detonation reaction engine

    NASA Technical Reports Server (NTRS)

    Lange, O. H.; Stein, R. J.; Tubbs, H. E.

    1968-01-01

    Reaction engine operates on the principles of a controlled condensed detonation rather than on the principles of gas expansion. The detonation results in reaction products that are expelled at a much higher velocity.

  3. Airbreathing/Rocket Single-Stage-to-Orbit Design Matrix

    NASA Technical Reports Server (NTRS)

    Hunt, James L.

    1995-01-01

    A definitive design/performance study was performed on a single-stage-to-orbit (SSTO) airbreathing propelled orbital vehicle with rocket propulsion augmentation in the Access to Space activities during 1993. A credible reference design was established, but by no means an optimum. The results supported the viability of SSTO airbreathing/rocket vehicles for operational scenarios and indicated compelling reasons to continue to explore the design matrix. This paper will (1) summarize the Access to Space design activity from the SSTO airbreathing/rocket perspective, (2) present an airbreathing/rocket SSTO design matrix established for continued optimization of the design space, and (3) focus on the compelling reasons for airbreathing vehicles in Access to Space scenarios.

  4. Autoignitions and detonations in engines and ducts.

    PubMed

    Bradley, Derek

    2012-02-13

    The origins of autoignition at hot spots are analysed and the pressure pulses that arise from them are related to knock in gasoline engines and to developing detonations in ducts. In controlled autoignition engines, autoignition is benign with little knock. There are several modes of autoignition and the existence of an operational peninsula, within which detonations can develop at a hot spot, helps to explain the performance of various engines. Earlier studies by Urtiew and Oppenheim of the development of autoignitions and detonations ahead of a deflagration in ducts are interpreted further, using a simple one-dimensional theory of the generation of shock waves ahead of a turbulent flame. The theory is able to indicate entry into the domain of autoignition in an 'explosion in the explosion'. Importantly, it shows the influence of the turbulent burning velocity, and particularly its maximum attainable value, upon autoignition. This value is governed by localized flame extinctions for both turbulent and laminar flames. The theory cannot show any details of the transition to a detonation, but regimes of eventually stable or unstable detonations can be identified on the operational peninsula. Both regimes exhibit transverse waves, triple points and a cellular structure. In the case of unstable detonations, transverse waves are essential to the continuing propagation. For hazard assessment, more needs to be known about the survival, or otherwise, of detonations that emerge from a duct into the same mixture at atmospheric pressure. PMID:22213665

  5. Air-Breathing Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    2000-01-01

    This photograph depicts an air-breathing rocket engine that completed an hour or 3,600 seconds of testing at the General Applied Sciences Laboratory in Ronkonkoma, New York. Referred to as ARGO by its design team, the engine is named after the mythological Greek ship that bore Jason and the Argonauts on their epic voyage of discovery. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced SpaceTransportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  6. Hypersonic airbreathing vehicle visions and enhancing technologies

    SciTech Connect

    Hunt, J.L.; Lockwood, M.K.; Petley, D.H.; Pegg, R.J.

    1997-01-01

    This paper addresses the visions for hypersonic airbreathing vehicles and the advanced technologies that forge and enhance the designs. The matrix includes space access vehicles (single-stage-to-orbit (SSTO), two-stage-to-orbit (2STO) and three-stage-to-orbit (3STO)) and endoatmospheric vehicles (airplanes{emdash}missiles are omitted). The characteristics, the performance potential, the technologies and the synergies will be discussed. A common design constraint is that all vehicles (space access and endoatmospheric) have enclosed payload bays. {copyright} {ital 1997 American Institute of Physics.}

  7. High temperature detonator

    DOEpatents

    Johnson, James O.; Dinegar, Robert H.

    1988-01-01

    A detonator assembly is provided which is usable at high temperatures about 300.degree. C. A detonator body is provided with an internal volume defining an anvil surface. A first acceptor explosive is disposed on the anvil surface. A donor assembly having an ignition element, an explosive material, and a flying plate, are placed in the body effective to accelerate the flying plate to impact the first acceptor explosive on the anvil for detonating the first acceptor explosive. A second acceptor explosive is eccentrically located in detonation relationship with the first acceptor explosive to thereafter effect detonation of a main charge.

  8. Detonation Wave Profile

    SciTech Connect

    Menikoff, Ralph

    2015-12-14

    The Zel’dovich-von Neumann-Doering (ZND) profile of a detonation wave is derived. Two basic assumptions are required: i. An equation of state (EOS) for a partly burned explosive; P(V, e, λ). ii. A burn rate for the reaction progress variable; d/dt λ = R(V, e, λ). For a steady planar detonation wave the reactive flow PDEs can be reduced to ODEs. The detonation wave profile can be determined from an ODE plus algebraic equations for points on the partly burned detonation loci with a specified wave speed. Furthermore, for the CJ detonation speed the end of the reaction zone is sonic. A solution to the reactive flow equations can be constructed with a rarefaction wave following the detonation wave profile. This corresponds to an underdriven detonation wave, and the rarefaction is know as a Taylor wave.

  9. Optimization of Air-Breathing Engine Concept

    NASA Technical Reports Server (NTRS)

    Patnaik, Surya N.; Lavelle, Thomas M.; Hopkins, Dale A.

    1996-01-01

    The design optimization of air-breathing propulsion engine concepts has been accomplished by soft-coupling the NASA Engine Performance Program (NEPP) analyzer with the NASA Lewis multidisciplinary optimization tool COMETBOARDS. Engine problems, with their associated design variables and constraints, were cast as nonlinear optimization problems with thrust as the merit function. Because of the large number of mission points in the flight envelope, the diversity of constraint types, and the overall distortion of the design space; the most reliable optimization algorithm available in COMETBOARDS, when used by itself, could not produce satisfactory, feasible, optimum solutions. However, COMETBOARDS' unique features-which include a cascade strategy, variable and constraint formulations, and scaling devised especially for difficult multidisciplinary applications-successfully optimized the performance of subsonic and supersonic engine concepts. Even when started from different design points, the combined COMETBOARDS and NEPP results converged to the same global optimum solution. This reliable and robust design tool eliminates manual intervention in the design of air-breathing propulsion engines and eases the cycle analysis procedures. It is also much easier to use than other codes, which is an added benefit. This paper describes COMETBOARDS and its cascade strategy and illustrates the capabilities of the combined design tool through the optimization of a high-bypass- turbofan wave-rotor-topped subsonic engine and a mixed-flow-turbofan supersonic engine.

  10. Transition to Detonation in Exploding Bridgewire Detonators

    SciTech Connect

    Roeski, F; Benterou, J; Lee, R; Roos, R

    2003-01-08

    We are investigating using breakout profile measurements and/or Fabry-Perot velocimeter measurements during early stages of initiation in Exploding Bridge Wire (EBW) detonators as a tool for understanding the physics of initiation of these devices and as a tool for monitoring aging effects. We believe any changes due to aging may be more readily observed in the very early stages of the initiation. We have developed a method that allows measurement of the detonation velocity, detonation profile and interface velocity as a function of distance from the bridgewire.

  11. Development of a laser ignited all secondary explosive DDT detonator

    SciTech Connect

    Woods, C.M.; Spangler, E.M.; Beckman, T.M.; Kramer, D.P.

    1992-09-01

    A hermetic, stand alone, laser-ignited deflagration-to-detonation transition (DDT) detonator has been developed. The detonator uses the secondary explosive HMX (cyclotetramethylenetetranitramine) and was originally developed for use with the US Navy`s Laser Initiated Transfer Energy subsystem (LITES). The design of the new detonator allows for its use with a variety of laser fire sets. A high density blend 6f 830 cm2/g HMX with 3% carbon black by weight was used for the ignition charge. Deflagration-to-detonation transition was achieved with a transition charge of undoped 830 cm2/g HMX. Using a 12-ms ND-YAG laser pulse coupled to the detonator via a 1-mm diameter optical fiber, the ignition threshold was determined to be approximately 30 mJ. Full detonating outputs were demonstrated by function testing in Navy detonation sensitivity fittings. Finally, the detonator was tested in several configurations with shielded mild detonating chord (SNMC) endrips to determine its capability for lighting transfer lines.

  12. Isothermal Gaseous Detonation Model

    NASA Astrophysics Data System (ADS)

    Prokhorov, E. S.

    2015-05-01

    We propose an isothermal gaseous detonation model taking into account the initial pressure of the explosive mixture that permits describing in a simplified form both the self-sustaining and the supercompressed and undercompressed detonation regimes. The exactness of this model has been estimated on the basis of a comparative analysis with the results of equilibrium calculations of the gas-dynamic parameters at the front of detonation waves.

  13. Air-Breathing Launch Vehicle Technology Being Developed

    NASA Technical Reports Server (NTRS)

    Trefny, Charles J.

    2003-01-01

    Of the technical factors that would contribute to lowering the cost of space access, reusability has high potential. The primary objective of the GTX program is to determine whether or not air-breathing propulsion can enable reusable single-stage-to-orbit (SSTO) operations. The approach is based on maturation of a reference vehicle design with focus on the integration and flight-weight construction of its air-breathing rocket-based combined-cycle (RBCC) propulsion system.

  14. Airbreathing Hypersonic Technology Vision Vehicles and Development Dreams

    NASA Technical Reports Server (NTRS)

    McClinton, C. R.; Hunt, J. L.; Ricketts, R. H.; Reukauf, P.; Peddie, C. L.

    1999-01-01

    Significant advancements in hypersonic airbreathing vehicle technology have been made in the country's research centers and industry over the past 40 years. Some of that technology is being validated with the X-43 flight tests. This paper presents an overview of hypersonic airbreathing technology status within the US, and a hypersonic technology development plan. This plan builds on the nation's large investment in hypersonics. This affordable, incremental plan focuses technology development on hypersonic systems, which could be operating by the 2020's.

  15. Hydrazine vapor detonations

    NASA Technical Reports Server (NTRS)

    Pedley, M. D.; Bishop, C. V.; Benz, F. J.; Bennett, C. A.; Mcclenagan, R. D.

    1988-01-01

    The detonation velocity and cell widths for hydrazine decomposition were measured over a wide range of temperatures and pressures. The detonation velocity in pure hydrazine was within 5 percent of the calculated C-J velocity. The detonation cell width measurements were interpreted using the Zeldovich-Doering-von Neumann model with a detailed reaction mechanism for hydrazine decomposition. Excellent agreement with experimental data for pure hydrazine was obtained using the empirical relation that detonation cell width was equal to 29 times the kinetically calculated reaction zone length.

  16. Detonation command and control

    DOEpatents

    Mace, Jonathan L.; Seitz, Gerald J.; Echave, John A.; Le Bas, Pierre-Yves

    2016-05-31

    The detonation of one or more explosive charges and propellant charges by a detonator in response to a fire control signal from a command and control system comprised of a command center and instrumentation center with a communications link there between. The fire control signal is selectively provided to the detonator from the instrumentation center if plural detonation control switches at the command center are in a fire authorization status, and instruments, and one or more interlocks, if included, are in a ready for firing status. The instrumentation and command centers are desirably mobile, such as being respective vehicles.

  17. Detonation command and control

    DOEpatents

    Mace, Jonathan L.; Seitz, Gerald J.; Echave, John A.; Le Bas, Pierre-Yves

    2015-11-10

    The detonation of one or more explosive charges and propellant charges by a detonator in response to a fire control signal from a command and control system comprised of a command center and instrumentation center with a communications link therebetween. The fire control signal is selectively provided to the detonator from the instrumentation center if plural detonation control switches at the command center are in a fire authorization status, and instruments, and one or more interlocks, if included, are in a ready for firing status. The instrumentation and command centers are desirably mobile, such as being respective vehicles.

  18. [Research on diagnosis of gas-liquid detonation exhaust based on double optical path absortion spectroscopy technique].

    PubMed

    Lü, Xiao-Jing; Li, Ning; Weng, Chun-Sheng

    2014-03-01

    The effect detection of detonation exhaust can provide measurement data for exploring the formation mechanism of detonation, the promotion of detonation efficiency and the reduction of fuel waste. Based on tunable diode laser absorption spectroscopy technique combined with double optical path cross-correlation algorithm, the article raises the diagnosis method to realize the on-line testing of detonation exhaust velocity, temperature and H2O gas concentration. The double optical path testing system is designed and set up for the valveless pulse detonation engine with the diameter of 80 mm. By scanning H2O absorption lines of 1343nm with a high frequency of 50 kHz, the on-line detection of gas-liquid pulse detonation exhaust is realized. The results show that the optical testing system based on tunable diode laser absorption spectroscopy technique can capture the detailed characteristics of pulse detonation exhaust in the transient process of detonation. The duration of single detonation is 85 ms under laboratory conditions, among which supersonic injection time is 5.7 ms and subsonic injection time is 19.3 ms. The valveless pulse detonation engine used can work under frequency of 11 Hz. The velocity of detonation overflowing the detonation tube is 1,172 m x s(-1), the maximum temperature of detonation exhaust near the nozzle is 2 412 K. There is a transitory platform in the velocity curve as well as the temperature curve. H2O gas concentration changes between 0-7% during detonation under experimental conditions. The research can provide measurement data for the detonation process diagnosis and analysis, which is of significance to advance the detonation mechanism research and promote the research of pulse detonation engine control technology. PMID:25208369

  19. 1-D detonability

    NASA Astrophysics Data System (ADS)

    Guirguis, R. H.

    1996-05-01

    The likelihood of an energetic material to detonate is expressed in terms of the thermicity σ.dλ/dt within the reaction zone of an incipient one-dimensional detonation wave propagating in the material. Ranking conventional explosives according to σ alone was found to reasonably parallel the ranking according to critical diameter.

  20. Bidirectional slapper detonator

    DOEpatents

    McCormick, Robert N.; Boyd, Melissa D.

    1984-01-01

    The disclosure is directed to a bidirectional slapper detonator. One embodiment utilizes a single bridge circuit to detonate a pair of opposing initiating pellets. A line generator embodiment uses a plurality of bridges in electrical series to generate opposing cylindrical wavefronts.

  1. On gas detonation limits

    SciTech Connect

    Nikolaev, Yu.A.; Gapanov, O.A.

    1995-11-01

    A one-dimensional model for a multiheaded detonation has been constructed with account for friction, heat losses, and the decay of gas velocity pulsations. The existence of detonation limits in narrow channels has been numerically shown. The calculation results are in satisfactory agreement with experimental data.

  2. Effects of Non-Uniform Fuel Distribution on Detonation Tube Performance

    NASA Technical Reports Server (NTRS)

    Perkins, H. Douglas; Sung, Chih-Jen

    2003-01-01

    A pulse detonation engine uses a series of high frequency intermittent detonation tubes to generate thrust. The process of filling the detonation tube with fuel and air for each cycle may yield non-uniform mixtures. Uniform mixing is commonly assumed when calculating detonation tube thrust performance. In this study, detonation cycles featuring idealized non-uniform H2/air mixtures were analyzed using a two-dimensional Navier-Stokes computational fluid dynamics code with detailed chemistry. Mixture non-uniformities examined included axial equivalence ratio gradients, transverse equivalence ratio gradients, and partially fueled tubes. Three different average test section equivalence ratios were studied; one stoichiometric, one fuel lean, and one fuel rich. All mixtures were detonable throughout the detonation tube. Various mixtures representing the same average test section equivalence ratio were shown to have specific impulses within 1% of each other, indicating that good fuel/air mixing is not a prerequisite for optimal detonation tube performance under the conditions investigated.

  3. Propagation Mechanism of Cylindrical Cellular Detonation

    NASA Astrophysics Data System (ADS)

    Han, Wen-Hu; Wang, Cheng; Ning, Jian-Guo

    2012-10-01

    We investigate the evolution of cylindrical cellular detonation with different instabilities. The numerical results show that with decreasing initial temperature, detonation becomes more unstable and the cells of the cylindrical detonation tend to be irregular. For stable detonation, a divergence of cylindrical detonation cells is formed eventually due to detonation instability resulting from a curved detonation front. For mildly unstable detonation, local overdriven detonation occurs. The detonation cell diverges and its size decreases. For highly unstable detonation, locally driven detonation is more obvious and the front is highly wrinkled. As a result, the diverging cylindrical detonation cell becomes highly irregular.

  4. Improved airbreathing launch vehicle performance with the use of rocket propulsion

    NASA Astrophysics Data System (ADS)

    Kauffman, H. G.; Grandhi, R. V.; Hankey, W. L.; Belcher, P. J.

    1991-04-01

    An efficient performance analysis method is developed to evaluate potential airbreathing/rocket propulsion systems for advanced technology single-stage-to-orbit launch vehicles. Evaluated are tradeoffs between airbreathing, rocket, and concurrent airbreathing/rocket propulsion in maximizing payload delivery to orbit for a given ascent flight trajectory. With the analysis method, several modes of airbreathing/rocket propulsion are compared to a baseline 'airbreather alone' propulsion system in terms of fuel/propellant required to attain orbital velocity. Concurrent airbreathing/rocket propulsion shows a reduction in fuel/propellant consumption over straight airbreather to rocket propulsion transition. The optimal switch point (staging) is identified for the transition from airbreathing to rocket propulsion.

  5. Development of detonation reaction engine

    NASA Technical Reports Server (NTRS)

    Lange, O. H.; Stein, R. J.; Tubbs, H. E.

    1968-01-01

    Reaction engine operates on the principle of a controlled condensed detonation. In this engine the gas products that are expelled from the engine to produce thrust are generated by the condensed detonation reaction. The engine is constructed of two basic sections consisting of a detonation wave generator section and a condensed detonation reaction section.

  6. Atomistic simulation of detonation initiation by ultra-short impact

    NASA Astrophysics Data System (ADS)

    Murzov, S. A.; Zhakhovsky, V. V.

    2015-11-01

    We present results of the classical molecular dynamics simulation of detonation initiation in simple AB model of a high explosive compressed by ultra-short shock wave (SW). The simplified reactive empirical bond order potential (REBO) defines interatomic forces in the AB model explosive made up of diatomic AB molecules. Simulation of ultra-short piston-driven compression of AB explosive with duration of picoseconds represents an indirect initiation via a thin metal foil irradiated by a femtosecond laser pulse. We studied transition of SW to a detonation wave (DW), including evolution of calculated pressure profile in a sample. A run distance to detonation of such AB explosive film, which is required for detonation initiation, was obtained. Variation of loading time and piston velocity gives a 2D region of transition from SW to DW. The influence of pores on detonation initiation threshold is discussed.

  7. Detonation Front Curvatures and Detonation Rates

    NASA Astrophysics Data System (ADS)

    Lauderbach, Lisa M.; Lorenz, K. Thomas; Lee, Edward L.; Souers, P. Clark

    2015-06-01

    We have normalized the LLNL library of detonation front curvatures by dividing lags by the edge lag and radii by the edge radius. We then fit the normalized data to the equation L = AR2 + BR8, where L is the normalized lag and R is the normalized radius. We attribute the quadratic term to thermal processes and the 8th-power term to shock processes. We compare the % of the quadratic term J at the edge with detonation rates obtained from the size effect. One class of results is made up of fine-grained, uniform explosives with large lags, where a low detonation rate leads to a high J and vice versa. This provides a rough way of estimating unknown rates if the unknown explosive is of high quality. The other, equally-large class contains rough-grained materials, often with small lags and small radii. These have curves that do not fit the equation but superfically often look quadratic. Some HMX and PETN curvatures even show a ``sombrero'' effect. Code models show that density differences of 0.03 g/cc in ram-pressed parts can cause pseudo-quadratic curves and even sombreros. Modeling is used to illustrate J at the lowest and highest possible detonation rates. This work performed under the auspices of the U. S. Department of Energy by Lawrence Livermore National Laboratory under Contract DE-AC52-07NA27344.

  8. Effects of Fuel Distribution on Detonation Tube Performance

    NASA Technical Reports Server (NTRS)

    Perkins, Hugh Douglas

    2002-01-01

    A pulse detonation engine (PDE) uses a series of high frequency intermittent detonation tubes to generate thrust. The process of filling the detonation tube with fuel and air for each cycle may yield non-uniform mixtures. Lack of mixture uniformity is commonly ignored when calculating detonation tube thrust performance. In this study, detonation cycles featuring idealized non-uniform H2/air mixtures were analyzed using the SPARK two-dimensional Navier-Stokes CFD code with 7-step H2/air reaction mechanism. Mixture non-uniformities examined included axial equivalence ratio gradients, transverse equivalence ratio gradients, and partially fueled tubes. Three different average test section equivalence ratios (phi), stoichiometric (phi = 1.00), fuel lean (phi = 0.90), and fuel rich (phi = 1.10), were studied. All mixtures were detonable throughout the detonation tube. It was found that various mixtures representing the same test section equivalence ratio had specific impulses within 1 percent of each other, indicating that good fuel/air mixing is not a prerequisite for optimal detonation tube performance.

  9. Preliminary Sizing of an Hypersonic Airbreathing Airliner

    NASA Astrophysics Data System (ADS)

    Ingenito, Antonella; Gulli, Stefano; Bruno, Claudio

    The purpose of this paper is to identify, for given technology levels (TRL) and mission requirements, those parameters that are critical for preliminary sizing of a hypersonic airbreathing airliner. Mission requirements will dictate a solution space of possible vehicle architecture capable of meeting cruise conditions as well as of taking-off (TO) and landing. In practice, once defined a range of cruise vehicle architectures, constraints are imposed (as to all passenger airliners), such as: 1. take off (=TO) and landing distance (so-called field length, FL): FL no longer than for the B-747-400, or 10000 ft; 2. completing TO with one engine off; 3. max acceleration at TO and climb-out (CO) = 0.4 g; 4. Hydrogen fuel (Meeting NOx emission limits (EINOx) is a further constraint not discussed in this paper). These constraints enable focusing on a realistic design out of the broad range of vehicles capable of performing the given mission. Thus a realistic vehicle must not only integrate aerodynamics and propulsion system; in fact, it is the result of many iterations in the design space, until performance and constraints are successfully achieved and met. The Gross Weight at Take Off (TOGW) was deliberately discarded as a constraint, based on Previous studies by Czysz. Typically, limiting from the beginning the TOGW leads to a vicious spiral where weight and propulsion system requirements keep growing, eventually denying convergence. In designing passenger airliners, in fact, it is the payload that is assumed fixed from the start, not the total weight. A parametric analysis of the hypersonic vehicle architecture is presented: in particular, optimal size, weight and geometrical shape are defined for different mission requirements. This analysis has shown that, it is possible to define a range of possible successful solutions for the European LAPCAT II project.

  10. Direct observation detonator operation

    NASA Astrophysics Data System (ADS)

    Hall, Charles R.

    2001-11-01

    The analysis of detonator-timing performance has involved the use of rotating-mirror cameras (RMC) used in the streak mode and high-speed film. Fiducial timing marks are applied to the film to provide temporal references. The use of a RMC for detonator analysis requires aligning the camera, performing an exposure test, capturing light from the detonation and then processing the film. This procedure can take up to an hour for two technicians. After the film is possessed another technician compares each light streak on the film with the fiducial timing marks also recorded on the film. Capturing light from a detonator and recording it directly to a digitizer can improve detonator-timing measurement in several ways. The digitized signals can then be directly analyzed with software. The direct recording method reduces the need for expensive rotating mirror cameras, film processing and subjective optical measurement comparison. Furthermore, an extensive support facility requiring several specialized technicians is reduced to a single technician in a modest laboratory. This technician is then capable of performing several tests an hour. Tests were preformed to measure light intensity at detonation. An optical method of capturing the light was designed using a remote microscope coupled to optical fiber to bring the light to an optical/electrical converter and a digitizer then records the signal. This system is presently used in parallel with a RMC. The results are compared for accuracy.

  11. Theory of gaseous detonations.

    PubMed

    Clavin, Paul

    2004-09-01

    The objective of the present paper is to review some developments that have occurred in detonation theory over the last ten years. They concern nonlinear dynamics of detonation fronts, namely patterns of pulsating and/or cellular fronts, selection of the cell size, dynamical self-quenching, direct (blast) or spontaneous initiation, and transition from deflagration to detonation. These phenomena are all well documented by experiments since the sixties but remained unexplained until recently. In the first part of the paper, the patterns of cellular detonations are described by an asymptotic solution to nonlinear hyperbolic equations (reactive Euler equations) in the form of unsteady (sometime chaotic) and multidimensional traveling-waves. In the second part, turning points of quasi-steady solutions are shown to correspond to critical conditions of fully unsteady problems, either for (direct or spontaneous) initiation or for spontaneous failure (self-quenching). Physical insights are tentatively presented rather than technical aspects. The challenge is to identify the physical mechanisms with their relevant parameters, and more specifically to explain how the length-scales involved in detonation dynamics are larger by two order of magnitude (at least) than the length-scale involved in the steady planar traveling-wave solution (detonation thickness). PMID:15446993

  12. Cable Discharge System for fundamental detonator studies

    NASA Technical Reports Server (NTRS)

    Peevy, Gregg R.; Barnhart, Steven G.; Brigham, William P.

    1994-01-01

    Sandia National Laboratories has recently completed the modification and installation of a cable discharge system (CDS) which will be used to study the physics of exploding bridgewire (EBW) detonators and exploding foil initiators (EFI or slapper). Of primary interest are the burst characteristics of these devices when subjected to the constant current pulse delivered by this system. The burst process involves the heating of the bridge material to a conductive plasma and is essential in describing the electrical properties of the bridgewire foil for use in diagnostics or computer models. The CDS described herein is capable of delivering up to an 8000 A pulse of 3 micron duration. Experiments conducted with the CDS to characterize the EBW and EFI burst behavior are also described. In addition, the CDS simultaneous VISAR capability permits updating the EFI electrical Gurney analysis parameters used in our computer simulation codes. Examples of CDS generated data for a typical EFI and EBW detonator are provided.

  13. Rapid detonation initiation by sparks in a short duct: a numerical study

    NASA Astrophysics Data System (ADS)

    Hu, Z. M.; Dou, H. S.; Khoo, B. C.

    2010-06-01

    Rapid onset of detonation can efficiently increase the working frequency of a pulse detonation engine (PDE). In the present study, computations of detonation initiation in a duct are conducted to investigate the mechanisms of detonation initiation. The governing equations are the Euler equations and the chemical kinetic model consists of 19 elementary reactions and nine species. Different techniques of initiation have been studied for the purpose of accelerating detonation onset with a relatively weak ignition energy. It is found that detonation ignition induced by means of multiple sparks is applicable to auto-ignition for a PDE. The interaction among shock waves, flame fronts and the strip of pre-compressed fresh (unburned) mixture plays an important role in rapid onset of detonation.

  14. Analysis of particle dynamics and heat transfer in detonation thermal spraying systems

    NASA Astrophysics Data System (ADS)

    Ramadan, K.; Butler, P. Barry

    2004-06-01

    A computational study of pulsed detonation thermal spraying is conducted using an axisymmetric two-dimensional transient gaseous detonation model. The variations of the particle velocity and temperature at impact on the target surface with the particle initial loading location are analyzed for different conditions. The geometry of the system and the loading locations of the particulate phase are key parameters in pulsed detonation thermal spraying. Since the process is extremely transient and the gas phase experiences a wide range of transient stages all on a timescale of a millisecond, the particle characteristics are strongly dependent on the instantaneous location in the gas stream. One cycle of detonation thermal spraying occurs on a time scale on the order of a millisecond due to the high gas velocities associated with detonation. Thus, a precise control of the process variable parameters is required to have a successful detonation coating process.

  15. Miniature plasma accelerating detonator and method of detonating insensitive materials

    SciTech Connect

    Bickes, R.W. Jr.; Kopczewski, M.R.; Schwarz, A.C.

    1986-11-11

    This patent describes a detonator assembly for initiating insensitive explosives or energetic materials. In the improvement described here the detonator assembly comprises: railgun accelerating means of a size sufficient to be used as a detonator for insensitive explosives or energetic materials in an amount of about 100 mg of explosives or less and capable of accelerating a plasma to detonation initiating velocities; and power supply means for supplying the power necessary to the railgun accelerating means to generate and accelerate the plasma.

  16. Computational analysis of hypersonic airbreathing aircraft flow fields

    NASA Technical Reports Server (NTRS)

    Dwoyer, Douglas L.; Kumar, Ajay

    1987-01-01

    The general problem of calculating the flow fields associated with hypersonic airbreathing aircraft is presented. Unique aspects of hypersonic aircraft aerodynamics are introduced and their demands on computational fluid dynamics are outlined. Example calculations associated with inlet/forebody integration and hypersonic nozzle design are presented to illustrate the nature of the problems considered.

  17. Prospects for future hypersonic air-breathing vehicles

    NASA Technical Reports Server (NTRS)

    Beach, H. L., Jr.; Blankson, Isaiah M.

    1991-01-01

    An overview of the technical progress achieved in key areas of hypersonic airbreathing vehicle development is presented. The context for hypersonic applications is discussed with emphasis placed on technology issues and requirements, particularly for propulsion and technology integration. Attention is given to CFD technology which allows the consideration of configurations and extrapolations to flight conditions that cannot be simulated on the ground.

  18. Research on filling process of fuel and oxidant during detonation based on absorption spectrum technology

    NASA Astrophysics Data System (ADS)

    Lv, Xiao-Jing; Li, Ning; Weng, Chun-Sheng

    2014-12-01

    Research on detonation process is of great significance for the control optimization of pulse detonation engine. Based on absorption spectrum technology, the filling process of fresh fuel and oxidant during detonation is researched. As one of the most important products, H2O is selected as the target of detonation diagnosis. Fiber distributed detonation test system is designed to enable the detonation diagnosis under adverse conditions in detonation process. The test system is verified to be reliable. Laser signals at different working frequency (5Hz, 10Hz and 20Hz) are detected. Change of relative laser intensity in one detonation circle is analyzed. The duration of filling process is inferred from the change of laser intensity, which is about 100~110ms. The peak of absorption spectrum is used to present the concentration of H2O during the filling process of fresh fuel and oxidant. Absorption spectrum is calculated, and the change of absorption peak is analyzed. Duration of filling process calculated with absorption peak consisted with the result inferred from the change of relative laser intensity. The pulse detonation engine worked normally and obtained the maximum thrust at 10Hz under experiment conditions. The results are verified through H2O gas concentration monitoring during detonation.

  19. 2-D Air-Breathing Lightcraft Engine Experiments in Hypersonic Conditions

    NASA Astrophysics Data System (ADS)

    Salvador, Israel I.; Myrabo, Leik N.; Minucci, Marco A. S.; de Oliveira, Antonio C.; Toro, Paulo G. P.; Chanes, José B.; Rego, Israel S.

    2011-11-01

    Experiments were performed with a 2-D, repetitively-pulsed (RP) laser Lightcraft model in hypersonic flow conditions. The main objective was the feasibility analysis for impulse generation with repetitively-pulsed air-breathing laser Lightcraft engines at hypersonic speeds. The future application of interest for this basic research endeavor is the laser launch of pico-, nano-, and micro-satellites (i.e., 0.1-100 kg payloads) into Low-Earth-Orbit, at low-cost and on-demand. The laser propulsion experiments employed a Hypersonic Shock Tunnel integrated with twin gigawatt pulsed Lumonics 620-TEA CO2 lasers (˜ 1 μs pulses), to produce the required test conditions. This hypersonic campaign was carried out at nominal Mach numbers ranging from 6 to 10. Time-dependent surface pressure distributions were recorded together with Schlieren movies of the flow field structure resulting from laser energy deposition. Results indicated laser-induced pressure increases of 0.7-0.9 bar with laser pulse energies of ˜ 170 J, on off-shroud induced breakdown condition, and Mach number of 7.

  20. In-Situ Continuous Detonation Velocity Measurements Using Fiber-optic Bragg Grating Sensors

    SciTech Connect

    Benterou, J; Udd, E; Wilkins, P; Roeske, F; Roos, E; Jackson, D

    2007-07-25

    In order to fully calibrate hydrocodes and dynamic chemistry burn models, initiation and detonation research requires continuous measurement of low order detonation velocities as the detonation runs up to full order detonation for a given density and initiation pressure pulse. A novel detector of detonation velocity is presented using a 125 micron diameter optical fiber with an integral chirped fiber Bragg grating as an intrinsic sensor. This fiber is embedded in the explosive under study and interrogated during detonation as the fiber Bragg grating scatters light back along the fiber to a photodiode, producing a return signal dependant on the convolution integral of the grating reflection bandpass, the ASE intensity profile and the photodetector response curve. Detonation velocity is measured as the decrease in reflected light exiting the fiber as the grating is consumed when the detonation reaction zone proceeds along the fiber sensor axis. This small fiber probe causes minimal perturbation to the detonation wave and can measure detonation velocities along path lengths tens of millimeters long. Experimental details of the associated equipment and preliminary data in the form of continuous detonation velocity records within nitromethane and PBX-9502 are presented.

  1. Motivations for laser detonator and firing system developments

    NASA Astrophysics Data System (ADS)

    Kennedy, James E.

    2006-08-01

    For ordnance system and testing applications in which safety is paramount, laser detonators and firing systems are strong candidates. Both low-power (1 W) and high-power (~1 MW) laser-driven explosive devices provide safety against stray current and electrostatic discharges, including lightning. This article addresses only one class of high-power laser-driven detonators that provide prompt detonation - the laser-driven analog of electrical exploding bridgewire (EBW) detonatorsm which we call a "laser EBW." Coupling of laser power into a plasma and then to the explosive powder will be described. Drawing upon current initiatives within DOE laboratories, this talk will emphasize similarities between high-power laser detonators and high-power electrical detonators in terms of firing power requirements and development of deonation. In explosive testing applications, laser detonators provide separation of noisy electrical firing systems from diagnostic sensors that may be embedded in an experimental assembly. Laser detonators can be made without any metallic content, and that is desirable for radiography experiments. Feasibility of reliable transmission of a firing pulse through optical fibers is a key element in applications for missile ordnance, warhead firing, and other mobile systems. The preparation and characterization of fibers, and their capabilities and limitations are also discussed briefly.

  2. Nonequilibrium detonation of composite explosives

    SciTech Connect

    Nichols III, A.L.

    1997-07-01

    The effect of nonequilibrium diffusional flow on detonation velocities in composite explosives is examined. Detonation conditions are derived for complete equilibrium, temperature and pressure equilibrium, and two forms of pressure equilibrium. Partial equilibria are associated with systems which have not had sufficient time for transport to smooth out the gradients between spatially separate regions. The nonequilibrium detonation conditions are implemented in the CHEQ equation of state code. We show that the detonation velocity decreases as the non-chemical degrees of freedom of the explosive are allowed to equilibrate. It is only when the chemical degrees of freedom are allowed to equilibrate that the detonation velocity increases.

  3. Clamp for detonating fuze

    NASA Technical Reports Server (NTRS)

    Holderman, E. J.

    1968-01-01

    Quick acting clamp provides physical support for a closely confined detonating fuse in an application requiring removal and replacement at frequent intervals during test. It can be designed with a base of any required strength and configuration to permit the insertion of an object.

  4. Innovative Airbreathing Propulsion Concepts for Access to Space

    NASA Technical Reports Server (NTRS)

    Whitlow, Jr., Woodrow; Blech, Richard A.; Blankson, Isaiah M.

    2001-01-01

    This paper will present technologies and concepts for novel aeropropulsion systems. These technologies will enhance the safety of operations, reduce life cycle costs, and contribute to reduced costs of air travel and access to space. One of the goals of the NASA program is to reduce the carbon-dioxide emissions of aircraft engines. Engine concepts that use highly efficient fuel cell/electric drive technologies in hydrogen-fueled engines will be presented in the proposed paper. Carbon-dioxide emissions will be eliminated by replacing hydrocarbon fuel with hydrogen, and reduce NOx emissions through better combustion process control. A revolutionary exoskeletal engine concept, in which the engine drum is rotated, will be shown. This concept has the potential to allow a propulsion system that can be used for subsonic through hypersonic flight. Dual fan concepts that have ultra-high bypass ratios, low noise, and low drag will be presented. Flow-controlled turbofans and control-configured turbofans also will be discussed. To increase efficiency, a system of microengines distributed along lifting surfaces and on the fuselage is being investigated. This concept will be presented in the paper. Small propulsion systems for affordable, safe personal transportation vehicles will be discussed. These low-oil/oilless systems use technologies that enable significant cost and weight reductions. Pulse detonation engine-based hybrid-cycle and combined-cycle propulsion systems for aviation and space access will be presented.

  5. Transcriptomic Analysis of Compromise Between Air-Breathing and Nutrient Uptake of Posterior Intestine in Loach (Misgurnus anguillicaudatus), an Air-Breathing Fish.

    PubMed

    Huang, Songqian; Cao, Xiaojuan; Tian, Xianchang

    2016-08-01

    Dojo loach (Misgurnus anguillicaudatus) is an air-breathing fish species by using its posterior intestine to breathe on water surface. So far, the molecular mechanism about accessory air-breathing in fish is seldom addressed. Five cDNA libraries were constructed here for loach posterior intestines form T01 (the initial stage group), T02 (mid-stage of normal group), T03 (end stage of normal group), T04 (mid-stage of air-breathing inhibited group), and T05 (the end stage of air-breathing inhibited group) and subjected to perform RNA-seq to compare their transcriptomic profilings. A total of 92,962 unigenes were assembled, while 37,905 (40.77 %) unigenes were successfully annotated. 2298, 1091, and 3275 differentially expressed genes (fn1, ACE, EGFR, Pxdn, SDF, HIF, VEGF, SLC2A1, SLC5A8 etc.) were observed in T04/T02, T05/T03, and T05/T04, respectively. Expression levels of many genes associated with air-breathing and nutrient uptake varied significantly between normal and intestinal air-breathing inhibited group. Intraepithelial capillaries in posterior intestines of loaches from T05 were broken, while red blood cells were enriched at the surface of intestinal epithelial lining with 241 ± 39 cells per millimeter. There were periodic acid-schiff (PAS)-positive epithelial mucous cells in posterior intestines from both normal and air-breathing inhibited groups. Results obtained here suggested an overlap of air-breathing and nutrient uptake function of posterior intestine in loach. Intestinal air-breathing inhibition in loach would influence the posterior intestine's nutrient uptake ability and endothelial capillary structure stability. This study will contribute to our understanding on the molecular regulatory mechanisms of intestinal air-breathing in loach. PMID:27457889

  6. Pulse

    MedlinePlus

    Heart rate; Heart beat ... The pulse can be measured at areas where an artery passes close to the skin. These areas include the: ... side of the foot Wrist To measure the pulse at the wrist, place the index and middle ...

  7. Airbreathing Hypersonic Systems Focus at NASA Langley Research Center

    NASA Technical Reports Server (NTRS)

    Hunt, James L.; Rausch, Vincent L.

    1998-01-01

    This paper presents the status of the airbreathing hypersonic airplane and space-access vehicle design matrix, reflects on the synergies and issues, and indicates the thrust of the effort to resolve the design matrix and to focus/advance systems technology maturation. Priority is given to the design of the vision operational vehicles followed by flow-down requirements to flight demonstrator vehicles and their design for eventual consideration in the Future-X Program.

  8. Aeronautical facilities catalogue. Volume 2: Airbreathing propulsion and flight simulators

    NASA Technical Reports Server (NTRS)

    Penaranda, F. E.; Freda, M. S.

    1985-01-01

    Volume two of the facilities catalogue deals with Airbreathing Propulsion and Flight Simulation Facilities. Data pertinent to managers and engineers are presented. Each facility is described on a data sheet that shows the facility's technical parameters on a chart and more detailed information in narratives. Facilities judged comparable in testing capability are noted and grouped together. Several comprehensive cross-indexes and charts are included.

  9. Optimization of Air-Breathing Propulsion Engine Concepts

    NASA Technical Reports Server (NTRS)

    Patnaik, Surya N.; Hopkins, Dale A.

    1997-01-01

    Air-breathing propulsion engines play an important role in the development of both civil and military aircraft Design optimization of such engines can lead to higher power, or more thrust for less fuel consumption. A multimission propulsion engine design can be modeled mathematically as a multivariable global optimization problem, with a sequence of subproblems, which are specific to the mission events defined through Mach number, altitude, and power setting combinations.

  10. Time Resolved X-ray Imaging of a Detonation Front

    NASA Astrophysics Data System (ADS)

    Vantine, Harry; Aufderheide, Maurice; Egan, Patrick; Morgan, David

    1998-03-01

    A one inch diameter cylinder of LX-14 was detonated at one end. The detonation was then made to propagate around an obstacle, in this case a one-half inch diameter lucite sphere. The detonation front was imaged at three times per shot with an array of three Scandiflash-450 Flash X-ray heads, spaced 2.1m from the explosive. The spectrum was 450 keV brehmstrahlung with a strong contribution from W K-radiation. The x-ray pulse width was approximately 20ns, and the tme between pulses was adjustable (typically 1-2 μs). Data were recorded on screened film (LANEX-fine and TMAT H) located 0.6m from the explosive. The film was digitized on a PDS microdensitometer for subsequent analysis. Tomagraphic reconstruction and comparison with CALE numerical simulations will be shown.

  11. Miniature plasma accelerating detonator and method of detonating insensitive materials

    DOEpatents

    Bickes, R.W. Jr.; Kopczewski, M.R.; Schwarz, A.C.

    1985-01-04

    The invention is a detonator for use with high explosives. The detonator comprises a pair of parallel rail electrodes connected to a power supply. By shorting the electrodes at one end, a plasma is generated and accelerated toward the other end to impact against explosives. A projectile can be arranged between the rails to be accelerated by the plasma. An alternative arrangement is to a coaxial electrode construction. The invention also relates to a method of detonating explosives. 3 figs.

  12. Miniature plasma accelerating detonator and method of detonating insensitive materials

    DOEpatents

    Bickes, Jr., Robert W.; Kopczewski, Michael R.; Schwarz, Alfred C.

    1986-01-01

    The invention is a detonator for use with high explosives. The detonator comprises a pair of parallel rail electrodes connected to a power supply. By shorting the electrodes at one end, a plasma is generated and accelerated toward the other end to impact against explosives. A projectile can be arranged between the rails to be accelerated by the plasma. An alternative arrangement is to a coaxial electrode construction. The invention also relates to a method of detonating explosives.

  13. Chemical Equilibrium Detonation

    NASA Astrophysics Data System (ADS)

    Bastea, Sorin; Fried, Laurence E.

