Pulse Detonation Engine as a Ramjet Replacement

A modular approach to the study of system performance and thermodynamic cycle efficiency of airbreathing pulse detonation engines (PDEs) is described. Each module represents a specific component of the engine, and its dynamic behavior is formulated using conservation laws in either one or two spatial dimensions. A framework is established for assessing quantitatively the influence of all known processes on engine dynamics. Various loss mechanisms limiting the PDE performance are identified. As a specific example, a supersonic PDE for high-altitude applications is studied comprehensively. The effects of chamber configuration and operating sequence on the engine propulsive efficiency are examined. The results demonstrate the existence of an optimum cycle frequency and valve close-up time for achieving maximum performance in terms of thrust and specific impulse. Furthermore, a choked convergent-divergent nozzle is required to render the PDE competitive with other airbreathing propulsion systems, such as gas-turbine and ramjet engines.

Nomenclature A e — nozzle exit area A t = nozzle throat area C* = characteristic velocity g = gravitational acceleration / = delivered impulse Isp — specific impulse Isp.d = delivered specific impulse m = propellant mass 777 = nozzle flow rate p = pressure p m = mass averaged pressure p ( = time averaged pressure p r = mean pressure for burning rate p = mean pressure r = burning rate / = time ta = action time tb = burn time /7 =time when p first becomes 0.05 pt w_ = propellant web ( ) = denotes a mean e — nozzle expansion ratio

Development of Transient Pulse Detonation Engine Cycle Analysis and Performance Prediction (PDE-CAPP) Code

We developed a rotary valve for a pulse detonation engine (PDE), and confirmed its basic characteristics and performance by the visualization experiment and the thrust measurement. We visualized a multi-shot of a pulse detonation rocket engine (PDRE) cycle at an operation frequency of 160 Hz by using a high-speed camera (time resolution: 3.33 μsec, space resolution: 0.4 mm), a schlieren method and square cross-section combustor. We confirmed the each process of one cycle of a PDE (propellant filling, detonation wave generation, and burned gas blowdown and purge process) with high time relolution. The DDT distance decreased under the premature ignition condition. With a passage width of 20 mm, the DDT distance decreased by 22% under the premature ignition condition to a minimum value (76 ± 8 mm). The DDT time from spark time reached a minimum value (69 ± 14 μsec) under the condition of a passage width of 10 mm and premature ignition. The detonation initiation time and the DDT distance were represented by the time until the flame expanded toward the tube-axis one-dimensionally from ignition (characteristic time). We also carried out thrust measurement using a single-tube PDRE and double-tube PDRE composed of a circular cross-section combustor and the newly developed valve. In the singletube PDRE system, we obtained a stable time-averaged thrust in a wide range of operation frequency (40 Hz - 160 Hz) and confirmed the increase of specific impulse due to a partialfill effect. At a maximum operation frequency of 159 Hz, we achieved a maximum propellant-based specific impulse of 232 sec and a maximum time-averaged thrust of 71 N. In the double-tube PDRE system, at a operation frequency per one tube of 65 Hz, we achieved a propellant-based specific impulse of 157 sec and a maximum time-averaged thrust of 104 N.

Performance analysis of precooled turbojet engine with a low-temperature endothermic fuel

Pulse detonation engines (PDE) have received significant attention in recent years due to their potential advantages over conventional propulsion systems, including high thermodynamics cycle efficiency, high thrust/weight and specific impulse. The fuel-air injection and mixing is one of important research issues. To optimize the fuel-air injection and mixing, a three-dimensional computation model of PDE is established. The mixing process for twin-fluid air-assisted atomizers for air supplied with tangential inlet and axial inlet is numerically analyzed by using the software Fluent as platform to solve the three-dimensional Navier-Stokes equations, and the relationship between drop size and air mass and gasoline/air ratio is arrived at. Numeric simulations results of the process with different inlets show that well mixing but bad concentration distribution occurs with tangential inlet especially along the central axis which makes PDE difficult to work in high frequency, while bad mixing but well-proportioned distribution occurs through axial inlet. And with the same gasoline/air ratio initialized, as the air mass increases, the droplet size decreases, and the result is also correct for the case in which the gasoline/air ratio decreases with the same air mass initialized. These results are found to be in good agreement with the limited experimental data that is currently available. The conclusions are available for the design of liquid-fueled PDE.

