Isothermal flow in a gas turbine combustor — a benchmark experimental study

An experimental investigation of the three-dimensional flow field within a water model of a can-type gas turbine combustion chamber is presented. Flow visualisation demonstrated that internal flow patterns simulated closely those expected in real combustors. The combustor comprised a swirl driven primary zone, annulus fed primary and dilution jets and an exit contraction nozzle. LDA measurements of the three mean velocity components and corresponding turbulence intensities were obtained to map out the flow development throughout the combustor. Besides providing information to aid understanding of the complex flow events inside combustors, the data are believed to be of sufficient quantity and quality to act as a benchmark test case for the assessment of the predictive accuracy of computational models for gas-turbine combustors.     

Introduction of turbulent model in a mixed finite volume/finite element method

The aim of this work is to present a new numerical method to compute turbulent flows in complex configurations. With this in view, a k-? model with wall functions has been introduced in a mixed finite volume/finite element method. The numerical method has been developed to deal with compressible flows but is also able to compute nearly incompressible flows. The physical model and the numerical method are first described, then validation results for an incompressible flow over a backward-facing step and for a supersonic flow over a compression ramp are presented. Comparisons are performed with experimental data and with other numerical results. These simulations show the ability of the present method to predict turbulent flows, and this method will be applied to simulate complex industrial flows (flow inside the combustion chamber of gas turbine engines). The main goal of this paper is not to test turbulence models, but to show that this numerical method is a solid base to introduce more sophisticated turbulence model.     

Analysis of Unsteady Motion with Respect to Noise Sources in a Gas Turbine Combustor: Isothermal Flow Case

In order to evaluate the direct and indirect contributions to the total combustion noise emission, a combustion chamber consisting of a swirl burner and an exit nozzle of Laval-shape, representing a gas turbine combustor, is investigated by means of experiments and large eddy simulation. Focused on the isothermal flow case first and encouraged by a good overall agreement between the LES and the experimental data for the flow field, a first characterisation of the flow with respect to noise sources is performed. To analyse acoustic properties of the flow, time and length scales are evaluated inside the combustor. Furthermore, the evidence for the existence of a precessing vortex core (PVC), typical for configurations with swirl, is revealed. Finally, the effect of the PVC on the flow inside the Laval nozzle is discussed.     

AEROELASTICITY ANALYSIS OF AN INDUSTRIAL GAS TURBINE COMBUSTOR USING A SIMPLIFIED COMBUSTION MODEL

Lean premixed industrial gas turbine combustors are susceptible to flame instabilities, resulting in large unsteady pressure waves that may cause the discharge nozzle to experience excessive vibration levels. A detailed aeroelasticity analysis, aimed at investigating possible structural failure mechanisms, was undertaken using a time-accurate unsteady flow representation, a simplified combustion disturbance and a structural model of the discharge nozzle. The computational domain included the lower part of the combustor geometry as well as the nozzle guide vanes (NGVs) at the HP turbine inlet. A pressure perturbation, representing the unsteadiness due to the combustion process, was applied below the tertiary fuel inlet and its frequency was set to each structural natural frequency in turn. The propagation of the pressure perturbation through the combustor nozzle, its reflection from the NGVs and further reflections were monitored using two different models. The first one, the so-called “open” system, ignored the reflections from the upper part of the combustion chamber while the second one, the “closed” system, assumed full reflection with an appropriate time shift. The calculations have shown that the imposed excitation could generate unsteady pressure shapes that were correlated with the “flap” modes of the discharge nozzle. In addition, an acoustic resonance condition was observed when the forcing pressure wave had a frequency close to 550 Hz, the experimentally observed failure frequency of the nozzle. The co-existence of these two factors, i.e., excitation/structural-mode match and the possibility of acoustic resonance, was thought to have the potential of producing very high vibration response.     