    Energetic materials are unique for having a strong exothermic reactivity, which has made them desirable for both military and commercial applications. The fundamental principles outlined in this chapter pertain to the study of detonation in both gas-phase and condensed-phase energetic materials, but our main focus will be on the condensed ones, particularly on high explosives (HEs). They share many properties with other classes of condensed energetic compounds such as propellants and pyrotechnics, but a detailed understanding of detonation is especially important for numerous HE applications. The usage and study of HE materials goes back more than a century, but many questions remain to be answered, e.g., on their reaction pathways at high pressures and temperatures, chemical properties, etc.

  14. Airbreathing Laser Propulsion Experiments with 1 {mu}m Terawatt Pharos IIILaser: Part 2

    SciTech Connect

    Myrabo, L. N.; Lyons, P. W.; Jones, R. A.; Liu, S.; Manka, C.

    2011-11-10

    This basic research study examines the physics of airbreathing laser propulsion at the extreme flux range of 1-2x10{sup 11} W/cm{sup 2}--within the air breakdown threshold for l {mu}m radiation--using the terawatt Pharos III neodymium-glass pulsed laser. Six different experimental setups were employed using a 34 mm line focus with 66 {mu}m focal waist, positioned near the flat impulse surface. The 2nd Campaign investigated impulse generation with the laser beam focused at grazing incidence across near horizontal target surfaces, with pulse energies ranging from 55 to 186 J, and pulse-widths of 2 to 30 ns FWHM. Laser generated impulse was measured with a horizontal Plexiglas registered ballistic pendulum equipped with either a steel target insert or 0.5 Tesla permanent magnet (NEIT-40), to quantify changes in the momentum coupling coefficient (C{sub M}). Part 2 of this 2-part paper covers Campaign no. 2 results including C{sub M} performance data, and long exposure color photos of LP plasma phenomena.

  15. Recent work on gaseous detonations

    NASA Astrophysics Data System (ADS)

    Nettleton, M. A.

    The paper reviews recent progress in the field of gaseous detonations, with sections on shock diffraction and reflection, the transition to detonation, hybrid, spherically-imploding, and galloping and stuttering fronts, their structure, their transmission and quenching by additives, the critical energy for initiation and detonation of more unusual fuels. The final section points out areas where our understanding is still far from being complete and contains some suggestions of ways in which progress might be made.

  16. Low voltage nonprimary explosive detonator

    DOEpatents

    Dinegar, Robert H.; Kirkham, John

    1982-01-01

    A low voltage, electrically actuated, nonprimary explosive detonator is disclosed wherein said detonation is achieved by means of an explosive train in which a deflagration-to-detonation transition is made to occur. The explosive train is confined within a cylindrical body and positioned adjacent to low voltage ignition means have electrical leads extending outwardly from the cylindrical confining body. Application of a low voltage current to the electrical leads ignites a self-sustained deflagration in a donor portion of the explosive train which then is made to undergo a transition to detonation further down the train.

  17. Seal Technology Development for Advanced Component for Airbreathing Engines

    NASA Technical Reports Server (NTRS)

    Snyder, Philip H.

    2008-01-01

    Key aspects of the design of sealing systems for On Rotor Combustion/Wave Rotor (ORC/WR) systems were addressed. ORC/WR systems generally fit within a broad class of pressure gain Constant Volume Combustors (CVCs) or Pulse Detonation Combustors (PDCs) which are currently being considered for use in many classes of turbine engines for dramatic efficiency improvement. Technology readiness level of this ORC/WR approaches are presently at 2.0. The results of detailed modeling of an ORC/WR system as applied to a regional jet engine application were shown to capture a high degree of pressure gain capabilities. The results of engine cycle analysis indicated the level of specific fuel consumption (SFC) benefits to be 17 percent. The potential losses in pressure gain due to leakage were found to be closely coupled to the wave processes at the rotor endpoints of the ORC/WR system. Extensive investigation into the sealing approaches is reported. Sensitivity studies show that SFC gains of 10 percent remain available even when pressure gain levels are highly penalized. This indicates ORC/WR systems to have a high degree of tolerance to rotor leakage effects but also emphasizes their importance. An engine demonstration of an ORC/WR system is seen as key to progressing the TRL of this technology. An industrial engine was judged to be a highly advantageous platform for demonstration of a first generation ORC/WR system. Prior to such a demonstration, the existing NASA pressure exchanger wave rotor rig was identified as an opportunity to apply both expanded analytical modeling capabilities developed within this program and to identify and fix identified leakage issues existing within this rig. Extensive leakage analysis of the rig was performed and a detailed design of additional sealing strategies for this rig was generated.

  18. Research in Hypersonic Airbreathing Propulsion at the NASA Langley Research Center

    NASA Technical Reports Server (NTRS)

    Kumar, Ajay; Drummond, J. Philip; McClinton, Charles R.; Hunt, James L.

    2001-01-01

    The NASA Langley Research Center has been conducting research for over four decades to develop technology for an airbreathing-propelled vehicle. Several other organizations within the United States have also been involved in this endeavor. Even though significant progress has been made over this period, a hypersonic airbreathing vehicle has not yet been realized due to low technology maturity. One of the major reasons for the slow progress in technology development has been the low level and cyclic nature of funding. The paper provides a brief historical overview of research in hypersonic airbreathing technology and then discusses current efforts at NASA Langley to develop various analytical, computational, and experimental design tools and their application in the development of future hypersonic airbreathing vehicles. The main focus of this paper is on the hypersonic airbreathing propulsion technology.

  19. Semiconductor bridge (SCB) detonator

    DOEpatents

    Bickes, Jr., Robert W.; Grubelich, Mark C.

    1999-01-01

    The present invention is a low-energy detonator for high-density secondary-explosive materials initiated by a semiconductor bridge igniter that comprises a pair of electrically conductive lands connected by a semiconductor bridge. The semiconductor bridge is in operational or direct contact with the explosive material, whereby current flowing through the semiconductor bridge causes initiation of the explosive material. Header wires connected to the electrically-conductive lands and electrical feed-throughs of the header posts of explosive devices, are substantially coaxial to the direction of current flow through the SCB, i.e., substantially coaxial to the SCB length.

  20. Semiconductor bridge (SCB) detonator

    DOEpatents

    Bickes, R.W. Jr.; Grubelich, M.C.

    1999-01-19

    The present invention is a low-energy detonator for high-density secondary-explosive materials initiated by a semiconductor bridge (SCB) igniter that comprises a pair of electrically conductive lands connected by a semiconductor bridge. The semiconductor bridge is in operational or direct contact with the explosive material, whereby current flowing through the semiconductor bridge causes initiation of the explosive material. Header wires connected to the electrically-conductive lands and electrical feed-throughs of the header posts of explosive devices, are substantially coaxial to the direction of current flow through the SCB, i.e., substantially coaxial to the SCB length. 3 figs.

  1. Detonation diffraction in gases

    SciTech Connect

    Pintgen, F.; Shepherd, J.E.

    2009-03-15

    We have experimentally investigated detonation diffraction out of a round tube into an unconfined half-space. The focus of our study is examining how the extent of detonation cellular instability influences the quantitative and qualitative features of diffraction. Detailed quantitative and qualitative measurements were obtained through simultaneous schlieren imaging, multiple-exposure chemiluminescence imaging, and planar laser-induced fluorescence imaging of OH molecules. Two types of stoichiometric mixtures, highly diluted H{sub 2}-O{sub 2}-Ar and H{sub 2}-N{sub 2}O, were studied in the sub-critical, critical and super-critical regime. These mixture types represent extreme cases in the classification of cellular instability with highly diluted H{sub 2}-O{sub 2}-Ar mixtures having very regular instability structures and H{sub 2}-N{sub 2}O having very irregular instability structures. The most striking differences between the mixtures occur in the sub-critical and critical regimes, for which the detonation fails to transition into the unconfined half-space. For the H{sub 2}-O{sub 2}-Ar mixture, the velocity on the center line was found to decay significantly slower than for the H{sub 2}-N{sub 2}O mixture. In case of the H{sub 2}-O{sub 2}-Ar mixture, it was evident from simultaneous schlieren-fluorescence images that the reaction front was coupled to the lead shock front up to 2.3 tube diameters from the exit plane. For the H{sub 2}-N{sub 2}O mixture, the reaction front velocity decreased to 60% of the corresponding Chapman-Jouguet value at 1.1 tube diameters from the tube exit plane. A geometric acoustic model showed that the observed differences in failure patterns are not caused by the differences in thermodynamic properties of the two mixtures but is linked to the larger effective activation energy and critical decay time in the H{sub 2}-N{sub 2}O mixture as compared to the H{sub 2}-O{sub 2}-Ar mixture. The re-initiation events appear similar for the two

  2. Detonator-activated ball shutter

    DOEpatents

    McWilliams, Roy A.; von Holle, William G.

    1983-01-01

    A detonator-activated ball shutter for closing an aperture in about 300.mu. seconds. The ball shutter containing an aperture through which light, etc., passes, is closed by firing a detonator which propels a projectile for rotating the ball shutter, thereby blocking passage through the aperture.

  3. Detonator-activated ball shutter

    DOEpatents

    McWilliams, R.A.; Holle, W.G. von.

    1983-08-16

    A detonator-activated ball shutter for closing an aperture in about 300[mu] seconds. The ball shutter containing an aperture through which light, etc., passes, is closed by firing a detonator which propels a projectile for rotating the ball shutter, thereby blocking passage through the aperture. 3 figs.

  4. Airbreathing Laser Propulsion Experiments with 1 {mu}m Terawatt Pharos III Laser: Part 1

    SciTech Connect

    Myrabo, L. N.; Lyons, P. W.; Jones, R. A.; Liu, S.; Manka, C.

    2011-11-10

    This basic research study examines the physics of airbreathing laser propulsion at the extreme flux range of 1-2x10{sup 11} W/cm{sup 2}--within the air breakdown threshold for l {mu}m radiation--using the terawatt PHAROS III neodymium-glass pulsed laser. Six different experimental setups were tested using a 34 mm line focus with 66 {mu}m focal waist, positioned near the flat impulse surface. The first campaign investigated impulse generation with the beam oriented almost normal to the target surface, with energies ranging from 23 to 376 J, and pulses of 5 to 30 ns FWHM. Air breakdown/ plasma dynamics were diagnosed with GOI cameras and color photography. Laser generated impulse was quantified with both vertical pendulums and piezoelectric pressure transducers using the standard performance metric, C{sub M}--the momentum coupling coefficient. Part 1 of this 2-part paper covers Campaign no. 1 results including laser plasma diagnostics, pressure gage and vertical pendulum data.

  5. Review of the PDWA Concept for Combustion Enhancement in a Supersonic Air-Breathing Combustor Environment

    NASA Technical Reports Server (NTRS)

    Canbier, Jean-Luc; Edwards, Thomas A. (Technical Monitor)

    1995-01-01

    This paper reviews the design of the Pulsed Detonation Wave Augmentor (PDWA) concept and the preliminary computational fluid dynamics studies that supported it. The PDWA relies on the rapid generation of detonation waves in a small tube, which are then injected into the supersonic stream of the main combustor. The blast waves thus generated are used to stimulate the mixing and combustion inside the main combustor. The mixing enhancement relies on various forms of the baroclinic interaction, where misaligned pressure and density gradients combine to produce vortical flow. By using unsteady shock waves, the concept also uses the Richtmyer-Meshkov effect to further increase the rate of mixing. By carefully designing the respective configurations of the combustor and the detonation tubes, one can also increase the penetration of the fuel into the supersonic air stream. The unsteady shocks produce lower stagnation pressure losses than steady shocks. Combustion enhancement can also be obtained through the transient shock-heating of the fuel-air interface, and the lowering of the ignition delay in these regions. The numerical simulations identify these processes, and show which configurations give the best results. Engineering considerations are also presented, and discuss the feasibility of the concept. Of primary importance are the enhancements in performance, the design simplicity, the minimization of the power, cost, and weight, and the methods to achieve very rapid cycling.

  6. A Pulsed Detonation Microthruster for Space Applications

    NASA Astrophysics Data System (ADS)

    Martel, E.; Brouillette, M.

    The majority of microsatellites in orbit do not have a propulsion system, which greatly limits their flexibility and their ability to perform complex tasks. Ambitious future missions will need the capability to produce the thrust required for either orbit change, drag makeup, station keeping or attitude control, and this usually requires more than one propulsion system [1].

  7. Aerodynamic characteristics of a series of airbreathing missile configurations

    NASA Technical Reports Server (NTRS)

    Hayes, C.

    1981-01-01

    The results of an experimental program conducted in order to compile a data base useful in the application of airbreathing propulsion to missiles are presented. The configurations investigated used two-dimensional or axisymmetric twin inlets located at three alternative circumferential positions: 90, 115 and 135 deg to the vertical centerline. The effects of a wing located above the inlets and of various tail configurations were investigated, with a view to longitudinal stability/control and lateral-directional stability characteristics. It is noted that of the three tail configurations tested, the 'X' tail showed (1) the most linear pitch-moment curve, (2) control effectiveness, and (3) positive lateral-directional stability.

  8. Fluid dynamic problems associated with air-breathing propulsive systems

    NASA Technical Reports Server (NTRS)

    Chow, W. L.

    1979-01-01

    A brief account of research activities on problems related to air-breathing propulsion is made in this final report for the step funded research grant NASA NGL 14-005-140. Problems include the aircraft ejector-nozzle propulsive system, nonconstant pressure jet mixing process, recompression and reattachment of turbulent free shear layer, supersonic turbulent base pressure, low speed separated flows, transonic boattail flow with and without small angle of attack, transonic base pressures, Mach reflection of shocks, and numerical solution of potential equation through hodograph transformation.

  9. Prospects for future hypersonic air-breathing vehicles

    NASA Technical Reports Server (NTRS)

    Beach, H. L., Jr.; Blankson, Isaiah M.

    1991-01-01

    The age of hypersonics is (almost) here. This is evident from the amount of activity in the United States, Europe, the USSR and Japan; this activity is a reflection of technical progress in key areas which will enable new vehicle systems, as well as renewed interest in the utilization of these systems. The current situation, at least in the United States, is the product of an interesting history which is briefly reviewed here. The context for hypersonic applications is discussed, but the emphasis is on hypersonic technology issues and needs, particularly for propulsion and technology integration. The paper concludes with prospects for accomplishing the objective of air-breathing hypersonic vehicle systems.

  10. Laser-driven hypersonic air-breathing propulsion simulator

    NASA Technical Reports Server (NTRS)

    Joshi, Prakash B.; Lo, Edmond Y.; Pugh, Evan R.

    1992-01-01

    A feasibility study is presented of simulating airbreathing propulsion on small scale hypersonic models using laser energy. The laser heat addition scheme allows simultaneous inlet and exhaust flows during wind tunnel testing of models with scramjet models. The proposed propulsion simulation concept has extended the Kantrowitz (1974) idea to propulsive wind tunnel models of hypersonic aircraft. Critical issues in aeropropulsive testing of models based on a ramjet power plant are addressed which include transfer of the correct amount of energy to the flowing gas, efficient absorption of laser energy into the gas, and test performance under tunnel reservoir conditions and at reasonable Reynolds numbers.

  11. Planar Reflection of Detonations Waves

    NASA Astrophysics Data System (ADS)

    Damazo, Jason; Shepherd, Joseph

    2012-11-01

    An experimental study examining normally reflected gaseous detonation waves is undertaken so that the physics of reflected detonations may be understood. Focused schlieren visualization is used to describe the boundary layer development behind the incident detonation wave and the nature of the reflected shock wave. Reflected shock wave bifurcation-which has received extensive study as it pertains to shock tube performance-is predicted by classical bifurcation theory, but is not observed in the present study for undiluted hydrogen-oxygen and ethylene-oxygen detonation waves. Pressure and thermocouple gauges are installed in the floor of the detonation tube so as to examine both the wall pressure and heat flux. From the pressure results, we observe an inconsistency between the measured reflected shock speed and the measured reflected shock strength with one dimensional flow predictions confirming earlier experiments performed in our laboratory. This research is sponsored by the DHS through the University of Rhode Island, Center of Excellence for Explosives Detection.

  12. Pulse

    MedlinePlus

    ... resting for at least 10 minutes. Take the exercise heart rate while you are exercising. ... pulse rate can help determine if the patient's heart is pumping. ... rate gives information about your fitness level and health.

  13. Planar Reflection of Gaseous Detonations

    NASA Astrophysics Data System (ADS)

    Damazo, Jason Scott

    Pipes containing flammable gaseous mixtures may be subjected to internal detonation. When the detonation normally impinges on a closed end, a reflected shock wave is created to bring the flow back to rest. This study built on the work of Karnesky (2010) and examined deformation of thin-walled stainless steel tubes subjected to internal reflected gaseous detonations. A ripple pattern was observed in the tube wall for certain fill pressures, and a criterion was developed that predicted when the ripple pattern would form. A two-dimensional finite element analysis was performed using Johnson-Cook material properties; the pressure loading created by reflected gaseous detonations was accounted for with a previously developed pressure model. The residual plastic strain between experiments and computations was in good agreement. During the examination of detonation-driven deformation, discrepancies were discovered in our understanding of reflected gaseous detonation behavior. Previous models did not accurately describe the nature of the reflected shock wave, which motivated further experiments in a detonation tube with optical access. Pressure sensors and schlieren images were used to examine reflected shock behavior, and it was determined that the discrepancies were related to the reaction zone thickness extant behind the detonation front. During these experiments reflected shock bifurcation did not appear to occur, but the unfocused visualization system made certainty impossible. This prompted construction of a focused schlieren system that investigated possible shock wave-boundary layer interaction, and heat-flux gauges analyzed the boundary layer behind the detonation front. Using these data with an analytical boundary layer solution, it was determined that the strong thermal boundary layer present behind the detonation front inhibits the development of reflected shock wave bifurcation.

  14. Static and Hypersonic Experimental Analysis of Impulse Generation in Air-Breathing Laser-Thermal Propulsion

    NASA Astrophysics Data System (ADS)

    Salvador, Israel Irone

    (98 to 161 mm in diameter), probably due to the more efficient delivery of laser-induced blast wave energy across the 2D model's larger impulse surface area. Next, the hypersonic campaign was carried out, subjecting the 2D model to nominal Mach numbers ranging from 6 to 10. Again, time-dependent surface pressure distributions were recorded together with Schlieren movies of the flow field structure resulting from laser energy deposition. These visualizations of inlet and absorption chamber flowfields, enabled the qualitative analysis of important phenomena impacting laser-propelled hypersonic airbreathing flight. The laser-induced breakdown took an elongated vertically-oriented geometry, occurring off-surface and across the inlet's mid-channel---quite different from the static case in which the energy was deposited very near the shroud under-surface. The shroud under-surface pressure data indicated laser-induced increases of 0.7-0.9 bar with laser pulse energies of ˜170 J, off-shroud induced breakdown condition, and Mach number of 7. The results of this research corroborate the feasibility of laser powered, airbreathing flight with infinite specific impulse (Isp=infinity): i.e., without the need for propellant injection at the laser focus. Additionally, it is shown that further reductions in inlet air working fluid velocity---with attendant increases in static pressure and density---is necessary to generate higher absorption chamber pressure and engine impulse. Finally, building on lessons learned from the present work, the future research plan is laid out for: a) the present 2D model with full inlet forebody, exploring higher laser pulse energies and multi-pulse phenomena; b) a smaller, redesigned 2D model; c) a 254 mm diameter axisymmetric Lightcraft model; and, d) a laser-electromagnetic accelerator model, designed around a 2-Tesla pulsed electromagnet contracted under the present program.

  15. Mathematical modeling of detonation initiation via flow cumulation effects

    NASA Astrophysics Data System (ADS)

    Semenov, I.; Utkin, P.; Akhmedyanov, I.

    2016-07-01

    The paper concerns two problems connected with the idea of gaseous detonation initiation via flow cumulation effects and convergence of relatively weak shock waves (SW). The first one is the three-dimensional (3D) numerical investigation of shock-to-detonation transition (SDT) in methane-air mixture in a tube with parabolic contraction followed by the tube section of narrow diameter and conical expansion. The second problem is the numerical study of the start-up of the model small-scale hydrogen electrochemical pulse detonation engine with the use of electrical discharge generating the toroidal SW. The investigation is performed by means of numerical simulation with the use of modern high-performance computing systems.

  16. Ferrite core coupled slapper detonator apparatus and method

    DOEpatents

    Boberg, Ralph E.; Lee, Ronald S.; Weingart, Richard C.

    1989-01-01

    Method and apparatus are provided for coupling a temporally short electric power pulse from a thick flat-conductor power cable into a thin flat-conductor slapper detonator circuit. A first planar and generally circular loop is formed from an end portion of the power cable. A second planar and generally circular loop, of similar diameter, is formed from all or part of the slapper detonator circuit. The two loops are placed together, within a ferrite housing that provides a ferrite path that magnetically couples the two loops. Slapper detonator parts may be incorporated within the ferrite housing. The ferrite housing may be made vacuum and water-tight, with the addition of a hermetic ceramic seal, and provided with an enclosure for protecting the power cable and parts related thereto.

  17. Ferrite core coupled slapper detonator apparatus and method

    DOEpatents

    Boberg, R.E.; Lee, R.S.; Weingart, R.C.

    1989-08-01

    Method and apparatus are provided for coupling a temporally short electric power pulse from a thick flat-conductor power cable into a thin flat-conductor slapper detonator circuit. A first planar and generally circular loop is formed from an end portion of the power cable. A second planar and generally circular loop, of similar diameter, is formed from all or part of the slapper detonator circuit. The two loops are placed together, within a ferrite housing that provides a ferrite path that magnetically couples the two loops. Slapper detonator parts may be incorporated within the ferrite housing. The ferrite housing may be made vacuum and water-tight, with the addition of a hermetic ceramic seal, and provided with an enclosure for protecting the power cable and parts related thereto. 10 figs.

  18. Screening studies of advanced control concepts for airbreathing engines

    NASA Technical Reports Server (NTRS)

    Ouzts, Peter J.; Lorenzo, Carl F.; Merrill, Walter C.

    1993-01-01

    The application of advanced control concepts to airbreathing engines may yield significant improvements in aircraft/engine performance and operability. Accordingly, the NASA Lewis Research Center has conducted screening studies of advanced control concepts for airbreathing engines to determine their potential impact on turbine engine performance and operability. The purpose of the studies was to identify concepts which offered high potential yet may incur high research and development risk. A target suite of proposed concepts was formulated by NASA and industry. These concepts were evaluated in a two phase study to quantify each concept's impact on desired engine characteristics. To aid in the evaluation, three target aircraft/engine combinations were considered: a military high performance fighter mission, a high speed civil transport mission, and a civil tiltrotor mission. Each of the advanced control concepts considered in the study were defined and described. The concept's potential impact on engine performance was determined. Relevant figures of merit on which to evaluate the concepts were also determined. Finally, the concepts were ranked with respect to the target aircraft/engine missions.

  19. Screening studies of advanced control concepts for airbreathing engines

    NASA Technical Reports Server (NTRS)

    Ouzts, Peter J.; Lorenzo, Carl F.; Merrill, Walter C.

    1992-01-01

    The application of advanced control concepts to airbreathing engines may yield significant improvements in aircraft/engine performance and operability. Accordingly, the NASA Lewis Research Center has conducted screening studies of advanced control concepts for airbreathing engines to determine their potential impact on turbine engine performance and operability. The purpose of the studies was to identify concepts which offered high potential yet may incur high research and development risk. A target suite of proposed concepts was formulated by NASA and industry. These concepts were evaluated in a two phase study to quantify each concept's impact on desired engine characteristics. To aid in the evaluation, three target aircraft/engine combinations were considered: a military high performance fighter mission, a high speed civil transport mission, and a civil tiltrotor mission. Each of the advanced control concepts considered in the study were defined and described. The concept's potential impact on engine performance was determined. Relevant figures of merit on which to evaluate the concepts were also determined. Finally, the concepts were ranked with respect to the target aircraft/engine missions.

  20. Advanced controls for airbreathing engines, volume 3: Allison gas turbine

    NASA Technical Reports Server (NTRS)

    Bough, R. M.

    1993-01-01

    The application of advanced control concepts to airbreathing engines may yield significant improvements in aircraft/engine performance and operability. Screening studies of advanced control concepts for airbreathing engines were conducted by three major domestic aircraft engine manufacturers to determine the potential impact of concepts on turbine engine performance and operability. The purpose of the studies was to identify concepts which offered high potential yet may incur high research and development risk. A target suite of proposed advanced control concepts was formulated and evaluated in a two-phase study to quantify each concept's impact on desired engine characteristics. To aid in the evaluation specific aircraft/engine combinations were considered: a Military High Performance Fighter mission, a High Speed Civil Transport mission, and a Civil Tiltrotor mission. Each of the advanced control concepts considered in the study are defined and described. The concept potential impact on engine performance was determined. Relevant figures of merit on which to evaluate the concepts are determined. Finally, the concepts are ranked with respect to the target aircraft/engine missions. A final report describing the screening studies was prepared by each engine manufacturer. Volume 3 of these reports describes the studies performed by the Allison Gas Turbine Division.

  1. Modeling the Effects of Turbulence in Rotating Detonation Engines

    NASA Astrophysics Data System (ADS)

    Towery, Colin; Smith, Katherine; Hamlington, Peter; van Schoor, Marthinus; TESLa Team; Midé Team

    2014-03-01

    Propulsion systems based on detonation waves, such as rotating and pulsed detonation engines, have the potential to substantially improve the efficiency and power density of gas turbine engines. Numerous technical challenges remain to be solved in such systems, however, including obtaining more efficient injection and mixing of air and fuels, more reliable detonation initiation, and better understanding of the flow in the ejection nozzle. These challenges can be addressed using numerical simulations. Such simulations are enormously challenging, however, since accurate descriptions of highly unsteady turbulent flow fields are required in the presence of combustion, shock waves, fluid-structure interactions, and other complex physical processes. In this study, we performed high-fidelity three dimensional simulations of a rotating detonation engine and examined turbulent flow effects on the operation, performance, and efficiency of the engine. Along with experimental data, these simulations were used to test the accuracy of commonly-used Reynolds averaged and subgrid-scale turbulence models when applied to detonation engines. The authors gratefully acknowledge the support of the Defense Advanced Research Projects Agency (DARPA).

  2. Detonation tube impulse in sub-atmospheric environments.

    SciTech Connect

    Cooper, Marcia A.; Shepherd, Joseph E.

    2005-04-01

    The thrust from a multi-cycle, pulse detonation engine operating at practical flight altitudes will vary with the surrounding environment pressure. We have carried out the first experimental study using a detonation tube hung in a ballistic pendulum arrangement within a large pressure vessel in order to determine the effect that the environment has on the single-cycle impulse. The air pressure inside the vessel surrounding the detonation tube varied between 100 and 1.4 kPa while the initial pressure of the stoichiometric ethylene-oxygen mixture inside the tube varied between 100 and 30 kPa. The original impulse model (Wintenberger et al., Journal of Propulsion and Power, Vol. 19, No. 1, 2002) was modified to predict the observed increase in impulse and blow down time as the environment pressure decreased below one atmosphere. Comparisons between the impulse from detonation tubes and ideal, steady flow rockets indicate incomplete expansion of the detonation tube exhaust, resulting in a 37% difference in impulse at a pressure ratio (ratio of pressure behind the Taylor wave to the environment pressure) of 100.

  3. [The Diagnostics of Detonation Flow External Field Based on Multispectral Absorption Spectroscopy Technology].

    PubMed

    Lü, Xiao-jing; Li, Ning; Weng, Chun-sheng

    2016-03-01

    Compared with traditional sampling-based sensing method, absorption spectroscopy technology is well suitable for detonation flow diagnostics, since it can provide with us fast response, nonintrusive, sensitive solution for situ measurements of multiple flow-field parameters. The temperature and concentration test results are the average values along the laser path with traditional absorption spectroscopy technology, while the boundary of detonation flow external field is unknown and it changes all the time during the detonation engine works, traditional absorption spectroscopy technology is no longer suitable for detonation diagnostics. The trend of line strength with temperature varies with different absorption lines. By increasing the number of absorption lines in the test path, more information of the non-uniform flow field can be obtained. In this paper, based on multispectral absorption technology, the reconstructed model of detonation flow external field distribution was established according to the simulation results of space-time conservation element and solution element method, and a diagnostic method of detonation flow external field was given. The model deviation and calculation error of the least squares method adopted were studied by simulation, and the maximum concentration and temperature calculation error was 20.1% and 3.2%, respectively. Four absorption lines of H2O were chosen and detonation flow was scanned at the same time. The detonation external flow testing system was set up for the valveless gas-liquid continuous pulse detonation engine with the diameter of 80 mm. Through scanning H2O absorption lines with a high frequency of 10 kHz, the on-line detection of detonation external flow was realized by direct absorption method combined with time-division multiplexing technology, and the reconstruction of dynamic temperature distribution was realized as well for the first time, both verifying the feasibility of the test method. The test results

  4. Detonation spreading in fine TATBs

    SciTech Connect

    Kennedy, J.E.; Lee, K.Y.; Spontarelli, T.; Stine, J.R.

    1998-12-31

    A test has been devised that permits rapid evaluation of the detonation-spreading (or corner-turning) properties of detonations in insensitive high explosives. The test utilizes a copper witness plate as the medium to capture performance data. Dent depth and shape in the copper are used as quantitative measures of the detonation output and spreading behavior. The merits of the test are that it is easy to perform with no dynamic instrumentation, and the test requires only a few grams of experimental explosive materials.

  5. Preparation of C60 by Detonation Technique

    NASA Astrophysics Data System (ADS)

    Wei, Xianfeng; Han, Yong; Long, Xinping

    2012-11-01

    A mixture of TNT (Trinitrotoluene) and natural graphite was detonated in a vacuum container which was immersed into cooling water; detonation products were collected for detecting. The results of mass spectroscopy, high performance liquid chromatography showed significant signals of C60, which proved that C60 could be synthesized by detonating the mixture of TNT/graphite and the detonation pressure was around 12.3 GPa and the detonation temperature was around 1985 K.

  6. Two phase detonation studies

    NASA Technical Reports Server (NTRS)

    Nicholls, J. A.; Pierce, T. H.; Miyajima, H.; Oza, R.; Patil, P.

    1974-01-01

    An experimental study of the passage of a shock wave over a burning fuel drop is described. This includes high speed framing photographs of the interaction taken at 500,000 frames per second. A theoretical prediction of the ignition of a fuel drop by a shock wave is presented and the results compared with earlier experimental work. Experimental attempts to generate a detonation in a liquid fuel drop (kerosene)-liquid oxidizer drop (hydrogen peroxide)-inert gas-environment are described. An appendix is included which gives the analytical prediction of power requirements for the drop generator to produce certain size drops at a certain mass rate. A bibliography is also included which lists all of the publications resulting from this research grant.

  7. New generation detonics

    SciTech Connect

    Souers, P.C.

    1996-12-15

    Modern theory is being used to accelerate the development of new high performance explosive molecules. Combining quantum chemistry calculations with synthesis of promising candidate molecules may enable the advance of the state of the art in this field by more than 50 years. We have established a high explosive performance prediction code by linking the thermochemical code CHEETAH with the ab initio electronic structure code GAUSSIAN and the molecular packing code MOLPAK. GAUSSIAN is first used to determine the shape of the molecule and its binding energy; the molecules are then packed together into a low energy configuration by MOLPAK. Finally, CHEETAH is used to transform the crystal energy and density into explosive performance measures such as detonation velocity, pressure, and energy. Over 70 target molecules have been created, and several of these show promise in combining performance, chemical stability, and ease of synthesis.

  8. Simulation of Acoustic Noise Generated by an Airbreathing, Beam-Powered Launch Vehicle

    NASA Astrophysics Data System (ADS)

    Kennedy, W. C.; Van Laak, P.; Scarton, H. A.; Myrabo, L. N.

    2005-04-01

    A simple acoustic model is developed for predicting the noise signature vs. power level for advanced laser-propelled lightcraft — capable of single-stage flights into low Earth orbit. This model predicts the noise levels generated by a pulsed detonation engine (PDE) during the initial lift-off and acceleration phase, for two representative `tractor-beam' lightcraft designs: a 1-place `Mercury' vehicle (2.5-m diameter, 900-kg); and a larger 5-place `Apollo' vehicle (5-m diameter, 5555-kg) — both the subject of an earlier study. The use of digital techniques to simulate the expected PDE noise signature is discussed, and three examples of fly-by noise signatures are presented. The reduction, or complete elimination of perceptible noise from such engines, can be accomplished by shifting the pulse frequency into the supra-audible or sub-audible range.

  9. Detonation of insensitive high explosives by a Q-switched ruby laser.

    NASA Technical Reports Server (NTRS)

    Yang, L. C.; Menichelli, V. J.

    1971-01-01

    Immediate longitudinal detonations have been observed in confined small-diameter columns of PETN, RDX, and tetryl by using a focused Q-switched ruby laser. The energy ranged from 0.8 to 4.0 J in a pulse width of 25 nsec. A 1000-A-thick aluminum film deposited on a glass window was used to generate a shock wave at the window-explosive interface when irradiated by the laser. In some cases, steady-state detonations were reached in less than .5 microsec with less than 10% variation in the detonation velocity.

  10. Double-Front Detonation Waves

    NASA Astrophysics Data System (ADS)

    Gubin, S. A.; Sumskoi, S. I.; Victorov, S. B.

    According to the theory of detonation, in a detonation wave there is a sound plane, named Chapman-Jouguet (CJ) plane. There are certain stationary parameters for this plane. In this work the possibility of the second CJ plane is shown. This second CJ plane is stationary as well. The physical mechanism of non-equilibrium transition providing the existence of the second CJ plane is presented. There is a non-equilibrium state, when the heat is removed from the reaction zone and the heat capacity decreases sharply. As a result of this non-equilibrium state, the sound velocity increases, and the local supersonic zone with second sonic plane (second CJ plane) appears. So the new mode of detonation wave is predicted. Equations describing this mode of detonation are presented. The exact analytical solution for the second CJ plane parameters is obtained. The example of double-front detonation in high explosive (TNT) is presented. In this double-front structure "nanodiamond-nanographite" phase transition takes place in condensed particles of detonation products.

  11. Aerodynamic characteristics of a series of airbreathing missile configurations

    NASA Technical Reports Server (NTRS)

    Hayes, C.

    1981-01-01

    Due to the interest in the application of airbreathing propulsion to missiles and the lack of a suitable data base, an experimental program has been conducted to contribute to such a data base. The configurations investigated were with twin-inlets, either two-dimensional or axisymmetric, each located at three circumferential locations. The effects of a wing located above the inlets and of tail configuration were investigated. Longitudinal stability and control and lateral-directional stability were included in the data obtained. This paper presents a summary of the program and some of the results obtained. Certain trends of the data, as well as problem areas, are discussed. Due to the large volume of data obtained, a detailed analysis is not presented.

  12. Integrated System Test of an Airbreathing Rocket (ISTAR)

    NASA Technical Reports Server (NTRS)

    Faulkner, Robert F.; Lyles, Garry (Technical Monitor)

    2001-01-01

    Rocket Based Combined Cycle (RBCC) propulsion system development and ground test is being conducted as part of the NASA Marshall Space Flight Center Integrated System Test of an Airbreathing Rocket (ISTAR) program. Rocketdyne, Aerojet and Pratt & Whitney have teamed as the Rocket Based Combined Cycle Consortium (RBC3) to work the propulsion system development. Each company offered unique RBCC propulsion concepts as candidates for the ISTAR propulsion system. A team of engine contractor, vehicle contractor and NASA representatives reviewed the concepts proposed by each company, reviewed the available data and selected the Aerojet RBCC propulsion system concept as the team propulsion system baseline for the ISTAR program. The ISTAR program is currently in a "Jumpstart" phase for development of the engine system leading to ground test of a thermally and power balanced RBCC propulsion system at Stennis Space Center in 2005. A parallel flight test demonstration of this propulsion system is anticipated to lead to first flight in the 2007 timeframe.

  13. 26th JANNAF Airbreathing Propulsion Subcommittee Meeting. Volume 1

    NASA Technical Reports Server (NTRS)

    Fry, Ronald S. (Editor); Gannaway, Mary T. (Editor)

    2002-01-01

    This volume, the first of four volumes, is a collection of 28 unclassified/unlimited-distribution papers which were presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 26th Airbreathing Propulsion Subcommittee (APS) was held jointly with the 38th Combustion Subcommittee (CS), 20th Propulsion Systems Hazards Subcommittee (PSHS), and 2nd Modeling and Simulation Subcommittee. The meeting was held 8-12 April 2002 at the Bayside Inn at The Sandestin Golf & Beach Resort and Eglin Air Force Base, Destin, Florida. Topics covered include: scramjet and ramjet R&D program overviews; tactical propulsion; space access; NASA GTX status; PDE technology; actively cooled engine structures; modeling and simulation of complex hydrocarbon fuels and unsteady processes; and component modeling and simulation.

  14. Instrumentation requirements from the user's view. [For airbreathing hypersonic engines

    SciTech Connect

    Harsha, P.T.

    1988-01-01

    The use of combustor diagnostics is considered from the point of view of demonstration of performance of an airbreathing hypersonic engine. The basic need is seen to be that of providing the data necessary to verify performance predictions for the engine as installed in the airplane. This necessitates the use of a diagnostics capability that can provide the inputs required by the computational analyses that will be used to assess this performance. Because of the cost of ground test facilities, a premium is placed on measurement technique reliability and redundancy of instrumentation. A mix of nonintrusive optical techniques and probe-based measurements is seen to be the best approach using current diagnostics capability; one such instrument mix is outlined for a ramjet/scramjet test program. 11 references.

  15. Air-breathing fuel cell stacks for portable power applications

    SciTech Connect

    Wilson, M.S.; DeCaro, D.; Neutzler, J.K.; Zawodzinski, C.; Gottesfeld, S.

    1996-10-01

    Increasing attention is being directed towards polymer electrolyte fuel cells as battery replacements because of their potentially superior energy densities and the possibility of `mechanical` refueling. On the low end of the power requirement scale (ca. 10 W), fuel cells can compete with primary and secondary batteries only if the fuel cell systems are simple, inexpensive, and reliable. Considerations of cost and simplicity (and minimal parasitic power) discourage the use of conventional performance enhancing subsystems (e.g., humidification, cooling, or forced-reactant flow). We are developing a stack design that is inherently self-regulating to allow effective operation without the benefit of such auxiliary components. The air cathode does not use forced flow to replenish the depleted oxygen. Instead, the oxygen in the air must diffuse into the stack from the periphery of the unit cells. For this reason the stack is described as `air-breathing.` This configuration limits the ability of water to escape which prevents the polymer electrolyte membranes from drying out, even at relatively high continuous operation temperatures (+60 degrees C). This results in stacks with reliable and stable performance. This air-breathing configuration assumes a unique stack geometry that utilizes circular flow-field plates with an annular hydrogen feed manifold and the single tie-bolt extending up through the central axis of the stack. With this geometry, the hydrogen supply to the unit cells is radially outward, and the air supply is from the periphery inward. This configuration has several advantages. The entire periphery is free to air access and allows greater heat conduction to enhance cooling. Furthermore, all of the components in the stack (e.g., the flow-fields, seals and membrane/electrode assemblies), are radially symmetrical, so part fabrication is simple and the entire system is potentially low-cost. Lastly, this configuration is compact and lightweight.

  16. 29 CFR 1926.908 - Use of detonating cord.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... kinks, or angles that direct the cord back toward the oncoming line of detonation. (g) All detonating... direction in which the detonation is to proceed. (j) Detonators for firing the trunkline shall not...

  17. Combining MHD Airbreathing and Fusion Rocket Propulsion for Earth-to-Orbit Flight

    SciTech Connect

    Froning, H. D. Jr; Yang, Yang; Momota, H.; Burton, E.; Miley, G. H.; Luo, Nie

    2005-02-06

    Previous studies have shown that Single-State-to-Orbit (SSTO) vehicle propellant can be reduced by Magnets-Hydro-Dynamic (MHD) processes that minimize airbreathing propulsion losses and propellant consumption during atmospheric flight. Similarly additional reduction in SSTO propellant is enabled by Inertial Electrostatic Confinement (IEC) fusion, whose more energetic reactions reduce rocket propellant needs. MHD airbreathing propulsion during an SSTO vehicle's initial atmospheric flight phase and IEC fusion propulsion during its final exo-atmospheric flight phase is therefore being explored. Accomplished work is not yet sufficient for claiming such a vehicle's feasibility. But takeoff and propellant mass for an MHD airbreathing and IEC fusion vehicle could be as much as 25 and 40 percent less than one with ordinary airbreathing and IEC fusion; and as much as 50 and 70 percent less than SSTO takeoff and propellant mass with MHD airbreathing and chemical rocket propulsion. Thus this unusual combined cycle engine shows great promise for performance gains beyond contemporary combined-cycle airbreathing engines.

  18. Detonation in TATB Hemispheres

    SciTech Connect

    Druce, B; Souers, P C; Chow, C; Roeske, F; Vitello, P; Hrousis, C

    2004-03-17

    Streak camera breakout and Fabry-Perot interferometer data have been taken on the outer surface of 1.80 g/cm{sup 3} TATB hemispherical boosters initiated by slapper detonators at three temperatures. The slapper causes breakout to occur at 54{sup o} at ambient temperatures and 42{sup o} at -54 C, where the axis of rotation is 0{sup o}. The Fabry velocities may be associated with pressures, and these decrease for large timing delays in breakout seen at the colder temperatures. At room temperature, the Fabry pressures appear constant at all angles. Both fresh and decade-old explosive are tested and no difference is seen. The problem has been modeled with reactive flow. Adjustment of the JWL for temperature makes little difference, but cooling to -54 C decreases the rate constant by 1/6th. The problem was run both at constant density and with density differences using two different codes. The ambient code results show that a density difference is probably there but it cannot be quantified.

  19. Stability of cosmological detonation fronts

    NASA Astrophysics Data System (ADS)

    Mégevand, Ariel; Membiela, Federico Agustín

    2014-05-01

    The steady-state propagation of a phase-transition front is classified, according to hydrodynamics, as a deflagration or a detonation, depending on its velocity with respect to the fluid. These propagation modes are further divided into three types, namely, weak, Jouguet, and strong solutions, according to their disturbance of the fluid. However, some of these hydrodynamic modes will not be realized in a phase transition. One particular cause is the presence of instabilities. In this work we study the linear stability of weak detonations, which are generally believed to be stable. After discussing in detail the weak detonation solution, we consider small perturbations of the interface and the fluid configuration. When the balance between the driving and friction forces is taken into account, it turns out that there are actually two different kinds of weak detonations, which behave very differently as functions of the parameters. We show that the branch of stronger weak detonations are unstable, except very close to the Jouguet point, where our approach breaks down.