The main thrust characteristics, such as thrust force, specific impulse, specific fuel consumption, and specific thrust, of a pulse detonation engine (PDE) with an air intake and nozzle in conditions of flight at a Mach number of 3 and various altitudes (from 8 to 28 km above sea level) are for the first time calculated with consideration given to the physicochemical characteristics of the oxidation and combustion of hydro-carbon fuel (propane), finite time of turbulent flame acceleration, and deflagration-to-detonation transition (DDT). In addition, a parametric analysis of the influence of the operation mode and design parameters of the PDE on its thrust characteristics in flight at a Mach number of 3 and an altitude of 16 km is performed, and the characteristics of engines with direct initiation of detonation and fast deflagration are compared. It is shown that a PDE of this design greatly exceeds an ideal ramjet engine in specific thrust, whereas regarding the specific impulse and specific fuel consumption, it is not inferior to the ideal ramjet.

The basic approach with the design of power plants is to first carry out a thermodynamic cycle analysis and then to vary certain cycle parameters such as the overall pressure ratio in order to determine the optimum or design point condition. One would then proceed to design the different components to match the process conditions. However, since component design has an impact on overall system performance, one cannot optimize the design of the components in isolation from the rest of the system. This calls for an iterative procedure where one has to move several times between the process and component levels to obtain an optimized integrated solution. Another problem faced by plant designers is that Computational Fluid Dynamics (CFD) codes that are increasingly used for detailed component design are slow and not well suited for optimization studies. They are not suited at all for the analysis of complete power plants. Furthermore, the main task of plant designers is not to do design point analyses but to analyze off-design performance, to do uncertainty analyses, to optimize the design and to characterize the dynamic behavior of the system for the purpose of controller design. An approach that has been used with great success for the design of the power conversion system of the Pebble Bed Modular Reactor (PBMR) is the systems CFD approach. The PBMR is a new High Temperature Gas-cooled Reactor (HTR) that is being developed in South Africa. The PBMR utilizes a direct closed recuperated Brayton cycle. Other cycles are also being investigated including various combined cycles. Systems CFD codes are based on the network approach and allow one to model the performance of large complex systems in an integrated fashion. Different levels of component models are provided for ranging from lumped models for components such as pumps to 1D, 2D or even 3D CFD models for components such as complex diffusers, heat exchangers and the pebble bed reactor. In this paper the systems CFD approach will be discussed including the most important component models. Various examples of the application of the systems CFD approach in the design of the PBMR plant will be given.

Results are presented from a calculation of the specific impulses that can be generated through the combustion of cryogenic CO and O2 over a range of fuel/oxidizer ratios, chamber pressures, nozzle expansion ratios, freestream pressures representative of Mars, and the limiting conditions of equilibrium and frozen nozzle flow. For an expansion ratio of 80 and 100-atm. chamber pressure, a specific impulse of 298 sec was obtained; this is comparable to the best solid rocket propellants.

Unlike conventional engines based on the Brayton cycle, detonation-based combustion devices can result in a net pressure gain, in addition to the temperature increase during the combustion process. Pulse detonation engines (PDEs) have been studied for more than seventy years. A brief overview of the fundamentals behind this engine concept is presented first, followed by the progress made up to a flight test of an aircraft using this engine. There is the potential to increase further the substantial performance gains made by the pulsed detonation engine (PDE) concept by moving from “pulsed” or “intermittent” detonations to a more “continuous” mode of operation with detonations. The rotating detonation engine (RDE) is an example of this approach. In this paper, an attempt is made to gather the key information presented in the previous publications in the open literature and to assess the potential for “rotating detonation-wave engines” (RDEs). Recent research is used to highlight the operation and performance of a generic rotating detonation-wave engine. The flow field within the engine is shown to be fairly complicated. However, it follows the “steady-state” thermal detonation cycle quite closely, unlike the PDE. Parametric studies varying the pressure conditions both upstream at the inlet plane and downstream through the back pressure reveal that the overall specific impulse is just a function of the ratio of the inlet stagnation pressure to the chamber back pressure for the generic device considered here. The importance of the inlet design in achieving reliable operation and performance is also highlighted. Finally, a brief overview of the global research efforts on this concept is presented.