Jet interaction at supersonic cross flow conditions

The thrust produced by lateral jet systems has been successfully used for several years to control the flight trajectory, i.e., the maneuverability of spacecraft in the high atmosphere and in orbit. Recently this technology has also been applied to projectiles and rockets flying in the low atmosphere from sea level up to more than 10 km. At ISL, investigations have been performed with a 90 mm caliber full-scale projectile in order to study a special side jet controlling system at flight speeds of about 1500 m/s, i.e., Mach number at altitudes of 1.5 and 7.5 km. The High Energy ISL Shock Tunnel facility is used as a ground testing facility in which the flow around the projectile is studied at fully duplicated flight conditions. In the test facility the projectile is fixed inside the test chamber and the atmospheric air is set in motion flowing around the projectile test model. The air flow is generated in the ISL Shock Tunnel STB which is equipped for this purpose with a divergent square nozzle with an exit side length of 184 mm. A lateral gas jet is produced by combusting a solid propellant in a combustion chamber, placed inside the projectile. The powder gases are blown out laterally via a nozzle, creating a complex flow field by the interaction of the lateral jet with the external cross flow. Differential interferometry is used to visualize the behavior of the external flow field distorted by the lateral jet outflow. Numerical simulations have been performed based on steady state computations using the conservation equations of mass, momentum and energy. This was done to theoretically predict the development of the flow field around the projectile under the influence of the side jet. As final result the lateral force acting on the projectile is given as force and moment amplification factors, K_F and K_M respectively.Received: 7 May 2002, Accepted: 12 March 2003, Published online: 16 May 2003An abridged version of this paper was presented at the 23rd Int. Symposium on Shock Waves at Fort Worth, Texas, from July 22 to 27, 2001     

Numerical modeling of the processes within the combustion chamber of a solid propellant rocket engine during the stabilization of the design operation mode

Comprehensive numerical modeling of the processes occurring in the combustion chamber of a solid propellant rocket engine during the stabilization of the design operation mode is performed. The self-consistent problem considered includes nonstationary operation of an ignition device, warmup and ignition of a solid propellant charge followed by its nonstationary burning, nonstationary threephase homogeneous-heterogeneous flow of the combustion products in the combustion chamber, in the nozzle, and behind the engine nozzle unit, engine depressurization, and nozzle unit plug blowing-out. The results of the calculations are presented.     

Flow of gas with a cellular vortical structure through a nozzle

The operation of rocket motors is often accompanied by the development of powerful secondary vortices in the combustion chamber [1–3], The superposition of the secondary vortices on the main flow leads to the formation of a cellular flow structure. Each of the cells represents a three-dimensional vortical circulation of the gas, and this causes a change in the working conditions of the nozzle. The model of helical motion [4] is used in this paper in considering the influence of the three-dimensional behavior of an adiabatic flow on the flow and traction characteristics of the nozzle.     

Experimental and Computational Investigations of Flow and Mixing in a Single-Annular Combustor Configuration

A complementary experimental and computational study of the flow and mixing in a single annular gas turbine combustor has been carried out. The object of the investigation is a generic mixing chamber model, representing an unfolded segment of a simplified Rich-Quick-Lean (RQL) combustion chamber operating under isothermal, non-reacting conditions at ambient pressure. Two configurations without and with secondary air injection were considered. To provide an appropriate reference database several planar optical measurement techniques (time-resolved flow visualisation, PIV, QLS) were used. The PIV measurements have been performed providing profiles of all velocity and Reynolds-stress components at selected locations within the combustor. Application of a two-layer hybrid LES/RANS (HLR) method coupling a near-wall k − ε RANS model with conventional LES in the core flow was the focus of the computational work. In addition to the direct comparison with the experimental results, the HLR performance is comparatively assessed with the results obtained by using conventional LES using the same (coarser) grid as HLR and two eddy-viscosity-based RANS models. The HLR model reproduced all important flow features, in particular with regard to the penetrating behaviour of the secondary air jets, their interaction with the swirled main flow, swirl-induced free recirculation zone evolution and associated precessing-vortex core phenomenon in good agreement with experimental findings.     

Subgrid Model Influence in Large Eddy Simulations of Non-reacting Flow in a Gas Turbine Combustor

Fuel efficiency improvement and harmful emission reduction are the paramount driving forces for development of gas turbine combustors. Lean-burn combustors can accomplish these goals, but require specific flow topologies to overcome their sensitivity to combustion instabilities. Large Eddy Simulations (LES) can accurately capture these complex and intrinsically unsteady flow fields, but estimating the appropriate numerical resolution and subgrid model(s) still remain challenges. This paper discusses the prediction of non-reacting flow fields in the DLR gas turbine model combustor using LES. Several important features of modern gas turbine combustors are present in this model combustor: multiple air swirlers and recirculation zones for flame stabilisation. Good overall agreement is obtained between LES outcomes and experimental results, both in terms of time-averaged and temporal RMS values. Findings of this study include a strong dependence of the opening angle of the swirling jet inside the combustion chamber on the subgrid viscosity, which acts mainly through the air mass flow split between the two swirlers in the DLR model combustor. This paper illustrates the ability of LES to obtain accurate flow field predictions in complex gas turbine combustors making use of open-source software and computational resources available to industry.     