  20. DETONATION PRESSURE MEASUREMENTS ON PETN

    SciTech Connect

    Green, L G; Lee, E L

    2006-06-23

    PETN is widely recognized as an example of nearly ideal detonation performance. The chemical composition is such that little or no carbon is produced in the detonation products. The reaction zone width is less than currently detectable. (<1 ns) Observations on PETN have thus become a baseline for EOS model predictions. It has therefore become important to characterize the detonation parameters as accurately as possible in order to provide the most exacting comparisons of EOS predictions with experimental results. We undertook a painstaking review of the detonation pressure measurements reported in an earlier work that was presented at the Fifth Detonation Symposium and found that corrections were required in determining the shock velocity in the PMMA witness material. We also refined the impedance calculation to account for the difference between the usual ''acoustic'' method and the more accurate Riemann integral. Our review indicates that the CJ pressures previously reported for full density PETN require an average lowering of about 6 percent. The lower densities require progressively smaller corrections. We present analysis of the records, supporting hydrodynamic simulations, the Riemann integral results, and EOS parameter values derived from the revised results.

  1. Point Measurement of Detonation Wave Speed

    NASA Astrophysics Data System (ADS)

    Lu, F. K.; Gupta, N. K. M.; Wilson, D. R.

    Accurate determination of the speed of a detonation wave is important for studies of detonation phenomena. Different types of sensors that measure pressure, ion and flame have been used for this purpose.

  2. Exhaust Gas Emissions from a Rotating Detonation-wave Engine

    NASA Astrophysics Data System (ADS)

    Kailasanath, Kazhikathra; Schwer, Douglas

    2015-11-01

    Rotating detonation-wave engines (RDE) are a form of continuous detonation-wave engines. They potentially provide further gains in performance than an intermittent or pulsed detonation-wave engine (PDE). The overall flow field in an idealized RDE, primarily consisting of two concentric cylinders, has been discussed in previous meetings. Because of the high pressures involved and the lack of adequate reaction mechanisms for this regime, previous simulations have typically used simplified chemistry models. However, understanding the exhaust species concentrations in propulsion devices is important for both performance considerations as well as estimating pollutant emissions. Progress towards addressing this need will be discussed in this talk. In this approach, an induction parameter model is used for simulating the detonation but a more detailed finite-chemistry model including NOx chemistry is used in the expansion flow region, where the pressures are lower and the uncertainties in the chemistry model are greatly reduced. Results show that overall radical concentrations in the exhaust flow are substantially lower than from earlier predictions with simplified models. The performance of a baseline hydrogen/air RDE increased from 4940 s to 5000 s with the expansion flow chemistry, due to recombination of radicals and more production of H2O, resulting in additional heat release. Work sponsored by the Office of Naval Research.

  3. Simulations of a Detonation Wave in Transverse Magnetic Fields

    NASA Astrophysics Data System (ADS)

    Cole, Lord; Karagozian, Ann; Cambier, Jean-Luc

    2010-11-01

    Numerical simulations of magneto-hydrodynamic (MHD) effects on detonation wave structures are performed, with applications to flow control and MHD power extraction in Pulse Detonation Engines (PDE) and their design variations. In contrast to prior studies of MHD interactions in PDEs,ootnotetextCambier, et al., AIAA-2008-4688 the effects of the finite relaxation length scale for ionization on the stability of the detonation wave are examined. Depending on the coupling parameters, the magnetic field can quench the detonation and effectively act as a barrier to its propagation. Conversely, an applied transient magnetic field can exert a force on a pre-ionized gas and accelerate it. The dynamics are subject to non-linear effects; a propagating transverse magnetic field will initially exert a small force if the gas has a low conductivity and the magnetic Reynolds number (Rem) is low. Nevertheless, the gas accelerated by the "piston" action of the field can pre-heat the ambient gas and increase its conductivity. As the wave progresses, Rem increases and the magnetic field becomes increasingly effective. The dynamics of this process are examined in detail with a high-order shock-capturing method and full kinetics of combustion and ionization. The complex chemical kinetics calculations are ported onto a GPU using the CUDA language, and computational performance is compared with standard CPU-based computations.

  4. Development and qualification testing of a laser-ignited, all-secondary (DDT) detonator

    NASA Technical Reports Server (NTRS)

    Blachowski, Thomas J.; Krivitsky, Darrin Z.; Tipton, Stephen

    1994-01-01

    The Indian Head Division, Naval Surface Warfare Center (IHDIV, NSWC) is conducting a qualification program for a laser-ignited, all-secondary (DDT) explosive detonator. This detonator was developed jointly by IHDIV, NSWC and the Department of Energy's EG&G Mound Applied Technologies facility in Miamisburg, Ohio to accept a laser initiation signal and produce a fully developed shock wave output. The detonator performance requirements were established by the on-going IHDIV, NSWC Laser Initiated Transfer Energy Subsystem (LITES) advanced development program. Qualification of the detonator as a component utilizing existing military specifications is the selected approach for this program. The detonator is a deflagration-to-detonator transfer (DDT) device using a secondary explosive, HMX, to generate the required shock wave output. The prototype development and initial system integration tests for the LITES and for the detonator were reported at the 1992 International Pyrotechnics Society Symposium and at the 1992 Survival and Flight Equipment National Symposium. Recent results are presented for the all-fire sensitivity and qualification tests conducted at two different laser initiation pulses.

  5. Time-resolved spectroscopic studies of detonating heterogeneous explosives. [HMX and HNS

    SciTech Connect

    Trott, W.M.; Renlund, A.M.

    1985-01-01

    Emission spectroscopy and pulsed-laser-excited Raman scattering methods have been applied to the study of detonating heterogeneous explosives, including PETN, HMX and HNS. Time-resolved spectra of emission from detonating HNS show the evolution of features due to electronically-excited radical species. For HNS, the CN(B-X) system near 388 nm has been studied at a wavelength resolution of 0.5 A. Boltzmann vibrational temperatures have been calculated by comparing the experimental data with computer-simulated spectra. These temperatures are consistent with the expected trend of detonation temperature as a function of charge density. Using 532-nm laser excitation, single-pulse Raman scattering measurements have been made at the free surface of detonating HMX and PETN samples. Monotonic attenuation of Raman scattering intensity over a 100-ns interval is observed after detonation front arrival at the free surface. Depletion of the Raman signal occurs prior to significant loss of the scattered laser light. The significance of the Raman measurements as a possible probe of reaction zone length in detonating explosives is discussed. 21 refs., 11 figs.

  6. Equation of state for detonation products. [Detonation products

    SciTech Connect

    Davis, W.C.

    1985-01-01

    The concepts of hydrodynamics and thermodynamics as they apply to equations of state for explosive products are collected and discussed. The physics behind the behavior of dense gases is considered. Some ideas about applications are presented. This paper is intended as an introduction to the subject of equation of state for detonation products. 7 references, 3 figures.

  7. Fluid dynamics in a Rotating-Detonation-Engine with micro-injectors

    NASA Astrophysics Data System (ADS)

    Schwer, Douglas

    2011-11-01

    Rotating detonation engines (RDE's) represent a natural extension of the extensively studied pulse detonation engines (PDE's) for obtaining propulsion from the high efficiency detonation cycle. RDE's require fuel and oxidizer under high pressure to be injected through micro-nozzles from one or two plenums (for premixed and non-premixed). This injection process is critically important to the stability and performance of the RDE. This paper studies the effect of this injection process on the detonation wave within the combustion chamber, with an emphasis on how the fluid dynamics are affected. Both two-dimensional and three-dimensional simulations are done using well proven numerical methods for both the combustion chamber and mixture plenums of an idealized RDE. This work is supported through NRL 6.1 Computational Physics Task Area

  8. Environmentally Benign Stab Detonators

    SciTech Connect

    Gash, A E

    2006-07-07

    The coupling of energetic metallic multilayers (a.k.a. flash metal) with energetic sol-gel synthesis and processing is an entirely new approach to forming energetic devices for several DoD and DOE needs. They are also practical and commercially viable manufacturing techniques. Improved occupational safety and health, performance, reliability, reproducibility, and environmentally acceptable processing can be achieved using these methodologies and materials. The development and fielding of this technology will enhance mission readiness and reduce the costs, environmental risks and the necessity of resolving environmental concerns related to maintaining military readiness while simultaneously enhancing safety and health. Without sacrificing current performance, we will formulate new impact initiated device (IID) compositions to replace materials from the current composition that pose significant environmental, health, and safety problems associated with functions such as synthesis, material receipt, storage, handling, processing into the composition, reaction products from testing, and safe disposal. To do this, we will advance the use of nanocomposite preparation via the use of multilayer flash metal and sol-gel technologies and apply it to new small IIDs. This work will also serve to demonstrate that these technologies and resultant materials are relevant and practical to a variety of energetic needs of DoD and DOE. The goal will be to produce an IID whose composition is acceptable by OSHA, EPA, the Clean Air Act, Clean Water Act, Resource Recovery Act, etc. standards, without sacrificing current performance. The development of environmentally benign stab detonators and igniters will result in the removal of hazardous and toxic components associated with their manufacturing, handling, and use. This will lead to improved worker safety during manufacturing as well as reduced exposure of Service personnel during their storage and or use in operations. The

  9. Airbreathing Propulsion System Analysis Using Multithreaded Parallel Processing

    NASA Technical Reports Server (NTRS)

    Schunk, Richard Gregory; Chung, T. J.; Rodriguez, Pete (Technical Monitor)

    2000-01-01

    In this paper, parallel processing is used to analyze the mixing, and combustion behavior of hypersonic flow. Preliminary work for a sonic transverse hydrogen jet injected from a slot into a Mach 4 airstream in a two-dimensional duct combustor has been completed [Moon and Chung, 1996]. Our aim is to extend this work to three-dimensional domain using multithreaded domain decomposition parallel processing based on the flowfield-dependent variation theory. Numerical simulations of chemically reacting flows are difficult because of the strong interactions between the turbulent hydrodynamic and chemical processes. The algorithm must provide an accurate representation of the flowfield, since unphysical flowfield calculations will lead to the faulty loss or creation of species mass fraction, or even premature ignition, which in turn alters the flowfield information. Another difficulty arises from the disparity in time scales between the flowfield and chemical reactions, which may require the use of finite rate chemistry. The situations are more complex when there is a disparity in length scales involved in turbulence. In order to cope with these complicated physical phenomena, it is our plan to utilize the flowfield-dependent variation theory mentioned above, facilitated by large eddy simulation. Undoubtedly, the proposed computation requires the most sophisticated computational strategies. The multithreaded domain decomposition parallel processing will be necessary in order to reduce both computational time and storage. Without special treatments involved in computer engineering, our attempt to analyze the airbreathing combustion appears to be difficult, if not impossible.

  10. An Overview of SBIR Phase 2 Airbreathing Propulsion Technologies

    NASA Technical Reports Server (NTRS)

    Nguyen, Hung D.; Steele, Gynelle C.; Bitler, Dean W.

    2014-01-01

    Technological innovation is the overall focus of NASA's Small Business Innovation Research (SBIR) program. The program invests in the development of innovative concepts and technologies to help NASA's mission directorates address critical research and development needs for agency projects. This report highlights innovative SBIR Phase II projects from 2007-2012 specifically addressing areas in Airbreathing Propulsion which is one of six core competencies at NASA Glenn Research Center. There are twenty technologies featured with emphasis on a wide spectrum of applications such as with a Turbo-Brayton cryocooler for aircraft superconducting systems, braided composite rotorcraft structures, engine air brake, combustion control valve, flexible composite driveshaft, and much more. Each article in this booklet describes an innovation, technical objective, and highlights NASA commercial and industrial applications. This report serves as an opportunity for NASA personnel including engineers, researchers, and program managers to learn of NASA SBIR's capabilities that might be crosscutting into this technology area. As the result, it would cause collaborations and partnerships between the small companies and NASA Programs and Projects resulting in benefit to both SBIR companies and NASA.

  11. Airframe Research and Technology for Hypersonic Airbreathing Vehicles

    NASA Technical Reports Server (NTRS)

    Glass, David E.; Merski, N. Ronald; Glass, Christopher E.

    2002-01-01

    The Hypersonics Investment Area (HIA) within NASA's Advanced Space Transportation Program (ASTP) has the responsibility to develop hypersonic airbreathing vehicles for access to space. The Airframe Research and Technology (AR and T) Project, as one of six projects in the HIA, will push the state-of-the-art in airframe and vehicle systems for low-cost, reliable, and safe space transportation. The individual technologies within the project are focused on advanced, breakthrough technologies in airframe and vehicle systems and cross-cutting activities that are the basis for improvements in these disciplines. Both low and medium technology readiness level (TRL) activities are being pursued. The key technical areas that will be addressed by the project include analysis and design tools, integrated vehicle health management (IVHM), composite (polymer, metal, and ceramic matrix) materials development, thermal/structural wall concepts, thermal protection systems, seals, leading edges, aerothermodynamics, and airframe/propulsion flowpath technology. Each of the technical areas or sub-projects within the Airframe R and T Project is described in this paper.

  12. Thermodynamic Cycle Analysis of Magnetohydrodynamic-Bypass Hypersonic Airbreathing Engines

    NASA Technical Reports Server (NTRS)

    Litchford, R. J.; Cole, J. W.; Bityurin, V. A.; Lineberry, J. T.

    2000-01-01

    The prospects for realizing a magnetohydrodynamic (MHD) bypass hypersonic airbreathing engine are examined from the standpoint of fundamental thermodynamic feasibility. The MHD-bypass engine, first proposed as part of the Russian AJAX vehicle concept, is based on the idea of redistributing energy between various stages of the propulsion system flow train. The system uses an MHD generator to extract a portion of the aerodynamic heating energy from the inlet and an MHD accelerator to reintroduce this power as kinetic energy in the exhaust stream. In this way, the combustor entrance Mach number can be limited to a specified value even as the flight Mach number increases. Thus, the fuel and air can be efficiently mixed and burned within a practical combustor length, and the flight Mach number operating envelope can be extended. In this paper, we quantitatively assess the performance potential and scientific feasibility of MHD-bypass engines using a simplified thermodynamic analysis. This cycle analysis, based on a thermally and calorically perfect gas, incorporates a coupled MHD generator-accelerator system and accounts for aerodynamic losses and thermodynamic process efficiencies in the various engin components. It is found that the flight Mach number range can be significantly extended; however, overall performance is hampered by non-isentropic losses in the MHD devices.

  13. Thermodynamic Cycle Analysis of Magnetohydrodynamic-Bypass Airbreathing Hypersonic Engines

    NASA Technical Reports Server (NTRS)

    Litchford, Ron J.; Bityurin, Valentine A.; Lineberry, John T.

    1999-01-01

    Established analyses of conventional ramjet/scramjet performance characteristics indicate that a considerable decrease in efficiency can be expected at off-design flight conditions. This can be explained, in large part, by the deterioration of intake mass flow and limited inlet compression at low flight speeds and by the onset of thrust degradation effects associated with increased burner entry temperature at high flight speeds. In combination, these effects tend to impose lower and upper Mach number limits for practical flight. It has been noted, however, that Magnetohydrodynamic (MHD) energy management techniques represent a possible means for extending the flight Mach number envelope of conventional engines. By transferring enthalpy between different stages of the engine cycle, it appears that the onset of thrust degradation may be delayed to higher flight speeds. Obviously, the introduction of additional process inefficiencies is inevitable with this approach, but it is believed that these losses are more than compensated through optimization of the combustion process. The fundamental idea is to use MHD energy conversion processes to extract and bypass a portion of the intake kinetic energy around the burner. We refer to this general class of propulsion system as an MHD-bypass engine. In this paper, we quantitatively assess the performance potential and scientific feasibility of MHD-bypass airbreathing hypersonic engines using ideal gasdynamics and fundamental thermodynamic principles.

  14. Detonation waves in relativistic hydrodynamics

    SciTech Connect

    Cissoko, M. )

    1992-02-15

    This paper is concerned with an algebraic study of the equations of detonation waves in relativistic hydrodynamics taking into account the pressure and the energy of thermal radiation. A new approach to shock and detonation wavefronts is outlined. The fluid under consideration is assumed to be perfect (nonviscous and nonconducting) and to obey the following equation of state: {ital p}=({gamma}{minus}1){rho} where {ital p}, {rho}, and {gamma} are the pressure, the total energy density, and the adiabatic index, respectively. The solutions of the equations of detonation waves are reduced to the problem of finding physically acceptable roots of a quadratic polynomial {Pi}({ital X}) where {ital X} is the ratio {tau}/{tau}{sub 0} of dynamical volumes behind and ahead of the detonation wave. The existence and the locations of zeros of this polynomial allow it to be shown that if the equation of state of the burnt fluid is known then the variables characterizing the unburnt fluid obey well-defined physical relations.

  15. Vapor phase explosions: elementary detonations?

    PubMed

    Fowles, G R

    1979-04-13

    Although liquid-vapor explosions are widely observed, there is no established explanation for their initiation and propagation. Thermodynamics admits the possibility that superheated liquids can support detonations analogous to those that occur in chemical explosives. For liquid methane superheated 50 K above its boiling point at 1 atmosphere, the energy of explosion is 2 to 3 percent of that of TNT. PMID:17738085

  16. Sensitized Liquid Hydrazine Detonation Studies

    NASA Technical Reports Server (NTRS)

    Rathgeber, K. A.; Keddy, C. P.; Bunker, R. L.

    1999-01-01

    Vapor-phase hydrazine (N2H4) is known to be very sensitive to detonation while liquid hydrazine is very insensitive to detonation, theoretically requiring extremely high pressures to induce initiation. A review of literature on solid and liquid explosives shows that when pure explosive substances are infiltrated with gas cavities, voids, and/or different phase contaminants, the energy or shock pressure necessary to induce detonation can decrease by an order of magnitude. Tests were conducted with liquid hydrazine in a modified card-gap configuration. Sensitization was attempted by bubbling helium gas through and/or suspending ceramic microspheres in the liquid. The hydrazine was subjected to the shock pressure from a 2 lb (0.9 kg) Composition C-4 explosive charge. The hydrazine was contained in a 4 in. (10.2 cm) diameter stainless steel cylinder with a 122 in(sup 3) (2 L) volume and sealed with a polyethylene cap. Blast pressures from the events were recorded by 63 high speed pressure transducers located on three radial legs extending from 4 to 115 ft (1.2 to 35.1 in) from ground zero. Comparison of the neat hydrazine and water baseline tests with the "sensitized" hydrazine tests indicates the liquid hydrazine did not detonate under these conditions.

  17. Performance Impact of Deflagration to Detonation Transition Enhancing Obstacles

    NASA Technical Reports Server (NTRS)

    Paxson, Daniel E.; Schauer, Frederick; Hopper, David

    2012-01-01

    A sub-model is developed to account for the drag and heat transfer enhancement resulting from deflagration-to-detonation (DDT) inducing obstacles commonly used in pulse detonation engines (PDE). The sub-model is incorporated as a source term in a time-accurate, quasi-onedimensional, CFD-based PDE simulation. The simulation and sub-model are then validated through comparison with a particular experiment in which limited DDT obstacle parameters were varied. The simulation is then used to examine the relative contributions from drag and heat transfer to the reduced thrust which is observed. It is found that heat transfer is far more significant than aerodynamic drag in this particular experiment.

  18. A conceptual design of an unmanned test vehicle using an airbreathing propulsion system

    NASA Technical Reports Server (NTRS)

    1992-01-01

    According to Aviation Week and Space Technology (Nov. 16, 1992), without a redefined approach to the problem of achieving single stage-to-orbit flight, the X-30 program is virtually assured of cancellation. One of the significant design goals of the X-30 program is to achieve single stage to low-earth orbit using airbreathing propulsion systems. In an attempt to avoid cancellation, the NASP Program has decided to design a test vehicle to achieve these goals. This report recommends a conceptual design of an unmanned test vehicle using an airbreathing propulsion system.

  19. Evaluation of some significant issues affecting trajectory and control management for air-breathing hypersonic vehicles

    NASA Technical Reports Server (NTRS)

    Hattis, Philip D.; Malchow, Harvey L.

    1992-01-01

    Horizontal takeoff airbreathing-propulsion launch vehicles require near-optimal guidance and control which takes into account performance sensitivities to atmospheric characteristics while satisfying physically-derived operational constraints. A generic trajectory/control analysis tool that deepens insight into these considerations has been applied to two versions of a winged-cone vehicle model. Information that is critical to the design and trajectory of these vehicles is derived, and several unusual characteristics of the airbreathing propulsion model are shown to have potentially substantial effects on vehicle dynamics.

  20. HERMES: A Model to Describe Deformation, Burning, Explosion, and Detonation

    SciTech Connect

    Reaugh, J E

    2011-11-22

    pressure that results from a more gradual increase. This disagrees with experiments, where explosives were subjected to a gradual rise in pressure and did not exhibit reaction. More recent models do distinguish between slow pressure rises and shocks, and have had some success in the describing the response of explosives to single and multiple shocks, and the increase of shock sensitivity with porosity, at least over a limited range. The original formulation is appropriate for sustained shocks, but further work is ongoing to describe the response to short pulses. The HERMES model combines features from these prior models. It describes burning and explosion in damaged reactant, and also will develop a detonation if the gradual rise in pressure from burning steepens into a strong-enough shock. The shock strength needed for detonation in a fixed run distance decreases with increasing porosity.

  1. Environmentally Benign Stab Detonators

    SciTech Connect

    Gash, A; Barbee, T; Simpson, R; Satcher, J; Walton, C

    2003-12-15

    This effort attempts to demonstrate that environmentally acceptable energetic sol-gel coated flash metal multilayer nanocomposites can be used to replace current impact initiated devices (IIDs) which have hazardous and toxic components. Successful completion of this project will result in IIDs that include innocuous compounds, have sufficient output energy for initiation, meet current military specifications, are small, cost competitive, and perform as well as or better than current devices. We expect flash metal multilayer and sol-gel to be generic technologies applicable to a wide range of devices, especially in small caliber ammunition and sub-munitions. We will replace the NOL-130 mixture with a nanocomposite that consists of a mechanically robust energetic multilayer foil that has been coated with a sol-gel energetic material. The exothermic reactions are activated in this nanocomposite are the transformation of the multilayer material to its respective intermetallic alloy and the thermite reaction, which is characterized by very high temperatures, a small pressure pulse, and hot particle ejection. The proposed materials and their reaction products consist of, but are not limited to aluminum, nickel, iron, aluminum oxide, titanium, iron oxide and boron. These materials have much more desirable environmental and health characteristics than the NOL-130 composition.

  2. Non-detonable explosive simulators

    DOEpatents

    Simpson, R.L.; Pruneda, C.O.

    1994-11-01

    A simulator which is chemically equivalent to an explosive, but is not detonable. The simulator has particular use in the training of explosives detecting dogs and calibrating sensitive analytical instruments. The explosive simulants may be fabricated by different techniques, a first involves the use of standard slurry coatings to produce a material with a very high binder to explosive ratio without masking the explosive vapor, and the second involves coating inert beads with thin layers of explosive molecules. 5 figs.

  3. Non-detonable explosive simulators

    DOEpatents

    Simpson, Randall L.; Pruneda, Cesar O.

    1994-01-01

    A simulator which is chemically equivalent to an explosive, but is not detonable. The simulator has particular use in the training of explosives detecting dogs and calibrating sensitive analytical instruments. The explosive simulants may be fabricated by different techniques, a first involves the use of standard slurry coatings to produce a material with a very high binder to explosive ratio without masking the explosive vapor, and the second involves coating inert beads with thin layers of explosive molecules.

  4. Improved detonation modeling with CHEETAH

    SciTech Connect

    Heller, A.

    1997-11-01

    A Livermore software program called CHEETAH, an important, even indispensable tool for energetic materials researchers worldwide, was made more powerful in the summer of 1997 with the release of CHEETAH 2.0, an advanced version that simulates a wider variety of detonations. Derived from more than 40 years of experiments on high explosives at Lawrence Livermore and Los Alamos national laboratories, CHEETAH predicts the results from detonating a mixture of specified reactants. It operates by solving thermodynamic equations to predict detonation products and such properties as temperature, pressure, volume, and total energy released. The code is prized by synthesis chemists and other researchers because it allows them to vary the starting molecules and conditions to optimize the desired performance properties. One of the Laboratory`s most popular computer codes, CHEETAH is used at more than 200 sites worldwide, including ones in England, Canada, Sweden, Switzerland, and France. Most sites are defense-related, although a few users, such as Japanese fireworks researchers, are in the civilian sector.

  5. Detonation interaction with an interface

    NASA Astrophysics Data System (ADS)

    Lieberman, D. H.; Shepherd, J. E.

    2007-09-01

    Detonation interaction with an interface was investigated, where the interface separated a combustible from an oxidizing or inert mixture. The ethylene-oxygen combustible mixture had a fuel-rich composition to promote secondary combustion with the oxidizer in the turbulent mixing zone (TMZ) that resulted from the interaction. Sharp interfaces were created by using a nitro-cellulose membrane to separate the two mixtures. The membrane was mounted on a wood frame and inserted in the experimental test section at a 45° angle to the bulk flow direction. The membrane was destroyed by the detonation wave. The interaction resulted in a transmitted and reflected wave at a node point similar to regular shock refraction. A detonation refraction analysis was carried out to compare with the measured shock angles. It was observed that the measured angle is consistently lower than the predicted value. An uncertainty analysis revealed possible explanations for this systematic variation pointing to factors such as the incident wave curvature and the role of the nitro-cellulose diaphragm. Analysis of the TMZ and Mach stem formed from the reflection of the transmitted shock wave off the solid boundary were carried out and found to justify the size and strength of these features as a function of the test gas composition. The role of secondary combustion in the TMZ was also investigated and found to have a small influence on the wave structure.

  6. Detonability of hydrocarbon fuels in air

    NASA Technical Reports Server (NTRS)

    Beeson, H. D.; Mcclenagan, R. D.; Bishop, C. V.; Benz, F. J.; Pitz, W. J.; Westbrook, C. K.; Lee, J. H. S.

    1991-01-01

    Studies were conducted of the detonation of gas-phase mixtures of n-hexane and JP-4, with oxidizers as varied as air and pure oxygen, measuring detonation velocities and cell sizes as a function of stoichiometry and diluent concentration. The induction length of a one-dimensional Zeldovich-von Neumann-Doering detonation was calculated on the basis of a theoretical model that employed the reaction kinetics of the hydrocarbon fuels used. Critical energy and critical tube diameter are compared for a relative measure of the heavy hydrocarbon fuels studied; detonation sensitivity appears to increase slightly with increasing carbon number.

  7. Radioactive Fallout from Terrorist Nuclear Detonations

    SciTech Connect

    Marrs, R E

    2007-05-03

    Responding correctly during the first hour after a terrorist nuclear detonation is the key to reducing casualties from a low-yield surface burst, and a correct response requires an understanding of the rapidly changing dose rate from fallout. This report provides an empirical formula for dose rate as a function of time and location that can guide the response to an unexpected nuclear detonation. At least one post-detonation radiation measurement is required if the yield and other characteristics of the detonation are unknown.

  8. Deflagrations and detonations in thermonuclear supernovae.

    PubMed

    Gamezo, Vadim N; Khokhlov, Alexei M; Oran, Elaine S

    2004-05-28

    We study a type Ia supernova explosion using three-dimensional numerical simulations based on reactive fluid dynamics. We consider a delayed-detonation model that assumes a deflagration-to-detonation transition. In contrast with the pure deflagration model, the delayed-detonation model releases enough energy to account for a healthy explosion, and does not leave carbon, oxygen, and intermediate-mass elements in central parts of a white dwarf. This removes the key disagreement between simulations and observations, and makes a delayed detonation the mostly likely mechanism for type Ia supernovae. PMID:15245271

  9. Two phase detonation studies conducted in 1971

    NASA Technical Reports Server (NTRS)

    Nicholls, J. A.

    1972-01-01

    A report is presented describing the research conducted on five phases: (1) ignition of fuel drops by a shock wave and passage of a shock wave over a burning drop, (2) the energy release pattern of a two-phase detonation with controlled drop sizes, (3) the attenuation of shock and detonation waves passing over an acoustic liner, (4) experimental and theoretical studies of film detonations, and (5) a simplified analytical model of a rotating two-phase detonation wave in a rocket motor.

  10. A summary of hydrogen-air detonation experiments

    SciTech Connect

    Guirao, C.M.; Knystautas, R.; Lee, J.H.

    1989-05-01

    Dynamic detonation parameters are reviewed for hydrogen-air-diluent detonations and deflagration-to-detonation transitions (DDT). These parameters include the characteristic chemical length scale, such as the detonation cell width, associated with the three-dimensional cellular structure of detonation waves, critical transmission conditions of confined detonations into unconfined environments, critical initiation energy for unconfined detonations, detonability limits, and critical conditions for DDT. The detonation cell width, which depends on hydrogen and diluent concentrations, pressure, and temperature, is an important parameter in the prediction of critical geometry-dependent conditions for the transmission of confined detonations into unconfined environments and the critical energies for the direct initiation of unconfined detonations. Detonability limits depend on both initial and boundary conditions and the limit has been defined as the onset of single head spin. Four flame propagation regimes have been identified and the criterion for DDT in a smooth tube is discussed. 108 refs., 28 figs., 5 tabs.

  11. Hydrogen-oxygen flame acceleration and transition to detonation in channels with no-slip walls for a detailed chemical reaction model.

    PubMed

    Ivanov, M F; Kiverin, A D; Liberman, M A

    2011-05-01

    The features of flame acceleration in channels with wall friction and the deflagration to detonation transition (DDT) are investigated theoretically and using high resolution numerical simulations of two-dimensional reactive Navier-Stokes equations, including the effects of viscosity, thermal conduction, molecular diffusion, and a detailed chemical reaction mechanism for hydrogen-oxygen gaseous mixture. It is shown that in a wide channel, from the beginning, the flame velocity increases exponentially for a short time and then flame acceleration decreases, ending up with the abrupt increase of the combustion wave velocity and the actual transition to detonation. In a thin channel with a width smaller than the critical value, the exponential increase of the flame velocity is not bounded and ends up with the transition to detonation. The transition to detonation occurs due to the pressure pulse, which is formed at the tip of the accelerating flame. The amplitude of the pressure pulse grows exponentially due to a positive feedback coupling between the pressure pulse and the heat released in the reaction. Finally, large amplitude pressure pulse steepens into a strong shock coupled with the reaction zone forming the overdriven detonation. The evolution from a temperature gradient to a detonation via the Zeldovich gradient mechanism and its applicability to the deflagration-to-detonation transition is investigated for combustible materials whose chemistry is governed by chain-branching kinetics. The results of the high resolution simulations are fully consistent with experimental observations of the flame acceleration and DDT. PMID:21728653

  12. Plasma-assisted ignition and deflagration-to-detonation transition.

    PubMed

    Starikovskiy, Andrey; Aleksandrov, Nickolay; Rakitin, Aleksandr

    2012-02-13

    Non-equilibrium plasma demonstrates great potential to control ultra-lean, ultra-fast, low-temperature flames and to become an extremely promising technology for a wide range of applications, including aviation gas turbine engines, piston engines, RAMjets, SCRAMjets and detonation initiation for pulsed detonation engines. The analysis of discharge processes shows that the discharge energy can be deposited into the desired internal degrees of freedom of molecules when varying the reduced electric field, E/n, at which the discharge is maintained. The amount of deposited energy is controlled by other discharge and gas parameters, including electric pulse duration, discharge current, gas number density, gas temperature, etc. As a rule, the dominant mechanism of the effect of non-equilibrium plasma on ignition and combustion is associated with the generation of active particles in the discharge plasma. For plasma-assisted ignition and combustion in mixtures containing air, the most promising active species are O atoms and, to a smaller extent, some other neutral atoms and radicals. These active particles are efficiently produced in high-voltage, nanosecond, pulse discharges owing to electron-impact dissociation of molecules and electron-impact excitation of N(2) electronic states, followed by collisional quenching of these states to dissociate the molecules. Mechanisms of deflagration-to-detonation transition (DDT) initiation by non-equilibrium plasma were analysed. For longitudinal discharges with a high power density in a plasma channel, two fast DDT mechanisms have been observed. When initiated by a spark or a transient discharge, the mixture ignited simultaneously over the volume of the discharge channel, producing a shock wave with a Mach number greater than 2 and a flame. A gradient mechanism of DDT similar to that proposed by Zeldovich has been observed experimentally under streamer initiation. PMID:22213667

  13. Survey of Aerothermodynamics Facilities Useful for the Design of Hypersonic Vehicles Using Air-Breathing Propulsion

    NASA Technical Reports Server (NTRS)

    Arnold, James O.; Deiwert, George S.

    1997-01-01

    This paper surveys the use of aerothermodynamic facilities which have been useful in the study of external flows and propulsion aspects of hypersonic, air-breathing vehicles. While the paper is not a survey of all facilities, it covers the utility of shock tunnels and conventional hypersonic blow-down facilities which have been used for hypersonic air-breather studies. The problems confronting researchers in the field of aerothermodynamics are outlined. Results from the T5 GALCIT tunnel for the shock-on lip problem are outlined. Experiments on combustors and short expansion nozzles using the semi-free jet method have been conducted in large shock tunnels. An example which employed the NASA Ames 16-Inch shock tunnel is outlined, and the philosophy of the test technique is described. Conventional blow-down hypersonic wind tunnels are quite useful in hypersonic air-breathing studies. Results from an expansion ramp experiment, simulating the nozzle on a hypersonic air-breather from the NASA Ames 3.5 Foot Hypersonic wind tunnel are summarized. Similar work on expansion nozzles conducted in the NASA Langley hypersonic wind tunnel complex is cited. Free-jet air-frame propulsion integration and configuration stability experiments conducted at Langley in the hypersonic wind tunnel complex on a small generic model are also summarized.

  14. Analysis of a Nuclear Enhanced Airbreathing Rocket for Earth to Orbit Applications

    NASA Technical Reports Server (NTRS)

    Adams, Robert B.; Landrum, D. Brian; Brown, Norman (Technical Monitor)

    2001-01-01

    The proposed engine concept is the Nuclear Enhanced Airbreathing Rocket (NEAR). The NEAR concept uses a fission reactor to thermally heat a propellant in a rocket plenum. The rocket is shrouded, thus the exhaust mixes with ingested air to provide additional thermal energy through combustion. The combusted flow is then expanded through a nozzle to provide thrust.

  15. Detonation diffraction through different geometries

    NASA Astrophysics Data System (ADS)

    Sorin, Rémy; Zitoun, Ratiba; Khasainov, Boris; Desbordes, Daniel

    2009-04-01

    We performed the study of the diffraction of a self-sustained detonation from a cylindrical tube (of inner diameter d) through different geometric configurations in order to characterise the transmission processes and to quantify the transmission criteria to the reception chamber. For the diffraction from a tube to the open space the transmission criteria is expressed by d c = k c · λ (with λ the detonation cell size and k c depending on the mixture and on the operture configuration, classically 13 for alkane mixtures with oxygen). The studied geometries are: (a) a sharp increase of diameter ( D/ d > 1) with and without a central obstacle in the diffracting section, (b) a conical divergent with a central obstacle in the diffracting section and (c) an inversed intermediate one end closed tube insuring a double reflection before a final diffraction between the initiator tube and the reception chamber. The results for case A show that the reinitiation process depends on the ratio d/ λ. For ratios below k c the re-ignition takes place at the receptor tube wall and at a fixed distance from the step, i.e. closely after the diffracted shock reflection shows a Mach stem configuration. For ratios below a limit ratio k lim (which depends on D/ d) the re-ignition distance increases with the decrease of d/λ. For both case A and B the introduction of a central obstacle (of blockage ratio BR = 0.5) at the exit of the initiator tube decreases the critical transmission ratio k c by 50%. The results in configuration C show that the re-ignition process depends both on d/ λ and the geometric conditions. Optimal configuration is found that provides the transmission through the two successive reflections (from d = 26 mm to D ch = 200 mm) at as small d/ λ as 2.2 whatever the intermediate diameter D is. This configuration provides a significant improvement in the detonation transmission conditions.

  16. Phase detonated shock tube (PFST)

    SciTech Connect

    Zerwekh, W.D.; Marsh, S.P.; Tan, Tai-Ho

    1993-07-01

    The simple, cylindrically imploding and axially driven fast shock tube (FST) has been a basic component in the high velocity penetrator (HVP) program. It is a powerful device capable of delivering a directed and very high pressure output that has been successfully employed to drive hypervelocity projectiles. The FST is configured from a hollow, high-explosive (HE) cylinder, a low-density Styrofoam core, and a one-point initiator at one end. A Mach stem is formed in the core as the forward-propagating, HE detonation wave intersects the reflected radial wave. This simple FST has been found to be a powerful pressure multiplier. Up to 1-Mbar output pressure can be obtained from this device. Further increase in the output pressure can be achieved by increasing the HE detonation velocity. The FST has been fine tuned to drive a thin plate to very high velocity under an impulse per unit area of about 1 Mbar{mu}s/cm{sup 2}. A 1.5-mm-thick stainless steel disk has been accelerated intact to 0.8 cm/{mu}s under a loading pressure rate of several Mbar/{mu}s. By making the plate curvature slightly convex at the loading side the authors have successfully accelerated it to almost 1.0 cm/{mu}s. The incorporation of a barrel at the end of the FST has been found to be important as confinement of the propellant gas by the barrel tends to accelerate the projectile to higher velocity. The desire to accelerate the plate above 1.0 cm/{mu}s provided the impetus to develop a more advanced fast shock tube to deliver a much higher output pressure. This report describes the investigation of a relatively simple air-lens phase-detonation system (PFST) with fifty percent higher phase-detonation velocity and a modest 2 Mbar output. Code calculations have shown that this PFST acceleration of a plate to about 1.2 cm/{mu}s can be achieved. The performance of these PFSTs has been evaluated and the details are discussed.

  17. Phase detonated shock tube (PFST)

    SciTech Connect

    Zerwekh, W.D.; Marsh, S.P.; Tan, Tai-Ho.

    1993-01-01

    The simple, cylindrically imploding and axially driven fast shock tube (FST) has been a basic component in the high velocity penetrator (HVP) program. It is a powerful device capable of delivering a directed and very high pressure output that has been successfully employed to drive hypervelocity projectiles. The FST is configured from a hollow, high-explosive (HE) cylinder, a low-density Styrofoam core, and a one-point initiator at one end. A Mach stem is formed in the core as the forward-propagating, HE detonation wave intersects the reflected radial wave. This simple FST has been found to be a powerful pressure multiplier. Up to 1-Mbar output pressure can be obtained from this device. Further increase in the output pressure can be achieved by increasing the HE detonation velocity. The FST has been fine tuned to drive a thin plate to very high velocity under an impulse per unit area of about 1 Mbar[mu]s/cm[sup 2]. A 1.5-mm-thick stainless steel disk has been accelerated intact to 0.8 cm/[mu]s under a loading pressure rate of several Mbar/[mu]s. By making the plate curvature slightly convex at the loading side the authors have successfully accelerated it to almost 1.0 cm/[mu]s. The incorporation of a barrel at the end of the FST has been found to be important as confinement of the propellant gas by the barrel tends to accelerate the projectile to higher velocity. The desire to accelerate the plate above 1.0 cm/[mu]s provided the impetus to develop a more advanced fast shock tube to deliver a much higher output pressure. This report describes the investigation of a relatively simple air-lens phase-detonation system (PFST) with fifty percent higher phase-detonation velocity and a modest 2 Mbar output. Code calculations have shown that this PFST acceleration of a plate to about 1.2 cm/[mu]s can be achieved. The performance of these PFSTs has been evaluated and the details are discussed.

  18. Tritium labeling of detonation nanodiamonds.

    PubMed

    Girard, Hugues A; El-Kharbachi, Abdelouahab; Garcia-Argote, Sébastien; Petit, Tristan; Bergonzo, Philippe; Rousseau, Bernard; Arnault, Jean-Charles

    2014-03-18

    For the first time, the radioactive labeling of detonation nanodiamonds was efficiently achieved using a tritium microwave plasma. According to our measurements, the total radioactivity reaches 9120 ± 120 μCi mg(-1), with 93% of (3)H atoms tightly bonded to the surface and up to 7% embedded into the diamond core. Such (3)H doping will ensure highly stable radiolabeled nanodiamonds, on which surface functionalization is still allowed. This breakthrough opens the way to biodistribution and pharmacokinetics studies of nanodiamonds, while this approach can be scalable to easily treat bulk quantities of nanodiamonds at low cost. PMID:24492594

  19. Reduced Model for Detonation Wave

    NASA Astrophysics Data System (ADS)

    Maillet, Jean-Bernard; Soulard, Laurent; Stoltz, Gabriel

    2007-06-01

    We present a mesoscopic model for reactive waves which extends the model proposed by G. Stoltz (G. Stoltz, Europhys. Lett. 76 (2006) 849). A complex molecule (or a group of molecules) is replaced by a single mesoparticle, evolving according to some Dissipative Particle Dynamics. Chemical reactions can be handled in a mean way by considering an additional variable per particle describing a rate of reaction. The evolution of this rate is governed by the kinetics of a reversible exothermic reaction. Numerical results show that the reactive wave behaves like a detonation wave.

  20. 14 CFR 33.47 - Detonation test.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Detonation test. 33.47 Section 33.47 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Reciprocating Aircraft Engines § 33.47 Detonation test. Each...

  1. Evaluation of the oblique detonation wave ramjet

    NASA Technical Reports Server (NTRS)

    Morrison, R. B.

    1978-01-01

    The potential performance of oblique detonation wave ramjets is analyzed in terms of multishock diffusion, oblique detonation waves, and heat release. Results are presented in terms of thrust coefficients and specific impulses for a range of flight Mach numbers of 6 to 16.

  2. 14 CFR 33.47 - Detonation test.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Detonation test. 33.47 Section 33.47 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Reciprocating Aircraft Engines § 33.47 Detonation test. Each...

