Numerical study on supersonic mixing and combustion with hydrogen injection upstream of a cavity flameholder

Characteristics of supersonic mixing and combustion with hydrogen injection upstream of a cavity flameholder are investigated numerically using hybrid RANS/LES (Reynolds-Averaged Navier–Stokes/Large-Eddy Simulation) method. Two types of inflow boundary layer are considered. One is a laminar-like boundary layer with inflow thickness of $\delta_{\inf } = 0.0$ and the other is a turbulent boundary layer with inflow thickness of $\delta_{\inf } = 2.5\,{\text{mm}}$ . The hybrid RANS/LES method acts as a DES (Detached Eddy Simulation) model for the laminar-like inflow condition and a wall-modeled LES for the turbulent inflow condition where the recycling/rescaling method is adopted. Although the turbulent inflow seems to have just minor influences on the supersonic cavity flow without fuel injection, its effects on the mixing and combustion processes are great. It is found that the unsteady turbulent structures in upstream incoming boundary layer interact with the injection jet, resulting in fluctuations of the upstream recirculation region and bow shock, and induce quick dispersion of the hydrogen fuel jet, which enhances the mixing as well as subsequent combustion.     

Calculation of hydrogen-air combustion behind detached shock wave for supersonic flow past a sphere

Many of the published theoretical studies of quasi-one-dimensional flows with combustion have been devoted to combustion in a nozzle, wake, or streamtube behind a normal shock wave [1–6].Recently, considerable interest has developed in the study of two-dimensional problems, specifically, the effective combustion of fuel in a supersonic air stream.In connection with experimental studies of the motion of bodies in combustible gas mixtures using ballistic facilities [7–9], the requirement has arisen for computer calculations of two-dimensional supersonic gas flow past bodies in the presence of combustion.In preceding studies [10–12] the present author has solved the steady-state problem under very simple assumptions concerning the structure of the combustion zone in a detonation wave.In the present paper we obtain a numerical solution of the problem of supersonic hydrogen-air flow past a sphere with account for the nonequilibrium nature of eight chemical reactions. The computations encompass only the subsonic and transonic flow regions.The author thanks G. G. Chernyi for valuable comments during discussion of the article.     

A computational study of supersonic disk-gap-band parachutes using Large-Eddy Simulation coupled to a structural membrane

Large-scale fluid-structure interaction simulations of compressible flows over flexible supersonic disk-gap-band parachutes are compared with matching experimental results. We utilize adaptive mesh refinement, large-eddy simulation of compressible flow coupled with a thin-shell structural finite-element model. The simulations are carried out in the regime where large canopy-area oscillations are present, around and above Mach 2, where strong nonlinear coupling between the system of bow shocks, turbulent wake and canopy is observed. Comparisons of drag history and its dependence on Mach number are discussed. Furthermore, it is observed that important dynamical features of this coupled system can only be reproduced when sufficient grid resolution is used. Lack of resolution resulted in incorrect flow-physics prediction and, consequently, incorrect fluid-structure interaction coupling.     

Ignition and combustion study of JP-8 fuel in a supersonic flowfield

The ignition and combustion process of fuels in a supersonic combustion chamber plays an important role in the design of hypersonic propulsion system. However, it is a quite complicated process, due to the large variation of inlet air velocity, temperature, oxygen concentration, and shocks in the supersonic combustion chamber. The purpose of this paper is to observe the ignition delay and combustion phenomenon of the JP-8 fuel droplets in a supersonic flowfield experimentally. A shock tube is used as a basic test facility to create a high-speed and high-temperature flowfield as a supersonic combustor. In the experiments, several test parameters are controlled, such as shock velocity, gas temperature, fuel droplet size and distance, initial fuel temperature, and oxygen concentration, etc. The test results show the influence of these parameters on ignition delay, ignition limitation, and detonation. The most important factor in the experiment is the initial fuel temperature effect, which is influenced by the altitude variation during a flight. Received 4 August 1995 / Accepted 12 December 1995     

Experimental study of the evolution of perturbations in the supersonic wake behind a flat plate

The evolution of perturbations in the wake of a flat plate (with symmetric wedge-shaped nose and tail) is studied experimentally at the Mach number M=2 for various plate thicknesses and tail lengths.Novosibirsk. Translated from Izvestiya Rossiiskoi Akademii Nauk, Mekhanika Zhidkosti i Gaza, No. 4, pp. 167–171, July–August, 1996.     

Flow behind the boundary layer separation point in a supersonic stream

The flow structure behind the separation point of a laminar boundary layer in a supersonic stream has been investigated. Analytic and numerical solutions are obtained for simple semiinfinite separation zones starting from the leading edge or a point on the smooth surface. The question of the pressure plateau in a separation zone of finite length is discussed and its value is calculated on the basis of asymptotic theory. The asymptotic theory of flow [1, 2] in the neighborhood of the separation point of the laminar boundary layer in a supersonic gas stream (region of free interaction) is employed. The local solution obtained is subsequently used to construct the flow pattern in the separation zone [3]. An analysis is made of the behavior of the solution for the free-interaction region on transition to the region of reverse flows. The results make it possible actually to compute (in the first approximation) the pressure in the plateau region after establishing the mathematical significance of this concept, previously introduced on the basis of the experimental results. At the same time relatively simple solutions are obtained for semiinfinite separation zones.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 3, pp. 19–25, May–June, 1971.     