Liquid flow in a simplex swirl nozzle

Pressure-swirl nozzles are widely used in applications such as combustion, painting, air-conditioning, and fire suppression. Understanding the effects of nozzle geometry and inlet flow conditions on liquid film thickness, discharge coefficient and spray angle is very important in nozzle design. The nozzle-internal flow is two-phase with a secondary flow which makes its detailed analysis rather complex. In the current work, the flow field inside a pressure-swirl nozzle is studied theoretically. Using the integral momentum method, the growth of the boundary layer from the nozzle entry to the orifice exit is investigated and the velocity through the boundary layer and the main body of the swirling liquid is calculated. A numerical modeling and a series of experiments have also been performed to validate the theoretical results. The effect of various geometrical parameters is studied and results are compared for viscous and inviscid cases. In addition, the condition in which the centrifugal force of the swirling flow overcomes the viscous force and induces an air core is predicted. The theoretical analysis discussed in this paper provides better criteria for the design and the performance analysis of nozzles.     

Magnetogasdynamic generation of electrical energy in models of aircraft with a high-speed jet engine

The possibility of generating electric power in a plane model of an integral high-speed hydrogen-burning jet engine by mounting a magnetogasdynamic (MHD) generator at the combustion chamber exit is discussed. Attention is concentrated on clarifying the effect of MHD energy extraction from the stream on the aircraft’s thrust characteristics. The internal and external flows are simulated numerically. The two-dimensional supersonic gasdynamic flow inside the engine (in the air-intake, combustion chamber, MHD generator, and nozzle) and the supersonic flow past the aircraft are described on the basis of the complete averaged system of Navier-Stokes equations (in the presence of turbulence), which includes MHD force and heat sources, a one-parameter turbulence model, the electrodynamic equations for an ideal segmented MHD generator, and the equations of the detailed chemical kinetics of hydrogen burning in air. The numerical solution is obtained by means of a computer program that uses a relaxation scheme and an implicit higher-order version of the Godunov method. It is shown that MHD electric power generation can be realized without disturbing the positive balance in the relation between the thrust and the drag of the aircraft with the engine operating with allowance for the MHD drag, but with some loss of effective thrust.     

斜爆轰发动机的推力性能理论分析

爆轰燃烧具有释热快、循环热效率高的特点. 斜爆轰发动机利用斜爆轰波进行燃烧组织, 在高超声速吸气式推进系统中具有重要地位. 以往研究主要关注斜爆轰波的起爆、驻定以及波系结构等, 缺少从整体层面出发对斜爆轰发动机开展推力性能分析. 本文将斜爆轰发动机内的流动和燃烧过程分解成进气压缩、燃料掺混、燃烧释热和排气膨胀4个基本模块并分别进行理论求解, 建立了斜爆轰发动机推力性能的理论分析模型. 在斜爆轰波系研究成果的基础上, 选取了过驱动斜爆轰、Chapman?Jouguet斜爆轰、过驱动正爆轰和斜激波诱导等容燃烧等4种燃烧模式来描述燃烧室内的燃烧释热过程, 并对比分析了不同燃烧模式对发动机比冲性能的影响. 此外, 还获得了不同来流参数、燃烧室参数和进排气参数等对发动机推力的影响规律, 发现来流马赫数和尾喷管的膨胀面积比是发动机理论燃料比冲的主要影响因素. 最后, 结合以往关于受限空间内斜爆轰波驻定特性等方面的研究成果, 提出了斜爆轰发动机燃烧室的设计方向.      

Velocity and Reynolds stresses in a precessing jet flow

 A novel fluid mixing device, described elsewhere, has been shown to have a dramatic effect on the combustion characteristics of a fuel jet. The main features of the flow are the deflection of the jet between 30° and 60° from the nozzle axis and its precession about that axis. Many of the factors governing the nozzle instabilities which drive the mixing in the external field are imprecisely defined. It is the aim of the present paper to examine, in isolation from the nozzle instabilities, the influence of precession on a deflected jet as it proceeds downstream from the nozzle exit. The fluid dynamically driven phenomena within the nozzle which cause the precession are in the present investigation replaced by a mechanical rotation of a nozzle from which is emerging a jet which is orientated at an angle from the nozzle axis. By this means the effect of precession on the deflected jet can be investigated independently of the phenomena which cause the precession. The experimental data reported here has been obtained from measurements made using a miniature, rapid response four-hole “Cobra” pitot probe in the field of the precessing jet. Phase-averaged three dimensional velocity components identify the large scale motions and overall flow patterns. The measured Reynolds stresses complement the velocity data and are found to be compatible with the higher entrainment rates of the jet found in earlier investigations. Received: 8 November 1995 / Accepted: 27 September 1996     