  3. 14 CFR 33.47 - Detonation test.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Detonation test. 33.47 Section 33.47 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Reciprocating Aircraft Engines § 33.47 Detonation test. Each...

  4. 14 CFR 33.47 - Detonation test.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Detonation test. 33.47 Section 33.47 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Reciprocating Aircraft Engines § 33.47 Detonation test. Each...

  5. Shock Initiation and Detonability of Liquid Nitroethane

    NASA Astrophysics Data System (ADS)

    Yoshinaka, Akio; Zhang, Fan

    2004-07-01

    Shock initiation of detonation in liquid nitroethane (NE) was investigated using a cylindrical donor-attenuator-receptor configuration. While NE is less sensitive than nitromethane (NM) and isopropyl nitrate (IPN), similar behaviour of direct initiation of detonation and SDT were observed at higher shock pressures. The critical shock initiation peak pressure was found to be above 9.2 GPa at 17±6 °C initial temperature. Critical diameter experiments were performed using steel and tungsten tubes of various wall thicknesses. The critical diameter was found to lie between 89 mm with 6.6 mm wall (no detonation) and 97 mm with 8.9 mm wall (detonation) in steel tubes, and between 51 mm and 63 mm in 6.35 mm wall tungsten tubes at 21±5 °C. The steady detonation velocity measured was 5.39 mm/μsec.

  6. Performance characterization of the NASA standard detonator

    SciTech Connect

    Tarbell, W.W.; Burke, T.L.; Solomon, S.E.

    1995-05-01

    The NASA Standard Detonator (NSD) is employed in support of a number of current applications, including the Space Shuttle. This effort was directed towards providing test results to characterize the output of this device for its use in a safe and arm device. As part of the investigation, flash X-ray was used to provide stop-motion photographs of the flying metal plate that is created by initiation of the detonator. This provided researchers with a better understanding of the shape and character of the high-velocity disk as it propagated across the gap between the detonator and next assembly. The second portion of the study used a velocity interferometer to evaluate the acceleration and velocity histories of the flying plate, providing a quantified assessment of the detonator`s ability to initiate the explosive in the next explosive.

  7. Molecular dynamics simulations of weak detonations.

    PubMed

    Am-Shallem, Morag; Zeiri, Yehuda; Zybin, Sergey V; Kosloff, Ronnie

    2011-12-01

    Detonation of a three-dimensional reactive nonisotropic molecular crystal is modeled using molecular dynamics simulations. The detonation process is initiated by an impulse, followed by the creation of a stable fast reactive shock wave. The terminal shock velocity is independent of the initiation conditions. Further analysis shows supersonic propagation decoupled from the dynamics of the decomposed material left behind the shock front. The dependence of the shock velocity on crystal nonlinear compressibility resembles solitary behavior. These properties categorize the phenomena as a weak detonation. The dependence of the detonation wave on microscopic potential parameters was investigated. An increase in detonation velocity with the reaction exothermicity reaching a saturation value is observed. In all other respects the model crystal exhibits typical properties of a molecular crystal. PMID:22304055

  8. Internal Detonation Velocity Measurements Inside High Explosives

    SciTech Connect

    Benterou, J; Bennett, C V; Cole, G; Hare, D E; May, C; Udd, E

    2009-01-16

    In order to fully calibrate hydrocodes and dynamic chemistry burn models, initiation models and detonation models of high explosives, the ability to continuously measure the detonation velocity within an explosive is required. Progress on an embedded velocity diagnostic using a 125 micron diameter optical fiber containing a chirped fiber Bragg grating is reported. As the chirped fiber Bragg grating is consumed by the moving detonation wave, the physical length of the unconsumed Bragg grating is monitored with a fast InGaAs photodiode. Experimental details of the associated equipment and data in the form of continuous detonation velocity records within PBX-9502 are presented. This small diameter fiber sensor has the potential to measure internal detonation velocities on the order of 10 mm/{micro}sec along path lengths tens of millimeters long.

  9. Initiation of Gaseous Detonation by Conical Projectiles

    NASA Astrophysics Data System (ADS)

    Verreault, Jimmy

    Initiation and stabilization of detonation by hypersonic conical projectiles launched into combustible gas mixtures is investigated. This phenomenon must be understood for the design and optimization of specific hypersonic propulsion devices, such as the oblique detonation wave engine and the ram accelerator. The criteria for detonation initiation by a projectile is also related to fundamental aspects of detonation research, such as the requirement for direct initiation of a detonation by a blast wave. Experimental results of this problem also offer useful references for validation of numerical and theoretical modeling. Projectiles with cone half angles varying from 15° to 60° were launched into stoichiometric mixtures of hydrogen/oxygen with 70% argon dilution at initial pressures between 10 and 200 kPa. The projectiles were launched from a combustion-driven gas gun at velocities up to 2.2 km/s (corresponding to 133% of the Chapman Jouguet velocity). Pictures of the flowfields generated by the projectiles were taken via Schlieren photography. Five combustion regimes were observed about the projectile ranging from prompt and delayed oblique detonation wave formation, combustion instabilities, a wave splitting, and an inert shock wave. Two types of transition from the prompt oblique detonation wave regime to the inert shock regime were observed. The first (the delayed oblique detonation wave regime) showed an inert shock attached to the tip of the projectile followed by a sharp kink at the onset of an oblique detonation wave; this regime occurred by decreasing the cone angle at high mixture pressures. The second (the combustion instabilities regime) exhibited large density gradients due to combustion ignition and quenching phenomena; this regime occurred by decreasing the mixture pressure at large cone angles. A number of theoretical models were considered to predict critical conditions for the initiation of oblique detonations. The Lee-Vasiljev model agreed

  10. Numerical Analysis of a Rotating Detonation Engine in the Relative Reference Frame

    NASA Technical Reports Server (NTRS)

    Paxson, Daniel E.

    2014-01-01

    A two-dimensional, computational fluid dynamic (CFD) simulation of a semi-idealized rotating detonation engine (RDE) is described. The simulation operates in the detonation frame of reference and utilizes a relatively coarse grid such that only the essential primary flow field structure is captured. This construction yields rapidly converging, steady solutions. Results from the simulation are compared to those from a more complex and refined code, and found to be in reasonable agreement. The performance impacts of several RDE design parameters are then examined. Finally, for a particular RDE configuration, it is found that direct performance comparison can be made with a straight-tube pulse detonation engine (PDE). Results show that they are essentially equivalent.

  11. Explosion and detonation characteristics of dimethyl ether.

    PubMed

    Mogi, Toshio; Horiguchi, Sadashige

    2009-05-15

    In this study, the explosion and detonation characteristics of dimethyl ether (DME) were experimentally investigated. A spherical pressure vessel with an internal volume of 180L was used as the explosion vessel. Therefore, tubes 10m in length with internal diameters of 25mm and 50mm were used as detonation tubes. In addition, we compared the characteristics of DME with those of propane since DME is considered as a substitute fuel for liquid petroleum gas (LPG). At room temperature and atmospheric pressure, the maximum explosive pressure increased tenfold. The explosion index (K(G) values), an indicator of the intensity of an explosion, was larger than that of propane, indicating that the explosion was intense. No experimental study has been conducted on the detonation behavior of DME so far, but this research confirmed a transition to detonation. The detonation characteristics were similar to the characteristics of the Chapman-Jouguet detonation, and the concentration range for detonation was from 5.5% to 9.0%. PMID:18774641

  12. Investigation of detonation initiation in aluminium suspensions

    NASA Astrophysics Data System (ADS)

    Veyssiere, B.; Khasainov, B. A.; Briand, A.

    2008-09-01

    Detonation initiation is investigated in aluminium/oxygen and aluminium/air mixtures. Critical conditions for initiation of spherical detonations are examined in analogy with the criteria defined for gaseous mixtures, which correlate critical parameters of detonation initiation to the characteristic size of the cellular structure. However, experimental data on the detonation cell size in these two-phase mixtures are very scarce, on account of the difficulty to perform large-scale experiments. Therefore, 2D numerical simulations of the detonation cellular structure have been undertaken, with the same combustion model for Al/air and Al/O2 mixtures. The cell size is found to be λ = 37.5 cm for a rich ( r = 1.61) aluminium-air mixture, and λ = 7.5 cm for a stoichiometric aluminium-oxygen mixture, which is in reasonable agreement with available experimental data. Calculations performed in large-scale configurations (up to 25 m in length and 1.5 m in lateral direction) suggest that the critical initiation energy and predetonation radius for direct initiation of the unconfined detonation in the aluminium-air mixture are, respectively, 10 kg of TNT and 8 m. Moreover, numerical simulations reveal that the structure of the detonation wave behind the leading front is even more complicated than in pure gaseous mixtures, due to two-phase flow effects.

  13. On the Existence of Pathological Detonation Waves

    SciTech Connect

    Tarver, C M

    2003-07-11

    Pathological detonation waves with velocities greater than Chapman-Jouguet (C-J) have been proposed theoretically but never observed experimentally in gaseous, liquid or solid explosives. Two types of pathological chemical reaction zones have been identified within the Zeldovich-von Neumann-Doring (ZND) model: an exothermic chemical decomposition with a mole decrease during from the von Neumann spike state to the C-J state and an exothermic reaction followed by an endothermic reaction (eigenvalue detonation). The high temperatures reached in detonation reaction zones cause sufficient radial and atom formation to insure overall mole increases in gaseous H{sub 2} + O{sub 2} detonations. Aluminized explosives exhibit a slight mole decrease when the solid aluminum particles are oxidized, but this does not negate the large mole increase that occurs during explosive decomposition. Porous solid explosives whose products form with more cold compression energy than that of the solid are an unlikely possibility for pathological detonation. Eigenvalue detonations have been postulated for H{sub 2} + Cl{sub 2} gas phase detonations and for plastic bonded solid explosives if endothermic binder decomposition follows exothermic explosive decomposition. Chemical kinetic and physical arguments are presented to eliminate these possible pathological detonations. In the case of H{sub 2} + Cl{sub 2}, highly vibrationally excited HCl molecules dissociate Cl{sub 2} molecules during the exothermic portion of the reaction zone rather than later in the flow process. In the plastic bonded explosives, the binders are located on the surfaces of explosive particles and thus are exposed to ''hot spots'' created by the three-dimensional Mach stem shock front. Any remaining binder material rapidly reacts in collisions with the high, vibrationally excited reaction products formed during explosive decomposition. Therefore eigenvalue detonations are extremely unlikely to occur in gaseous, liquid or

  14. Calibration of the Langley 8-Foot High Temperature Tunnel for Hypersonic Airbreathing Propulsion Testing

    NASA Technical Reports Server (NTRS)

    Huebner, Lawrence D.; Rock, Kenneth E.; Voland, Randall T.; Wieting, Allan R.

    1996-01-01

    The NASA Langley 8-Foot High Temperature Tunnel has recently been modified to produce a unique testing capability for hypersonic airbreathing propulsion systems. Prior to these modifications, the facility was used primarily for aerothermal loads and structural verification testing at true flight total enthalpy conditions for Mach numbers between 6 and 7. One of the recent modifications was an oxygen replenishment system which allows operating airbreathing propulsion systems to be tested at true flight total enthalpies. Following the modifications to the facility, calibration runs were performed at total enthalpies corresponding to flight Mach numbers of 6.3 and 6.8 to establish the flow characteristics of the facility with its new capabilities. The results of this calibration, as well as modifications to tunnel combustor hardware prior to calibration to improve tunnel flow quality, are described in this paper.

  15. An Air-Breathing Launch Vehicle Concept for Single-Stage-to-Orbit

    NASA Technical Reports Server (NTRS)

    Trefny, Charles J.

    1999-01-01

    The "Trailblazer" is a 300-lb payload, single-stage-to-orbit launch vehicle concept that uses air-breathing propulsion to reduce the required propellant fraction. The integration of air-breathing propulsion is done considering performance, structural and volumetric efficiency, complexity, and design risk. The resulting configuration is intended to be viable using near-term materials and structures. The aeropropulsion performance goal for the Trailblazer launch vehicle is an equivalent effective specific impulse (I*) of 500 sec. Preliminary analysis shows that this requires flight in the atmosphere to about Mach 10, and that the gross lift-off weight is 130,000 lb. The Trailblazer configuration and proposed propulsion system operating modes are described. Preliminary performance results are presented, and key technical issues are highlighted. An overview of the proposed program plan is given.

  16. Propulsion Airframe Integration Test Techniques for Hypersonic Airbreathing Configurations at NASA Langley Research Center

    NASA Technical Reports Server (NTRS)

    Witte, David W.; Huebner, Lawrence D.; Trexler, Carl A.; Cabell, Karen F.; Andrews, Earl H., Jr.

    2003-01-01

    The scope and significance of propulsion airframe integration (PAI) for hypersonic airbreathing vehicles is presented through a discussion of the PAI test techniques utilized at NASA Langley Research Center. Four primary types of PAI model tests utilized at NASA Langley for hypersonic airbreathing vehicles are discussed. The four types of PAI test models examined are the forebody/inlet test model, the partial-width/truncated propulsion flowpath test model, the powered exhaust simulation test model, and the full-length/width propulsion flowpath test model. The test technique for each of these four types of PAI test models is described, and the relevant PAI issues addressed by each test technique are illustrated through the presentation of recent PAI test data.

  17. Eigenvalue Detonation of Combined Effects Aluminized Explosives

    NASA Astrophysics Data System (ADS)

    Capellos, C.; Baker, E. L.; Nicolich, S.; Balas, W.; Pincay, J.; Stiel, L. I.

    2007-12-01

    Theory and performance for recently developed combined—effects aluminized explosives are presented. Our recently developed combined-effects aluminized explosives (PAX-29C, PAX-30, PAX-42) are capable of achieving excellent metal pushing, as well as high blast energies. Metal pushing capability refers to the early volume expansion work produced during the first few volume expansions associated with cylinder and wall velocities and Gurney energies. Eigenvalue detonation explains the observed detonation states achieved by these combined effects explosives. Cylinder expansion data and thermochemical calculations (JAGUAR and CHEETAH) verify the eigenvalue detonation behavior.

  18. Numerical Simulation of Aluminum Dust Detonations with Different Product Phases

    NASA Astrophysics Data System (ADS)

    Teng, H. H.; Jiang, Z. L.

    Detonation waves are waves of supersonic combustion induced by strong coupling shock and heat release. Detonation research has attracted much attention in recent years owing to its potential applications in hypersonic propulsion. Aluminum (Al) particle detonation is a type of dust detonation, and its research is important in the prevention of industrial explosions. Al dust detonations for flake and spherical particles have been studied , which is found to be very sensitive to the specific area[1].

  19. Printable sensors for explosive detonation

    SciTech Connect

    Griffith, Matthew J. Cooling, Nathan A.; Elkington, Daniel C.; Belcher, Warwick J.; Dastoor, Paul C.; Muller, Elmar

    2014-10-06

    Here, we report the development of an organic thin film transistor (OTFT) based on printable solution processed polymers and employing a quantum tunnelling composite material as a sensor to convert the pressure wave output from detonation transmission tubing (shock tube) into an inherently amplified electronic signal for explosives initiation. The organic electronic detector allows detection of the signal in a low voltage operating range, an essential feature for sites employing live ordinances that is not provided by conventional electronic devices. We show that a 30-fold change in detector response is possible using the presented detector assembly. Degradation of the OTFT response with both time and repeated voltage scans was characterised, and device lifetime is shown to be consistent with the requirements for on-site printing and usage. The integration of a low cost organic electronic detector with inexpensive shock tube transmission fuse presents attractive avenues for the development of cheap and simple assemblies for precisely timed initiation of explosive chains.

  20. Air-breathing hypersonic cruise - Prospects for Mach 4-7 waverider aircraft

    NASA Technical Reports Server (NTRS)

    Blankson, Isaiah M.

    1992-01-01

    In the Mach 4-7 range, waverider aircraft are considered as candidates for both short- and long-range cruise missions, as hypersonic missiles, and as high L/D highly maneuverable craft. The potential for near- and far-term application of airbreathing engines to the waverider vehicle missions and concepts is presented. Attention is focused on the cruise mission and attempts are made to compare and contrast it with the accelerator mission.

  1. Minimum-fuel ascent to orbit using air-breathing propulsion

    NASA Technical Reports Server (NTRS)

    Van Buren, Mark A.; Mease, Kenneth D.

    1989-01-01

    Single-stage vehicles using air-breathing propulsion hold promise for more economical delivery of payloads to orbit. The characterization of minimum-fuel trajectories over the range of possible engine and aerodynamic performance of such vehicles provides useful feedback to engine and vehicle designers and paves the way for the development of guidance logic. The minimum-fuel trajectory problem is formulated, propulsion system and aerodynamic models are presented, a numerical solution approach is described, and some preliminary results are discussed.

  2. Impact of aeroelasticity on propulsion and longitudinal flight dynamics of an air-breathing hypersonic vehicle

    NASA Technical Reports Server (NTRS)

    Raney, David L.; Mcminn, John D.; Pototzky, Anthony S.; Wooley, Christine L.

    1993-01-01

    Many air-breathing hypersonic aerospacecraft design concepts incorporate an elongated fuselage forebody acting as the aerodynamic compression surface for a hypersonic combustion module, or scram jet. This highly integrated design approach creates the potential for an unprecedented form of aero-propulsive-elastic interaction in which deflections of the vehicle fuselage give rise to propulsion transients, producing force and moment variations that may adversely impact the rigid body flight dynamics and/or further excite the fuselage bending modes. To investigate the potential for such interactions, a math model was developed which included the longitudinal flight dynamics, propulsion system, and first seven elastic modes of a hypersonic air-breathing vehicle. Perturbation time histories from a simulation incorporating this math model are presented that quantify the propulsive force and moment variations resulting from aeroelastic vehicle deflections. Root locus plots are presented to illustrate the effect of feeding the propulsive perturbations back into the aeroelastic model. A concluding section summarizes the implications of the observed effects for highly integrated hypersonic air-breathing vehicle concepts.

  3. The Attenuation of a Detonation Wave by an Aircraft Engine Axial Turbine Stage

    NASA Technical Reports Server (NTRS)

    VanZante, Dale; Envia, Edmane; Turner, Mark G.

    2007-01-01

    A Constant Volume Combustion Cycle Engine concept consisting of a Pulse Detonation Combustor (PDC) followed by a conventional axial turbine was simulated numerically to determine the attenuation and reflection of a notional PDC pulse by the turbine. The multi-stage, time-accurate, turbomachinery solver TURBO was used to perform the calculation. The solution domain consisted of one notional detonation tube coupled to 5 vane passages and 8 rotor passages representing 1/8th of the annulus. The detonation tube was implemented as an initial value problem with the thermodynamic state of the tube contents, when the detonation wave is about to exit, provided by a 1D code. Pressure time history data from the numerical simulation was compared to experimental data from a similar configuration to verify that the simulation is giving reasonable results. Analysis of the pressure data showed a spectrally averaged attenuation of about 15 dB across the turbine stage. An evaluation of turbine performance is also presented.

  4. Functional performance of the T-6A Texan (JPATS) CFIS laser detonator

    NASA Astrophysics Data System (ADS)

    Blachowski, Thomas J.; Thom, Travis

    2010-08-01

    The Indian Head Division, Naval Surface Warfare Center (IHD NSWC) CAD Engineering Division is conducting a program to evaluate the laser and energetic components which comprise the Canopy Fracturing Initiation System (CFIS). This system is currently installed on the T-6A Texan II or JPATS (Joint Primary Aircraft Training System) aircraft. The T-6A Texan II is the first aircraft used by the military to train future pilots. The CFIS is an element of the pilot emergency escape system which weakens the canopy in the path of the ejection seat. The CFIS is comprised of three differing laser configurations (Internal, External, and Seat Motion) which generate a pulse that is distributed through a fiber optic energy transmission system. This pulse, in turn, initiates one of the system's explosive components, a detonator (specifically, the CCU-158/A Laser Initiated Detonator). This detonator transfers the signal to the remaining energetic components that, in turn, function to weaken their respective canopies. All of the CFIS laser types are flashlamp-pumped, neodymium glass lasers which are located at various positions in the aircraft cockpit area. This paper builds on the previous SPIE papers (2008 - Conference 7070 and 2009 - Conference 7434, respectively) and presents the initial functional test results for the CFIS Laser Detonator. These functional test results provide the technical support to justify the useful lifetime of this energetic component while being installed in the T-6A Texan II aircraft under operational conditions.

  5. Net ion fluxes in the facultative air-breather Hoplosternum littorale (tamoata) and the obligate air-breather Arapaima gigas (pirarucu) exposed to different Amazonian waters.

    PubMed

    Baldisserotto, Bernardo; Copatti, Carlos E; Gomes, Levy C; Chagas, Edsandra C; Brinn, Richard P; Roubach, Rodrigo

    2008-12-01

    Fishes that live in the Amazon environment may be exposed to several kinds of water: black water (BW), acidic black water (pH 3.5) (ABW) and white water (WW), among others. The aim of the present study was to analyze net ion fluxes in the facultative air-breather Hoplosternum littorale (tamoata) and the obligate air-breather Arapaima gigas (pirarucu) exposed to different types of water. Fishes were acclimated in well water and later placed in individual chambers containing one type of water for ion flux measurements. After 4 h, the water in the chambers was replaced by a different type of water. The transfer of both species to ABW (independent of previous water exposure) increased net ion loss. Tamoatas transferred from ABW to BW or WW presented a net ion influx, but pirarucus showed only small changes on net ion efflux. These results allow us to conclude that tamoatas and pirarucus present differences in terms of ion regulation but that the general aspects of the ion flux are similar: (1) exposure to ABW led to net ion loss; (2) transfer from BW to WW or vice-versa induced only minor changes on net ion fluxes. These observations demonstrate that any osmoregulatory difficulties encountered by either species during changes between these latter two waters can be easily overcome. PMID:18958598

  6. Propagation of detonations in hydrazine vapor

    NASA Technical Reports Server (NTRS)

    Heinrich, H. J.

    1985-01-01

    In the range of greater hydrazine vapor pressure, detonation speed depends exclusively on the extent of the ammonia decomposition in the second reaction stage. As vapor pressure decreases, the ammonia disintegration speed becomes increasingly slower and the reaction reached in the reaction zone increasingly decreases until finally, in the vapor pressure range between 53 and 16 Torr, the contribution of the second stage to detonation propagation disappears, and only the first stage remains active. Since the disintegration speed of the hydrazine in this pressure range has decreased markedly as well, no level, but rather only spinning, detonations occur. Temporary separations of the impact front and the reaction zone in the process lead to fluctuations of the detonation speed.

  7. Detonation wave augmentation of gas turbines

    NASA Technical Reports Server (NTRS)

    Wortman, A.

    1984-01-01

    The results of a feasibility study that examined the effects of using detonation waves to augment the performance of gas turbines are reported. The central ideas were to reduce compressor requirements and to maintain high performance in jet engines. Gasdynamic equations were used to model the flows associated with shock waves generated by the detonation of fuel in detonator tubes. Shock wave attenuation to the level of Mach waves was found possible, thus eliminating interference with the compressor and the necessity of valves and seals. A preliminary parametric study of the performance of a compressor working at a 4:1 ratio in a conceptual design of a detonation wave augmented jet engine in subsonic flight indicated a clear superiority over conventional designs in terms of fuel efficiency and thrust.

  8. Parametric study of double cellular detonation structure

    NASA Astrophysics Data System (ADS)

    Khasainov, B.; Virot, F.; Presles, H.-N.; Desbordes, D.

    2013-05-01

    A parametric numerical study is performed of a detonation cellular structure in a model gaseous explosive mixture whose decomposition occurs in two successive exothermic reaction steps with markedly different characteristic times. Kinetic and energetic parameters of both reactions are varied in a wide range in the case of one-dimensional steady and two-dimensional (2D) quasi-steady self-supported detonations. The range of governing parameters of both exothermic steps is defined where a "marked" double cellular structure exists. It is shown that the two-level cellular structure is completely governed by the kinetic parameters and the local overdrive ratio of the detonation front propagating inside large cells. Furthermore, since it is quite cumbersome to use detailed chemical kinetics in unsteady 2D case, the proposed work should help to identify the mixtures and the domain of their equivalence ratio where double detonation structure could be observed.

  9. Study of ignition and detonation of hydrocarbon-air mixtures with detailed and reduced chemical mechanisms

    NASA Astrophysics Data System (ADS)

    Varatharajan, Balachandar

    Uncertainties in the chemical-kinetic processes that take place in detonations lead to difficulties in obtaining fundamental knowledge about detonations and in facilitating investigations of practical devices like Pulse Detonation Engines. This research is focused on reducing the chemical-kinetic uncertainties and developing simplified chemical-kinetic descriptions for use in detonation studies. The fuels investigated are acetylene, ethylene and JP-10. Conditions addressed cover initial (post-shock) temperatures between 1000 K and 2500 K, pressures between 0.5 bar and 100 bar and equivalence ratios between 0.5 and 2. An existing detailed mechanism is extended to 175 steps among 37 chemical species by evaluating the rates of several additional reactions relevant for acetylene, ethylene and JP-10 combustion. This mechanism is tested extensively with data from shock-tube studies and flame-speed measurements. Based on the detailed mechanism, short mechanisms are derived for ignition and detonation of acetylene and ethylene in air. Application of steady-state and partial-equilibrium approximations leads to further systematic reduction. A seven-step reduced mechanism is obtained for acetylene detonations, four of which are important during the induction stage and the remaining three are important for the slower carbon-monoxide oxidation and radical-recombination processes that follow the induction stage. The theory of chain-branching thermal explosions is developed using activation-energy asymptotics and is applied for acetylene ignition, leading to an expression for ignition time. For ethylene, the strong dependence of the chemistry on initial temperatures and pressures complicates analysis and leads to identification of separate reduced mechanisms for high and low temperatures. Expressions for ignition time in terms of the elementary reaction rates are also derived. For JP-10 ignition, a reaction set including 27 additional reactions is proposed, which involves overall

  10. Deflagration to detonation transition in thermonuclear supernovae

    SciTech Connect

    Khokhlov, A.M.; Oran, E.S.; Wheeler, J.C.

    1996-12-03

    The authors derive the criteria for deflagration to detonation transition (DDT) in a Type Ia supernova. The theory is based on the two major assumptions: (i) detonation is triggered via the Zeldovich gradient mechanism inside a region of mixed fuel and products, (ii) the mixed region is produced by a turbulent mixing of fuel and products either inside an active deflagration front or during the global expansion and subsequent contraction of an exploding white dwarf. The authors determine the critical size of the mixed region required to initiate a detonation in a degenerate carbon oxygen mixture. This critical length is much larger than the width of the reaction front of a Chapman-Jouguet detonation. However, at densities greater than = 5 x 10{sup 6} g/cc, it is much smaller than the size of a white dwarf. They derive the critical turbulent intensity required to create the mixed region inside an active deflagration front in which a detonation can form. They conclude that the density rho sub sigma at which a detonation can form in a carbon-oxygen white dwarf is low, approximately less than 2 to 5 x 10{sup 6} g/cc, but greater than 5 x 10{sup 6} g/cc.

  11. Prechamber initiation of detonation in gaseous mixtures

    NASA Astrophysics Data System (ADS)

    Bivol, G. Yu; Golovastov, S. V.; Golub, V. V.

    2015-11-01

    A process of deflagration-to-detonation transition in propane-butane-oxygen and acetylene-oxygen mixtures, in an open channel with a circular cross section with a diameter of 3 mm, was investigated experimentally. Detonation initiation was carried out by burning the mixture in the prechamber connected to the channel. The prechamber was considered as an extended source for the initiation of the detonation of a finite volume. To measure the velocity of a flame front, photodiodes, installed along the axis of the channel, were used. To determine the boundary conditions at the entrance to the channel, a piezoelectric pressure transducer was used. The influence of the dimensions of the prechamber, equivalence ratio and fuel on the pressure profile, and evolution of the flame front along the axis of the channel are presented. It was shown that, the dynamics of the flame front and shock waves in the channel can occur in different scenarios depending on the geometry of the prechamber and equivalence ratio. Two limit effects of the prechamber detonation initiation in the channel have been analyzed. The pre-detonation distances and the minimal energy of direct initiation of the detonation were determined.

  12. Effect of Resolution on Propagating Detonation Wave

    SciTech Connect

    Menikoff, Ralph

    2014-07-10

    Simulations of the cylinder test are used to illustrate the effect of mesh resolution on a propagating detonation wave. For this study we use the xRage code with the SURF burn model for PBX 9501. The adaptive mesh capability of xRage is used to vary the resolution of the reaction zone. We focus on two key properties: the detonation speed and the cylinder wall velocity. The latter is related to the release isentrope behind the detonation wave. As the reaction zone is refined (2 to 15 cells for cell size of 62 to 8μm), both the detonation speed and final wall velocity change by a small amount; less than 1 per cent. The detonation speed decreases with coarser resolution. Even when the reaction zone is grossly under-resolved (cell size twice the reaction-zone width of the burn model) the wall velocity is within a per cent and the detonation speed is low by only 2 per cent.

  13. Initiation and Detonation Physics on Millimeter Scales

    SciTech Connect

    Philllips, D F; Benterou, J J; May, C A

    2012-03-20

    The LLNL Detonation Science Project has a major interest in understanding the physics of detonation on a millimeter scale. This report summarizes the rate stick experiment results of two high explosives. The GO/NO-GO threshold between varying diameters of ultra-fine TATB (ufTATB) and LX-16 were recorded on an electronic streak camera and analyzed. This report summarizes the failure diameters of rate sticks for ufTATB and LX-16. Failure diameter for the ufTATB explosive, with densities at 1.80 g/cc, begin at 2.34 mm (not maintaining detonation velocity over the entire length of the rate stick). ufTATB rate sticks at the larger 3.18 mm diameter maintain a constant detonation velocity over the complete length. The PETN based and LLNL developed explosive, LX-16, with densities at 1.7 g/cc, shows detonation failure between 0.318 mm and 0.365 mm. Additional tests would be required to narrow this failure diameter further. Many of the tested rate sticks were machined using a femtosecond laser focused into a firing tank - in case of accidental detonation.

  14. The propagation mechanism of cellular detonation

    NASA Astrophysics Data System (ADS)

    Lee, J. H. S.

    The present lecture examined the available experimental and numerical results on cellular detonations. It is concluded that apart from special mixtures where the detonation is only "weakly unstable", the classical shock ignition and thermal explosion mechanism cannot describe the physical and chemical processes in the highly complex reaction zone of unstable detonations in general. Turbulence will play an important role in both the ignition and the combustion mechanism in the reaction zone of highly unstable detonations. Vorticity and turbulence generation from shock-shock, shock-vortex, shock-density interactions and the baroclinic torque mechanism are considered important in contrast to the velocity gradient shear flow mechanism of turbulence production in incompressible flows. It is recommended that the experimental determination of the hydrodynamic thickness, its correlation with chemical, thermodynamic and transport properties of the mixture and the formulation of a turbulence model to describe the steady mean flow properties of cellular detonation structure are important problems in detonation research for the immediate future.

  15. 2-D Airbreathing Lightcraft Engine Experiments in Quiescent Conditions

    NASA Astrophysics Data System (ADS)

    Salvador, Israel I.; Myrabo, Leik N.; Minucci, Marco A. S.; de Oliveira, Antonio C.; Toro, Paulo G. P.; Chanes, José B.; Rego, Israel S.

    2011-11-01

    Ground-breaking laser propulsion (LP) experiments were performed under quiescent conditions with a 25 cm wide, two-dimensional Lightcraft model using a Lumonics TEA-622 CO2 laser emitting ˜ 1 μs pulses. In preparation for subsequent hypersonic experiments, this static test campaign was conducted at ambient pressures of 0.06, 0.15, 0.30 and 1 bar with laser pulse energies of 150 to 230 J. Time-variant pressure distributions, generated over engine "absorption chamber" walls, were integrated to obtain total impulse and momentum coupling coefficients (Cm) representative of a single propulsion cycle. Schlieren visualization of laser-induced air breakdown and expanding blast waves was also accomplished. Surprisingly, the Cm results of 600-3000 Ns/MJ were 2.5x to 5x greater than previous results from smaller Lightcraft models; this suggests that higher static Cm performance can likely be achieved in larger scale LP engines. This research collaboration, forged between the USAF and Brazilian Air Force, was carried out at the Henry T. Nagamatsu Laboratory of Aerothermodynamics and Hypersonics in Brazil.

  16. One Year Term Review as a Participating Guest in the Detonator and Detonation Physics Group

    SciTech Connect

    Lefrancois, A; Roeske, F; Tran, T; Lee, R S

    2006-02-06

    The one year stay was possible after a long administrative process, because of the fact that this was the first participating guest of B division as a foreign national in HEAF (High Explosives Application Facility) with the Detonator/Detonation Physics Group.

  17. Detonability of H/sub 2/-air-diluent mixtures

    SciTech Connect

    Tieszen, S.R.; Sherman, M.P.; Benedick, W.B.; Berman, M.

    1987-06-01

    This report describes the Heated Detonation Tube (HDT). Detonation cell width and velocity results are presented for H/sub 2/-air mixtures, undiluted and diluted with CO/sub 2/ and H/sub 2/O for a range of H/sub 2/ concentration, initial temperature and pressure. The results show that the addition of either CO/sub 2/ or H/sub 2/O significantly increases the detonation cell width and hence reduces the detonability of the mixture. The results also show that the detonation cell width is reduced (detonability is increased) for increased initial temperature and/or pressure.

  18. Deflagration to detonation transition in combustible gas mixtures

    SciTech Connect

    Smirnov, N.N.; Panfilov, I.I.

    1995-04-01

    This paper presents the results of a computational investigation of the process of deflagration to detonation transition in a combustible gas mixture. The type of combustion (i.e., deflagration or detonation) supported by a two-step reaction scheme is studied as a function of the activation energies. It is shown that both a deflagration to detonation transition and a deflagration wave that lags behind a leading shock are possible. Two types of deflagration to detonation transitions are found theoretically: initiation of detonation from the flame zone and initiation of detonation along a contact discontinuity in the compressed gas near the primary shock wave.

  19. Gaseous detonation synthesis and characterization of nano-oxide

    NASA Astrophysics Data System (ADS)

    Yan, Honghao; Wu, Linsong; Li, Xiaojie; Wang, Xiaohong

    2015-07-01

    Gaseous detonation is a new method of heating the precursor of nanomaterials into gas, and integrating it with combustible gas as mixture to be detonated for the synthesis of nanomaterials. In this paper, the mixed gas of oxygen and hydrogen is used as the source for detonation, to synthesize nano TiO2, nano SiO2 and nano SnO2 through gaseous detonation method, characterization and analysis of the products, it was found that the products from gaseous detonation method were of high purity, good dispersion, smaller particle size and even distribution. It also shows that for the synthesis of nano-oxides, gaseous detonation is universal.

  20. High order hybrid numerical simulations of two dimensional detonation waves

    NASA Technical Reports Server (NTRS)

    Cai, Wei

    1993-01-01

    In order to study multi-dimensional unstable detonation waves, a high order numerical scheme suitable for calculating the detailed transverse wave structures of multidimensional detonation waves was developed. The numerical algorithm uses a multi-domain approach so different numerical techniques can be applied for different components of detonation waves. The detonation waves are assumed to undergo an irreversible, unimolecular reaction A yields B. Several cases of unstable two dimensional detonation waves are simulated and detailed transverse wave interactions are documented. The numerical results show the importance of resolving the detonation front without excessive numerical viscosity in order to obtain the correct cellular patterns.

  1. Standardized piezoelectric polymer (PVDF) gauge for detonator response measurement

    SciTech Connect

    Moore, L.M.; Graham, R.A.; Reed, R.P.; Lee, L.M.; Bauer, F.; Warren, T.W.

    1989-01-01

    Time-resolved measurements of pressure profiles from the detonation of explosive devices have proven to be essential for detailed study of device performance. At present such measurements are routinely carried out with laser velocity interferometer (VISAR) systems, electromagnetic particle velocity gauges or piezoresistant (Manganin) gauges. One of the most promising new gauges to appear in recent years is the piezoelectric polymer gauge (PVDF) which has been standardized for general use based on the materials processing techniques developed by Bauer. The special problems presented by small size piezoelectric polymer gauges and nonplanar impacts are studied over a range of impact conditions. The response of 1 mm by 1 mm active area PVDF gauges under precisely controlled compressed-gas gun impacts shows highly reproducible results to pressures of 20 GPa. For the PVDF gauge which is placed within 12 microns of the impact surface and under approximately planar loading with small detonator flyer plates, PVDF gauge signals appear to be reduced by about 15% compared to the compressed-gas gun loading response. For highly nonplanar flyer impacts the PVDF signals are reduced by about 35%. In all stress environments, high quality, time-resolved current pulses are observed. 13 refs., 10 figs., 2 tabs.

  2. Ignition of a Thermonuclear Detonation Wave in the Focus of Two Magnetically Insulated Transmission Lines

    NASA Astrophysics Data System (ADS)

    Winterberg, F.

    2003-04-01

    For the ignition of a thermonuclear detonation wave assisted by a strong magnetic field, it is proposed to use two concentrically nested magnetically insulated transmission lines, the inner one transmitting a high- voltage lower-current-, and the outer one a high-current lower-voltage- electromagnetic pulse drawn from two Marx generators. The concept has the potential of large thermonuclear gains with an input energy conceivably as small as 105 J.

  3. Jet initiation of deflagration and detonation

    NASA Astrophysics Data System (ADS)

    Krok, James Christopher

    We have constructed a facility for the study of jet-initiated deflagration and detonation in hydrogen-air-steam mixtures. The facility is built around two pressure vessels. Mixtures of hydrogen, oxygen and nitrogen are spark-ignited in the driver vessel, generating a hot mixture of combustion products. The pressure rise ruptures a diaphragm, venting the products into the receiver vessel through nozzles of 12.7-92 mm diameter. The receiver is filled with hydrogen-air and hydrogen-oxygen mixtures diluted with either nitrogen or steam. The deflagration tests studied the lean and maximum-dilution limits of hydrogen-air mixtures ignited by a hydrogen-steam jet. The lean limit of 6% hydrogen was comparable to other studies. The maximum dilution limit for steam was 60%. This is higher than the limit found in spark/glow plug ignition experiments. Shock oscillations in the receiver increased with nozzle size. Further tests studied the initiation of detonation in both hydrogen-air and stoichiometric hydrogen-oxygen-diluent mixtures. In terms of jet diameter, D, and receiver detonation cell size, lambda, we found initiation limits of 2detonation. Mixtures with steam dilution were prone to DDT near the detonation limit, as the slower flame speed allows shock reflection and pressurization to occur before the reactants are consumed. Tests with nitrogen dilution had no DDT regime. Because of DDT and shock focusing, peak pressures were highest in mixtures that were slightly less sensitive than the detonation threshold. Schlieren movies confirmed the formation of a detonation near the nozzle exit.

  4. Jet initiation of deflagration and detonation

    NASA Astrophysics Data System (ADS)

    Krok, James Christopher

    We have constructed a facility for the study of jet-initiated deflagration and detonation in hydrogen-air-steam mixtures. The facility is built around two pressure vessels. Mixtures of hydrogen, oxygen and nitrogen are spark-ignited in the driver vessel, generating a hot mixture of combustion products. The pressure rise ruptures a diaphragm, venting the products into the receiver vessel through nozzles of 12.7-92 mm diameter. The receiver is filled with hydrogen-air and hydrogen-oxygen mixtures diluted with either nitrogen or steam.The deflagration tests studied the lean and maximum-dilution limits of hydrogen-air mixtures ignited by a hydrogen-steam jet. The lean limit of 6% hydrogen was comparable to other studies. The maximum dilution limit for steam was 60%. This is higher than the limit found in spark/glow plug ignition experiments. Shock oscillations in the receiver increased with nozzle size.Further tests studied the initiation of detonation in both hydrogen-air and stoichiometric hydrogen-oxygen-diluent mixtures. In terms of jet diameter, D, and receiver detonation cell size, [?], we found initiation limits of 2 < D[?] < 7, where other experiments required a D[?] of 11 or more. We propose that the D[?] model does not adequately characterize jet initiation, as it does not reflect the conditions in the driver.The tests indicated that shock focusing plays an important role, promoting strong secondary explosions with or without prompt initiation of detonation. Mixtures with steam dilution were prone to DDT near the detonation limit, as the slower flame speed allows shock reflection and pressurization to occur before the reactants are consumed. Tests with nitrogen dilution had no DDT regime. Because of DDT and shock focusing, peak pressures were highest in mixtures that were slightly less sensitive than the detonation threshold. Schlieren movies confirmed the formation of a detonation near the nozzle exit.

  5. Detonation diffraction from an annular channel

    NASA Astrophysics Data System (ADS)

    Meredith, James; Ng, Hoi Dick; Lee, John H. S.

    2010-12-01

    In this study, gaseous detonation diffraction from an annular channel was investigated with a streak camera and the critical pressure for transmission of the detonation wave was obtained. The annular channel was used to approximate an infinite slot resulting in cylindrically expanding detonation waves. Two mixtures, stoichiometric acetylene-oxygen and stoichiometric acetylene-oxygen with 70% Ar dilution, were tested in a 4.3 and 14.3 mm channel width ( W). The undiluted and diluted mixtures were found to have values of the critical channel width over the cell size around 3 and 12 respectively. Comparing these results to values of the critical diameter ( d c ), in which a spherical detonation occurs, a value of critical d c / W c near 2 is observed for the highly diluted mixture. This value corresponds to the geometrical factor of the curvature term between a spherical and cylindrical diverging wave. Hence, the result is in support of Lee's proposed mechanism [Lee in Dynamics of Exothermicity, pp. 321, Gordon and Breach, Amsterdam, 1996] for failure due to diffraction based on curvature in stable mixtures such as those highly argon diluted with very regular detonation cellular patterns.

  6. Structure and properties of detonation soot particles

    SciTech Connect

    MalKOV, I.Y.; Titiov, V.M.