The present study addresses the influence of convective heat losses on the flow field and performance of pulse detonation engines (PDE). We investigate the simplest PDE configuration of a straight detonation tube open at one end. The computations are carried out for stoichiometric hydrogen-oxygen, propane-oxygen and ethylene-oxygen mixtures at standard conditions using the method of characteristics for non-isentropic flows using a polytropic gas assumption. The thermodynamic foundation of this gas assumption is investigated. It is found that the model provides a very good approximation for an equilibrium flow. Previous heat flux measurements obtained immediately behind hydrogen-air and hydrogen-oxygen detonations are used to calibrate the assumed convective heat flux model. The calculations agree very well with experimental pressure and velocity profiles at various locations in the tube. The integrated impulse is also found in very good agreement with existing experimental results. The present study clarifies the previous observations of specific impulse deficits with increasing tube length. It is shown that the non-dimensional tube length L / D, where D is tube diameter, governs the amount of losses, the rate of pressure decay at the thrust wall, and hence the performance of PDE’s. The present study reveals that approximately 20% specific impulse losses are expected in typical tube geometries of L / D = 50, the losses varying quasi-linearly with L / D. Computations for ethylene-oxygen using the same heat loss model but including the exact thermo-chemical properties of the detonation products enabled validation of the polytropic gas calculations. Moreover, these finite rate calculations permitted comparisons with experimental temperature measurements, where excellent agreement is found. NOMENCLATURE

Aluminum-doped strontium ferrites for a two-step solar thermochemical air separation cycle: Thermodynamic characterization and cycle analysis

A simplified flowpath analysis of a single-tube airbreathing pulse detonation engine is described. The configuration consists of a steady supersonic inlet, a large plenum, a valve, and a straight detonation tube (no exit nozzle). The interaction of the filling process with the detonation is studied, and it is shown how the flow in the plenum is coupled with the flow in the detonation tube. This coupling results in total pressure losses and pressure oscillations in the plenum caused by the unsteadiness of the flow. Moreover, the filling process generates a moving flow into which the detonation has to initiate and propagate. An analytical model is developed for predicting the flow and estimating performance based on an open-system control volume analysis and gasdynamics. The existing single-cycle impulse model is extended to include the effect of filling velocity on detonation tube impulse. Based on this, the engine thrust is found to be the sum of the contributions of detonation tube impulse, momentum, and pressure terms. Performance calculations for pulse detonation engines operating with stoichiometric hydrogen–air and JP10–air are presented and compared to the performance of the ideal ramjet over a range of Mach numbers.

A technique is introduced in which a supersonic flow is compressed in a supersonic inlet by a high pressure jet. This fluidic compression technique is analyzed in the present work using computational techniques. It is shown that by adjusting the pressure of the compression jet, the flow can be kept near design conditions in a fixed-geometry configuration. Results show that this method increases the minimum flowpath area, maintains shock-onlip over a range of Mach numbers, and reduces shock wave/boundary layer interactions. The potential advantages include lower spillage at low Mach numbers, reduced tendency to unstart, and the ability to have near-design operation over a wide range of flight Mach numbers without resorting to variable-geometry machinery. In addition, for combined-cycle engines it yields a low-blockage ftowpath for ejector and ramjet modes while maintaining high compression in scramjet mode. Global analysis is used to provide qualitative performance estimates. Grid doubling is used to analyze spatial resolution.

Spin-stabilized projectiles with liquid payloads may experience different types of flight instabilities caused by the fluid motion in the payload cylinder. The first type is known to occur in low-viscosity fluids, i.e., at high Reynolds numbers, owing to resonance with inertial waves at critical frequencies. The second type originates from a forced secondary flow at arbitrary frequency, and is most pronounced for fluids of high viscosity, i.e., relatively low Reynolds numbers. For cylinders completely filled with a single fluid, these instabilities were analyzed by eigenfunction expansion developed by Selmi, Li, and Herbert [Phys. Fluids A 4, 1998 (1992)]. The method permits unified analysis of both types of instability, since it can be used for flows at moderate as well as high Reynolds numbers. Often in practice, cylinders are made to include a central rod to alter resonance properties or are partially filled during production, to ensure safety as the liquid payload may expand under different conditions. In this paper, the eigenfunction approach is extended to analyze the moments caused by the flow in a spinning and nutating cylinder, containing a partial fill or a central rod. The analysis shows that the fill ratio (defined as the ratio of the volume of the fluid to the volume of the cylinder) affects resonance with inertial waves. The inviscid flow equations are solved analytically to provide criteria for the onset of resonance in the two configurations.