Detonation combustion of hydrogen in a Laval nozzle under rarefied atmosphere conditions

Detonation combustion of a hydrogen-air mixture entering an axisymmetric convergent-divergent nozzle at a supersonic velocity is considered under atmospheric conditions at altitudes up to 24 km. The investigation is carried out on the basis of the two-dimensional gasdynamic Euler equations for a multicomponent reacting gas. The limiting altitude ensuring detonation combustion in a Laval nozzle of given geometry is numerically established for freestream Mach numbers 6 and 7. The possibility of the laser initiation of detonation in a supersonic flow of a stoichiometric, preliminarily heated hydrogen-air mixture is experimentally studied. The investigation is carried out in a shock tube under conditions simulating a supersonic flow in the nozzle throat region.     

Large eddy simulations of isothermal confined swirling flow in an industrial gas-turbine

The paper describes the results of a computational study of the strongly swirling isothermal flow in the combustion chamber of an industrial gas turbine. The flow field characteristics are computed using large eddy simulation in conjunction with a dynamic version of the Smagorinsky model for the sub-grid-scale stresses. Grid refinement studies demonstrate that the results are essentially grid independent. The LES results are compared with an extensive set of measurements and the agreement with these is overall good. The method is shown to be capable of reproducing the observed precessing vortex and central vortex cores and the profiles of mean and rms velocities are found to be captured to a good accuracy. The overall flow structure is shown to be virtually independent of Reynolds number.     

Numerical solution of the direct-problem of mixed nozzle flow

A large part of the known results of Laval nozzle theory relates to the inverse problem, in which the velocity distribution on some line (usually the axis of symmetry) is given rather than the nozzle contour. Many important properties of transonic flows have been disclosed as a result of numerous studies, whose basic results were presented together with an extensive bibliography in Ryzhov's monograph [1]. The solution of the inverse problem has recently been used not only to analyze the qualitative characteristics but also to construct nozzles with rather marked variation of the slope of the generator, which are of practical interest. In this connection we note the work of Pirumov [2] and also the studies of Hopkins and Hill [3, 4]. The latter authors, in addition to the classical Laval nozzle, studied several nozzle schemes with a centerbody. Pirumov used a specially developed numerical method for the solution of the inverse problem (we note that in the subsonic part of the nozzle the corresponding Cauchy problem is incorrect), while Hopkins and Hill used a series expansion which was preceded by a change of variables.There are considerably fewer studies devoted to the solution of the direct problem of mixed nozzle flow. Numerical methods have been used by Alikhashkin, Favorskii, and Chushkin [5], Favorskii [6], and Danilov [7], with the method of integral relations being used in the first two studies. Finally, there has recently been extensive development of the method of expansion in powers of ^1/2, where is the ratio of the radius (or half-width of the nozzle to the radius of curvature of the wall, calculated at the throat section. Such expansions have been used by Hall [8] and Kliegel and Quan [9] to study flow in classical Laval nozzles, and by Moore [10] and Moore and Hall [11] to study flow in nozzles with a centerbody. We note that the ^1/2-expansion method is suitable only in those cases in which the wall radii of curvature are large.In the following the asymptotic method is used to solve the direct problem of mixed flow in nozzles. This reduces the very complex boundary value problem for an elliptic-hyperbolic system of equations with two unknown variables to the Cauchy problem (more precisely, to a mixed problem with initial conditions in a bounded two-dimensional region and boundary conditions which are independent of the third variable) for a hyperbolic system with three unknown variables. The integration of the equations describing the two-dimensional (plane of axisymmetric) nonsteady flow was accomplished with the aid of the Godunov-Zabrodin-Prokopov difference scheme [12]. Several types of nozzles with centerbody are calculated as well as the classical Laval nozzle. The contours of the subsonic parts of the nozzles were either closed (finite combustion chamber) or open (nozzle joins an infinite cylindrical tube). In the first case the flow is provided by three-dimensional mass and energy sources which are introduced at some fixed part of the combustion chamber. In the second case there are no mass and energy sources, but a boundary condition is established at a plane perpendicular to the nozzle axis and located at a finite distance from the throat section, and this condition becomes the flow uniformity condition as this plane moves away to infinity.The authors wish to thank I. Yu. Brailovskii for valuable advice in the selection of the difference scheme, U. G. Pirumov for the kind offer of the results of his calculations, and A. M. Konkina and L. P. Frolova for assistance in the calculations.     