    1996-05-01

    The influence of TNT/RDX (50/50) detonation parameters and conservation conditions of detonation products during their expansion in hermetic detonation chamber on structure and phase composition of the detonation carbon has been considered. Systematic studies made it possible to establish the real structure of detonation carbon depending on experimental conditions. It has been shown that both during explosion in a chamber and thermal annealing in vacuum the nanoparticles of diamond have the tendency to transform not into graphite particles, as was assumed earlier, but into onionlike structures of fullerene series, composed of closed concentric carbon shells, the so-called carbon onions. The nanometer carbon particles have been obtained which comprise a diamond nucleus surrounded by a graphite-like mantle composed of quasi-spherical carbon shells which are the intermediate products of annealing of nanodiamond. The influence of initial sizes of the diamond particles and temperature on the annealing of diamond has been studied. {copyright} {ital 1996 American Institute of Physics.}

  7. Modeling Hemispheric Detonation Experiments in 2-Dimensions

    SciTech Connect

    Howard, W M; Fried, L E; Vitello, P A; Druce, R L; Phillips, D; Lee, R; Mudge, S; Roeske, F

    2006-06-22

    Experiments have been performed with LX-17 (92.5% TATB and 7.5% Kel-F 800 binder) to study scaling of detonation waves using a dimensional scaling in a hemispherical divergent geometry. We model these experiments using an arbitrary Lagrange-Eulerian (ALE3D) hydrodynamics code, with reactive flow models based on the thermo-chemical code, Cheetah. The thermo-chemical code Cheetah provides a pressure-dependent kinetic rate law, along with an equation of state based on exponential-6 fluid potentials for individual detonation product species, calibrated to high pressures ({approx} few Mbars) and high temperatures (20000K). The parameters for these potentials are fit to a wide variety of experimental data, including shock, compression and sound speed data. For the un-reacted high explosive equation of state we use a modified Murnaghan form. We model the detonator (including the flyer plate) and initiation system in detail. The detonator is composed of LX-16, for which we use a program burn model. Steinberg-Guinan models5 are used for the metal components of the detonator. The booster and high explosive are LX-10 and LX-17, respectively. For both the LX-10 and LX-17, we use a pressure dependent rate law, coupled with a chemical equilibrium equation of state based on Cheetah. For LX-17, the kinetic model includes carbon clustering on the nanometer size scale.

  8. Characterization Of High Explosives Detonations Via Laser-Induced Plasmas

    SciTech Connect

    Villa-Aleman, E.

    2015-10-08

    One objective of the Department of Energy’s National Security Administration is to develop technologies that can help the United States government to detect foreign nuclear weapons development activities. The realm of high explosive (HE) experiments is one of the key areas to assess the nuclear ambitions of a country. SRNL has participated in the collection of particulates from HE experiments and characterized the material with the purpose to correlate particulate matter with HE. Since these field campaigns are expensive, on-demand simulated laboratory-scale explosion experiments are needed to further our knowledge of the chemistry and particle formation in the process. Our goal is to develop an experimental test bed in the laboratory to test measurement concepts and correlate particle formation processes with the observables from the detonation fireball. The final objective is to use this knowledge to tailor our experimental setups in future field campaigns. The test bed uses pulsed laser-induced plasmas to simulate micro-explosions, with the intent to study the temporal behavior of the fireball observed in field tests. During FY15, a plan was prepared and executed which assembled two laser ablation systems, procured materials for study, and tested a Step-Scan Fourier Transform Infrared Spectrometer (SS-FTIR). Designs for a shadowgraph system for shock wave analysis, design for a micro-particulate collector from ablated pulse were accomplished. A novel spectroscopic system was conceived and a prototype system built for acquisition of spectral/temporal characterization of a high speed event such as from a high explosive detonation. Experiments and analyses will continue into FY16.

  9. The QED engine spectrum - Fusion-electric propulsion for air-breathing to interstellar flight

    NASA Technical Reports Server (NTRS)

    Bussard, Robert W.; Jameson, Lorin W.

    1993-01-01

    A new inertial-electrostatic-fusion direct electric power source can be used to drive a relativistic e-beam to heat propellant. The resulting system is shown to yield specific impulse and thrust/mass ratio 2-3 orders of magnitude larger than from other advanced propulsion concepts. This QED system can be applied to aerospace vehicles from air-breathing to near-interstellar flight. Examples are given for Earth/Mars flight missions, that show transit times of 40 d with 20 percent payload in single-stage vehicles.

  10. Hypersonic airbreathing missile concepts under study at NASA Langley Research Center

    NASA Technical Reports Server (NTRS)

    Hunt, J. L.; Johnston, P. J.; Cubbage, J. M.; Dillon, J. L.; Richie, C. B.; Marcum, D. C., Jr.; Carlson, C. H.

    1982-01-01

    The design and performance of several tactical and strategic hypersonic airbreathing missile concepts under study at the NASA Langley Research Center are discussed from an evolutionary perspective. A mid- and chin inlet missile design, constrained to the Navy's vertical box launcher, was investigated; a performance comparison is presented that is favorable to the mid-inlet approach. Parasol wing, confined flow field, and spatula-like cruise missile configurations were examined with strategic applications in mind. The preliminary results are encouraging with respect to aerodynamic and volumetric efficiency and choice of engine integration schemes.

  11. Cubic PdNP-based air-breathing cathodes integrated in glucose hybrid biofuel cells

    NASA Astrophysics Data System (ADS)

    Faggion Junior, D.; Haddad, R.; Giroud, F.; Holzinger, M.; Maduro de Campos, C. E.; Acuña, J. J. S.; Domingos, J. B.; Cosnier, S.

    2016-05-01

    Cubic Pd nanoparticles (PdNPs) were synthesized using ascorbic acid as a reducing agent and were evaluated for the catalytic oxygen reduction reaction. PdNPs were confined with multiwalled carbon nanotube (MWCNT) dispersions to form black suspensions and these inks were dropcast onto glassy carbon electrodes. Different nanoparticle sizes were synthesized and investigated upon oxygen reduction capacities (onset potential and electrocatalytic current densities) under O2 saturated conditions at varying pH values. Strong evidence of O2 diffusion limitation was demonstrated. In order to overcome oxygen concentration and diffusion limitations in solution, we used a gas diffusion layer to create a PdNP-based air-breathing cathode, which delivered -1.5 mA cm-2 at 0.0 V with an onset potential of 0.4 V. This air-breathing cathode was combined with a specially designed phenanthrolinequinone/glucose dehydrogenase-based anode to form a complete glucose/O2 hybrid bio-fuel cell providing an open circuit voltage of 0.554 V and delivering a maximal power output of 184 +/- 21 μW cm-2 at 0.19 V and pH 7.0.Cubic Pd nanoparticles (PdNPs) were synthesized using ascorbic acid as a reducing agent and were evaluated for the catalytic oxygen reduction reaction. PdNPs were confined with multiwalled carbon nanotube (MWCNT) dispersions to form black suspensions and these inks were dropcast onto glassy carbon electrodes. Different nanoparticle sizes were synthesized and investigated upon oxygen reduction capacities (onset potential and electrocatalytic current densities) under O2 saturated conditions at varying pH values. Strong evidence of O2 diffusion limitation was demonstrated. In order to overcome oxygen concentration and diffusion limitations in solution, we used a gas diffusion layer to create a PdNP-based air-breathing cathode, which delivered -1.5 mA cm-2 at 0.0 V with an onset potential of 0.4 V. This air-breathing cathode was combined with a specially designed phenanthrolinequinone

  12. The Design of Future Airbreathing Engine Systems within an Intelligent Synthesis Environment

    NASA Technical Reports Server (NTRS)

    Malone, J. B.; Housner, J. M.; Lytle, J. K.

    1999-01-01

    This paper describes a new Initiative proposed by the National Aeronautics and Space Administration (NASA). The purpose of this initiative is to develop a future design environment for engineering and science mission synthesis for use by NASA scientists and engineers. This new initiative is called the Intelligent Synthesis Environment (ISE). The paper describes the mission of NASA, future aerospace system characteristics, the current engineering design process, the ISE concept, and concludes with a description of possible ISE applications for the decision of air-breathing propulsion systems.

  13. Comparisons of rocket and air-breathing vehicle concepts for earth-to-orbit transportarion

    NASA Astrophysics Data System (ADS)

    Dorrington, G. E.

    1990-07-01

    To illustrate that there is ample room for improvement in earth-to-orbit reliability, transportation cost and environmental cleanliness, some future European launch vehicle concepts are presented. Varying assumptions of technology level and operational strategy offer a wide range of system/subsystem options for consideration. Specific examples cited include: the advanced reusable single-stage VTOL all-rocket vehicles, the all-liquid hydrogen-oxygen variants of Ariane 5, and the advanced one-and-a-half-stage horizontal take-off air-breathing vehicles.

  14. Study of Forebody Injection and Mixing with Application to Hypervelocity Airbreathing Propulsion

    NASA Technical Reports Server (NTRS)

    Axdahl, Erik; Kumar, Ajay; Wilhite, Alan

    2012-01-01

    The use of premixed, shock-induced combustion in the context of a hypervelocity, airbreathing vehicle requires effective injection and mixing of hydrogen fuel and air on the vehicle forebody. Three dimensional computational simulations of fuel injection and mixing from flush-wall and modified ramp and strut injectors are reported in this study. A well-established code, VULCAN, is used to conduct nonreacting, viscous, turbulent simulations on a flat plate at conditions relevant to a Mach 12 flight vehicle forebody. In comparing results of various fuel injection strategies, it is found that strut injection provides the greatest balance of performance between mixing efficiency and stream thrust potential.

  15. Stability and control characteristics of an air-breathing missile configuration having a forward located inlet

    NASA Technical Reports Server (NTRS)

    Sawyer, W. C.; Hayes, C.

    1976-01-01

    An investigation was made to determine the aerodynamic characteristics of an airbreathing missile configuration having a forward located inlet. Control was provided by cruciform tail surfaces. Aerodynamic data were obtained for the body-tail configuration alone and with planar or cruciform wings. At Mach numbers from 2.86 to 4.63, the model had internal flow. At Mach numbers from 1.70 to 2.86, the internal duct was closed, and an inlet fairing was installed to simulate boost conditions.

  16. Aerodynamic characteristics of a series of single-inlet air-breathing missile configurations

    NASA Technical Reports Server (NTRS)

    Hayes, C.

    1983-01-01

    A series of air-breathing missile configurations was investigated to provide a data base for the design of such missiles. The model could be configured with either a single axisymmetric or a two dimensional inlet located at the bottom of the body. Two tail configurations were investigated: a tri-tail and an X-tail. The tail surfaces could be deflected to provide pitch control. A wing could be located above the inlet on the center line of the model. Tests were made at supersonic Mach numbers with the inlet open and internal flow, and at subsonic-transonic Mach numbers with the internal duct closed and no internal flow.

  17. Laser-shocked energetic materials with metal additives: evaluation of detonation performance

    NASA Astrophysics Data System (ADS)

    Gottfried, Jennifer; Bukowski, Eric

    A focused, nanosecond-pulsed laser with sufficient energy to exceed the breakdown threshold of a material generates a laser-induced plasma with high peak temperatures, pressures, and shock velocities. Depending on the laser parameters and material properties, nanograms to micrograms of material is ablated, atomized, ionized and excited in the laser-induced plasma. The subsequent shock wave expansion into the air above the sample has been monitored using high-speed schlieren imaging in a recently developed technique, laser-induced air shock from energetic materials (LASEM). The estimated detonation velocities using LASEM agree well with published experimental values. A comparison of the measured shock velocities for various energetic materials including RDX, DNTF, and LLM-172 doped with Al or B to the detonation velocities predicted by CHEETAH for inert or active metal participation demonstrates that LASEM has potential for predicting the early time participation of metal additives in detonation events. The LASEM results show that reducing the amount of hydrogen present in B formulations increases the resulting detonation velocities

  18. Chemical Kinetics in the expansion flow field of a rotating detonation-wave engine

    NASA Astrophysics Data System (ADS)

    Kailasanath, Kazhikathra; Schwer, Douglas

    2014-11-01

    Rotating detonation-wave engines (RDE) are a form of continuous detonation-wave engines. They potentially provide further gains in performance than an intermittent or pulsed detonation-wave engine (PDE). The overall flow field in an idealized RDE, primarily consisting of two concentric cylinders, has been discussed in previous meetings. Because of the high pressures involved and the lack of adequate reaction mechanisms for this regime, previous simulations have typically used simplified chemistry models. However, understanding the exhaust species concentrations in propulsion devices is important for both performance considerations as well as estimating pollutant emissions. A key step towards addressing this need will be discussed in this talk. In this approach, an induction parameter model is used for simulating the detonation but a more detailed finite-chemistry model is used in the expansion flow region, where the pressures are lower and the uncertainties in the chemistry model are greatly reduced. Results show that overall radical concentrations in the exhaust flow are substantially lower than from earlier predictions with simplified models. The performance of a baseline hydrogen/air RDE increased from 4940 s to 5000 s with the expansion flow chemistry, due to recombination of radicals and more production of H2O, resulting in additional heat release.

  19. Some observations on the initiation and onset of detonation.

    PubMed

    Thomas, Geraint

    2012-02-13

    The results of experimental studies during which transition to detonation events occurred are presented. These observations and their interpretation are then discussed, and the conditions for the onset of detonation are described, with particular attention paid to the nature of the phenomena of deflagration-to-detonation transition. The resulting implications for predicting detonation evolution using computational fluid dynamic methods in practical applications are also discussed. PMID:22213666

  20. Precursor detonation wave development in ANFO due to aluminum confinement

    SciTech Connect

    Jackson, Scott I; Klyanda, Charles B; Short, Mark

    2010-01-01

    Detonations in explosive mixtures of ammonium-nitrate-fuel-oil (ANFO) confined by aluminum allow for transport of detonation energy ahead of the detonation front due to the aluminum sound speed exceeding the detonation velocity. The net effect of this energy transport on the detonation is unclear. It could enhance the detonation by precompressing the explosive near the wall. Alternatively, it could decrease the explosive performance by crushing porosity required for initiation by shock compression or destroying confinement ahead of the detonation. At present, these phenomena are not well understood. But with slowly detonating, non-ideal high explosive (NIHE) systems becoming increasing prevalent, proper understanding and prediction of the performance of these metal-confined NIHE systems is desirable. Experiments are discussed that measured the effect of this ANFO detonation energy transported upstream of the front by a 76-mm-inner-diameter aluminum confining tube. Detonation velocity, detonation-front shape, and aluminum response are recorded as a function of confiner wall thickness and length. Detonation shape profiles display little curvature near the confining surface, which is attributed to energy transported upstream modifying the flow. Average detonation velocities were seen to increase with increasing confiner thickness, while wavefront curvature decreased due to the stiffer, subsonic confinement. Significant radial sidewall tube motion was observed immediately ahead of the detonation. Axial motion was also detected, which interfered with the front shape measurements in some cases. It was concluded that the confiner was able to transport energy ahead of the detonation and that this transport has a definite effect on the detonation by modifying its characteristic shape.

  1. Gas detonation gun for thermal spraying

    SciTech Connect

    Kadyrov, E.; Kadyrov, V.

    1995-08-01

    High-velocity oxy-fuel and gas detonation are competing spray coating processes well known for providing premium quality coatings with low porosity and high adhesion. They are favored for applications in environments of extreme wear, heat, and aggressive corrosion. Nevertheless, they both have limitations. For the HVOF process, these include excessive gas consumption, high rate of heat transfer to the sprayed substrate, and the short life of the (supersonic) Laval nozzle. On the other hand, the traditional gas detonation gun also has drawbacks, and the purpose of this article is to outline some factors that led to the design of an improved gas detonation coating process called Demeton, produced by Demeton USA Inc., Garden City Park, N.Y.

  2. Detonation wave profiles in HMX based explosives

    SciTech Connect

    Gustavsen, R.L.; Sheffield, S.A.; Alcon, R.R.

    1997-11-01

    Detonation wave profiles have been measured in several HMX based plastic bonded explosives including PBX9404, PBX9501, and EDC-37, as well as two HMX powders (coarse and fine) pressed to 65% of crystal density. The powders had 120 and 10 {micro}m average grain sizes, respectively. Planar detonations were produced by impacting the explosive with projectiles launched in a 72-mm bore gas gun. Impactors, impact velocity, and explosive thickness were chosen so that the run distance to detonation was always less than half the explosive thickness. For the high density plastic bonded explosives, particle velocity wave profiles were measured at an explosive/window interface using two VISAR interferometers. PMMA windows with vapor deposited aluminum mirrors were used for all experiments. Wave profiles for the powdered explosives were measured using magnetic particle velocity gauges. Estimates of the reaction zone parameters were obtained from the profiles using Hugoniots of the explosive and window.

  3. Multistage reaction pathways in detonating high explosives

    SciTech Connect

    Li, Ying; Kalia, Rajiv K.; Nakano, Aiichiro; Nomura, Ken-ichi; Vashishta, Priya

    2014-11-17

    Atomistic mechanisms underlying the reaction time and intermediate reaction products of detonating high explosives far from equilibrium have been elusive. This is because detonation is one of the hardest multiscale physics problems, in which diverse length and time scales play important roles. Here, large spatiotemporal-scale reactive molecular dynamics simulations validated by quantum molecular dynamics simulations reveal a two-stage reaction mechanism during the detonation of cyclotrimethylenetrinitramine crystal. Rapid production of N{sub 2} and H{sub 2}O within ∼10 ps is followed by delayed production of CO molecules beyond ns. We found that further decomposition towards the final products is inhibited by the formation of large metastable carbon- and oxygen-rich clusters with fractal geometry. In addition, we found distinct unimolecular and intermolecular reaction pathways, respectively, for the rapid N{sub 2} and H{sub 2}O productions.

  4. Detonation Properties of Ammonium Dinitramide (ADN)

    NASA Astrophysics Data System (ADS)

    Wätterstam, A.; Östmark, H.; Helte, A.; Karlsson, S.

    1999-06-01

    Ammonium Dinitramide, ADN, has a potential as an oxidizer for underwater high explosives. Pure ADN has a large reaction-zone length and shows a strong non-ideal behaviour. The work presented here is an extension of previous work.(Sensitivity and Performance Characterization of Ammonium Dinitramide (ADN). Presented at 11th International Detonation Symposium, Snowmass, CO, 1998.) Experiments for determining the detonation velocity as a function of inverse charge radius and density, reaction-zone length and curvature, and the detonation pressure are presented. Measurements of pressure indicates that no, or weak von-Neumann spike exists, suggesting an immediate chemical decomposition. Experimental data are compared with predicted using thermochemical codes and ZND-theory.

  5. Detonation Performance Testing of LX-19

    NASA Astrophysics Data System (ADS)

    Vincent, Samuel; Aslam, Tariq; Jackson, Scott

    2015-06-01

    CL-20 was developed at the Naval Surface Weapons Center at China Lake, CA in the mid 80's. Being less sensitive than PETN, but considerably more powerful than HMX, it is the highest energy and density compound known among organic chemicals. LX-19 was developed at LLNL in the early 90's. It is a high-energy plastic bonded explosive, composed of 95.8 wt% CL-20 and 4.2 wt% Estane binder, and is similar to LX-14 (composed of HMX and Estane), but with greater sensitivity characteristics with use of the more energetic CL-20 explosive. We report detonation performance results for unconfined cylindrical rate sticks of LX-19. The experimental diameter effects are shown, along with detonation front shapes, and reaction zone profiles for different test diameters. This data is critical for calibration to Detonation Shock Dynamics (DSD). LA-UR-15-20672.

  6. Eigenvalue Detonation of Combined Effects Aluminized Explosives

    NASA Astrophysics Data System (ADS)

    Capellos, Christos; Baker, Ernest; Balas, Wendy; Nicolich, Steven; Stiel, Leonard

    2007-06-01

    This paper reports on the development of theory and performance for recently developed combined effects aluminized explosives. Traditional high energy explosives used for metal pushing incorporate high loading percentages of HMX or RDX, whereas blast explosives incorporate some percentage of aluminum. However, the high blast explosives produce increased blast energies, with reduced metal pushing capability due to late time aluminum reaction. Metal pushing capability refers to the early volume expansion work produced during the first few volume expansions associated with cylinder wall velocities and Gurney energies. Our Recently developed combined effects aluminized explosives (PAX-29C, PAX-30, PAX-42) are capable of achieving excellent metal pushing and high blast energies. Traditional Chapman-Jouguet detonation theory does not explain the observed detonation states achieved by these combined effects explosives. This work demonstrates, with the use of cylinder expansion data and thermochemical code calculations (JAGUAR and CHEETAH), that eigenvalue detonation theory explains the observed behavior.

  7. 30 CFR 75.1311 - Transporting explosives and detonators.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 30 Mineral Resources 1 2012-07-01 2012-07-01 false Transporting explosives and detonators. 75.1311... SAFETY AND HEALTH MANDATORY SAFETY STANDARDS-UNDERGROUND COAL MINES Explosives and Blasting § 75.1311 Transporting explosives and detonators. (a) When explosives and detonators are to be transported...

  8. 30 CFR 56.6400 - Compatibility of electric detonators.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... 30 Mineral Resources 1 2014-07-01 2014-07-01 false Compatibility of electric detonators. 56.6400... Electric Blasting § 56.6400 Compatibility of electric detonators. All electric detonators to be fired in a round shall be from the same manufacturer and shall have similar electrical firing characteristics....

  9. 30 CFR 56.6400 - Compatibility of electric detonators.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 30 Mineral Resources 1 2012-07-01 2012-07-01 false Compatibility of electric detonators. 56.6400... Electric Blasting § 56.6400 Compatibility of electric detonators. All electric detonators to be fired in a round shall be from the same manufacturer and shall have similar electrical firing characteristics....

  10. 30 CFR 56.6400 - Compatibility of electric detonators.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 30 Mineral Resources 1 2011-07-01 2011-07-01 false Compatibility of electric detonators. 56.6400... Electric Blasting § 56.6400 Compatibility of electric detonators. All electric detonators to be fired in a round shall be from the same manufacturer and shall have similar electrical firing characteristics....

  11. 30 CFR 56.6400 - Compatibility of electric detonators.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 30 Mineral Resources 1 2013-07-01 2013-07-01 false Compatibility of electric detonators. 56.6400... Electric Blasting § 56.6400 Compatibility of electric detonators. All electric detonators to be fired in a round shall be from the same manufacturer and shall have similar electrical firing characteristics....

  12. 30 CFR 57.6400 - Compatibility of electric detonators.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 30 Mineral Resources 1 2010-07-01 2010-07-01 false Compatibility of electric detonators. 57.6400... Electric Blasting-Surface and Underground § 57.6400 Compatibility of electric detonators. All electric detonators to be fired in a round shall be from the same manufacturer and shall have similar...

  13. 30 CFR 57.6400 - Compatibility of electric detonators.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... 30 Mineral Resources 1 2014-07-01 2014-07-01 false Compatibility of electric detonators. 57.6400... Electric Blasting-Surface and Underground § 57.6400 Compatibility of electric detonators. All electric detonators to be fired in a round shall be from the same manufacturer and shall have similar...

  14. 30 CFR 57.6400 - Compatibility of electric detonators.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 30 Mineral Resources 1 2012-07-01 2012-07-01 false Compatibility of electric detonators. 57.6400... Electric Blasting-Surface and Underground § 57.6400 Compatibility of electric detonators. All electric detonators to be fired in a round shall be from the same manufacturer and shall have similar...

  15. 30 CFR 56.6400 - Compatibility of electric detonators.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 30 Mineral Resources 1 2010-07-01 2010-07-01 false Compatibility of electric detonators. 56.6400... Electric Blasting § 56.6400 Compatibility of electric detonators. All electric detonators to be fired in a round shall be from the same manufacturer and shall have similar electrical firing characteristics....

  16. 30 CFR 57.6400 - Compatibility of electric detonators.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 30 Mineral Resources 1 2013-07-01 2013-07-01 false Compatibility of electric detonators. 57.6400... Electric Blasting-Surface and Underground § 57.6400 Compatibility of electric detonators. All electric detonators to be fired in a round shall be from the same manufacturer and shall have similar...

  17. 30 CFR 57.6400 - Compatibility of electric detonators.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 30 Mineral Resources 1 2011-07-01 2011-07-01 false Compatibility of electric detonators. 57.6400... Electric Blasting-Surface and Underground § 57.6400 Compatibility of electric detonators. All electric detonators to be fired in a round shall be from the same manufacturer and shall have similar...

  18. 30 CFR 57.6402 - Deenergized circuits near detonators.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... Electric Blasting-Surface and Underground § 57.6402 Deenergized circuits near detonators. Electrical distribution circuits within 50 feet of electric detonators at the blast site shall be deenergized. Such circuits need not be deenergized between 25 to 50 feet of the electric detonators if stray current...

  19. 30 CFR 57.6402 - Deenergized circuits near detonators.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... Electric Blasting-Surface and Underground § 57.6402 Deenergized circuits near detonators. Electrical distribution circuits within 50 feet of electric detonators at the blast site shall be deenergized. Such circuits need not be deenergized between 25 to 50 feet of the electric detonators if stray current...

  20. 30 CFR 57.6402 - Deenergized circuits near detonators.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... Electric Blasting-Surface and Underground § 57.6402 Deenergized circuits near detonators. Electrical distribution circuits within 50 feet of electric detonators at the blast site shall be deenergized. Such circuits need not be deenergized between 25 to 50 feet of the electric detonators if stray current...

  1. 30 CFR 56.6402 - Deenergized circuits near detonators.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... Electric Blasting § 56.6402 Deenergized circuits near detonators. Electrical distribution circuits within 50 feet of electric detonators at the blast site shall be deenergized. Such circuits need not be deenergized between 25 to 50 feet of the electric detonators if stray current tests, conducted as...

  2. 30 CFR 56.6402 - Deenergized circuits near detonators.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... Electric Blasting § 56.6402 Deenergized circuits near detonators. Electrical distribution circuits within 50 feet of electric detonators at the blast site shall be deenergized. Such circuits need not be deenergized between 25 to 50 feet of the electric detonators if stray current tests, conducted as...

  3. 30 CFR 56.6402 - Deenergized circuits near detonators.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... Electric Blasting § 56.6402 Deenergized circuits near detonators. Electrical distribution circuits within 50 feet of electric detonators at the blast site shall be deenergized. Such circuits need not be deenergized between 25 to 50 feet of the electric detonators if stray current tests, conducted as...

  4. 30 CFR 75.1311 - Transporting explosives and detonators.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 30 Mineral Resources 1 2010-07-01 2010-07-01 false Transporting explosives and detonators. 75.1311... Transporting explosives and detonators. (a) When explosives and detonators are to be transported underground... transported by any cars or vehicles— (1) The cars or vehicles shall be marked with warnings to identify...

  5. 30 CFR 75.1311 - Transporting explosives and detonators.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 30 Mineral Resources 1 2013-07-01 2013-07-01 false Transporting explosives and detonators. 75.1311... SAFETY AND HEALTH MANDATORY SAFETY STANDARDS-UNDERGROUND COAL MINES Explosives and Blasting § 75.1311 Transporting explosives and detonators. (a) When explosives and detonators are to be transported...

  6. 30 CFR 75.1311 - Transporting explosives and detonators.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 30 Mineral Resources 1 2011-07-01 2011-07-01 false Transporting explosives and detonators. 75.1311... SAFETY AND HEALTH MANDATORY SAFETY STANDARDS-UNDERGROUND COAL MINES Explosives and Blasting § 75.1311 Transporting explosives and detonators. (a) When explosives and detonators are to be transported...

  7. 30 CFR 75.1311 - Transporting explosives and detonators.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... 30 Mineral Resources 1 2014-07-01 2014-07-01 false Transporting explosives and detonators. 75.1311... SAFETY AND HEALTH MANDATORY SAFETY STANDARDS-UNDERGROUND COAL MINES Explosives and Blasting § 75.1311 Transporting explosives and detonators. (a) When explosives and detonators are to be transported...

  8. Feasibility and parameter study of a detonation diffuser

    NASA Astrophysics Data System (ADS)

    Stevens, Christopher A.

    This research includes an investigation of the mechanisms of diffraction and reinitiation that enable a detonation diffuser. It describes a set of geometric parameters necessary to design a diffuser for a given detonable mixture and initial channel height. Predetonators with channel height less than the critical height are ineffective because detonations in small channels decouple into separate shock and combustion fronts when the channel height increases. A detonation diffuser allows the channel height to increase by utilizing the decoupled shock wave to reinitiate detonation. In the diffuser, a detonation initially decouples into separate shock and combustion fronts, and then the decoupled shock front reflects from an oblique surface initiating a secondary detonation that survives the expansion. This research investigated the three regions of a detonation diffuser: the initial diffraction, the reflecting surface, and the second diffraction corner. Schlieren video of two-dimensional diffracting detonations recorded the position of the detonation, decoupled shock front and flame front. Observations of the decoupled shocks reflecting from surfaces showed that a 45° reflecting surface must be placed less than 80 mm downstream of the initial diffraction corner to initiate a secondary detonation in more than 91% of repeated trials. Observations of the interaction of diffracting detonations with multiple obstacles revealed that the best performance (smallest separation, and highest Mach number) occurred when the decoupled shock reflected from four separate obstacles at approximately the same time.

  9. Slag characterization and removal using pulse detonation for coal gasification

    SciTech Connect

    Huque, Z.; Mei, D.; Biney, P.O.; Zhou, J.

    1997-06-25

    This report is a preliminary economic analysis of the Sherburne Co. plant North State Power. This analysis is made with the cooperation of Joe Brojberg (senior analysis engineer of NSP) and Steve Bension (Slag and ash specialist, President of Microbeam Technologies Incorporated (MTI) of North Dakota) and Paul Johnson of Diamond Power Speciality.

  10. A Target Indirect Thrust Measurement Method of Pulse Detonation Engine

    NASA Astrophysics Data System (ADS)

    Huang, Xiqiao; Xiong, Yuefei; Li, Chao; Zheng, Longxi; Li, Qing

    2015-05-01

    An indirect thrust measurement method based on impulse of a target plate was developed, and a new thrust measurement system (TMS) was successfully designed and constructed. A series of multi-cycle experiments on thrust measurement were conducted to investigate the feasibility of this method with the newly-built indirect TMS. The thrust measurement of PDE was made at different plate target axial positions and operating frequencies. All the experiments were conducted using gasoline as fuel and air as oxidant. The experimental results implied that the thrust of PDE by using the indirect impulse method was a function of the target plate axial position, and there existed an optimum measurement position for PDE with a diameter of 60 mm. The optimum target plate position located at 3.33. According to the experimental results, the thrusts obtained by using indirect TMS were less than the actual values, and so the observed value of thrust was modified in order to make the thrust more reliable. A relative accurate calibration formula depending on the operating frequency was found.

  11. Modeling Initiation in Exploding Bridgewire Detonators

    SciTech Connect

    Hrousis, C A

    2005-05-18

    One- and two-dimensional models of initiation in detonators are being developed for the purpose of evaluating the performance of aged and modified detonator designs. The models focus on accurate description of the initiator, whether it be an EBW (exploding bridgewire) that directly initiates a high explosive powder or an EBF (exploding bridgefoil) that sends an inert flyer into a dense HE pellet. The explosion of the initiator is simulated using detailed MHD equations of state as opposed to specific action-based phenomenological descriptions. The HE is modeled using the best available JWL equations of state. Results to date have been promising, however, work is still in progress.

  12. Statistical Hot Spot Model for Explosive Detonation

    SciTech Connect

    Nichols, III, A L

    2005-07-14

    The Non-local Thermodynamic Equilibrium Statistical Hot Spot Model (NLTE SHS), a new model for explosive detonation, is described. In this model, the formation, ignition, propagation, and extinction of hot spots is explicitly modeled. The equation of state of the explosive mixture is treated with a non-local equilibrium thermodynamic assumption. A methodology for developing the parameters for the model is discussed, and applied to the detonation velocity diameter effect. Examination of these results indicates where future improvements to the model can be made.

  13. Statistical Hot Spot Model for Explosive Detonation

    SciTech Connect

    Nichols III, A L

    2004-05-10

    The Non-local Thermodynamic Equilibrium Statistical Hot Spot Model (NLTE SHS), a new model for explosive detonation, is described. In this model, the formation, ignition, propagation, and extinction of hot spots is explicitly modeled. The equation of state of the explosive mixture is treated with a nonlocal equilibrium thermodynamic assumption. A methodology for developing the parameters for the model is discussed, and applied to the detonation velocity diameter effect. Examination of these results indicates where future improvements to the model can be made.

  14. Detonation duct gas generator demonstration program

    NASA Technical Reports Server (NTRS)

    Wortman, A.; Othmer, P.; Rostafinski, W.

    1992-01-01

    An experimental demonstration is presented for the generation of detonation waves that move periodically across high speed channel flow; these waves can compress the outflow from a low pressure compressor, and thereby both reduce the compressor requirements associated with conventional gas turbines and enhance thermodynamic efficiency through isochoric energy addition. By generating transient transverse waves, rather than standing waves, shock-wave losses are reduced by an order of magnitude; the result is a Humphrey cycle augmenting the basic Brayton-cycle gas turbine. Attention is presently given to results from an experimental detonation duct.

  15. Pressure Feedback in Rotating Detonation Engines

    NASA Astrophysics Data System (ADS)

    Schwer, Douglas; Kailasanath, K.

    2012-11-01

    Rotating detonation engines (RDEs) represent a unique method for obtaining propulsion from the high efficiency detonation cycle. In order for the RDE to be a practical propulsive device, engines must be capable of running efficiently at low pressure ratios, however, this type of injection typically results in a large amount of pressure feedback into the injection system. This paper examines different aspects of the pressure feedback phenomena, and investigates approaches to injecting fresh mixture that reduce the amount of feedback. This work is supported by the Office of Naval Research through NRL 6.1 Computational Physics Task Area.

  16. Performance Validation Approach for the GTX Air-Breathing Launch Vehicle

    NASA Technical Reports Server (NTRS)

    Trefny, Charles J.; Roche, Joseph M.

    2002-01-01

    The primary objective of the GTX effort is to determine whether or not air-breathing propulsion can enable a launch vehicle to achieve orbit in a single stage. Structural weight, vehicle aerodynamics, and propulsion performance must be accurately known over the entire flight trajectory in order to make a credible assessment. Structural, aerodynamic, and propulsion parameters are strongly interdependent, which necessitates a system approach to design, evaluation, and optimization of a single-stage-to-orbit concept. The GTX reference vehicle serves this purpose, by allowing design, development, and validation of components and subsystems in a system context. The reference vehicle configuration (including propulsion) was carefully chosen so as to provide high potential for structural and volumetric efficiency, and to allow the high specific impulse of air-breathing propulsion cycles to be exploited. Minor evolution of the configuration has occurred as analytical and experimental results have become available. With this development process comes increasing validation of the weight and performance levels used in system performance determination. This paper presents an overview of the GTX reference vehicle and the approach to its performance validation. Subscale test rigs and numerical studies used to develop and validate component performance levels and unit structural weights are outlined. The sensitivity of the equivalent, effective specific impulse to key propulsion component efficiencies is presented. The role of flight demonstration in development and validation is discussed.

  17. Affordable Flight Demonstration of the GTX Air-Breathing SSTO Vehicle Concept

    NASA Technical Reports Server (NTRS)

    Krivanek, Thomas M.; Roche, Joseph M.; Riehl, John P.; Kosareo, Daniel N.

    2002-01-01

    The rocket based combined cycle (RBCC) powered single-stage-to-orbit (SSTO) reusable launch vehicle has the potential to significantly reduce the total cost per pound for orbital payload missions. To validate overall system performance, a flight demonstration must be performed. This paper presents an overview of the first phase of a flight demonstration program for the GTX SSTO vehicle concept. Phase 1 will validate the propulsion performance of the vehicle configuration over the supersonic and hypersonic airbreathing portions of the trajectory. The focus and goal of Phase 1 is to demonstrate the integration and performance of the propulsion system flowpath with the vehicle aerodynamics over the air-breathing trajectory. This demonstrator vehicle will have dual mode ramjet/scramjets, which include the inlet, combustor, and nozzle with geometrically scaled aerodynamic surface outer mold lines (OML) defining the forebody, boundary layer diverter, wings, and tail. The primary objective of this study is to demonstrate propulsion system performance and operability including the ram to scram transition, as well as to validate vehicle aerodynamics and propulsion airframe integration. To minimize overall risk and development cost the effort will incorporate proven materials, use existing turbomachinery in the propellant delivery systems, launch from an existing unmanned remote launch facility, and use basic vehicle recovery techniques to minimize control and landing requirements. A second phase would demonstrate propulsion performance across all critical portions of a space launch trajectory (lift off through transition to all-rocket) integrated with flight-like vehicle systems.

  18. Survey of Aerothermodynamics Facilities Useful for the Design of Hypersonic Vehicles Using Air-Breathing Propulsion

    NASA Technical Reports Server (NTRS)

    Arnold, James O.; Deiwert, G. S.

    1997-01-01

    The dream of producing an air-breathing, hydrogen fueled, hypervelocity aircraft has been before the aerospace community for decades. However, such a craft has not yet been realized, even in an experimental form. Despite the simplicity and beauty of the concept, many formidable problems must be overcome to make this dream a reality. This paper summarizes the aero/aerothermodynamic issues that must be addressed to make the dream a reality and discusses how aerothermodynamics facilities and their modem companion, real-gas computational fluid dynamics (CFD), can help solve the problems blocking the way to realizing the dream. The approach of the paper is first to outline the concept of an air-breathing hypersonic vehicle and then discuss the nose-to-tail aerothermodynamics issues and special aerodynamic problems that arise with such a craft. Then the utility of aerothermodynamic facilities and companion CFD analysis is illustrated by reviewing results from recent United States publications wherein these problems have been addressed. Papers selected for the discussion have k e n chosen such that the review will serve to survey important U.S. aero/aerothermodynamic real gas and conventional wind tunnel facilities that are useful in the study of hypersonic, hydrogen propelled hypervelocity vehicles.

  19. Power Reduction of the Air-Breathing Hall-Effect Thruster

    NASA Astrophysics Data System (ADS)

    Kim, Sungrae

    Electric propulsion system is spotlighted as the next generation space propulsion system due to its benefits; one of them is specific impulse. While there are a lot of types in electric propulsion system, Hall-Effect Thruster, one of electric propulsion system, has higher thrust-to-power ratio and requires fewer power supplies for operation in comparison to other electric propulsion systems, which means it is optimal for long space voyage. The usual propellant for Hall-Effect Thruster is Xenon and it is used to be stored in the tank, which may increase the weight of the thruster. Therefore, one theory that uses the ambient air as a propellant has been proposed and it is introduced as Air-Breathing Hall-Effect Thruster. Referring to the analysis on Air-Breathing Hall-Effect Thruster, the goal of this paper is to reduce the power of the thruster so that it can be applied to real mission such as satellite orbit adjustment. To reduce the power of the thruster, two assumptions are considered. First one is changing the altitude for the operation, while another one is assuming the alpha value that is electron density to ambient air density. With assumptions above, the analysis was done and the results are represented. The power could be decreased to 10s˜1000s with the assumptions. However, some parameters that do not satisfy the expectation, which would be the question for future work, and it will be introduced at the end of the thesis.

  20. Ammonia as a respiratory gas in water and air-breathing fishes.

    PubMed

    Randall, David J; Ip, Yuen K

    2006-11-01

    Ammonia is produced in the liver and excreted as NH(3) by diffusion across the gills. Elevated ammonia results in an increase in gill ventilation, perhaps via stimulation of gill oxygen chemo-receptors. Acidification of the water around the fish by carbon dioxide and acid excretion enhances ammonia excretion and constitutes "environmental ammonia detoxification". Fish have difficulties in excreting ammonia in alkaline water or high concentrations of environmental ammonia, or when out of water. The mudskipper, Periphthalmodon schlosseri, is capable of active NH(4)(+) transport, maintaining low internal levels of ammonia. To prevent a back flux of NH(3), these air-breathing fish can increase gill acid excretion and reduce the membrane NH(3) permeability by modifying the phospholipid and cholesterol compositions of their skin. Several air-breathing fish species can excrete ammonia into air through NH(3) volatilization. Some fish detoxify ammonia to glutamine or urea. The brains of some fish can tolerate much higher levels of ammonia than other animals. Studies of these fish may offer insights into the nature of ammonia toxicity in general. PMID:16731054

  1. Cubic PdNP-based air-breathing cathodes integrated in glucose hybrid biofuel cells.

    PubMed

    Faggion Junior, D; Haddad, R; Giroud, F; Holzinger, M; Maduro de Campos, C E; Acuña, J J S; Domingos, J B; Cosnier, S

    2016-05-21

    Cubic Pd nanoparticles (PdNPs) were synthesized using ascorbic acid as a reducing agent and were evaluated for the catalytic oxygen reduction reaction. PdNPs were confined with multiwalled carbon nanotube (MWCNT) dispersions to form black suspensions and these inks were dropcast onto glassy carbon electrodes. Different nanoparticle sizes were synthesized and investigated upon oxygen reduction capacities (onset potential and electrocatalytic current densities) under O2 saturated conditions at varying pH values. Strong evidence of O2 diffusion limitation was demonstrated. In order to overcome oxygen concentration and diffusion limitations in solution, we used a gas diffusion layer to create a PdNP-based air-breathing cathode, which delivered -1.5 mA cm(-2) at 0.0 V with an onset potential of 0.4 V. This air-breathing cathode was combined with a specially designed phenanthrolinequinone/glucose dehydrogenase-based anode to form a complete glucose/O2 hybrid bio-fuel cell providing an open circuit voltage of 0.554 V and delivering a maximal power output of 184 ± 21 μW cm(-2) at 0.19 V and pH 7.0. PMID:27142300

  2. Detonation of hydrogen-air mixtures. [PWR; BWR

    SciTech Connect

    Lee, J.H.S.; Knystautas, R.; Benedick, W.B.

    1983-01-01

    The detonation of a hydrogen-air cloud subsequent to an accidental release of hydrogen into ambient surroundings cannot be totally ruled out in view of the relative sensitivity of the hydrogen-air system. The present paper investigates the key parameters involved in hydrogen-air detonations and attempts to establish quantitative correlations between those that have important practical implications. Thus, for example, the characteristic length scale lambda describing the cellular structure of a detonation front is measured for a broad range of hydrogen-air mixtures and is quantitatively correlated with the key dynamic detonation properties such as detonability, transmission and initiation.

  3. Ethylene-air detonation in water spray

    NASA Astrophysics Data System (ADS)

    Jarsalé, G.; Virot, F.; Chinnayya, A.

