Jeong-Yeol Choi

Propulsive Performance of Airbreathing Pulse Detonation Engines

The propulsive performance of airbreathing pulse detonation engines at selected flight conditions is evaluated by means of a combined analytical/numerical analysis. The work treats the conservation equations in axisymmetric coordinates and takes into account finite-rate chemistry and variable thermophysical properties for a stoichiometric hydrogen/air mixture. In addition, an analytical model accounting for the state changes of the working fluid in pulse detonation engine operation is established to predict the engine performance in an idealized situation. The system under consideration includes a supersonic inlet, an air manifold, a valve, a detonation tube, and a convergent-divergent nozzle. Both internal and external modes of valve operation are implemented. Detailed flow evolution is explored, and various performance loss mechanisms are identified and quantified. The influences of all known effects (such as valve operation timing, filling fraction of reactants, nozzle configuration, and flight condition) on the engine propulsive performance are investigated systematically. A performance map is established over the flight Mach number of 1.2-3.5. Results indicate that the pulse detonation engine outperforms ramjet engines for all the flight conditions considered herein. The benefits of pulse detonation engines are significant at low-supersonic conditions, but gradually decrease with increasing flight Mach number.

Thrust Chamber Dynamics and Propulsive Performance of Single-Tube Pulse Detonation Engines

This pape r deals with the modeling and simulation of the thrust chamber dynamics in an airbreathing pulse detonation engine (PDE). The system under consideration includes a supersonic inlet, an air manifold, a rotary valve, a single -tube combustor , and a convergent -divergent nozzle. The analysis accommodates the full conservation equations in two - dimensional coordinates, along with a calibrated one - progress -variable chemical reaction scheme for a stoichiometric hydrogen/air mixture. The combustion and flow dynamics involved in typical PDE operations are carefully examined. In addition, a flow -path based performance prediction model is established to estimate the theoretical limit of the engine propulsive performance. Various performance loss mechanisms, including refilling process, mismatch of nozzle exit flow conditions with the ambient state, nozzle flow divergence, and internal flow dynamics, are identified and quantified. The internal flow loss, which mainly arises from the shock waves within the chamber, play s a dominant role in degrading the PDE performance. The effects of engine operating parameters and nozzle configurations on the system dynamics are also studied in depth. Results indicate the existence of an optimum operating frequency for achieving a be st performance margin. For a given cycle period and purge time, the performance increases with decreasing valve close -up time in most cases. On the other hand, a larger purge time decreases the specific thrust but increases the specific impulse for a giv en cycle period and valve close -up time. The nozzle throat area affects both the flow expansion process and chamber dynamics, thereby exerting a much more significant influence than the other nozzle geometrical parameters.

Numerical Investigation of Transverse Hydrogen Jet into Supersonic Crossflow Using Detached-Eddy Simulation

A three-dimensional unsteady reacting flowfield that is generated by transverse hydrogen injection into a supersonic mainstream is numerically investigated using detached-eddy simulation and a finite-rate chemistry model. Grid refinement with the grid-convergence-index concept is applied to the instantaneous flowfield for assessing the grid resolution and solution convergence.Validation is performed for the jet penetration height, and the predicted result is in good agreement with experimental trends. The results indicate that jet vortical structures are generated as the interacting counter-rotating vortices become alternately detached in the upstream recirculation region. Although the numerical OH distribution reproduces the experimental OH–planar-laser-induced fluorescence well, there are some disparities in the ignition delay times due to the restricted availability of experimental and numerical data. The effects of the turbulence model on combustion are identified by a comparative analysis of the Reynolds-averaged Navier–Stokes and detached-eddy simulation approaches. Their effects are quantified by the production ofH2O, which is the primary species of hydrogen combustion.

Numerical Study of Scram Accelerator Starting Characteristics

A numerical study is carried out to investigate the ignition and the detonation initiation process in a scram accelerator operating at a superdetonative mode. To simulate the scram accelerator launching process, a conical projectile is considered, injected with an initial velocity of 2500 m/s from the 1 atm air into a 25 atm 2H 2 + O 2 + mN 2 mixture. As a dilution gas, nitrogen is selected and assumed to be inert. The time accurate solutions of Reynolds averaged Navier-Stokes equations for chemically reacting flows are obtained by using a point-implicit method and an upwind-biased third-order scheme with a steady-state solution for airflow as an initial condition. To examine combustion characteristics and ram-accelerator operation limits, mixture compositions are varied from 2H 2 + O 2 + 3.76N 2 to 2H 2 + O 2 + 9N 2 by changing the amount of N 2 . The flowfield results show the detailed ignition mechanism, the initiation process of the oblique detonation, and the staring characteristics of the scram accelerator. The results also identify clearly the combustion characteristics of the operational failures at lower and upper dilution limits that have been observed in experiments

Numerical Study of Mixing Enhancement by Shock Waves in Model Scramjet Engine

Anumerical study hasbeenconductedto investigatetheeffectofshockwaveson thesupersonichydrogen ‐airjet e ame stabilized in a Mach 2.5 circularcross-section combustor. Thenumerical model utilizes multispecies Navier ‐ Stokes equationswith detailed chemical reaction modelsand employsa k‐! shear stresstransport model. A wedge is mounted on the side wall of the combustor in order to e nd the interaction of the oblique shock waves with the hydrogen‐air jetlike e ame. The interaction between the shock waves and the mixing layer is classie ed according to the increasing tendency of the growth rate of the mixing layer downstream of the shock waves. It is found that the shock wavescreatea radially inward/outward aire owto thee ame and elongatea e ame-holding recirculation zone, and thus fuel ‐air mixing is enhanced signie cantly, resulting in improved combustion efe ciency. Also, the overall performance is investigated by changing the shock position and considering the mixing/combustion efe ciency and total pressure loss in a model scramjet combustor. Because there exists a tradeoff between the enhanced mixing/combustion efe ciency and the decreased total pressure recovery, it is suggested that the optimized shock position needs to be determined in order to obtain the maximum overall combustor performance using the overall performance index.