Stability of the combustion process in elastic combustion chambers

High-frequency instability phenomena in rigid combustion chambers have been studied theoretically in [1–3]. This phenomenon is attributed to the interaction between the combustion processes and combustion-product fluctuations in the chamber. One of the possible mechanisms of formation of high-frequency instability is examined in [3], where the combustion rate is represented in the form of a retarded pressure functional. In this case, the problem is reduced to studying the stability of a certain distributed self-oscillating time-lag system.If the oscillation frequencies of the combustion products are comparable to the natural vibrations of the shell which forms the combustion chamber, then it is natural to expect that the elasticity of the chamber walls will affect the combustion process. Coupled effects of acoustoelastic instability can arise, in whose development the vibrations of the chamber wall play a substantial role. These effects are particularly undesirable from the point of view of the vibrational stability of combustion chambers.In this paper, a theory of high-frequency instability of stationary combustion is developed with allowance for elastic deformations of the combustion chamber walls. The theory is based on the mechanism of vibrational combustion [1–3], according to which the combustion front is assumed to the concentrated, while the velocity jump at the front is expressed through a retarded pressure functional. It is assumed that the combustion product flow is one-dimensional and isentropic and that the chamber is cylindrical. The deformations of the chamber are described via the moment theory of shells. The existence is revealed of additional instability regions produced by the interaction between the elastic vibrations of the chamber walls and the acoustic oscillations of the combustion products. The influence of the relation between the elastic and acoustic frequencies and of the structural damping factor in the combustion chamber walls on the stability of the stationary combustion process is examined. The problem discussed is treated as a mathematical model for more complex asymmetric problems in which the elastic and acoustic frequencies can be of the same order.     

Numerical modeling of detonation combustion of hydrogen-air mixtures in a convergent-divergent nozzle

The flow of igniting hydrogen-air mixtures entering an axisymmetric convergent-divergent nozzle at a supersonic velocity is considered. A possibility of stabilizing detonation combustion is numerically investigated at different freestream Mach numbers with account for nonuniform distribution of hydrogen concentration at the nozzle entry. The investigation is performed on the basis of the two-dimensional gasdynamic Euler equations for a multicomponent reacting gas. A detailed model of chemical reactions is used. The calculated thrust is compared with the drag of a conical housing containing the supersonic nozzle considered.     

Acoustically Induced Flashback in a Staged Swirl-Stabilized Combustor

This paper describes a joint experimental and numerical investigation of the interaction between thermoacoustics and flashback mechanisms in a swirled turbulent burner. An academic air/propane combustor terminated by a choked nozzle is operated up to 2.5 bars. Experiments show that the flame can stabilize either within the combustion chamber or flashback inside the injection duct, intermittently or permanently. The present study focuses on the mechanisms leading to flashback: this phenomenon can occur naturally, depending on the swirl level which can be adjusted in the experiment by introducing axial flow through the upstream inlet. It can also be triggered by acoustic waves, either through acoustic forcing or self-excited thermoacoustic instability. Flashback is difficult to study experimentally, but it can be investigated numerically using LES: in a first configuration, the outlet of the chamber is treated as a non-reflecting surface through which harmonic waves can be introduced. In this case, a 20 kPa acoustic forcing is sufficient to trigger permanent flashback after a few cycles. When the LES computational domain includes the choked nozzle used experimentally, no forcing is needed for flashback to occur. Self-excited oscillations reach high levels rapidly, leading to flame flashback, as observed experimentally. These results also suggest a simple method to avoid flashback by using fuel staging, which is then tested successfully in both LES and experiments.     

燃烧轻气炮多级渐扩型燃烧室流场特性数值研究

为了分析多级渐扩型燃烧室结构对燃烧轻气炮氢氧燃烧特性的影响, 通过计算流体力学方法, 分别对采用传统圆柱型燃烧室和多级渐扩型燃烧室的燃烧轻气炮氢氧燃烧发射过程进行数值模拟。对比结果表明, 多级渐扩型燃烧室结构能够明显地减小燃烧室压力波动幅度, 提高氢氧燃烧稳定性; 多级渐扩型燃烧室内形成回流区, 可以减小气流轴向运动速度; 火焰扩展形态与渐扩型结构相吻合, 燃烧反应区表面变化平稳; 多级渐扩型燃烧室结构对氢氧火焰传播过程和压力波动现象有着重要影响。