    2016-07-01

    Detonation experiments are conducted in a 52 mm square channel with an ethylene-air gaseous mixture with dispersed liquid water droplets. The tests were conducted with a fuel-air equivalence ratio ranging from 0.9 to 1.1 at atmospheric pressure. An ultrasonic atomizer generates a polydisperse liquid water spray with droplet diameters of 8.5-12 μm, yielding an effective density of 100-120 g/m3 . Pressure signals from seven transducers and cellular structure are recorded for each test. The detonation structure in the two-phase mixture exhibits a gaseous-like behaviour. The pressure profile in the expansion fan is not affected by the addition of water. A small detonation velocity deficit of up to 5 % was measured. However, the investigation highlights a dramatic increase in the cell size (λ ) associated with the increase in the liquid water mass fraction in the two-phase mixture. The detonation structure evolves from a multi-cell to a half-cell mode. The analysis of the decay of the post-shock pressure fluctuations reveals that the ratio of the hydrodynamic thickness over the cell size (x_{{HT}}/{λ } ) remains quite constant, between 5 and 7. A slight decrease of this ratio is observed as the liquid water mass fraction is increased, or the ethylene-air mixture is made leaner.

  4. Detonation propagation in a high loss configuration

    SciTech Connect

    Jackson, Scott I; Shepherd, Joseph E

    2009-01-01

    This work presents an experimental study of detonation wave propagation in tubes with inner diameters (ID) comparable to the mixture cell size. Propane-oxygen mixtures were used in two test section tubes with inner diameters of 1.27 mm and 6.35 mm. For both test sections, the initial pressure of stoichiometric mixtures was varied to determine the effect on detonation propagation. For the 6.35 mm tube, the equivalence ratio {phi} (where the mixture was {phi} C{sub 3}H{sub 8} + 50{sub 2}) was also varied. Detonations were found to propagate in mixtures with cell sizes as large as five times the diameter of the tube. However, under these conditions, significant losses were observed, resulting in wave propagation velocities as slow as 40% of the CJ velocity U{sub CJ}. A review of relevant literature is presented, followed by experimental details and data. Observed velocity deficits are predicted using models that account for boundary layer growth inside detonation waves.

  5. Detonation duct gas generator demonstration program

    NASA Technical Reports Server (NTRS)

    Wortman, Andrew; Brinlee, Gayl A.; Othmer, Peter; Whelan, Michael A.

    1991-01-01

    The feasibility of the generation of detonation waves moving periodically across high speed channel flow is experimentally demonstrated. Such waves are essential to the concept of compressing requirements and increasing the engine pressure compressor with the objective of reducing conventional compressor requirements and increasing the engine thermodynamic efficiency through isochoric energy addition. By generating transient transverse waves, rather than standing waves, shock wave losses are reduced by an order of magnitude. The ultimate objective is to use such detonation ducts downstream of a low pressure gas turbine compressor to produce a high overall pressure ratio thermodynamic cycle. A 4 foot long, 1 inch x 12 inch cross-section, detonation duct was operated in a blow-down mode using compressed air reservoirs. Liquid or vapor propane was injected through injectors or solenoid valves located in the plenum or the duct itself. Detonation waves were generated when the mixture was ignited by a row of spark plugs in the duct wall. Problems with fuel injection and mixing limited the air speeds to about Mach 0.5, frequencies to below 10 Hz, and measured pressure ratios of about 5 to 6. The feasibility of the gas dynamic compression was demonstrated and the critical problem areas were identified.

  6. Detonation and deflagration properties of pyrotechnic mixtures

    SciTech Connect

    Tanaka, Katsumi

    1996-07-01

    Theoretical calculation of detonation and deflagration properties of pyrotechnic mixtures have been performed including report charges and display charges. Calculation were performed with the KHT (Kihara-Hikta-Tanaka) code. KHT results are compared with a modified version of the TIGER code which allows calculation with 900 gaseous and 600 condensed product species at high pressure. Detonation properties computed by KHT and BKWS (Becker-Kistiakowskii-Wilson) give favorable agreement with experimental results of detonation velocity measurements. Hydrodynamic computation by one dimensional Lagrangian hydrodynamic code using the isentrope given by KHT constant volume explosion, indicated that experimental results for blast wave measurement for 30kg and 50kg of report charge were an incomplete reaction. Underwater detonation experiments with explosive charge of 25g, however, indicates a more energetic nature than the KHT prediction. This scale effect indicates complicated slow reactions and a number of condensed phase deflagration products of powder mixtures such as aluminum or titanium with oxidizers such as potassium perchlorate or nitrate salts as suggested by Hobbs et al.

  7. Hydrogen detonation and detonation transition data from the High-Temperature Combustion Facility

    SciTech Connect

    Ciccarelli, G.; Boccio, J.L.; Ginsberg, T.; Finfrock, C.; Gerlach, L.; Tagawa, H.; Malliakos, A.

    1995-12-31

    The BNL High-Temperature Combustion Facility (HTCF) is an experimental research tool capable of investigating the effects of initial thermodynamic state on the high-speed combustion characteristic of reactive gas mixtures. The overall experimental program has been designed to provide data to help characterize the influence of elevated gas-mixture temperature (and pressure) on the inherent sensitivity of hydrogen-air-steam mixtures to undergo detonation, on the potential for flames accelerating in these mixtures to transition into detonations, on the effects of gas venting on the flame-accelerating process, on the phenomena of initiation of detonations in these mixtures by jets of hot reactant product,s and on the capability of detonations within a confined space to transmit into another, larger confined space. This paper presents results obtained from the completion of two of the overall test series that was designed to characterize high-speed combustion phenomena in initially high-temperature gas mixtures. These two test series are the intrinsic detonability test series and the deflagration-to-detonation (DDT) test series. A brief description of the facility is provided below.

  8. Hydrogen detonation and detonation transition data from the High-Temperature Combustion Facility

    SciTech Connect

    Ciccarelli, G.; Boccio, J.L.; Ginsberg, T.; Finfrock, C.

    1996-03-01

    The BNL High-Temperature Combustion Facility (HTCF) is an experimental research tool capable of investigating the effects of initial thermodynamic state on the high-speed combustion characteristic of reactive gas mixtures. The overall experimental program has been designed to provide data to help characterize the influence of elevated gas-mixture temperature (and pressure) on the inherent sensitivity of hydrogen-air-steam mixtures to undergo detonation, on the potential for flames accelerating in these mixtures to transition into detonations, on the effects of gas venting on the flame-accelerating process, on the phenomena of initiation of detonations in these mixtures by jets of hot reactant products, and on the capability of detonations within a confined space to transmit into another, larger confined space. This paper presents results obtained from the completion of two of the overall test series that was designed to characterize high-speed combustion phenomena in initially high-temperature gas mixtures. These two test series are the intrinsic detonability test series and the deflagration-to-detonation (DDT) test series. A brief description of the facility is provided below.

  9. Detonation and Transition to Detonation in Horizontal Water-Filled Pipes

    NASA Astrophysics Data System (ADS)

    Bitter, Neal P.; Shepherd, Joseph E.

    2012-11-01

    Detonations and deflagration-to-detonation transition (DDT) are experimentally studied in horizontal pipes which are partially filled with water. The gas layer above the water is stoichiometric hydrogen-oxygen at 1 bar. The detonation wave produces oblique shock waves in the water, which focus at the bottom of the pipe due to the curvature of the walls. This results in peak pressures at the bottom of the pipe that are 4-6 times greater than the peak detonation pressure. Such pressure amplification is measured for water depths of 0.25, 0.5, 0.75, 0.87, and 0.92 pipe diameters. Focusing of the oblique shock waves is studied further by measuring the circumferential variation of pressure when the water depth is 0.5 pipe diameters, and reasonable agreement with theoretical modeling is found. Failure of the detonation waves was not observed, even for water depths as high as 0.92 pipe diameters. Transition to detonation also occurred at every water height, and transition distance did not vary significantly with water height.

  10. Combined effects of vortex flow and the Shchelkin spiral dimensions on characteristics of deflagration-to-detonation transition

    NASA Astrophysics Data System (ADS)

    Asato, K.; Miyasaka, T.; Watanabe, Y.; Tanabashi, K.

    2013-07-01

    The objective of this investigation was to achieve the shortest possible distance for the deflagration-to-detonation transition (DDT) for a pulse detonation engine without losing the engine's simplicity. The effects of rapid flame propagation, rotating velocity, and Shchelkin spiral dimensions in the vortex flow (VF) on DDT characteristics were examined. A VF field was established in the ignition and transition domains of a detonation tube using a VF-type injector. The flame propagation in the ignition domain was observed by a high-speed video camera and the detonation propagation process was observed by measuring the pressure and ionization current in the detonation tube. The DDT distance in the VF could be shortened by 50-57 % of that in the counterflow by optimizing the VF and the Shchelkin spiral dimensions. The shortening effect became remarkable as the rotating velocity increased. The governing factors for this effect are considered to be the formation of an area of higher energy density in the ignition domain of the tube, as well as flame acceleration due to rapid flame propagation in the VF and turbulence promotion near the tube wall by the rotating velocity and the Shchelkin spiral in the transition domain of the tube. However, the combined effects of VF and the Shchelkin spiral on the shortening of the DDT distance were deteriorated in very strong turbulence. It is necessary to optimize the rotating velocity and dimensions of the Shchelkin spiral to achieve the shortest possible distance of DDT in the VF.

  11. Synchro-ballistic recording of detonation phenomena

    SciTech Connect

    Critchfield, R.R.; Asay, B.W.; Bdzil, J.B.; Davis, W.C.; Ferm, E.N.; Idar, D.J.

    1997-09-01

    Synchro-ballistic use of rotating-mirror streak cameras allows for detailed recording of high-speed events of known velocity and direction. After an introduction to the synchro-ballistic technique, this paper details two diverse applications of the technique as applied in the field of high-explosives research. In the first series of experiments detonation-front shape is recorded as the arriving detonation shock wave tilts an obliquely mounted mirror, causing reflected light to be deflected from the imaging lens. These tests were conducted for the purpose of calibrating and confirming the asymptotic Detonation Shock Dynamics (DSD) theory of Bdzil and Stewart. The phase velocities of the events range from ten to thirty millimeters per microsecond. Optical magnification is set for optimal use of the film`s spatial dimension and the phase velocity is adjusted to provide synchronization at the camera`s maximum writing speed. Initial calibration of the technique is undertaken using a cylindrical HE geometry over a range of charge diameters and of sufficient length-to-diameter ratio to insure a stable detonation wave. The final experiment utilizes an arc-shaped explosive charge, resulting in an asymmetric detonation-front record. The second series of experiments consists of photographing a shaped-charge jet having a velocity range of two to nine millimeters per microsecond. To accommodate the range of velocities it is necessary to fire several tests, each synchronized to a different section of the jet. The experimental apparatus consists of a vacuum chamber to preclude atmospheric ablation of the jet tip with shocked-argon back lighting to produce a shadow-graph image.

  12. Mitigation of explosion bubble pulsation caused by the deep underwater detonation of a tapered charge. Final report 21 Apr-31 Oct 80

    SciTech Connect

    Thomsen, J.M.; Ruhl, S.F.

    1980-10-31

    The dynamics of the explosion bubble growth and collapse due to the underwater detonation of a 5.3-m-long tapered charge containing 454 kg of TNT are examined calculationally. The burst was 305 m below the sea surface in deep water. The calculational method, using a coupled Eulerian-Lagrangian finite-difference hydrodynamics computer code, was made credible by successfully calculating the explosion bubble dynamics of spherical TNT charges detonated at the same depth of burst. Results show the tapered charge explosion bubble exhibits characteristics that are very similar to an equal-weight spherical charge over the first bubble period. This implies that bubble pulses will be produced by tapered charge detonations at this depth and deeper. The bubble pulses emitted by deep tapered charge detonations are not characteristic of the underwater nuclear detonation that these charges are designed to simulate. In fact, for conventional charges, bubble reloading of models is severe, and can lead to their unplanned collapse. Three methods of tapered charge bubble pulse mitigation were examined: (1) injection of propellant gas into the bubble to prevent/cushion collapse, (2) dissipation of bubble energy through enhancement of turbulence at the first minimum by creating an asymmetric bubble, and (3) using an explosive that produces condensible/soluble detonation products, thereby allowing more rapid transfer of bubble energy to the water and preventing the creation of a substantial bubble pulse. Results show that 2 and 3 are promising methods of bubble pulse mitigation. Additional work must be done to quantify the effectiveness of these methods and produce specific test designs.

  13. Effects of a Vortex Flow on Characteristics of Deflagration-to-Detonation Transition

    NASA Astrophysics Data System (ADS)

    Asato, Katsuo; Miyasaka, Takeshi; Watanabe, Yuta; Ishikawa, Soushin; Tanabashi, Kouki

    The effects of a vortex flow (VF) on the characteristics of deflagration-to-detonation transition (DDT) were examined in order to achieve the shortest distance of DDT for a pulse detonation engine (PDE). The DDT distances in a vortex flow were shortened by 15-47 % than those in a counterflow. The shortening effect becomes remarkable as the rotating velocity increases. Formation of the area of higher energy density in the ignition domain of the tube, and flame acceleration due to rapid flame propagation in the vortex flow and promotion of turbulence near the tube wall by the rotating velocity in the transition domain of the tube are considered to be the governing factors in shortening the DDT distance.

  14. Experimental Investigation of Detonation Re-initiation Mechanisms Following a Mach Reflection of a Quenched Detonation

    NASA Astrophysics Data System (ADS)

    Bhattacharjee, Rohit Ranjan

    Detonation waves are supersonic combustion waves that have a multi-shock front structure followed by a spatially non-uniform reaction zone. During propagation, a de-coupled shock-flame complex is periodically re-initiated into an overdriven detonation following a transient Mach reflection process. Past researchers have identified mechanisms that can increase combustion rates and cause localized hot spot re-ignition behind the Mach shock. But due to the small length scales and stochastic behaviour of detonation waves, the important mechanisms that can lead to re-initiation into a detonation requires further clarification. If a detonation is allowed to diffract behind an obstacle, it can quench to form a de-coupled shock-flame complex and if allowed to form a Mach reflection, re-initiation of a detonation can occur. The use of this approach permits the study of re-initiation mechanisms reproducibly with relatively large length scales. The objective of this study is to experimentally elucidate the key mechanisms that can increase chemical reaction rates and sequentially lead to re-initiation of a de-coupled shock-flame complex into an overdriven detonation wave following a Mach reflection. All experiments were carried out in a thin rectangular channel using a stoichiometric mixture of oxy-methane. Three different types of obstacles were used - a half-cylinder, a roughness plate along with the half-cylinder and a full-cylinder. Schlieren visualization was achieved by using a Z-configuration setup, a high speed camera and a high intensity light source. Results indicate that forward jetting of the slip line behind the Mach stem can potentially increase combustion rates by entraining hot burned gas into unburned gas. Following ignition and jet entrainment, a detonation wave first appears along the Mach stem. The transverse wave can form a detonation wave due to rapid combustion of unburned gas which may be attributed to shock interaction with the unburned gas

  15. Ultrashort-pulse laser generated nanoparticles of energetic materials

    DOEpatents

    Welle, Eric J.; Tappan, Alexander S.; Palmer, Jeremy A.

    2010-08-03

    A process for generating nanoscale particles of energetic materials, such as explosive materials, using ultrashort-pulse laser irradiation. The use of ultrashort laser pulses in embodiments of this invention enables one to generate particles by laser ablation that retain the chemical identity of the starting material while avoiding ignition, deflagration, and detonation of the explosive material.

  16. Proton radiography of PBX 9502 detonation shock dynamics confinement sandwich test

    SciTech Connect

    Aslam, Tariq D; Jackson, Scott I; Morris, John S

    2009-01-01

    Recent results utilizing proton radiography (P-Rad) during the detonation of the high explosive PBX 9502 are presented. Specifically, the effects of confinement of the detonation are examined in the LANL detonation confinement sandwich geometry. The resulting detonation velocity and detonation shock shape are measured. In addition, proton radiography allows one to image the reflected shocks through the detonation products. Comparisons are made with detonation shock dynamics (DSD) and reactive flow models for the lead detonation shock and detonation velocity. In addition, predictions of reflected shocks are made with the reactive flow models.

  17. Numerical simulation of spinning detonation in circular section channels

    NASA Astrophysics Data System (ADS)

    Levin, V. A.; Manuylovich, I. S.; Markov, V. V.

    2016-06-01

    Numerical simulation of three-dimensional structures of gas detonation in circular section channels that emerge due to the instability when the one-dimensional flow is initiated by energy supply at the closed end of the channel is performed. It is found that in channels with a large diameter, an irregular three-dimensional cellular detonation structure is formed. Furthermore, it is found that in channels with a small diameter circular section, the initially plane detonation wave is spontaneously transformed into a spinning detonation wave, while passing through four phases. A critical value of the channel diameter that divides the regimes with the three-dimensional cellular detonation and spinning detonation is determined. The stability of the spinning detonation wave under perturbations occurring when the wave passes into a channel with a greater (a smaller) diameter is investigated. It is found that the spin is preserved if the diameter of the next channel (into which the wave passes) is smaller (respectively, greater) than a certain critical value. The computations were performed on the Lomonosov supercomputer using from 0.1 to 10 billions of computational cells. All the computations of the cellular and spinning detonation were performed in the whole long three-dimensional channel (up to 1 m long) rather than only in its part containing the detonation wave; this made it possible to adequately simulate and investigate the features of the transformation of the detonation structure in the process of its propagation.

  18. Detonation diffraction in combustible high-speed flows

    NASA Astrophysics Data System (ADS)

    Gui, Mingyue; Fan, Baochun; Li, Baoming

    2016-03-01

    Detonation propagating in a T-shaped tube with quiescent and moving hydrogen/oxygen/argon mixtures is numerically examined based on the Euler equations with detailed finite-rate chemistry using the fifth-order weighted essentially non-oscillatory scheme. When diffracted in a quiescent combustible mixture, the detonation wave propagating from the bottom of the T-shaped tube is influenced by the corner rarefaction waves and decays into a non-reacting shock. Subsequently, the decoupled shock reflects irregularly from the top wall. Through several reflections back and forth between the top and bottom walls, a planar detonation is finally re-established. When the combustible mixture in the horizontal part flows from the left to the right, the detonation products ejected from the vertical tube will retard the flow, generating a compression flow upstream and a rarefaction flow downstream. The disturbed detonation on the left side is stronger than that on the right side. The final planar detonation in the upstream direction propagates faster than the Chapman-Jouguet (CJ) detonation with compressed, fine cellular structures, whereas the detonation in the downstream direction propagates more slowly than the CJ detonation with elongated, coarse cellular structures. The details of the transient behavior of diffracting detonation in high-speed flows are discussed.

  19. Hypersonic airbreathing vehicle conceptual design (focus on aero-space plane)

    NASA Technical Reports Server (NTRS)

    Hunt, James L.; Martin, John G.

    1989-01-01

    The airbreathing single stage to orbit (SSTO) vehicle design environment is variable-rich, intricately networked and sensitivity intensive. As such, it represents a tremondous technology challenge. Creating a viable design will require sophisticated configuration/synthesis and the synergistic integration of advanced technologies across the discipline spectrum. In design exercises, reductions in the fuel weight-fraction requirements projected for an orbital vehicle concept can result from improvements in aerodynamics/controls, propulsion efficiencies and trajectory optimization; also, gains in the fuel weight-fraction achievable for such a concept can result from improvements in structural design, heat management techniques, and material properties. As these technology advances take place, closure on a viable vehicle design will be realizable.

  20. Continuous high order sliding mode controller design for a flexible air-breathing hypersonic vehicle.

    PubMed

    Wang, Jie; Zong, Qun; Su, Rui; Tian, Bailing

    2014-05-01

    This paper investigates the problem of tracking control with uncertainties for a flexible air-breathing hypersonic vehicle (FAHV). In order to overcome the analytical intractability of this model, an Input-Output linearization model is constructed for the purpose of feedback control design. Then, the continuous finite time convergence high order sliding mode controller is designed for the Input-Output linearization model without uncertainties. In addition, a nonlinear disturbance observer is applied to estimate the uncertainties in order to compensate the controller and disturbance suppression, where disturbance observer and controller synthesis design is obtained. Finally, the synthesis of controller and disturbance observer is used to achieve the tracking for the velocity and altitude of the FAHV and simulations are presented to illustrate the effectiveness of the control strategies. PMID:24534328

  1. Robust tracking control for an air-breathing hypersonic vehicle with input constraints

    NASA Astrophysics Data System (ADS)

    Gao, Gang; Wang, Jinzhi; Wang, Xianghua

    2014-12-01

    The focus of this paper is on the design and simulation of robust tracking control for an air-breathing hypersonic vehicle (AHV), which is affected by high nonlinearity, uncertain parameters and input constraints. The linearisation method is employed for the longitudinal AHV model about a specific trim condition, and then considering the additive uncertainties of three parameters, the linearised model is just in the form of affine parameter dependence. From this point, the linear parameter-varying method is applied to design the desired controller. The poles for the closed-loop system of the linearised model are placed into a desired vertical strip, and the quadratic stability of the closed-loop system is guaranteed. Input constraints of the AHV are addressed by additional linear matrix inequalities. Finally, the designed controller is evaluated on the nonlinear AHV model and simulation results demonstrate excellent tracking performance with good robustness.

  2. Air-breathing aerospace plane development essential: Hypersonic propulsion flight tests

    NASA Technical Reports Server (NTRS)

    Mehta, Unmeel B.

    1994-01-01

    Hypersonic air-breathing propulsion utilizing scramjets can fundamentally change transatmospheric accelerators for low earth-to-orbit and return transportation. The value and limitations of ground tests, of flight tests, and of computations are presented, and scramjet development requirements are discussed. It is proposed that near full-scale hypersonic propulsion flight tests are essential for developing a prototype hypersonic propulsion system and for developing computational-design technology so that it can be used for designing this system. In order to determine how these objectives should be achieved, some lessons learned from past programs are presented. A conceptual two-stage-to-orbit (TSTO) prototype/experimental aerospace plane is recommended as a means of providing access-to-space and for conducting flight tests. A road map for achieving these objectives is also presented.

  3. Mach 6 flow field surveys beneath the forebody of an airbreathing missile

    NASA Technical Reports Server (NTRS)

    Johnson, P. J.; Hunt, J. L.

    1986-01-01

    Wall static, local stream static, and pitot pressure surveys were made on the windward side of a hypersonic airbreathing missile at full-scale length Reynolds numbers. In the inviscid part of the flow field, the experimental massflow ratios agreed with trends predicted by a three-dimensional method-of-characteristics solution. At a longitudinal station 3.5 diameters downstrea of the nose, the boundary layer was transitional or turbulent at zero incidence but became laminar as the angle of attack increased. The bell-shaped distribution of the boundary layer across the width of the body affected the mass flow distribution out to the bow shock and decreased the mass flow available the engine inlet.

  4. Multi-Disciplinary Design Optimization of Hypersonic Air-Breathing Vehicle

    NASA Astrophysics Data System (ADS)

    Wu, Peng; Tang, Zhili; Sheng, Jianda

    2016-06-01

    A 2D hypersonic vehicle shape with an idealized scramjet is designed at a cruise regime: Mach number (Ma) = 8.0, Angle of attack (AOA) = 0 deg and altitude (H) = 30kms. Then a multi-objective design optimization of the 2D vehicle is carried out by using a Pareto Non-dominated Sorting Genetic Algorithm II (NSGA-II). In the optimization process, the flow around the air-breathing vehicle is simulated by inviscid Euler equations using FLUENT software and the combustion in the combustor is modeled by a methodology based on the well known combination effects of area-varying pipe flow and heat transfer pipe flow. Optimization results reveal tradeoffs among total pressure recovery coefficient of forebody, lift to drag ratio of vehicle, specific impulse of scramjet engine and the maximum temperature on the surface of vehicle.

  5. Planar array stack design aided by rapid prototyping in development of air-breathing PEMFC

    NASA Astrophysics Data System (ADS)

    Chen, Chen-Yu; Lai, Wei-Hsiang; Weng, Biing-Jyh; Chuang, Huey-Jan; Hsieh, Ching-Yuan; Kung, Chien-Chih

    The polymer electrolyte membrane fuel cell (PEMFC) is one of the most important research topics in the new and clean energy area. The middle or high power PEMFCs can be applied to the transportation or the distributed power system. But for the small power application, it is needed to match the power requirement of the product generally. On the other hand, the direct methanol fuel cell (DMFC) is one of the most common type that researchers are interested in, but recently the miniature or the micro-PEMFCs attract more attention due to their advantages of high open circuit voltage and high power density. The objective of this study is to develop a new air-breathing planar array fuel cell stacked from 10 cells made by rapid prototyping technology which has potential for fast commercial design, low cost manufacturing, and even without converters/inverters for the system. In this paper, the main material of flow field plates is acrylonitrile-butadiene-styrene (ABS) which allows the fuel cell be mass-manufactured by plastic injection molding technology. The rapid prototyping technology is applied to construct the prototype and verify the practicability of the proposed stack design. A 10-cell air-breathing miniature PEMFC stack with a volume of 6 cm × 6 cm × 0.9 cm is developed and tested. Its segmented membrane electrode assembly (MEA) is designed with the active surface area of 1.3 cm × 1.3 cm in each individual MEA. The platinum loading at anode and cathode are 0.2 mg cm -2 and 0.4 mg cm -2, respectively. Results show that the peak power densities of the parallel connected and serial connected stack are 99 mW cm -2 at 0.425 V and 92 mW cm -2 at 4.25 V, respectively under the conditions of 70 °C relative saturated humidity (i.e., dew point temperature), ambient temperature and free convection air. Besides, the stack performance is increased under forced convection. If the cell surface air is blown by an electric fan, the peak power densities of parallel connected and

  6. Performance analysis of a turbofan as a part of an airbreathing propulsion system for space shuttles

    NASA Astrophysics Data System (ADS)

    Steinebach, D. A.; Kuehl, W.; Gallus, H. E.

    1993-04-01

    This paper presents the results of the design and performance analysis of airbreathing engines for aerospace planes. The analysis is illustrated by introducing an exemplary twin-shaft turbofan engine with post-combustion and bypass-combustion. Some modules of the performance analysis algorithm such as inlet pressure recovery or real gas effects are also presented. The jet engine is designed in view of increasing temperatures at high flight Mach numbers. Hence, the engine design data are dependent on the characteristics of the available materials as well as on the trajectory of the aerospace plane. The results illustrate the strong influence of the real gas effects on the engine thrust particularly in the case of over-stoichiometric combustion of hydrogen. Turbofan engines offer the following advantages in comparison with equivalent turbojet engines: higher thrust performance in supersonic flight range and lower fuel consumption due to operation management of post-combustion and bypass-combustion.

  7. Hypersonic propulsion flight tests as essential to air-breathing aerospace plane development

    NASA Technical Reports Server (NTRS)

    Mehta, U.

    1995-01-01

    Hypersonic air-breathing propulsion utilizing scramjets can fundamentally change transatmospheric acclerators for transportation from low Earth orbits (LEOs). The value and limitations of ground tests, of flight tests, and of computations are presented, and scramjet development requirements are discussed. Near-full-scale hypersonic propulsion flight tests are essential for developing a prototype hypersonic propulsion system and for developing computation-design technology that can be used in designing that system. In order to determine how these objectives should be achieved, some lessons learned from past programs are presented. A conceptual two-stage-to-orbit (TSTO) prototype/experimental aerospace plane is recommended as a means of providing access-to-space and for conducting flight tests. A road map for achieving these objectives is also presented.

  8. Air-breathing aerospace plane development essential: Hypersonic propulsion flight tests

    NASA Technical Reports Server (NTRS)

    Mehta, Unmeel B.

    1995-01-01

    Hypersonic airbreathing propulsion utilizing scramjets can change transatmospheric accelerators for low earth-to-orbit and return transportation. The value and limitation of ground tests, of flight tests, and of computations are presented, and scramjet development requirements are discussed. It is proposed that near full-scale hypersonic propulsion flight tests are essential for developing computational design technology so that it can be used for designing this system. In order to determine how these objectives should be achieved, some lessons learned from past programs are presented. A conceptual two-stage-to-orbit (TSTO) prototype/experimental aerospace plane is recommended as a means of providing access-to-space and for conducting flight tests.

  9. Ascent performance of an air-breathing horizontal-takeoff launch vehicle

    NASA Technical Reports Server (NTRS)

    Powell, Richard W.; Shaughnessy, John D.; Cruz, Christopher I.; Naftel, J. C.

    1991-01-01

    Simulations are conducted to investigate a proposed NASA launch vehicle that is fully reusable, takes off horizontally, and uses airbreathing propulsion in a single stage. The propulsion model is based on a cycle analysis method, and the vehicle is assumed to be a rigid structure with distributed fuel, operating under a range of atmospheric conditions. The program to optimize simulated trajectories (POST) is modified to include a predictor-corrector guidance capability and then used to generate the trajectories. Significant errors are encountered during the unpowered coast phase due to uncertainty in the atmospheric density profile. The amount of ascent propellant needed is shown to be directly related to the thrust-vector angle and the location of the center of gravity of the vehicle because of the importance of aim-drag losses to total ideal velocity.

  10. Simulations of Heterogeneous Detonations and Post Detonation Turbulent Mixing and Afterburning

    NASA Astrophysics Data System (ADS)

    Menon, Suresh; Gottiparthi, Kalyana

    2011-06-01

    Most metal-loaded explosives and thermobaric explosives exploit the afterburning of metals to maintain pressure and temperature conditions.The use of such explosives in complex environment can result in post detonation flow containing many scales of vortical motion, flow jetting and shear, as well as plume-surface interactions due to flow impingement and wall flows. In general, all these interactions can lead to highly turbulent flow fields even if the initial ambient conditions were quiescent. Thus, turbulent mixing can dominate initial mixing and impact the final afterburn. We conduct three-dimensional numerical simulations of the propagation of detonation resulting from metal-loaded (inert or reacting) explosives and analyze the afterburn process as well as the generation of multiple scales of mixing in the post detonation flow field. Impact of the detonation and post-detonation flow field on solid surface is also considered for a variety of initial conditions. Comparison with available data is carried out to demonstrate validity of the simulation method. Supported by Defense Threat Reduction Agency

  11. Turbojet-type engines for the airbreathing propulsion of reusable winged launchers

    NASA Astrophysics Data System (ADS)

    Duparcq, J. L.; Hermant, E.; Scherrer, D.

    Combined propulsion systems for hypersonic application have become new challenges for industrial and research organizations. In France, SNECMA and SEP, which have just joined together for a common effort on hypersonics within Hyperspace, and ONERA have been involved, under CNES (French space agency) contracts, in the assessment of new propulsion concepts for reusable winged launchers (SSTO or TSTO). As potential solutions for the airbreathing propulsion, some turbojet-type engines are presented: —the twin spool turbojet or turbofan with reheat —the turbojet with reheat —the twin-duct turbojet ramjet —the precooled turbojet with reheat. All these engines have been sized for a flight Mach number under seven with a cryogenic fuel (liquid hydrogen). Mainly due to total temperature and pressure encountered along the trajectory, the systems will have to withstand severe physical constraints. Coupled with performance and size requirements, like specific thrust and maximum air capture area, these operating conditions have been taken into account in order to select each engine cycle and technical arrangement. Performance and mass criteria make it possible to compare these systems and to emphasize their distinctive features among the propulsion concepts envisioned for the future reusable winged launchers (including airbreathing combined engines under study in France). The first step of the final selection, leading to the best adaptation between the engine and the vehicle, will then be tackled. This will be particularly enhanced by the analysis of potential advantages or technical difficulties, like thrust-to-weight ratio or needs of variable geometry and heat exchangers. The twin-duct turbojet ramjet, for example, is probably one of the best candidates for the first stages of propulsion of a reusable winged launcher.

  12. Cascade Optimization Strategy for Aircraft and Air-Breathing Propulsion System Concepts

    NASA Technical Reports Server (NTRS)

    Patnaik, Surya N.; Lavelle, Thomas M.; Hopkins, Dale A.; Coroneos, Rula M.

    1996-01-01

    Design optimization for subsonic and supersonic aircraft and for air-breathing propulsion engine concepts has been accomplished by soft-coupling the Flight Optimization System (FLOPS) and the NASA Engine Performance Program analyzer (NEPP), to the NASA Lewis multidisciplinary optimization tool COMETBOARDS. Aircraft and engine design problems, with their associated constraints and design variables, were cast as nonlinear optimization problems with aircraft weight and engine thrust as the respective merit functions. Because of the diversity of constraint types and the overall distortion of the design space, the most reliable single optimization algorithm available in COMETBOARDS could not produce a satisfactory feasible optimum solution. Some of COMETBOARDS' unique features, which include a cascade strategy, variable and constraint formulations, and scaling devised especially for difficult multidisciplinary applications, successfully optimized the performance of both aircraft and engines. The cascade method has two principal steps: In the first, the solution initiates from a user-specified design and optimizer, in the second, the optimum design obtained in the first step with some random perturbation is used to begin the next specified optimizer. The second step is repeated for a specified sequence of optimizers or until a successful solution of the problem is achieved. A successful solution should satisfy the specified convergence criteria and have several active constraints but no violated constraints. The cascade strategy available in the combined COMETBOARDS, FLOPS, and NEPP design tool converges to the same global optimum solution even when it starts from different design points. This reliable and robust design tool eliminates manual intervention in the design of aircraft and of air-breathing propulsion engines where it eases the cycle analysis procedures. The combined code is also much easier to use, which is an added benefit. This paper describes COMETBOARDS

  13. Geometry Modeling and Adaptive Control of Air-Breathing Hypersonic Vehicles

    NASA Astrophysics Data System (ADS)

    Vick, Tyler Joseph

    Air-breathing hypersonic vehicles have the potential to provide global reach and affordable access to space. Recent technological advancements have made scramjet-powered flight achievable, as evidenced by the successes of the X-43A and X-51A flight test programs over the last decade. Air-breathing hypersonic vehicles present unique modeling and control challenges in large part due to the fact that scramjet propulsion systems are highly integrated into the airframe, resulting in strongly coupled and often unstable dynamics. Additionally, the extreme flight conditions and inability to test fully integrated vehicle systems larger than X-51 before flight leads to inherent uncertainty in hypersonic flight. This thesis presents a means to design vehicle geometries, simulate vehicle dynamics, and develop and analyze control systems for hypersonic vehicles. First, a software tool for generating three-dimensional watertight vehicle surface meshes from simple design parameters is developed. These surface meshes are compatible with existing vehicle analysis tools, with which databases of aerodynamic and propulsive forces and moments can be constructed. A six-degree-of-freedom nonlinear dynamics simulation model which incorporates this data is presented. Inner-loop longitudinal and lateral control systems are designed and analyzed utilizing the simulation model. The first is an output feedback proportional-integral linear controller designed using linear quadratic regulator techniques. The second is a model reference adaptive controller (MRAC) which augments this baseline linear controller with an adaptive element. The performance and robustness of each controller are analyzed through simulated time responses to angle-of-attack and bank angle commands, while various uncertainties are introduced. The MRAC architecture enables the controller to adapt in a nonlinear fashion to deviations from the desired response, allowing for improved tracking performance, stability, and

  14. From combustion and detonation to nitrogen oxides

    NASA Astrophysics Data System (ADS)

    Ivanov, M. F.; Kiverin, A. D.; Klumov, B. A.; Fortov, V. E.

    2014-03-01

    This paper looks at Ya B Zeldovich's ideas on the combustion and detonation physics of gaseous mixtures and how they evolved as work in this field progressed. The paper demonstrates the fundamental role of Zeldovich's concept of spontaneous combustion waves in studying transient initiation processes for various combustion regimes and in determining the energy and concentration inflammation limits for combustible gaseous mixtures. It shows how his notion that flame front stretching crucially influences flame acceleration in channels explains in a new way the deflagration-to-detonation transition in highly reactive gaseous mixtures. Most of the presented results were obtained by simulations, allowing Zeldovich's ideas to be extended to the combustion of real gaseous mixtures, where chemical reactions and gasdynamical flows add hugely to the complexity of the problem. The paper concludes by using Zeldovich's mechanism to assess the amount of nitrogen oxide produced by a lightning discharge.

  15. Spectrum acquisition of detonation based on CMOS

    NASA Astrophysics Data System (ADS)

    Li, Yan; Bai, Yonglin; Wang, Bo; Liu, Baiyu; Xue, Yingdong; Zhang, Wei; Gou, Yongsheng; Bai, Xiaohong; Qin, Junjun; Xian, Ouyang

    2010-10-01

    The detection of high-speed dynamic spectrum is the main method to acquire transient information. In order to obtain the large amount spectral data in real-time during the process of detonation, a CMOS-based system with high-speed spectrum data acquisition is designed. The hardware platform of the system is based on FPGA, and the unique characteristic of CMOS image sensors in the rolling shutter model is used simultaneously. Using FPGA as the master control chip of the system, not only provides the time sequence for CIS, but also controls the storage and transmission of the spectral data. In the experiment of spectral data acquisition, the acquired information is transmitted to the host computer through the CameraLink bus. The dynamic spectral curve is obtained after the subsequent processing. The experimental results demonstrate that this system is feasible in the acquisition and storage of high-speed dynamic spectrum information during the process of detonation.

  16. Detonation in shocked homogeneous high explosives

    SciTech Connect

    Yoo, C.S.; Holmes, N.C.; Souers, P.C.

    1995-11-01

    We have studied shock-induced changes in homogeneous high explosives including nitromethane, tetranitromethane, and single crystals of pentaerythritol tetranitrate (PETN) by using fast time-resolved emission and Raman spectroscopy at a two-stage light-gas gun. The results reveal three distinct steps during which the homogeneous explosives chemically evolve to final detonation products. These are (1) the initiation of shock compressed high explosives after an induction period, (2) thermal explosion of shock-compressed and/or reacting materials, and (3) a decay to a steady-state representing a transition to the detonation of uncompressed high explosives. Based on a gray-body approximation, we have obtained the CJ temperatures: 3800 K for nitromethane, 2950 K for tetranitromethane, and 4100 K for PETN. We compare the data with various thermochemical equilibrium calculations. In this paper we will also show a preliminary result of single-shot time-resolved Raman spectroscopy applied to shock-compressed nitromethane.

  17. Smooth blasting with the electronic delay detonator

    SciTech Connect

    Yamamoto, Masaaki; Ichijo, Toshiyuki; Tanaka, Yoshiharu

    1995-12-31

    The authors utilized electronic detonators (EDs) to investigate the effect of high detonator delay accuracy on overbreak, remaining rock damage, and surface smoothness, in comparison with that of long-period delay detonators (0.25 sec interval) PDs. The experiments were conducted in a deep mine, in a test site region composed of very hard granodiorite with a seismic wave velocity of about 6.0 km/sec and a uniaxial compressive strength, uniaxial tensile strength, and Young`s modulus of 300 MPa, 12 MPa, and 73 GPa, respectively. The blasting design was for a test tunnel excavation of 8 m{sup 2} in cross section, with an advance per round of 2.5 m. Five rounds were performed, each with a large-hole cut and perimeter holes in a 0.4-m spacing charged with 20-mm-diameter water gel explosive to obtain low charge concentration. EDs were used in the holes on the perimeter of the right half, and PDs were used in all other holes. Following each shot, the cross section was measured by laser to determine amount of overbreak and surface smoothness. In situ seismic prospecting was used to estimate the depth of damage in the remaining rock, and the damage was further investigated by boring into both side walls.

  18. Kinetic information from detonation front curvature

    SciTech Connect

    Souers, P. C., LLNL

    1998-06-15

    The time constants for time-dependent modeling may be estimated from reaction zone lengths, which are obtained from two sources One is detonation front curvature, where the edge lag is close to being a direct measure The other is the Size Effect, where the detonation velocity decreases with decreasing radius as energy is lost to the cylinder edge A simple theory that interlocks the two effects is given A differential equation for energy flow in the front is used, the front is described by quadratic and sixth-power radius terms The quadratic curvature comes from a constant power source of energy moving sideways to the walls Near the walls, the this energy rises to the total energy of detonation and produces the sixth-power term The presence of defects acting on a short reaction zone can eliminate the quadratic part while leaving the wall portion of the cuvature A collection of TNT data shows that the reaction zone increases with both the radius and the void fraction

  19. A gasdynamic gun driven by gaseous detonation

    NASA Astrophysics Data System (ADS)

    Li, Jinping; Chen, Hong; Zhang, Shizhong; Zhang, Xiaoyuan; Yu, Hongru

    2016-01-01

    A gasdynamic gun driven by gaseous detonation was developed to address the disadvantages of the insufficient driving capability of high-pressure gas and the constraints of gunpowder. The performance of this gasdynamic gun was investigated through experiments and numerical simulations. Much more powerful launching capability was achieved by this gun relative to a conventional high-pressure gas gun, owing to the use of the chemical energy of the driver gas. To achieve the same launching condition, the initial pressure required for this gun was an order of magnitude lower than that for a gun driven by high-pressure H2. Because of the presence of the detonation, however, a more complex internal ballistic process of this gun was observed. Acceleration of projectiles for this gun was accompanied by a series of impulse loads, in contrast with the smooth acceleration for a conventional one, which indicates that this gun should be used conditionally. The practical feasibility of this gun was verified by experiments. The experiments demonstrated the convenience of taking advantage of the techniques developed for detonation-driven shock tubes and tunnels.

  20. A gasdynamic gun driven by gaseous detonation.

    PubMed

    Li, Jinping; Chen, Hong; Zhang, Shizhong; Zhang, Xiaoyuan; Yu, Hongru

    2016-01-01

    A gasdynamic gun driven by gaseous detonation was developed to address the disadvantages of the insufficient driving capability of high-pressure gas and the constraints of gunpowder. The performance of this gasdynamic gun was investigated through experiments and numerical simulations. Much more powerful launching capability was achieved by this gun relative to a conventional high-pressure gas gun, owing to the use of the chemical energy of the driver gas. To achieve the same launching condition, the initial pressure required for this gun was an order of magnitude lower than that for a gun driven by high-pressure H2. Because of the presence of the detonation, however, a more complex internal ballistic process of this gun was observed. Acceleration of projectiles for this gun was accompanied by a series of impulse loads, in contrast with the smooth acceleration for a conventional one, which indicates that this gun should be used conditionally. The practical feasibility of this gun was verified by experiments. The experiments demonstrated the convenience of taking advantage of the techniques developed for detonation-driven shock tubes and tunnels. PMID:26827358

  1. Characterizing detonator output using dynamic witness plates

    SciTech Connect

    Murphy, Michael John; Adrian, Ronald J

    2009-01-01

    A sub-microsecond, time-resolved micro-particle-image velocimetry (PIV) system is developed to investigate the output of explosive detonators. Detonator output is directed into a transparent solid that serves as a dynamic witness plate and instantaneous shock and material velocities are measured in a two-dimensional plane cutting through the shock wave as it propagates through the solid. For the case of unloaded initiators (e.g. exploding bridge wires, exploding foil initiators, etc.) the witness plate serves as a surrogate for the explosive material that would normally be detonated. The velocity-field measurements quantify the velocity of the shocked material and visualize the geometry of the shocked region. Furthermore, the time-evolution of the velocity-field can be measured at intervals as small as 10 ns using the PIV system. Current experimental results of unloaded exploding bridge wire output in polydimethylsiloxane (PDMS) witness plates demonstrate 20 MHz velocity-field sampling just 300 ns after initiation of the wire.