Unsteady-State Simulation of Model Ram Accelerator in Expansion Tube

Steady- and unsteady-state numerical simulations have been carried out to investigate the ram accelerator flowfield that had been studied experimentally using an expansion tube facility at Stanford University. Navier-Stokes equations for chemically reactive flows were used for the modeling with a detailed hydrogen-air combustion mechanism. The governing equations were analyzed using a fully implicit and time-accurate total variation diminishing scheme. As a result, steady-state simulation reveals that the near-wall combustion regions are induced by aerodynamic heating in the separated flow region. This result agrees well with experiments in the case of the 2H 2 + O 2 + 17N 2 mixture but fails to reproduce the centerline combustion in the case of the 2H 2 + O 2 + 12N 2 mixture. To investigate the reason for this disagreement in the flow establishment process, unsteady-state simulations have been carried out, and the results show the detailed process of flow stabilization. The centerline combustion is revealed to be an intermediate process during flow stabilization. It is induced behind a Mach stem formed by the intersection of strong oblique shock waves at an early stage of the flow stabilization process

Internal Flow Dynamics in a Valveless Airbreathing Pulse Detonation Engine

The internal flow dynamics in a valveless airbreathing pulse detonation engine operating on ethylene fuel is studied numerically. The system involves no mechanical valves in the air flowpath, and the isolation between the inlet and combustor is achieved through gas-dynamic means. The valve operation timing for the fuel injection and initiator is determined based on the local flow conditions. The analysis accommodates the full conservation equations in axisymmetric coordinates and takes into account simplified finite-rate chemistry and variable properties for an ethylene/air/oxygen system. The detailed flow evolution and detonation dynamics during the limit-cycle operation is explored systematically. The calculated pressure history and propulsive performance agree well with experimental data. A sensitivity study of operation timing is also conducted to further elucidate the system dynamics and to provide guidelines for engine design optimization.

Dimensional analysis of the effect of flow conditions on shock-induced combustion

A dimensional analysis is attempted to elucidate the effects of the flow conditions on the regines of shock-induced combustion around a blunt body in stoichiometric hydrogen-oxygen mixtures. At a given Mach number, the effects of body size, inflow pressure, and inflow temperature are investigated. The analysis is based on a series of numerical simulations employing a fully implicit method and an upwind scheme. From the results of the numerical simulations with various flow conditions, different regimes of shock-induced combustion are observed ranging from decoupled shock-deflagration to overdriven detonation. From a viewpoint of nondimensionalization, the two flow-field-configuration variables (the shock standoff distance and the induction distance) and the two dimensionless parameters (the first Damkohler number and the heat release parameter) are redefined along the stagnation streamline. The configuration variables and the dimensionless parameters are evaluated from the flow-field data to analyze quantitatively the global features of a shock-induced combustion flow field. The effects of flow conditions are identified in terms of the configuration variables and the dimensionless parameters. The correlations between the configuration variables and the dimensionless parameters are also investigated to understand the effects of flow conditions within a frame of reference.

Dynamics Combustion Characteristics in Scramjet Combustors with Transverse Fuel Injection

A comprehensive DES quality numerical analysis has been carried out for reacting flows in constant-area and divergent scramjet combustor configurations with and without a cavity. Transverse injection of hydrogen is considered over a broad range of injection pressure. The corresponding equivalence ratio of the overall fuel/air mixture ranges from 0.167 to 0.50. The work features detailed resolution of the flow and flame dynamics in the combustor, which was not typically available in most of the previous studies. In particular, the oscillatory flow characteristics are captured at a scale sufficient to identify the underlying physical mechanisms. Much of the flow unsteadiness is related not only to the cavity, but also to the intrinsic unsteadiness in the flowfield. The interactions between the unsteady flow and flame evolution may cause a large excursion of flow oscillation. The roles of the cavity, injection pressure, and heat release in determining the flow dynamics are examined systematically.

Application of approximate chemical Jacobians for constant volume reaction and shock-induced combustion

Abstract The present study suggests an approximate Jacobian, the Gauss–Seidel partial Jacobian, derived from the concept of preconditioned time differencing methods, and includes it into the LU-SGS (Lower Upper Symmetric Gauss–Seidel) scheme. The validity of the Gauss–Seidel partial Jacobian is demonstrated by calculating a constant volume reaction and by comparing the results with those of other approximate Jacobians. Then, the performance of the LU-SGS scheme is examined by stability analyses and computations of shock-induced combustions. The results show that the LU-SGS scheme with the Gauss–Seidel partial Jacobian is as stable and accurate as the full Jacobian and about 5–30% faster than those with the full Jacobian, while the scheme with other approximate Jacobians produces satisfactory results only at small time step sizes.