Calculation of interaction of a turbulent near-wake behind a step with a supersonic jet

Existing computational methods [1–5] do not enable one to calculate complex flows behind steps, accounting for nonuniformity of the incident supersonic flow and the effect of compression and expansion waves arriving in the near-wake region. For example, computational methods based on the methods of [1] or [2] are used mainly in uniform supersonic flow ahead of the base edge and, for the most part, cannot be used to calculate flow in annular nozzles with irregular conditions. An exception is reference [6], which investigated flow in an annular nozzle behind a cylindrical center-body. The present paper suggests a method, based on references [7, 8] for calculating the base pressure behind two-dimensional and three-dimensional steps, washed by a supersonic jet.Translated from Zhurnal Prikladnoi Mekhaniki i Tekhnicheskoi Fiziki, No. 6, pp. 43–51, November– December 1977.     

Initiation and development of separated flow behind a circular cylinder in a supersonic stream

On the basis of the two-dimensional Navier-Stokes equations, the initiation and development of the separated flow behind a thermally insulated circular cylinder in a supersonic perfect-gas stream is investigated in relation to the Reynolds number. It is shown that the entire Re-range can be subdivided into a number of intervals with their own characteristic features. In particular, the conditions for the generation and development of global separation are established. Moscow. Translated from Izvestiya Rossiiskoi Akademii Nauk, Mekhanika Zhidkosti i Gaza, No. 6, pp. 27–36, November–December, 1998. The study was carried out with the support of the Russian Foundation for Basic Research (project No. 95-01-01129a).     

Measurement and imaging of supersonic combustion in a model scramjet engine

Results of tests performed in a free-piston shock tunnel on a model scramjet engine are presented. Two conditions which differed in Mach number were tested. Flow at the lower Mach number condition was achieved using a variable-angle diffuser. Shadowgraph images and floor static pressure measurements were obtained, the latter used as the basis of a finite-difference calculation of flow properties in the scramjet. Received 9 May 1998 / Accepted 30 September 1998     

A numerical study of supersonic turbulence

We report on simulations of nonstationary supersonic isotropic turbulent flow using the Piecewise-Parabolic Method on uniform grids of 2048² in two dimensions and 256³ in three dimensions. Intersecting shock waves initiate the transfer of energy from long to short wavelengths. Weak shocks survive for many acoustic times _ac. In two dimensions eddies merge over many _ac. In three dimensions vortex sheets break-up into short vortex filaments within two _ac. Entropy fluctuations, produced by strong shocks, are stretched into filaments over several _ac. These filaments persist and are mirrored in the density. We observe three temporal phases: onset, with the initial formation of shocks; quasi-supersonic, with strong density contrasts; and post-supersonic, with a slowly decaying root mean square Mach number. Compressive modes quickly establish a k ^–2 velocity power spectrum. In three dimensions solenoidal modes build-up during the supersonic phase delineating through time a Kolmogorov k ^–5/3 envelope and leaving a self-similarly decaying k ^–0.9 spectrum at lower wave numbers.This work was supported at the University of Minnesota by grants DE-FG02-87ER25035 from the Office of Energy Research of the Department of Energy, AST-8611404 from the National Science Foundation, and by equipment grants from Sun Microsystems, Gould Electronics, Seagate Technology, and the Air Force Office of Scientific Research (AFOSR-86-0239). Partial support for this work has also come from the Army Research Office Contract Number DAAL03-89-C-0038 funding the Army High Performance Computing Research Center (AHPCRC) at the University of Minnesota. At Nice this work was supported under DRET Contract 500-276, under the GdR CNRS-SPI Mécanique des Fluides, and under two special grants from the Observatoire de la Côte d'Azur.     

OH-PLIF visualisation of radical farming supersonic combustion flows

Experimental investigations employing Planar Laser-induced fluorescence visualisation of the qualitative distribution of the OH radical (OH-PLIF), coupled with surface pressure measurements, have been made of flow in a generic, nominally two-dimensional inlet-injection radical farming supersonic combustion scramjet engine model. The test flows were provided by a hypersonic shock tunnel, and covered total enthalpies corresponding to the flight Mach number range 8.7–11.8 and approximately 150 kPa dynamic pressure. The surface pressure measurements displayed radical farming behaviour, that is a series of adjacent high and low pressure regions corresponding to successive shock/expansion structures, with no significant combustion-induced pressure rise until the second structure. OH-PLIF imaging between the first two structures provides the first direct experimental evidence of significant OH radical concentrations upstream of the ignition point in this mode of scramjet operation and shows that combustion reactions were occurring in highly localised regions in a complex turbulent and poorly micromixed fuel/air mixing layer confined to the fuel injection side of the combustor.     

Development of a hybrid-fuzzy air temperature controller for a supersonic combustion test facility

 Hydrogen gas is burned in air to raise and maintain the stagnation temperature of a supersonic combustion test facility to a desired setpoint. In order to reach the desired operating conditions for stagnation temperature, there are three phases to the hydrogen control; H₂ ignition at facility start-up, H₂ ramp-up while the facility is ramped-up, and H₂ iteration to achieve the desired temperature setpoint. Each phase incorporates a different type of control. Fuzzy logic is used to design a computer based supervisory controller that recognizes the different phases of operation and chooses the appropriate control method. Received: 28 November 1999/Accepted: 2 November 2000