  2. Multistage reaction pathways in detonating high explosives

    NASA Astrophysics Data System (ADS)

    Li, Ying; Kalia, Rajiv; Nakano, Aiichiro; Vashishta, Priya; CACS Collaboration; ALCF Team

    2015-06-01

    Atomistic mechanisms underlying the reaction time and intermediate reaction products of detonating high explosives far from equilibrium have been elusive. This is because detonation is one of the hardest multiscale physics problems, in which diverse length and time scales play important roles. Here, large spatiotemporal-scale reactive molecular dynamics simulations validated by quantum molecular dynamics simulations reveal a two-stage reaction mechanism during the detonation of cyclotrimethylenetrinitramine crystal. Rapid production of N2 and H2O within 10 ps is followed by delayed production of CO molecules beyond ns. We found that further decomposition towards the final products is inhibited by the formation of large metastable carbon- and oxygen-rich clusters with fractal geometry. In addition, we found distinct uni-molecular and intermolecular reaction pathways, respectively, for the rapid N2 and H2O productions. This work was supported by the Office of Naval Research Grant No. N000014-12-1-0555 and the Basic Research Program of Defense Threat Reduction Agency (DTRA) Grant No. HDTRA1-08-1-0036. All the simulations were performed at USC and Argonne LCF.

  3. Detonation of Meta-stable Clusters

    SciTech Connect

    Kuhl, Allen; Kuhl, Allen L.; Fried, Laurence E.; Howard, W. Michael; Seizew, Michael R.; Bell, John B.; Beckner, Vincent; Grcar, Joseph F.

    2008-05-31

    We consider the energy accumulation in meta-stable clusters. This energy can be much larger than the typical chemical bond energy (~;;1 ev/atom). For example, polymeric nitrogen can accumulate 4 ev/atom in the N8 (fcc) structure, while helium can accumulate 9 ev/atom in the excited triplet state He2* . They release their energy by cluster fission: N8 -> 4N2 and He2* -> 2He. We study the locus of states in thermodynamic state space for the detonation of such meta-stable clusters. In particular, the equilibrium isentrope, starting at the Chapman-Jouguet state, and expanding down to 1 atmosphere was calculated with the Cheetah code. Large detonation pressures (3 and 16 Mbar), temperatures (12 and 34 kilo-K) and velocities (20 and 43 km/s) are a consequence of the large heats of detonation (6.6 and 50 kilo-cal/g) for nitrogen and helium clusters respectively. If such meta-stable clusters could be synthesized, they offer the potential for large increases in the energy density of materials.

  4. Methods for proving the equivalency of detonator performance

    SciTech Connect

    Munger, Alan C; Akinci, Adrian A; Thomas, Keith A; Clarke, Steve A; Martin, Eric S; Murphy, Michael J

    2009-01-01

    One of the challenges facing engineers is developing newer, safer detonators that are equivalent to devices currently in use. There is no clear consensus on an exact method for drawing equivalence of detonators. This paper summarizes our current efforts to develop diagnostics addressing various aspects of detonator design to better quantify and prove equivalency. We consider various optical techniques to quantify the output pressure and output wave shape. The development of a unique interpretation of streak camera breakouts, known as the apparent center of initiation, will be discussed as a metric for detonation wave shape. Specific examples apply these techniques to the comparison of a new laser-driven detonator with an existing exploding bridgewire (EBW) detonator. Successes and short-comings of the techniques will be discussed.

  5. Numerical simulation of spinning detonation in square tube

    NASA Astrophysics Data System (ADS)

    Tsuboi, Nobuyuki; Asahara, Makoto; Eto, Keitaro; Hayashi, A. Koichi

    2008-09-01

    A single spinning detonation wave propagating in a square tube is simulated three-dimensionally with the detailed chemical reaction mechanism for hydrogen/air mixture proposed by Petersen and Hanson. The spinning detonation is composed of a transverse detonation rotating around the wall normal to the tube axis, triple lines propagating partially out of phase, and a short pressure trail. The formation of an unburned gas pocket behind the detonation front was not observed in the present simulations because the rotating transverse detonation completely consumed the unburned gas. The calculated profiles of instantaneous OH mass fraction have a keystone shape behind the detonation front. The numerical results for the pitch and track angle on the tube wall agree well with the experimental results.

  6. Insensitive detonator apparatus for initiating large failure diameter explosives

    DOEpatents

    Perry, III, William Leroy

    2015-07-28

    A munition according to a preferred embodiment can include a detonator system having a detonator that is selectively coupled to a microwave source that functions to selectively prime, activate, initiate, and/or sensitize an insensitive explosive material for detonation. The preferred detonator can include an explosive cavity having a barrier within which an insensitive explosive material is disposed and a waveguide coupled to the explosive cavity. The preferred system can further include a microwave source coupled to the waveguide such that microwaves enter the explosive cavity and impinge on the insensitive explosive material to sensitize the explosive material for detonation. In use the preferred embodiments permit the deployment and use of munitions that are maintained in an insensitive state until the actual time of use, thereby substantially preventing unauthorized or unintended detonation thereof.

  7. Effects of high sound speed confiners on ANFO detonations

    NASA Astrophysics Data System (ADS)

    Kiyanda, Charles; Jackson, Scott; Short, Mark

    2011-06-01

    The interaction between high explosive (HE) detonations and high sound speed confiners, where the confiner sound speed exceeds the HE's detonation speed, has not been thoroughly studied. The subsonic nature of the flow in the confiner allows stress waves to travel ahead of the main detonation front and influence the upstream HE state. The interaction between the detonation wave and the confiner is also no longer a local interaction, so that the confiner thickness now plays a significant role in the detonation dynamics. We report here on larger scale experiments in which a mixture of ammonium nitrate and fuel oil (ANFO) is detonated in aluminium confiners with varying charge diameter and confiner thickness. The results of these large-scale experiments are compared with previous large-scale ANFO experiments in cardboard, as well as smaller-scale aluminium confined ANFO experiments, to characterize the effects of confiner thickness.

  8. Dynamics of detonations and explosions: Detonations; International Colloquium on Dynamics of Explosions and Reactive Systems, 12th, University of Michigan, Ann Arbor, July 23-28, 1989, Technical Papers

    SciTech Connect

    Kuhl, A.L.; Leyer, J.-C.; Borisov, A.A.; Sirignano, W.A.

    1991-01-01

    The present volume on the dynamics of gaseous detonations, detonation initiation and transmission, multiphase detonations, and nonideal detonations and boundary effects, discusses the detonability of hydrocarbon fuels in air, the detonation of cryogenic gaseous hydrogen-oxygen mixtures, chemical kinetics-detonation structure correlations for gaseous explosives, the initiation of hydrogen-air detonations by turbulent fluorine-air jets, and the initiation of a detonation wave due to multistage self-ignition. Also discussed are the limit criterion for detonation in circular cubes, oblique detonation at hypersonic velocities, the mechanisms of detonation propagation in porous structures, surface detonations and indirect ignition processes, the detonation of unconfined large-scale fuel-spray/air clouds, the detonation structure of corn starch particles-oxygen mixtures, and the lean detonation limit of sensitized kerosene sprays in air.

  9. Explosives malfunction from sympathetic detonation to shock desensitization

    SciTech Connect

    Katsabanis, P.D.; Yeung, C.; Fitz, G.; Heater, R.

    1994-12-31

    Explosives malfunction due to shock waves is a serious concern for successful blasting results. Malfunction can range from sympathetic detonation to desensitization and modification of firing times of conventional pyrotechnic detonators. Decked charges consisting of commercial emulsion explosives having a detonator and a primer were placed in 10cm diameter blastholes and their performance was recorded. Due to the limited length of the holes the events were mainly sympathetic detonations although desensitization was also recorded. Pressure measurements along the stemming column showed that shock waves produced by an explosive have a significant amplitude even at relatively large distances away from the detonating explosive. It was found that 2m away from a detonating charge the pressures in the stemming material were above 0.1 GPa indicating that there is potential for primers and detonators to malfunction. Parallel charges consisting of a commercial emulsion explosive with a diameter of 32mm were confined in 2mm thick steel tubes and initiation was attempted using detonators having a delay interval of 25ms. The charges were placed in sand and the velocity of detonation of the acceptor charge was recorded using a continuous resistance probe system. Carbon resistors were also placed in the same position as the acceptor charge to examine the dynamic pressures that were applied to the charge. Sympathetic detonation, complete desensitization, partial desensitization and properly sequenced detonations were observed as the distance between charges was increased from 76 mm to 305 mm. Delay detonators were also tested in a similar to the last configuration. Modification of firing times was observed at distances between 150 and 360 mm.

  10. THE DETONATION MECHANISM OF THE PULSATIONALLY ASSISTED GRAVITATIONALLY CONFINED DETONATION MODEL OF Type Ia SUPERNOVAE

    SciTech Connect

    Jordan, G. C. IV; Graziani, C.; Weide, K.; Norris, J.; Hudson, R.; Lamb, D. Q.; Fisher, R. T.; Townsley, D. M.; Meakin, C.; Reid, L. B.

    2012-11-01

    We describe the detonation mechanism composing the 'pulsationally assisted' gravitationally confined detonation (GCD) model of Type Ia supernovae. This model is analogous to the previous GCD model reported in Jordan et al.; however, the chosen initial conditions produce a substantively different detonation mechanism, resulting from a larger energy release during the deflagration phase. The resulting final kinetic energy and {sup 56}Ni yields conform better to observational values than is the case for the 'classical' GCD models. In the present class of models, the ignition of a deflagration phase leads to a rising, burning plume of ash. The ash breaks out of the surface of the white dwarf, flows laterally around the star, and converges on the collision region at the antipodal point from where it broke out. The amount of energy released during the deflagration phase is enough to cause the star to rapidly expand, so that when the ash reaches the antipodal point, the surface density is too low to initiate a detonation. Instead, as the ash flows into the collision region (while mixing with surface fuel), the star reaches its maximally expanded state and then contracts. The stellar contraction acts to increase the density of the star, including the density in the collision region. This both raises the temperature and density of the fuel-ash mixture in the collision region and ultimately leads to thermodynamic conditions that are necessary for the Zel'dovich gradient mechanism to produce a detonation. We demonstrate feasibility of this scenario with three three-dimensional (3D), full star simulations of this model using the FLASH code. We characterized the simulations by the energy released during the deflagration phase, which ranged from 38% to 78% of the white dwarf's binding energy. We show that the necessary conditions for detonation are achieved in all three of the models.

  11. Effects of diameter and temperature on XTX-8004 detonation velocity

    SciTech Connect

    Campos, C.A.

    1980-10-01

    This study was performed to determine the dependence of XTX-8004 steady detonation velocity on charge diameter and temperature. The tests were performed for four different diameters at three temperatures using a standard 4-track detonation velocity block and corresponding printed circuit ionization switch plate. The explosive was loaded in the detonation velocity block to a nominal density of 1.553 g/cc. Measurements obtained from two samples per temperature indicate the critical diameter is less than 0.178 cm. A relationship between detonation velocity and density due to temperature was established using experimental measurements.

  12. Detonation shock dynamics calibration for non-ideal HE: ANFO

    SciTech Connect

    Short, Mark; Salyer, Terry R; Aslam, Tariq D; Kiyanda, Charles B; Morris, John S; Zimmerley, Tony

    2009-01-01

    Linear D{sub n}-{kappa} detonation shock dynamics (DSD) filling forms are obtained for four ammonium nitrate-fuel oil (ANFO) mixtures involving variations in the ammonium nitrate prill properties and ANFO stoichiometries. The detonation of ammonium nitrate-fuel oil (ANFO) mixtures is considered to be highly nonideal involving long reaction zones ({approx} several cms), low detonation energies and large failure diameters ({approx} 10s-100s cms). A number of experimental programs have been undertaken to understand ANFO detonation properties as a function of the AN properties [1]-[7]. Given the highly heterogeneous nature of ANFO mixtures (typical high explosive (HE) grade AN prills are porous with a range of diameters) a predictive reactive flow simulation of ANFO detonation will present significant challenges. At Los Alamos, a simulation capability has been developed for predicting the propagation of detonation in non-ideal HE and the work conducted on surrounding materials via a combination of a detonation shock dynamics (DSD) approach and a modified programmed burn method known as the pseudo-reaction-zone (or PRZ) method that accounts for the long detonation reaction zone. In the following, linear D{sub n}-{kappa} DSD fitting forms are obtained for four ammonium nitrate-fuel oil mixtures involving variation in the ammonium nitrate prill properties and ANFO stoichiometries. A detonation shock dynamics calibration for ANFO consisting of regular porous HE grade AN in a 94/6 wt.% AN to FO mix has been obtained in [7].

  13. Safety and performance enhancement circuit for primary explosive detonators

    DOEpatents

    Davis, Ronald W.

    2006-04-04

    A safety and performance enhancement arrangement for primary explosive detonators. This arrangement involves a circuit containing an energy storage capacitor and preset self-trigger to protect the primary explosive detonator from electrostatic discharge (ESD). The circuit does not discharge into the detonator until a sufficient level of charge is acquired on the capacitor. The circuit parameters are designed so that normal ESD environments cannot charge the protection circuit to a level to achieve discharge. When functioned, the performance of the detonator is also improved because of the close coupling of the stored energy.

  14. A library of prompt detonation reaction zone data

    SciTech Connect

    Souers, P. C., LLNL

    1998-06-01

    Tables are given listing literature data that allows calculation of sonic reaction zones at or near steady-state for promptly detonating explosive cylinders. The data covers homogeneous, heterogeneous, composite, inorganic and binary explosives and allows comparison across the entire explosive field. Table 1 lists detonation front curvatures. Table 2 lists Size Effect data, i.e. the change of detonation velocity with cylinder radius. Table 3 lists failure radii and detonation velocities. Table 4 lists explosive compositions. A total of 51 references dating back into the 1950`s are given. Calculated reaction zones, radii of curvature and growth rate coefficients are listed.

  15. Hydroxyapatite Reinforced Coatings with Incorporated Detonationally Generated Nanodiamonds

    SciTech Connect

    Pramatarova, L.; Pecheva, E.; Hikov, T.; Fingarova, D.; Dimitrova, R.; Spassov, T.; Krasteva, N.; Mitev, D.

    2010-01-21

    We studied the effect of the substrate chemistry on the morphology of hydroxyapatite-detonational nanodiamond composite coatings grown by a biomimetic approach (immersion in a supersaturated simulated body fluid). When detonational nanodiamond particles were added to the solution, the morphology of the grown for 2 h composite particles was porous but more compact then that of pure hydroxyapatite particles. The nanodiamond particles stimulated the hydroxyapatite growth with different morphology on the various substrates (Ti, Ti alloys, glasses, Si, opal). Biocompatibility assay with MG63 osteoblast cells revealed that the detonational nanodiamond water suspension with low and average concentration of the detonational nanodiamond powder is not toxic to living cells.

  16. Hydroxyapatite Reinforced Coatings with Incorporated Detonationally Generated Nanodiamonds

    NASA Astrophysics Data System (ADS)

    Pramatarova, L.; Pecheva, E.; Dimitrova, R.; Spassov, T.; Krasteva, N.; Hikov, T.; Fingarova, D.; Mitev, D.

    2010-01-01

    We studied the effect of the substrate chemistry on the morphology of hydroxyapatite-detonational nanodiamond composite coatings grown by a biomimetic approach (immersion in a supersaturated simulated body fluid). When detonational nanodiamond particles were added to the solution, the morphology of the grown for 2 h composite particles was porous but more compact then that of pure hydroxyapatite particles. The nanodiamond particles stimulated the hydroxyapatite growth with different morphology on the various substrates (Ti, Ti alloys, glasses, Si, opal). Biocompatibility assay with MG63 osteoblast cells revealed that the detonational nanodiamond water suspension with low and average concentration of the detonational nanodiamond powder is not toxic to living cells.

  17. Detonation Shock Dynamics Calibration for Non-Ideal HE: ANFO

    NASA Astrophysics Data System (ADS)

    Short, Mark; Salyer, Terry

    2009-06-01

    The detonation of ammonium nitrate (AN) and fuel-oil (FO) mixtures (ANFO) is significantly influenced by the properties of the AN (porosity, particle size, coating) and fuel-oil stoichiometry. We report on a new series of rate-stick experiments in cardboard confinement that highlight detonation front speed and curvature dependence on AN/FO stoichiometry and AN particle properties. Standard detonation velocity-curvature calibrations to the experimental data will be presented, as well as higher-order time-dependent detonation shock dynamics calibrations.

  18. Molecular-dynamics investigation of the desensitization of detonable material

    NASA Astrophysics Data System (ADS)

    Rice, Betsy M.; Mattson, William; Trevino, Samuel F.

    1998-05-01

    A molecular-dynamics investigation of the effects of a diluent on the detonation of a model crystalline explosive is presented. The diluent, a heavy material that cannot exothermally react with any species of the system, is inserted into the crystalline explosive in two ways. The first series of simulations investigates the attenuation of the energy of a detonation wave in a pure explosive after it encounters a small layer of crystalline diluent that has been inserted into the lattice of the pure explosive. After the shock wave has traversed the diluent layer, it reenters the pure explosive. Unsupported detonation is not reestablished unless the energy of the detonation wave exceeds a threshold value. The second series of simulations investigates detonation of solid solutions of different concentrations of the explosive and diluent. For both types of simulations, the key to reestablishing or reaching unsupported detonation is the attainment of a critical number density behind the shock front. Once this critical density is reached, the explosive molecules make a transition to an atomic phase. This is the first step in the reaction mechanism that leads to the heat release that sustains the detonation. The reactive fragments formed from the atomization of the heteronuclear reactants subsequently combine with new partners, with homonuclear product formation exothermally favored. The results of detonation of the explosive-diluent crystals are consistent with those presented in an earlier study on detonation of pure explosive [B. M. Rice, W. Mattson, J. Grosh, and S. F. Trevino, Phys. Rev. E 53, 611 (1996)].

  19. LASERS: Iodine laser pumped by light from a shock front created by detonating an explosive

    NASA Astrophysics Data System (ADS)

    Arzhanov, V. P.; Borovich, Boris L.; Zuev, V. S.; Kazanskiĭ, V. M.; Katulin, V. A.; Kirillov, G. A.; Kormer, S. B.; Kuratov, Yu V.; Kuryapin, A. I.; Nosach, O. Yu; Sinitsyn, M. V.; Stoĭlov, Yu Yu

    1992-02-01

    The results are presented of investigations, performed in 1965-1966, of a pulsed photodissociation iodine laser utilizing CF3I and C3F7I molecules and pumped by light from a shock front created by detonating an explosive charge. Such lasers were found to possess a unique combination of high output energy and high pulse power. Two types of laser were studied. In one of them the active medium was pumped by light from a shock wave in xenon, and in the other a shock wave propagated through a mixture of the active medium and a rare gas. The energy characteristics of the second type of laser were found to be considerably higher than those of the first type. The laser pulse radiation enegy reached ~100 J with an average power of ~15 MW.

  20. The exhaust flow field of a rotating detonation-wave engine

    NASA Astrophysics Data System (ADS)

    Kailasanath, Kazhikathra; Schwer, Douglas

    2012-11-01

    Rotating detonation-wave engines (RDE) are a form of continuous detonation-wave engine. They potentially provide further gains than an intermittent or pulsed detonation-wave engine (PDE). However, significantly less work has been done on this concept when compared to the PDE. Last year, we presented the details of the injection system on the overall flow field in an RDE. In this talk, we focus on the effects of adding an exhaust plenum to this idealized RDE. While the overall exhaust flow shows that a recirculation zone sets up behind the RDE when a plenum is added, the net effect on the flow field within the RDE and on performance is found to be small. However, the slight modification to the flow field may impact the design of suitable nozzles for this device. This is explored further with the addition of a simple conical nozzle. This nozzle reduces the size of the recirculation zone and also reduces the temperature in the plume but has little effect on the flow field inside the RDE. Work Sponsored by ONR through NRL 6.1 Computational Physics Task Area.

  1. Synchronous initiation of optical detonators by Q-switched solid laser sources

    NASA Astrophysics Data System (ADS)

    Goujon, J.; Musset, O.; Marchand, A.; Bigot, C.

    2008-10-01

    The initiation of pyrotechnic substances by a laser light has been studied for more than 30 years. But until recently the use of this technology for defence applications encountered three main technical problems: the volume and the mass of lasers, the linear loss of optical fibres and their possible damage caused by the transport of strong laser power. Recent technical progress performed in the field of electrical and optical devices are now very promising for future opto-pyrotechnic functional chains. The objective of this paper is to present a demonstrator developed in order to initiate in a synchronous way four optical detonators and to measure the dispersion of their functioning times. It includes four compact Q-switched Nd:Cr:GSGG solid laser sources, pumped by flash lamp (energy ~110mJ, FWHM ~8.5 ns), two ultra-fast electro-optical selectors (based on RTP crystals) used to steer the laser beam and six optical fibre lines to transmit the laser pulses to the optical detonators. The set-up integrates also complex control and safety systems, as well as cameras allowing an optimal alignment of optical fibres. Experiments led us to initiate in a synchronous way four detonators with a mean scattering of 50 ns. The perspectives in this domain of initiation concern mainly the miniaturization and the hardening to the environments of electrical and optical components.

  2. The influence of detonation cell size and regularity on the propagation of gaseous detonations in granular materials

    NASA Astrophysics Data System (ADS)

    Slungaard, T.; Engebretsen, T.; Sønju, O. K.

    This paper presents results from an experimental study of transmission of gaseous detonation waves through various granular filters. Spherical glass beads of 4 and 8 mm diameter and crushed rock of 7.5 mm volume averaged diameter were used as filter material. Varying the initial pressure of the detonating gas mixture controlled the detonation cell size. Dilution with argon was used to vary the detonation cell regularity. The complete range from almost no detonation velocity deficit to complete extinction of the combustion wave was observed. The existing correlation for gaseous detonation velocity deficit V/VCJ = [1-0.35 (d_ c/dps)] +/- 0.1 where dc is the critical diameter for the gaseous detonation and dps is the pore size, is found to be applicable for both smooth spherical particles and irregular crushed rock when considering irregular detonation structures. Soot films and pressure measurements show that as the detonation cell size is increased, reinitiation of a reanular filter until it finally no longer occurs at V/VCJ ~ 0.4--0.45. Complete extinction of the combustion wave occurs at V/VCJ ~ 0.25--0.3. These two limits appear to be about the same for irregular and regular detonation cell structures. For irregular structures without argon dilution, dc/dps ~ 50 can be found for detonation wave failure, and dc/dps ~ 100 can be found for complete extinction of the combustion wave. For argon dilution these limits are changed to dc/dps ~ 10 and dc/dps ~ 40, respectively. The data are a bit scarce as a basis for proposing a new correlation for regular structures, but as a first approximation V/VCJ =[0.8--0.35log(dc/dps)] +/- 0.1 is suggested for regular structures. The detonation or combustion wave is found to approach a constant velocity in the granular filter if not extinguished.

  3. A membraneless air-breathing hydrogen biofuel cell based on direct wiring of thermostable enzymes on carbon nanotube electrodes.

    PubMed

    Lalaoui, Noémie; de Poulpiquet, Anne; Haddad, Raoudha; Le Goff, Alan; Holzinger, Michael; Gounel, Sébastien; Mermoux, Michel; Infossi, Pascale; Mano, Nicolas; Lojou, Elisabeth; Cosnier, Serge

    2015-05-01

    A biocathode was designed by the modification of a carbon nanotube (CNT) gas-diffusion electrode with bilirubin oxidase from Bacillus pumilus, achieving high current densities up to 3 mA cm(-2) for the reduction of O2 from air. A membraneless air-breathing hydrogen biofuel cell was designed by combination of this cathode with a functionalized CNT-based hydrogenase anode. PMID:25845356

  4. The delayed-detonation model of Type Ia supernovae. 2: The detonation phase

    NASA Technical Reports Server (NTRS)

    Arnett, David; Livne, Eli

    1994-01-01

    The investigation, by use of two-dimensional numerical hydrodynamics simulations, of the 'delayed detonation' mechanism of Khokhlov for the explosion of Type Ia supernovae is continued. Previously we found that the deflagration is insufficient to unbind the star. Expansion shuts off the flame; much of this small production of iron group nuclei occurs at lower densities, which reduces the electron-capture problem. Because the degenerate star has an adiabatic exponent only slightly above 4/3, the energy released by deflagration drives a pulsation of large amplitude. During the first expansion phase, adiabatic cooling shuts off the burning, and a Rayleigh-Taylor instability then gives mixing of high-entropy ashes with low-entropy fuel. During the first contraction phase, compressional heating reignites the material. The burning was allowed to develop into a detonation in these nonspherical models. The detonation grows toward spherical symmetry at late times. At these densities (rho approx. 10(exp 7) to 10(exp 8) g cm(exp -3)), either Ni-56 or nuclei of the Si-Ca group are the dominant products of the burning. The bulk yields are sensitive to the density of the star when the transition to detonation occurs. The relevance of the abundances, velocities, mixing, and total energy release to the theory and interpretation of Type Ia supernovae is discussed.

  5. I(sup STAR), NASA's Next Step in Air-Breathing Propulsion for Space Access

    NASA Technical Reports Server (NTRS)

    Hutt, John J.; McArthur, Craig; Cook, Stephen (Technical Monitor)

    2001-01-01

    The United States' National Aeronautics and Space Administration (NASA) has established a strategic plan for future activities in space. A primary goal of this plan is to make drastic improvements in the cost and safety of earth to low-earth-orbit transportation. One approach to achieving this goal is through the development of highly reusable, highly reliable space transportation systems analogous to the commercial airline system. In the year 2000, NASA selected the Rocket Based Combined Cycle (RBCC) engine as the next logical step towards this goal. NASA will develop a complete flight-weight, pump-fed engine system under the Integrated System Test of an Airbreathing Rocket (I(sup STAR)) Project. The objective of this project is develop a reusable engine capable of self-powering a vehicle through the air-augmented rocket, ramjet and scramjet modes required in all RBCC based operational vehicle concepts. The project is currently approved and funded to develop the engine through ground test demonstration. Plans are in place to proceed with flight demonstration pending funding approval. The project is in formulation phase and the Preliminary Requirements Review has been completed. The engine system and vehicle have been selected at the conceptual level. The I(sup STAR) engine concept is based on an air-breathing flowpath downselected from three configurations evaluated in NASA's Advanced Reusable Technology contract. The selected flowpath features rocket thrust chambers integrated into struts separating modular flowpath ducts, a variable geometry inlet, and a thermally choked throat. The engine will be approximately 220 inches long and 79 inches wide and fueled with a hydrocarbon fuel using liquid oxygen as the primary oxidizer candidate. The primary concept for the pump turbine drive is pressure-fed catalyzed hydrogen peroxide. In order to control costs, the flight demonstration vehicle will be launched from a B-52 aircraft. The vehicle concept is based on the Air

  6. SN 2012hn: a tidal detonation event?

    NASA Astrophysics Data System (ADS)

    Maccarone, Thomas

    2013-09-01

    We propose for a 30 kilosecond observation of SN 2102hn, a Ca-rich gap transient. These enigmatic objects, with properties intermediate between those of classical novae and Type Ia supernovae, can be well-explained by tidal detonations of low mass white dwarfs by intermediate mass black holes. In such a case, fall-back accretion of the tidal debris would power an X-ray source for which we propose to search. Because supermassive black holes will swallow white dwarfs whole, a successful outcome to this proposal would both explain the Ca-rich gap transients *and* establish the existence of intermediate mass black holes.

  7. Shock-to-Detonation Transition simulations

    SciTech Connect

    Menikoff, Ralph

    2015-07-14

    Shock-to-detonation transition (SDT) experiments with embedded velocity gauges provide data that can be used for both calibration and validation of high explosive (HE) burn models. Typically, a series of experiments is performed for each HE in which the initial shock pressure is varied. Here we describe a methodology for automating a series of SDT simulations and comparing numerical tracer particle velocities with the experimental gauge data. Illustrative examples are shown for PBX 9502 using the HE models implemented in the xRage ASC code at LANL.

  8. Detonation Reaction Zones in Condensed Explosives

    SciTech Connect

    Tarver, C M

    2005-07-14

    Experimental measurements using nanosecond time resolved embedded gauges and laser interferometric techniques, combined with Non-Equilibrium Zeldovich--von Neumann--Doring (NEZND) theory and Ignition and Growth reactive flow hydrodynamic modeling, have revealed the average pressure/particle velocity states attained in reaction zones of self-sustaining detonation waves in several solid and liquid explosives. The time durations of these reaction zone processes is discussed for explosives based on pentaerythritol tetranitrate (PETN), nitromethane, octahydro-1,3,5,7-tetranitro-1,3,5,7-tetrazocine (HMX), triaminitrinitrobenzene(TATB) and trinitrotoluene (TNT).

  9. Impact waves and detonation. Part I

    NASA Technical Reports Server (NTRS)

    Becker, R

    1929-01-01

    Among the numerous thermodynamic and kinetic problems that have arisen in the application of the gaseous explosive reaction as a source of power in the internal combustion engine, the problem of the mode or way by which the transformation proceeds and the rate at which the heat energy is delivered to the working fluid became very early in the engine's development a problem of prime importance. The work of Becker here given is a notable extension of earlier investigations, because it covers the entire range of the explosive reaction in gases - normal detonation and burning.

  10. DDT and detonation waves in dust-air mixtures

    NASA Astrophysics Data System (ADS)

    Zhang, F.; Grönig, H.; van de Ven, A.

    This paper summarizes the studies of DDT and stable detonation waves in dust-air mixtures at the Stosswellenlabor of RWTH Aachen. The DDT process and propagation mechanism for stable heterogeneous dust detonations in air are essentially the same as in the oxygen environment studied previously. The dust DDT process in tubes is composed of a reaction compression stage followed by a reaction shock stage as the pre-detonation process. The transverse waves that couple the shock wave and the chemical energy release are responsible for the propagation of a stable dust-air detonation. However, the transverse wave spacing of dust-air mixtures is much larger. Therefore, DDT and propagation of a stable detonation in most industrial and agricultural, combustible dust-air mixtures require a tube that has a large diameter between 0.1 m and 1 m and a sufficient length-diameter ratio beyond 100, when an appropriately strong initiation energy is used. Two dust detonation tubes, 0.14 m and 0.3 m in diameter, were used for observation of the above-mentioned results in cornstarch, anthraquinone and aluminum dust suspended in air. Smoked-foil technique was also used to measure the cellular structure of dust detonations in the 0.3 m detonation tube.

  11. Computation of a diverging Comp-B detonation

    SciTech Connect

    Bukiet, B.G.

    1989-01-01

    The expansion which occurs in diverging detonations weakens the wave and yields pressures and densities below those occurring in planar geometry. We study the problem of a spherically expanding detonation of Comp-B. The effect of varying the order of reaction as well as the rate law parameters (using an Arrhenius burn model) is studied. 14 refs., 3 figs.

  12. Gas-phase detonation propagation in mixture composition gradients.

    PubMed

    Kessler, D A; Gamezo, V N; Oran, E S

    2012-02-13

    The propagation of detonations through several fuel-air mixtures with spatially varying fuel concentrations is examined numerically. The detonations propagate through two-dimensional channels, inside of which the gradient of mixture composition is oriented normal to the direction of propagation. The simulations are performed using a two-component, single-step reaction model calibrated so that one-dimensional detonation properties of model low- and high-activation-energy mixtures are similar to those observed in a typical hydrocarbon-air mixture. In the low-activation-energy mixture, the reaction zone structure is complex, consisting of curved fuel-lean and fuel-rich detonations near the line of stoichiometry that transition to decoupled shocks and turbulent deflagrations near the channel walls where the mixture is extremely fuel-lean or fuel-rich. Reactants that are not consumed by the leading detonation combine downstream and burn in a diffusion flame. Detonation cells produced by the unstable reaction front vary in size across the channel, growing larger away from the line of stoichiometry. As the size of the channel decreases relative to the size of a detonation cell, the effect of the mixture composition gradient is lessened and cells of similar sizes form. In the high-activation-energy mixture, detonations propagate more slowly as the magnitude of the mixture composition gradient is increased and can be quenched in a large enough gradient. PMID:22213660

  13. Modelling of detonation cellular structure in aluminium suspensions

    NASA Astrophysics Data System (ADS)

    Briand, A.; Veyssiere, B.; Khasainov, B. A.

    2010-12-01

    Heterogeneous detonations involving aluminium suspensions have been studied for many years for industrial safety policies, and for military and propulsion applications. Owing to their weak detonability and to the lack of available experimental results on the detonation cellular structure, numerical simulations provide a convenient way to improve the knowledge of such detonations. One major difficulty arising in numerical study of heterogeneous detonations involving suspensions of aluminium particles in oxidizing atmospheres is the modelling of aluminium combustion. Our previous two-step model provided results on the effect on the detonation cellular structure of particle diameter and characteristic chemical lengths. In this study, a hybrid model is incorporated in the numerical code EFAE, combining both kinetic and diffusion regimes in parallel. This more realistic model provides good agreement with the previous two-step model and confirms the correlations found between the detonation cell width, and particle diameter and characteristic lengths. Moreover, the linear dependence found between the detonation cell width and the induction length remains valid with the hybrid model.

  14. 30 CFR 57.6402 - Deenergized circuits near detonators.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... Electric Blasting-Surface and Underground § 57.6402 Deenergized circuits near detonators. Electrical... 30 Mineral Resources 1 2011-07-01 2011-07-01 false Deenergized circuits near detonators. 57.6402 Section 57.6402 Mineral Resources MINE SAFETY AND HEALTH ADMINISTRATION, DEPARTMENT OF LABOR METAL...

  15. 30 CFR 56.6402 - Deenergized circuits near detonators.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 30 Mineral Resources 1 2011-07-01 2011-07-01 false Deenergized circuits near detonators. 56.6402... NONMETAL MINE SAFETY AND HEALTH SAFETY AND HEALTH STANDARDS-SURFACE METAL AND NONMETAL MINES Explosives Electric Blasting § 56.6402 Deenergized circuits near detonators. Electrical distribution circuits...

  16. 30 CFR 57.6402 - Deenergized circuits near detonators.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... Electric Blasting-Surface and Underground § 57.6402 Deenergized circuits near detonators. Electrical... 30 Mineral Resources 1 2010-07-01 2010-07-01 false Deenergized circuits near detonators. 57.6402 Section 57.6402 Mineral Resources MINE SAFETY AND HEALTH ADMINISTRATION, DEPARTMENT OF LABOR METAL...

  17. 30 CFR 56.6402 - Deenergized circuits near detonators.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 30 Mineral Resources 1 2010-07-01 2010-07-01 false Deenergized circuits near detonators. 56.6402... NONMETAL MINE SAFETY AND HEALTH SAFETY AND HEALTH STANDARDS-SURFACE METAL AND NONMETAL MINES Explosives Electric Blasting § 56.6402 Deenergized circuits near detonators. Electrical distribution circuits...

  18. 33 CFR 154.820 - Fire, explosion, and detonation protection.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... 33 Navigation and Navigable Waters 2 2012-07-01 2012-07-01 false Fire, explosion, and detonation... Systems § 154.820 Fire, explosion, and detonation protection. (a) A vapor control system with a single... oxygen analyzer that samples the vapor concentration continuously at a point not more than 6 meters...

  19. 33 CFR 154.820 - Fire, explosion, and detonation protection.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... 33 Navigation and Navigable Waters 2 2011-07-01 2011-07-01 false Fire, explosion, and detonation... Systems § 154.820 Fire, explosion, and detonation protection. (a) A vapor control system with a single... oxygen analyzer that samples the vapor concentration continuously at a point not more than 6 meters...

  20. 33 CFR 154.820 - Fire, explosion, and detonation protection.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 33 Navigation and Navigable Waters 2 2013-07-01 2013-07-01 false Fire, explosion, and detonation... Systems § 154.820 Fire, explosion, and detonation protection. (a) A vapor control system with a single... oxygen analyzer that samples the vapor concentration continuously at a point not more than 6 meters...

  1. 33 CFR 154.820 - Fire, explosion, and detonation protection.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 33 Navigation and Navigable Waters 2 2010-07-01 2010-07-01 false Fire, explosion, and detonation... Systems § 154.820 Fire, explosion, and detonation protection. (a) A vapor control system with a single... oxygen analyzer that samples the vapor concentration continuously at a point not more than 6 meters...

  2. Influence of cathode opening size and wetting properties of diffusion layers on the performance of air-breathing PEMFCs

    NASA Astrophysics Data System (ADS)

    Schmitz, A.; Tranitz, M.; Eccarius, S.; Weil, A.; Hebling, C.

    Air-breathing PEMFCs consist of an open cathodic side to allow an entirely passive supply of oxygen by diffusion. Furthermore, a large fraction of the produced water is removed by evaporation from the open cathode. Gas diffusion layers (GDLs) and the opening size of the cathode have a crucial influence on the performance of an air-breathing PEMFC. In order to assure an unobstructed supply of oxygen the water has to be removed efficiently and condensation in the GDL has to be avoided. On the other hand good humidification of the membrane has to be achieved to obtain high protonic conductivity. In this paper the influence of varying cathodic opening sizes (33%, 50% and 80% opening ratios) and of GDLs with different wetting properties are analysed. GDLs with hydrophobic and hydrophilic properties are prepared by coating of untreated GDLs (Toray ® carbon paper TGP-H-120, thickness of 350 μm). The air-breathing PEMFC test samples are realised using printed circuit board (PCB) technology. The cell samples were characterised over the entire potential range (0-0.95 V) by extensive measurements of the current density, the temperature and the cell impedance at 1 kHz. Additionally, measurements of the water balance were carried out at distinct operation points. The best cell performance was achieved with the largest opening ratio (80%) and an untreated GDL. At the maximum power point, this cell sample achieved a power density of 100 mW cm -2 at a moderate cell temperature of 43 °C. Furthermore, it could be shown that GDLs with hydrophilic or intense hydrophobic properties do not improve the performance of an air-breathing PEMFC. Based on the extensive characterisations, two design rules for air-breathing PEMFCs could be formulated. Firstly, it is crucial to maximise the cathode opening as far as an appropriate compression pressure of the cell assembly and therewith low contact resistance can be assured. Secondly, it is advantageous to use an untreated, slightly hydrophobic

  3. Energy Absorption Structure of Laser Supported Detonation Wave

    NASA Astrophysics Data System (ADS)

    Wang, Bin; Yamaguchi, Toshikazu; Hatai, Keigo; Komurasaki, Kimiya; Arakawa, Yoshihiro

    2010-05-01

    In Repetitive Pulsed (RP) laser propulsion, when the high energy laser beam is focused in the thruster, Laser Supported Detonation (LSD) wave is generated. This LSD wave converts the laser energy to the enthalpy of the blast wave, which will then apply impulse to the wall of the thruster. Therefore, the energy absorption structure and sustaining condition of LSD wave are important to be understood, which was still not clear though some visualized experiments have been conducted by Ushio et al. before. In this paper, 2-wavelength Mach-Zehnder interferometry is brought to investigate the electron density distribution of LSD area. At the same time, the temperature of the laser induced plasma is measured by an emission spectroscopy experiment, and calculated based on the assumption of local thermal equilibrium. The results show that in LSD, the electron density has a peak (as high as 2×1024[m-3]) behind the shock wave. The irradiated laser can be entirely absorbed before reaching the position of this peak. As a result, a new peak is always generating in front of the old one and this propagating has the same velocity as that of the blast wave. In this way, high heating ratio is sustained right after the shock front. However, as the laser pulse energy becomes lower, the propagating peak cannot catch up with the blast wave anymore, which leads to a termination of the LSD wave. From this study, it is found that for sustaining the LSD wave, a sufficiently thin laser absorption layer is necessary.

  4. Numerical analysis of thermonuclear detonation in dense plasma

    NASA Astrophysics Data System (ADS)

    Avronin, Y. N.; Bunatyan, A. A.; Gadzhiyev, A. D.; Mustafin, K. A.; Nurbakov, A. S.; Pisarev, V. N.; Feoktistov, L. P.; Frolov, V. D.; Shibarshov, L. I.

    1985-01-01

    The propagation of thermonuclear combustion from the region heated to thermonuclear temperatures by an external source to the remaining part of the target was investigated. The target was a tube of inert material (gold, lead, beryllium, etc.) filled with a deuterium-tritium mixture. It was determined analytically that thermonuclear combustion can propagate from a small portion of a nonspherical target to the remainder of the target and that a steady-state thermonuclear detonation wave can be formed. The role of various physical processes in thermonuclear detonation was investigated. Shock wave is the main mechanism underlying detonation propagation. The detonation rate and intensity of the thermonuclear reaction is influenced by the leakage of heat due to transvere heat conductivity. The critical diameter for thermonuclear detonation was determined approximately for a plasma filament encased in a housing with twice the density of the fuel.

  5. Multidimensional detonation propagation modeled via nonlinear shock wave superposition

    NASA Astrophysics Data System (ADS)

    Higgins, Andrew; Mehrjoo, Navid

    2010-11-01

    Detonation waves in gases are inherently multidimensional due to their cellular structure, and detonations in liquids and heterogeneous solids are often associated with instabilities and stochastic, localized reaction centers (i.e., hot spots). To explore the statistical nature of detonation dynamics in such systems, a simple model that idealizes detonation propagation as an ensemble of interacting blast waves originating from spatially random point sources has been proposed. Prior results using this model exhibited features that have been observed in real detonating systems, such as anomalous scaling between axisymmetric and two-dimensional geometries. However, those efforts used simple linear superposition of the blast waves. The present work uses a model of blast wave superposition developed for multiple-source explosions (the LAMB approximation) that incorporates the nonlinear interaction of shock waves analytically, permitting the effect of a more physical model of blast wave interaction to be explored. The results are suggestive of a universal behavior in systems of spatially randomized energy sources.

  6. Effect of prill structure on detonation performance of ANFO

    SciTech Connect

    Salyer, Terry R; Short, Mark; Kiyanda, Charles B; Morris, John S; Zimmerly, Tony

    2010-01-01

    While the effects of charge diameter, fuel mix ratio, and temperature on ANFO detonation performance are substantial, the effects of prill type are considerable as well as tailorable. Engineered AN prills provide a means to improve overall performance, primarily by changing the material microstructure through the addition of features designed to enhance hot spot action. To examine the effects of prill type (along with fuel mix ratio and charge diameter) on detonation performance, a series of precision, large-scale, ANFO front-curvature rate-stick tests was performed. Each shot used standard No. 2 diesel for the fuel oil and was essentially unconfined with cardboard confinement. Detonation velocities and front curvatures were measured while actively maintaining consistent shot temperatures. Based on the experimental results, DSD calibrations were performed to model the detonation performance over a range of conditions, and the overall effects of prill microstructure were examined and correlated with detonation performance.

  7. Light detonation wave in a cylindrical Z-pinch

    NASA Astrophysics Data System (ADS)

    Yusupaliev, U.; Sysoev, N. N.; Shuteev, S. A.; Elenskii, V. G.

    2015-09-01

    A secondary compression wave previously observed by other researchers in a cylindrical Z-pinch has been identified in this work as a light detonation wave. It appears on the inner surface of a discharge chamber under the action of the intense ultraviolet radiation from a plasma pinch at the stage of its maximum compression. The condition of the light detonation wave has been determined experimentally. The dependence of its Mach number on a generalized dimensionless variable has been determined taking into account the conservation laws for the light detonation wave including the pressure of the gas, expenses on the formation of the surface plasma, and the energy of ionization of the gas involved in the wave. An analogy with the laser-supported detonation wave created by intense laser radiation has been revealed. The indicated dependence is within the error of measurement in agreement with the experimental data for light detonation waves created by both methods.

  8. Detonation engine fed by acetylene-oxygen mixture

    NASA Astrophysics Data System (ADS)

    Smirnov, N. N.; Betelin, V. B.; Nikitin, V. F.; Phylippov, Yu. G.; Koo, Jaye

    2014-11-01

    The advantages of a constant volume combustion cycle as compared to constant pressure combustion in terms of thermodynamic efficiency has focused the search for advanced propulsion on detonation engines. Detonation of acetylene mixed with oxygen in various proportions is studied using mathematical modeling. Simplified kinetics of acetylene burning includes 11 reactions with 9 components. Deflagration to detonation transition (DDT) is obtained in a cylindrical tube with a section of obstacles modeling a Shchelkin spiral; the DDT takes place in this section for a wide range of initial mixture compositions. A modified ka-omega turbulence model is used to simulate flame acceleration in the Shchelkin spiral section of the system. The results of numerical simulations were compared with experiments, which had been performed in the same size detonation chamber and turbulent spiral ring section, and with theoretical data on the Chapman-Jouguet detonation parameters.

  9. Measuring In-Situ Mdf Velocity Of Detonation

    DOEpatents

    Horine, Frank M.; James, Jr., Forrest B.

    2005-10-25

    A system for determining the velocity of detonation of a mild detonation fuse mounted on the surface of a device includes placing the device in a predetermined position with respect to an apparatus that carries a couple of sensors that sense the passage of a detonation wave at first and second spaced locations along the fuse. The sensors operate a timer and the time and distance between the locations is used to determine the velocity of detonation. The sensors are preferably electrical contacts that are held spaced from but close to the fuse such that expansion of the fuse caused by detonation causes the fuse to touch the contact, causing an electrical signal to actuate the timer.

  10. On the Initiation Mechanism in Exploding Bridgewire and Laser Detonators

    NASA Astrophysics Data System (ADS)

    Stewart, D. Scott; Thomas, Keith A.; Clarke, S.; Mallett, H.; Martin, E.; Martinez, M.; Munger, A.; Saenz, Juan

    2006-07-01

    Since its invention by Los Alamos during the Manhattan Project era the exploding bridgewire detonator (EBW) has seen tremendous use and study. Recent development of a laser-powered device with detonation properties similar to an EBW is reviving interest in the basic physics of the deflagration-to-detonation (DDT) process in both of these devices. Cutback experiments using both laser interferometry and streak camera observations are providing new insight into the initiation mechanism in EBWs. These measurements are being correlated to a DDT model of compaction to detonation and shock to detonation developed previously by Xu and Stewart. The DDT model is incorporated into a high-resolution, multi-material model code for simulating the complete process. Model formulation and the modeling issues required to describe the test data will be discussed.

  11. Jaguar Procedures for Detonation Behavior of Explosives Containing Boron

    NASA Astrophysics Data System (ADS)

    Stiel, L. I.; Baker, E. L.; Capellos, C.

    2009-12-01

    The Jaguar product library was expanded to include boron and boron containing products by analysis of Available Hugoniot and static volumetric data to obtain constants of the Murnaghan relationships for the components. Experimental melting points were also utilized to obtain the constants of the volumetric relationships for liquid boron and boron oxide. Detonation velocities for HMX—boron mixtures calculated with these relationships using Jaguar are in closer agreement with literature values at high initial densities for inert (unreacted) boron than with the completely reacted metal. These results indicate that the boron does not react near the detonation front or that boron mixtures exhibit eigenvalue detonation behavior (as shown by some aluminized explosives), with higher detonation velocities at the initial points. Analyses of calorimetric measurements for RDX—boron mixtures indicate that at high boron contents the formation of side products, including boron nitride and boron carbide, inhibits the detonation properties of the formulation.

  12. Detonation characteristics of dimethyl ether and ethanol-air mixtures

    NASA Astrophysics Data System (ADS)

    Diakow, P.; Cross, M.; Ciccarelli, G.

    2015-05-01

    The detonation cell structure in dimethyl ether vapor and ethanol vapor-air mixtures was measured at atmospheric pressure and initial temperatures in the range of 293-373 K. Tests were carried out in a 6.2-m-long, 10-cm inner diameter tube. For more reactive mixtures, a series of orifice plates were used to promote deflagration-to-detonation transition in the first half of the tube. For less reactive mixtures prompt detonation initiation was achieved with an acetylene-oxygen driver. The soot foil technique was used to capture the detonation cell structure. The measured cell size was compared to the calculated one-dimensional detonation reaction zone length. For fuel-rich dimethyl ether mixtures the calculated reaction zone is highlighted by a temperature gradient profile with two maxima, i.e., double heat release. The detonation cell structure was interpreted as having two characteristic sizes over the full range of mixture compositions. For mixtures at the detonation propagation limits the large cellular structure approached a single-head spin, and the smaller cells approached the size of the tube diameter. There is little evidence to support the idea that the two cell sizes observed on the foils are related to the double heat release predicted for the rich mixtures. There was very little influence of initial temperature on the cell size over the temperature range investigated. A double heat release zone was not predicted for ethanol-air detonations. The detonation cell size for stoichiometric ethanol-air was found to be similar to the size of the small cells for dimethyl ether. The measured cell size for ethanol-air did not vary much with composition in the range of 30-40 mm. For mixtures near stoichiometric it was difficult to discern multiple cell sizes. However, near the detonation limits there was strong evidence of a larger cell structure similar to that observed in dimethyl ether air mixtures.

  13. Characterization of detonator performance using photonic Doppler velocimetry

    NASA Astrophysics Data System (ADS)

    Maisey, Matthew P.; Bowden, Mike D.

    2008-08-01

    Detonators are used to convert electrical or other energy into an explosive output. This output can then be used to initiate further explosive charges. To aid in the development of explosive systems, it is important to characterize the output of detonators, in particularly the pressure produced. Recent advances over the last five years in high-speed digitizing oscilloscopes and high-bandwidth photodiodes, driven primarily by the telecommunications industry, have enabled the development of a new type of interferometer for measuring high velocities, such as those found in detonics experiments. The Photonic Doppler Velocimeter (PDV) can be visualized as a fiber-based Michelson interferometer. The light from a single-mode fiber laser at 1550 nm is passed through a circulator, which acts to separate bi-directional light. The beam is then reflected via free-space optics off the surface of interest, and then focused back into the same fiber. This reflected light is then mixed with an approximately equal amount of non-reflected light, and the resulting interference is recorded using a high-bandwidth photodiode and oscilloscope. In contrast to more traditional Velocimetry techniques such as VISAR, only a single data channel is required. We have used our PDV system to investigate the performance of optical and electrical detonators. The detonators examined are the commercially available RISI RP-80, and an AWE DOI (Direct Optical Initiation) detonator. The RP-80 is an exploding bridgewire (EBW) detonator, utilizing Pentaerythritol Tetranitrate as the initiating explosive and a RDX output pellet. The DOI detonator uses an aluminum flyer to initiate a Hexanitrostilbene (HNS) pellet. Both detonators are canned in aluminum and the velocity of the can was measured, and from this, the output pressure of the detonator has been determined. This is compared to calculated values.

  14. On the influence of low initial pressure and detonation stochastic nature on Mach reflection of gaseous detonation waves

    NASA Astrophysics Data System (ADS)

    Wang, C. J.; Guo, C. M.

    2014-09-01

    The two-dimensional, time-dependent and reactive Navier-Stokes equations were solved to obtain an insight into Mach reflection of gaseous detonation in a stoichiometric hydrogen-oxygen mixture diluted by 25 % argon. This mixture generates a mode-7 detonation wave under an initial pressure of 8.00 kPa. Chemical kinetics was simulated by an eight-species, forty-eight-reaction mechanism. It was found that a Mach reflection mode always occurs for a planar detonation wave or planar air shock wave sweeping over wedges with apex angles ranging from to . However, for cellular detonation waves, regular reflection always occurs first, which then transforms into Mach reflection. This phenomenon is more evident for detonations ignited under low initial pressure. Low initial pressure may lead to a curved wave front, that determines the reflection mode. The stochastic nature of boundary shape and transition distance, during deflagration-to-detonation transition, leads to relative disorder of detonation cell location and cell shape. Consequently, when a detonation wave hits the wedge apex, there appears a stochastic variation of triple point origin and variation of the angle between the triple point trajectory and the wedge surface. As the wedge apex angle increases, the distance between the triple point trajectory origin and the wedge apex increases, and the angle between the triple point trajectory and the wedge surface decreases exponentially.

  15. Bonfire-safe low-voltage detonator

    DOEpatents

    Lieberman, Morton L.

    1990-01-01

    A column of explosive in a low-voltage detonator which makes it bonfire-safe includes a first layer of an explosive charge of CP, or a primary explosive, and a second layer of a secondary organic explosive charge, such as PETN, which has a degradation temperature lower than the autoignition temperature of the CP or primary explosives. The first layer is composed of a pair of increments disposed in a bore of a housing of the detonator in an ignition region of the explosive column and adjacent to and in contact with an electrical ignition device at one end of the bore. The second layer is composed of a plurality of increments disposed in the housing bore in a transition region of the explosive column next to and in contact with the first layer on a side opposite from the ignition device. The first layer is loaded under a sufficient high pressure, 25 to 40 kpsi, to achieve ignition, whereas the second layer is loaded under a sufficient low pressure, about 10 kpsi, to allow occurrence of DDT. Each increment of the first and second layers has an axial length-to-diameter ratio of one-half.

  16. Bonfire-safe low-voltage detonator

    DOEpatents

    Lieberman, M.L.

    1988-07-01

    A column of explosive in a low-voltage detonator which makes it bonfire-safe includes a first layer of an explosive charge of CP, or a primary explosive, and a second layer of a secondary organic explosive charge, such as PETN, which has a degradation temperature lower than the autoignition temperature of the CP or primary explosives. The first layer is composed of a pair of increments disposed in a bore of a housing of the detonator in an ignition region of the explosive column and adjacent to and in contact with an electrical ignition device at one end of the bore. The second layer is composed of a plurality of increments disposed in the housing bore in a transition region of the explosive column next to and in contact with the first layer on a side opposite from the ignition device. The first layer is loaded under a sufficient high pressure, 25 to 40 kpsi, to achieve ignition, whereas the second layer is loaded under a sufficient low pressure, about 10 kpsi, to allow occurrence of DDT. Each increment of the first and second layers has an axial length-to-diameter ratio of one-half. 2 figs.

  17. Spark-safe low-voltage detonator

    DOEpatents

    Lieberman, Morton L.

    1989-01-01

    A column of explosive in a low-voltage detonator which makes it spark-safe ncludes an organic secondary explosive charge of HMX in the form of a thin pad disposed in a bore of a housing of the detonator in an ignition region of the explosive column and adjacent to an electrical ignition device at one end of the bore. The pad of secondary charge has an axial thickness within the range of twenty to thirty percent of its diameter. The explosive column also includes a first explosive charge of CP disposed in the housing bore in the ignition region of the explosive column next to the secondary charge pad on a side opposite from the ignition device. The first CP charge is loaded under sufficient pressure, 25 to 40 kpsi, to provide mechanical confinement of the pad of secondary charge and physical coupling thereof with the ignition device. The explosive column further includes a second explosive charge of CP disposed in the housing bore in a transition region of the explosive column next to the first CP charge on a side opposite from the pad of secondary charge. The second CP charge is loaded under sufficient pressure, about 10 kpsi, to allow occurrence of DDT. The first explosive CP charge has an axial thickness within the range of twenty to thirty percent of its diameter, whereas the second explosive CP charge contains a series of increments (nominally 4) each of which has an axial thickness-to-diameter ratio of one to two.

  18. Numerical solution of under-resolved detonations

    NASA Astrophysics Data System (ADS)

    Tosatto, Luca; Vigevano, Luigi

    2008-02-01

    A new fractional-step method is proposed for the numerical solution of high speed reacting flows, where the chemical time scales are often much smaller than the fluid dynamical time scales. When the problem is stiff, because of insufficient spatial/temporal resolution, a well-known spurious numerical phenomenon occurs in standard finite volume schemes: the incorrect calculation of the speed of propagation of discontinuities. The new method is first illustrated considering a one-dimensional scalar hyperbolic advection/reaction equation with stiff source term, which may be considered as a model problem to under-resolved detonations. During the reaction step, the proposed scheme replaces the cell average representation with a two-value reconstruction, which allows us to locate the discontinuity position inside the cell during the computation of the source term. This results in the correct propagation of discontinuities even in the stiff case. The method is proved to be second-order accurate for smooth solutions of scalar equations and is applied successfully to the solution of the one-dimensional reactive Euler equations for Chapman-Jouguet detonations.

  19. The dynamics of unsteady detonation in ozone

    SciTech Connect

    Aslam, Tariq D; Powers, Joseph M

    2008-01-01

    An ultra-fine, sub-micron discrete grid is used to capture the unsteady dynamics of a one-dimensional detonation in an inviscid O - O{sub 2} - O{sub 3} mixture. The ultra-fine grid is necessary to capture the length scales revealed by a complementary analysis of the steady detonation wave structure. For the unsteady calculations, shock-fitting coupled with a high order spatio-temporal discretization scheme combine to render numerical corruption negligible. As a result, mathematically verified solutions for a mixture initially of all O{sub 3} at one atmosphere and 298.15 K have been obtained; the solutions are converging at a rate much faster than the sub-first order convergence rate of all shock-capturing schemes. Additionally, the model has been validated against limited experimental data. Transient calculations show that strongly overdriven waves are stable and moderately overdriven waves unstable. New limit cycle behavior is revealed, and the first high resolution bifurcation diagram for etonation with detailed kinetics is found.

  20. Prompt Reaction of Aluminum in Detonating Explosives

    NASA Astrophysics Data System (ADS)

    Sandusky, Harold

    2005-07-01

    The potential of aluminum reaction to boost detonation energy has been studied for decades, most recently spurred by the availability of nanometer-sized particles. A review of the literature provides perspective for a recent study with the small-scale shock reactivity test. In this test, <1/2-g samples in confinement are shock loaded on one end, and the output at the other end dents a soft witness block. One test series had 0.3 g of HMX mixed with various forms of aluminum added in amounts of up to 25% of the total sample mass, with the deepest dent for H-5 aluminum occuring at 15%. Test results on ammonium perchlorate mixed with H-5 aluminum were consistent with the peak in detonation velocity reported in Combustion and Flame by Price in 1973 on similar mixtures. One outcome of this study is a new interpretation for the participation of aluminum in large scale gap tests on plastic-bonded explosives, which was discussed by Bernecker at this meeting in 1987.

  1. Propulsion integration of hypersonic air-breathing vehicles utilizing a top-down design methodology

    NASA Astrophysics Data System (ADS)

    Kirkpatrick, Brad Kenneth

    In recent years, a focus of aerospace engineering design has been the development of advanced design methodologies and frameworks to account for increasingly complex and integrated vehicles. Techniques such as parametric modeling, global vehicle analyses, and interdisciplinary data sharing have been employed in an attempt to improve the design process. The purpose of this study is to introduce a new approach to integrated vehicle design known as the top-down design methodology. In the top-down design methodology, the main idea is to relate design changes on the vehicle system and sub-system level to a set of over-arching performance and customer requirements. Rather than focusing on the performance of an individual system, the system is analyzed in terms of the net effect it has on the overall vehicle and other vehicle systems. This detailed level of analysis can only be accomplished through the use of high fidelity computational tools such as Computational Fluid Dynamics (CFD) or Finite Element Analysis (FEA). The utility of the top-down design methodology is investigated through its application to the conceptual and preliminary design of a long-range hypersonic air-breathing vehicle for a hypothetical next generation hypersonic vehicle (NHRV) program. System-level design is demonstrated through the development of the nozzle section of the propulsion system. From this demonstration of the methodology, conclusions are made about the benefits, drawbacks, and cost of using the methodology.

  2. Biannular Airbreathing Nozzle Rig (BANR) facility checkout and plug nozzle performance test data

    NASA Astrophysics Data System (ADS)

    Cummings, Chase B.

    2010-09-01

    The motivation for development of a supersonic business jet (SSBJ) platform lies in its ability to create a paradigm shift in the speed and reach of commercial, private, and government travel. A full understanding of the performance capabilities of exhaust nozzle configurations intended for use in potential SSBJ propulsion systems is critical to the design of an aircraft of this type. Purdue University's newly operational Biannular Airbreathing Nozzle Rig (BANR) is a highly capable facility devoted to the testing of subscale nozzles of this type. The high accuracy, six-axis force measurement system and complementary mass flowrate measurement capabilities of the BANR facility make it rather ideally suited for exhaust nozzle performance appraisal. Detailed accounts pertaining to methods utilized in the proper checkout of these diagnostic capabilities are contained herein. Efforts to quantify uncertainties associated with critical BANR test measurements are recounted, as well. Results of a second hot-fire test campaign of a subscale Gulfstream Aerospace Corporation (GAC) axisymmetric, shrouded plug nozzle are presented. Determined test article performance parameters (nozzle thrust efficiencies and discharge coefficients) are compared to those of a previous test campaign and numerical simulations of the experimental set-up. Recently acquired data is compared to published findings pertaining to plug nozzle experiments of similar scale and operating range. Suggestions relating to the future advancement and improvement of the BANR facility are provided. Lessons learned with regards to test operations and calibration procedures are divulged in an attempt to aid future facility users, as well.

  3. Air-breathing direct formic acid microfluidic fuel cell with an array of cylinder anodes

    NASA Astrophysics Data System (ADS)

    Zhu, Xun; Zhang, Biao; Ye, Ding-Ding; Li, Jun; Liao, Qiang

    2014-02-01

    An air-breathing direct formic acid membraneless microfluidic fuel cell using graphite cylinder arrays as the anode is proposed. The three dimensional anode volumetrically extends the reactive surface area and improves fuel utilization. The effects of spacer configuration, fuel and electrolyte concentration as well as reactant flow rate on the species transport and cell performance are investigated. The dynamic behavior of generated CO2 bubbles is visualized and its effect on current generation is discussed. The results show that the absence of two spacers adjacent to the cathode surface improves the cell performance by reducing the proton transfer resistance. The CO2 gas bubbles are constrained within the anode array and expelled by the fluid flow periodically. Proper reactant concentration and flow rate are crucial for cell operation. At optimum conditions, a maximum current density of 118.3 mA cm-3 and a peak power density of 21.5 mW cm-3 are obtained. In addition, benefit from the volumetrically stacked anodes and enhanced fuel transfer, the maximum single pass fuel utilization rate reaches up to 87.6% at the flow rate of 1 mL h-1.

  4. Computational modeling of alkaline air-breathing microfluidic fuel cells with an array of cylinder anodes

    NASA Astrophysics Data System (ADS)

    Ye, Ding-Ding; Zhang, Biao; Zhu, Xun; Sui, Pang-Chieh; Djilali, Ned; Liao, Qiang

    2015-08-01

    A three-dimensional computational model is developed for an alkaline air-breathing microfluidic fuel cell (AMFC) with an array of cylinder anodes. The model is validated against experimental data from an in-house prototype AMFC. The distributions of fluid velocity, fuel concentration and current density of the fuel cell are analyzed in detail. The effect of reactant flow rate on the cell performance and electrode potentials is also studied. The model results suggest that fuel crossover is minimized by the fast electrolyte flow in the vicinity of the cathode. The current production of each anode is uneven and is well correlated with internal ohmic resistance. Fuel transfer limitation occurs at low flow rates (<100 μL min-1) but diminishes at high flow rates. The model results also indicate that cathode potential reversal takes place at combined low flow rate and high current density conditions, mainly due to the improved overpotential downstream where fuel starvation occurs. The anode reaction current distribution is found to be relatively uniform, which is a result of a compensating mechanism that improves the current production of the bottom anodes downstream.

  5. Uncertainty analysis and robust trajectory linearization control of a flexible air-breathing hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Pu, Zhiqiang; Tan, Xiangmin; Fan, Guoliang; Yi, Jianqiang

    2014-08-01

    Flexible air-breathing hypersonic vehicles feature significant uncertainties which pose huge challenges to robust controller designs. In this paper, four major categories of uncertainties are analyzed, that is, uncertainties associated with flexible effects, aerodynamic parameter variations, external environmental disturbances, and control-oriented modeling errors. A uniform nonlinear uncertainty model is explored for the first three uncertainties which lumps all uncertainties together and consequently is beneficial for controller synthesis. The fourth uncertainty is additionally considered in stability analysis. Based on these analyses, the starting point of the control design is to decompose the vehicle dynamics into five functional subsystems. Then a robust trajectory linearization control (TLC) scheme consisting of five robust subsystem controllers is proposed. In each subsystem controller, TLC is combined with the extended state observer (ESO) technique for uncertainty compensation. The stability of the overall closed-loop system with the four aforementioned uncertainties and additional singular perturbations is analyzed. Particularly, the stability of nonlinear ESO is also discussed from a Liénard system perspective. At last, simulations demonstrate the great control performance and the uncertainty rejection ability of the robust scheme.

  6. The Sensitivity of Precooled Air-Breathing Engine Performance to Heat Exchanger Design Parameters

    NASA Astrophysics Data System (ADS)

    Webber, H.; Bond, A.; Hempsell, M.

    The issues relevant to propulsion design for Single Stage To Orbit (SSTO) vehicles are considered. In particular two air- breathing engine concepts involving precooling are compared; SABRE (Synergetic Air-Breathing and Rocket Engine) as designed for the Skylon SSTO launch vehicle, and a LACE (Liquid Air Cycle Engine) considered in the 1960's by the Americans for an early generation spaceplane. It is shown that through entropy minimisation the SABRE has made substantial gains in performance over the traditional LACE precooled engine concept, and has shown itself as the basis of a viable means of realising a SSTO vehicle. Further, it is demonstrated that the precooler is a major source of thermodynamic irreversibility within the engine cycle and that further reduction in entropy can be realised by increasing the heat transfer coefficient on the air side of the precooler. If this were to be achieved, it would improve the payload mass delivered to orbit by the Skylon launch vehicle by between 5 and 10%.

  7. Migration history of air-breathing fishes reveals Neogene atmospheric circulation patterns

    NASA Astrophysics Data System (ADS)

    Böhme, M.

    2004-05-01

    The migration history of an air-breathing fish group (Channidae; snakehead fishes) is used for reconstructing Neogene Eurasian precipitation and atmospheric circulation patterns. The study shows that snakeheads are sensitive indicators of summer precipitation maxima in subtropical and temperate regions, and are present regularly if the wettest month exceeds 150 mm precipitation and 20 °C mean temperature. The analysis of 515 fossil freshwater fish deposits of the past 50 m.y. from Africa and Eurasia shows two continental-scale migration events from the snakeheads' center of origin in the south Himalayan region, events that can be related to changes in the Northern Hemisphere circulation pattern. The first migration, ca. 17.5 Ma, into western and central Eurasia may have been caused by a northward shift of the Intertropical Convergence Zone that brought western Eurasia under the influence of trade winds that produced a zonal and meridional precipitation gradient in Europe. During the second migration, between 8 and 4 Ma, into Africa and East Asia, snakeheads reached their present-day distribution. This migration could have been related to the intensification of the Asian monsoon that brought summer precipitation to their migratory pathways in East Africa Arabia and East Asia.

  8. Affordable Flight Demonstration of the GTX Air-Breathing SSTO Vehicle Concept

    NASA Technical Reports Server (NTRS)

    Krivanek, Thomas M.; Roche, Joseph M.; Riehl, John P.; Kosareo, Daniel N.

    2003-01-01

    The rocket based combined cycle (RBCC) powered single-stage-to-orbit (SSTO) reusable launch vehicle has the potential to significantly reduce the total cost per pound for orbital payload missions. To validate overall system performance, a flight demonstration must be performed. This paper presents an overview of the first phase of a flight demonstration program for the GTX SSTO vehicle concept. Phase 1 will validate the propulsion performance of the vehicle configuration over the supersonic and hypersonic air- breathing portions of the trajectory. The focus and goal of Phase 1 is to demonstrate the integration and performance of the propulsion system flowpath with the vehicle aerodynamics over the air-breathing trajectory. This demonstrator vehicle will have dual mode ramjetkcramjets, which include the inlet, combustor, and nozzle with geometrically scaled aerodynamic surface outer mold lines (OML) defining the forebody, boundary layer diverter, wings, and tail. The primary objective of this study is to demon- strate propulsion system performance and operability including the ram to scram transition, as well as to validate vehicle aerodynamics and propulsion airframe integration. To minimize overall risk and develop ment cost the effort will incorporate proven materials, use existing turbomachinery in the propellant delivery systems, launch from an existing unmanned remote launch facility, and use basic vehicle recovery techniques to minimize control and landing requirements. A second phase would demonstrate propulsion performance across all critical portions of a space launch trajectory (lift off through transition to all-rocket) integrated with flight-like vehicle systems.

  9. Navier-Stokes predictions of dynamic stability derivatives for air-breathing hypersonic vehicle

    NASA Astrophysics Data System (ADS)

    Liu, Xu; Liu, Wei; Zhao, Yunfei

    2016-01-01

    Dynamic derivatives are important parameters for designing vehicle trajectory and attitude control system that directly decide the divergence behavior of vibration of the aircraft open-loop system under interference. After calibration model validation, the dynamic behavior of air-breathing hypersonic vehicle WR-A is characterized. The unsteady flow field of aircraft forced simple harmonic vibration (SHV) is simulated using N-S equation. The direct damping derivatives, cross derivatives, acceleration derivatives and rotary derivatives of WR-A under different frequencies, amplitudes and positions of centroid are obtained. Research demonstrates that the proportion of acceleration derivatives, which represents the flow time lag effect, in the direct damping derivatives can be as high as 40% but is opposite to the damping derivative value symbols in some cases, contributing to dynamic instability. Numerical simulation on large-amplitude forced vibration of WR-A indicates that the aerodynamic behavior predicted by the dynamic derivative model agrees well with unsteady calculations. The inlet performance parameter derivatives are solved using the Etkin theory. The inlet performance parameters under large-amplitude vibration are successfully predicted using the dynamic derivative model. This offers a guideline for characterizing the dynamic internal flow field and unsteady inlet performance.

  10. Interactions between Flight Dynamics and Propulsion Systems of Air-Breathing Hypersonic Vehicles

    NASA Astrophysics Data System (ADS)

    Dalle, Derek J.

    The development and application of a first-principles-derived reduced-order model called MASIV (Michigan/AFRL Scramjet In Vehicle) for an air-breathing hypersonic vehicle is discussed. Several significant and previously unreported aspects of hypersonic flight are investigated. A fortunate coupling between increasing Mach number and decreasing angle of attack is shown to extend the range of operating conditions for a class of supersonic inlets. Detailed maps of isolator unstart and ram-to-scram transition are shown on the flight corridor map for the first time. In scram mode the airflow remains supersonic throughout the engine, while in ram mode there is a region of subsonic flow. Accurately predicting the transition between these two modes requires models for complex shock interactions, finite-rate chemistry, fuel-air mixing, pre-combustion shock trains, and thermal choking, which are incorporated into a unified framework here. Isolator unstart occurs when the pre-combustion shock train is longer than the isolator, which blocks airflow from entering the engine. Finally, cooptimization of the vehicle design and trajectory is discussed. An optimal control technique is introduced that greatly reduces the number of computations required to optimize the simulated trajectory.

  11. Design Evolution and Performance Characterization of the GTX Air-Breathing Launch Vehicle Inlet

    NASA Technical Reports Server (NTRS)

    DeBonis, J. R.; Steffen, C. J., Jr.; Rice, T.; Trefny, C. J.

    2002-01-01

    The design and analysis of a second version of the inlet for the GTX rocket-based combine-cycle launch vehicle is discussed. The previous design did not achieve its predicted performance levels due to excessive turning of low-momentum comer flows and local over-contraction due to asymmetric end-walls. This design attempts to remove these problems by reducing the spike half-angle to 10- from 12-degrees and by implementing true plane of symmetry end-walls. Axisymmetric Reynolds-Averaged Navier-Stokes simulations using both perfect gas and real gas, finite rate chemistry, assumptions were performed to aid in the design process and to create a comprehensive database of inlet performance. The inlet design, which operates over the entire air-breathing Mach number range from 0 to 12, and the performance database are presented. The performance database, for use in cycle analysis, includes predictions of mass capture, pressure recovery, throat Mach number, drag force, and heat load, for the entire Mach range. Results of the computations are compared with experimental data to validate the performance database.

  12. Evaluation of an Ejector Ramjet Based Propulsion System for Air-Breathing Hypersonic Flight

    NASA Technical Reports Server (NTRS)

    Thomas, Scott R.; Perkins, H. Douglas; Trefny, Charles J.

    1997-01-01

    A Rocket Based Combined Cycle (RBCC) engine system is designed to combine the high thrust to weight ratio of a rocket along with the high specific impulse of a ramjet in a single, integrated propulsion system. This integrated, combined cycle propulsion system is designed to provide higher vehicle performance than that achievable with a separate rocket and ramjet. The RBCC engine system studied in the current program is the Aerojet strutjet engine concept, which is being developed jointly by a government-industry team as part of the Air Force HyTech program pre-PRDA activity. The strutjet is an ejector-ramjet engine in which small rocket chambers are embedded into the trailing edges of the inlet compression struts. The engine operates as an ejector-ramjet from takeoff to slightly above Mach 3. Above Mach 3 the engine operates as a ramjet and transitions to a scramjet at high Mach numbers. For space launch applications the rockets would be re-ignited at a Mach number or altitude beyond which air-breathing propulsion alone becomes impractical. The focus of the present study is to develop and demonstrate a strutjet flowpath using hydrocarbon fuel at up to Mach 7 conditions.

  13. Correlation and analysis of oil flow data for an air-breathing missile model

    NASA Technical Reports Server (NTRS)

    Stoy, S. L.; Dillon, J. L.; Roman, A. P.

    1985-01-01

    This paper will present the results of an oil flow investigation on an airbreathing missile model. This oil flow study examined the flow around the model, which can be configured with both axisymmetric and two-dimensional inlets. Flow visualization analyses were conducted for both types of geometries by examining the surface flow patterns made visible by the oil flows for Mach numbers of 2.5 and 3.95. The analysis has shown the extent of flow spillage around the inlet which has helped explain the force and moment data collected during previous testing of the model. The oil flow data has also been used to develop guidelines for modeling the location of the crossflow separation line along inlet fairings. Finally, the oil flow analysis has been used to identify unique features of the boattail flow. These boattail flow characteristics have been correlated with previous oil flow analysis of noncircular body models. This paper demonstrates the use of this type of oil flow analysis in developing missile flow field analysis and aerodynamic predictions ranging from impact angle methods through Navier-Stokes methods.

  14. Hypersonic Airbreathing Propulsion: An Aerodynamics, Aerothermodynamics, and Acoustics Competency White Paper

    NASA Technical Reports Server (NTRS)

    Drummond, J. Philip; Cockrell, Charles E., Jr.; Pellett, Gerald L.; Diskin, Glenn S.; Auslender, Aaron H.; Exton, Reginald J.; Guy, R. Wayne; Hoppe, John C.; Puster, Richard L.; Rogers, R. Clayton

    2002-01-01

    This White Paper examines the current state of Hypersonic Airbreathing Propulsion at the NASA Langley Research Center and the factors influencing this area of work and its personnel. Using this knowledge, the paper explores beyond the present day and suggests future directions and strategies for the field. Broad views are first taken regarding potential missions and applications of hypersonic propulsion. Then, candidate propulsion systems that may be applicable to these missions are suggested and discussed. Design tools and experimental techniques for developing these propulsion systems are then described, and approaches for applying them in the design process are considered. In each case, current strategies are reviewed and future approaches that may improve the techniques are considered. Finally, the paper concentrates on the needs to be addressed in each of these areas to take advantage of the opportunities that lay ahead for both the NASA Langley Research Center and the Aerodynamic Aerothermodynamic, and Aeroacoustics Competency. Recommendations are then provided so that the goals set forth in the paper may be achieved.

  15. Aerospace applications of pulsed plasmas

    NASA Astrophysics Data System (ADS)

    Starikovskiy, Andrey

    2012-10-01

    The use of a thermal equilibrium plasma for combustion control dates back more than a hundred years to the advent of internal combustion (IC) engines and spark ignition systems. The same principles are still applied today to achieve high efficiency in various applications. Recently, the potential use of nonequilibrium plasma for ignition and combustion control has garnered increasing interest due to the possibility of plasma-assisted approaches for ignition and flame stabilization. During the past decade, significant progress has been made toward understanding the mechanisms of plasma chemistry interactions, energy redistribution and the nonequilibrium initiation of combustion. In addition, a wide variety of fuels have been examined using various types of discharge plasmas. Plasma application has been shown to provide additional combustion control, which is necessary for ultra-lean flames, high-speed flows, cold low-pressure conditions of high-altitude gas turbine engine (GTE) relight, detonation initiation in pulsed detonation engines (PDE) and distributed ignition control in homogeneous charge-compression ignition (HCCI) engines, among others. The present paper describes the current understanding of the nonequilibrium excitation of combustible mixtures by electrical discharges and plasma-assisted ignition and combustion. Nonequilibrium plasma demonstrates an ability to control ultra-lean, ultra-fast, low-temperature flames and appears to be an extremely promising technology for a wide range of applications, including aviation GTEs, piston engines, ramjets, scramjets and detonation initiation for pulsed detonation engines. To use nonequilibrium plasma for ignition and combustion in real energetic systems, one must understand the mechanisms of plasma-assisted ignition and combustion and be able to numerically simulate the discharge and combustion processes under various conditions.

  16. Developmental transcriptome analysis and identification of genes involved in formation of intestinal air-breathing function of Dojo loach, Misgurnus anguillicaudatus.

    PubMed

    Luo, Weiwei; Cao, Xiaojuan; Xu, Xiuwen; Huang, Songqian; Liu, Chuanshu; Tomljanovic, Tea

    2016-01-01

    Dojo loach, Misgurnus anguillicaudatus is a freshwater fish species of the loach family Cobitidae, using its posterior intestine as an accessory air-breathing organ. Little is known about the molecular regulatory mechanisms in the formation of intestinal air-breathing function of M. anguillicaudatus. Here high-throughput sequencing of mRNAs was performed from six developmental stages of posterior intestine of M. anguillicaudatus: 4-Dph (days post hatch) group, 8-Dph group, 12-Dph group, 20-Dph group, 40-Dph group and Oyd (one-year-old) group. These six libraries were assembled into 81300 unigenes. Totally 40757 unigenes were annotated. Subsequently, 35291 differentially expressed genes (DEGs) were scanned among different developmental stages and clustered into 20 gene expression profiles. Finally, 15 key pathways and 25 key genes were mined, providing potential targets for candidate gene selection involved in formation of intestinal air-breathing function in M. anguillicaudatus. This is the first report of developmental transcriptome of posterior intestine in M. anguillicaudatus, offering a substantial contribution to the sequence resources for this species and providing a deep insight into the formation mechanism of its intestinal air-breathing function. This report demonstrates that M. anguillicaudatus is a good model for studies to identify and characterize the molecular basis of accessory air-breathing organ development in fish. PMID:27545457

  17. Developmental transcriptome analysis and identification of genes involved in formation of intestinal air-breathing function of Dojo loach, Misgurnus anguillicaudatus

    PubMed Central

    Luo, Weiwei; Cao, Xiaojuan; Xu, Xiuwen; Huang, Songqian; Liu, Chuanshu; Tomljanovic, Tea

    2016-01-01

    Dojo loach, Misgurnus anguillicaudatus is a freshwater fish species of the loach family Cobitidae, using its posterior intestine as an accessory air-breathing organ. Little is known about the molecular regulatory mechanisms in the formation of intestinal air-breathing function of M. anguillicaudatus. Here high-throughput sequencing of mRNAs was performed from six developmental stages of posterior intestine of M. anguillicaudatus: 4-Dph (days post hatch) group, 8-Dph group, 12-Dph group, 20-Dph group, 40-Dph group and Oyd (one-year-old) group. These six libraries were assembled into 81300 unigenes. Totally 40757 unigenes were annotated. Subsequently, 35291 differentially expressed genes (DEGs) were scanned among different developmental stages and clustered into 20 gene expression profiles. Finally, 15 key pathways and 25 key genes were mined, providing potential targets for candidate gene selection involved in formation of intestinal air-breathing function in M. anguillicaudatus. This is the first report of developmental transcriptome of posterior intestine in M. anguillicaudatus, offering a substantial contribution to the sequence resources for this species and providing a deep insight into the formation mechanism of its intestinal air-breathing function. This report demonstrates that M. anguillicaudatus is a good model for studies to identify and characterize the molecular basis of accessory air-breathing organ development in fish. PMID:27545457

  18. Graphene nanosheets produced via controlled detonation of hydrocarbons

    NASA Astrophysics Data System (ADS)

    Nepal, Arjun

    We demonstrated that gram quantities of pristine graphene nanosheets (GNs) can be produced via detonation of a hydrocarbon. This one-step and catalyst-free method is eco-friendly and economical for the production of GNs. The hydrocarbons detonated were C2H2, C 2H4, C3H8 and CH4 in the presence of O2. The carbon products obtained from the detonation were analyzed by XRD, TEM, XPS and Raman spectroscopy. Depending upon the ratio of O2 to C2H2, the GNs of size up to ˜ 250 nm, SSA up to ˜ 200 m2/g and yield up to 70% with 2-3 layers' stack have been obtained so far. N2O was determined as a good alternative to O2 as an oxidizer to produce GNs by detonating C2H2 with it. A two-color pyrometer was designed and calibrated to measure the temperature of the detonation of hydrocarbons. The measured detonation temperatures were in between 2700 K and 4300 K. Along with the high detonation temperature, the composition of precursor hydrocarbon was observed to be crucial as well to determine its suitability to detonate with oxidizer to produce GNs. The hydrocarbons C2H2 and C2H4 were determined as the suitable precursors to produce GNs whereas detonation of C3H8 yields mere amorphous carbon soot and CH4 gives no solid carbon while detonated with O2. It has been proposed that the hydrocarbons with C/H≥0.5 are suitable for GNs production by detonation method. Highly oxidized graphene nanosheets (OGNs) were produced by solution-based oxidation of GNs prepared via a controlled detonation of acetylene at O 2/C2H2=0.8. The produced OGNs were about 250 nm in size and hydrophilic in nature. The C/O ratio was dramatically reduced from 49:1 in the pristine GNs to about 1:1 in OGNs, as determined by X-ray photoelectron spectroscopy. This C/O in OGNs is the least ever found in all oxidized graphitic materials that have been reported. Thus, the OGNs produced from the detonated GNs with such high degree of oxidation herein yields a novel and promising material for future applications.

  19. 33 CFR 154.822 - Detonation arresters, flame arresters, and flame screens.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... 33 Navigation and Navigable Waters 2 2010-07-01 2010-07-01 false Detonation arresters, flame... BULK Vapor Control Systems § 154.822 Detonation arresters, flame arresters, and flame screens. (a) Each detonation arrester required by this part must: (1) Be capable of arresting a detonation from either side...

  20. On the Initiation Mechanism in Exploding Bridgewire and Laser Detonators

    NASA Astrophysics Data System (ADS)

    Stewart, D. Scott; Thomas, K.; Saenz, J.

    2005-07-01

    Since its invention by Los Alamos during the Manhattan Project era the exploding bridgewire detonator (EBW) has seen tremendous use and study. Recent development of a laser-powered device with detonation properties similar to an EBW is reviving interest in the basic physics of the Deflagration-to-Detonation (DDT) process in both of these devices,[1]. Cutback experiments using both laser interferometry and streak camera observations are providing new insight into the initiation mechanism in EBWs. These measurements are being correlated to a DDT model of compaction to detonation and shock to detonation developed previously by Xu and Stewart, [2]. The DDT model is incorporated into a high-resolution, multi-material model code for simulating the complete process. Model formulation and predictions against the test data will be discussed. REFS. [1] A. Munger, J. Kennedy, A. Akinci, and K. Thomas, "Dev. of a Laser Detonator" 30th Int. Pyrotechnics Seminar, Fort Collins, CO, (2004). [2] Xu, S. and Stewart, D. S. Deflagration to detonation transition in porous energetic materials: A model study. J. Eng. Math., 31, 143-172 (1997)