Combustion Flow Field Research Articles

The Brief Introduction of Different Laser Diagnostics Methods Used in Aeroengine Combustion Research

Combustion test diagnosis has always been one of the most important technologies for the development of aerospace engineering. The traditional methods of measurement have been unable to meet the requirements of accurate capture of the flow field in the development process of the aeroengine combustor. Therefore, the development of high-precision measurement and diagnostic techniques to meet the needs of the aeroengine combustor design is imperative. Laser diagnostics techniques developed quickly in the past several years. They are used to measure the parameters of the combustion flow field such as velocity, temperature, and components concentration with high space and time resolution and brought no disturbance. Planar laser-induced fluorescence, coherent anti-Stokes Raman scattering, tunable diode laser absorption spectroscopy, and Raman scattering were introduced systemically in this paper. After analysis of their own advantages and disadvantages, the authors considered validated Raman scattering system and Tunable Diode Laser Absorption Tomography are more suitable for research activities on aeroengine combustion systems.

The model constant A of the eddy dissipation model

The eddy dissipation model has been used extensively in turbulent combustion modelling because of its simplicity, good convergence, and reasonable accuracy. The model constant A with the standard value 4 has been used in numerous CFD applications without verification and validation. A systematic investigation of the model constant A has been carried out here. Eleven turbulent diffusion flames including round jet flames and opposed jet flames are CFD simulated and the simulation results are compared with experimental data. It is revealed that the flame temperature predicted with A = 4 deviates from experimental data severely. An optimal A value can be found to match experimental data for each flame; this value depends on fuel, chemistry, and the turbulence strength of combustion flow field; it varies between 0.77 and 25 for the 11 flames. A correlation of optimal A value with turbulent Reynolds number is presented; a modification to the original eddy dissipation model is also presented and validated.

THE BRIEF INTRODUCTION OF DIFFERENT LASER DIAGNOSTICS METHODS USED IN AERO-ENGINE COMBUSTION RESEARCH

Combustion test diagnose has always been one of the most important technologies for the development of aerospace engineering. Laser diagnostics techniques developed quickly in the past several years. They are used to measure the parameters of the combustion flow field such as velocity, temperature, components concentration with high space and time resolution and brought no disturbance. Planar Laser-Induced Fluorescence, Coherent Anti-Stokes Raman Scattering, Tunable Diode Laser Absorption Spectroscopy and Raman Scattering were introduced systemically in this paper. After analysis their own advantages and disadvantages, it is believed that Raman Scattering system is more suitable for research activities on aero-engine combustion systems.

Effect of Slot Injection and Effusion Array on the Liner Heat Transfer Coefficient of a Scaled Lean-Burn Combustor With Representative Swirling Flow

International standards regarding polluting emissions from civil aircraft engines are becoming gradually even more stringent. Nowadays, the most prominent way to meet the target of reducing NOx emissions in modern aero-engine combustors is represented by lean-burn technology. Swirl injectors are usually employed to provide the dominant flame stabilization mechanism coupled to high-efficiency fuel atomization solutions. These systems generate very complex flow structures, such as recirculations, vortex breakdown, and processing vortex core, which affect the distribution and therefore the estimation of heat loads on the gas side of the liner as well as the interaction with the cooling system flows. The main purpose of the present work is to provide detailed measurements of heat transfer coefficient (HTC) on the gas side of a scaled combustor liner highlighting the impact of the cooling flows injected through a slot system and an effusion array. Furthermore, for a deeper understanding of the interaction phenomena between gas and cooling flows, a standard two-dimensional (2D) particle image velocimetry (PIV) technique has been employed to characterize the combustor flow field. The experimental arrangement has been developed within EU project LEMCOTEC and consists of a nonreactive three sectors planar rig installed in an open-loop wind tunnel. Three swirlers, replicating the real geometry of a GE Avio partially evaporated and rapid mixing (PERM) injector technology, are used to achieve representative swirled flow conditions in the test section. The effusion geometry is composed by a staggered array of 1236 circular holes with an inclination of 30 deg, while the slot exit has a constant height of 5 mm. The experimental campaign has been carried out using a thermochromic liquid crystals (TLCs) steady-state technique with a thin Inconel heating foil and imposing several cooling flow conditions in terms of slot coolant consumption and effusion pressure drop. A data reduction procedure has been developed to take into account the nonuniform heat generation and the heat loss across the liner plate. Results in terms of 2D maps and averaged distributions of HTC have been supported by flow field measurements with 2D PIV technique focussed on the corner recirculation region.

Simulation of Two-Dimensional Scramjet Combustor Reacting Flow Field Using Reynolds Averaged Navier-Stokes WENO Solver

AbstractThe non-equilibrium chemical reacting combustion flows of a proposed long slender scramjet system were numerically studied by solving the turbulent Reynolds averaged Navier-Stokes (RANS) equations. The Spalart-Allmaras one equation turbulence model is used which produces better results for near wall and boundary layer flow field problems. The lower-upper symmetric Gauss-Seidel implicit scheme, which enables results converge efficiently under steady state condition, is combined with the weighted essentially non-oscillatory (WENO) scheme to yield an accurate simulation tool for scramjet combustion flow field analysis. Using the WENO schemes high-order accuracy and its non-oscillatory solution at flow discontinuities, better resolution of the hypersonic flow problems involving complex shock-shock/shock-boundary layer interactions inside the flow path, can be achieved. Two types of scramjet combustor with cavity-based and strut-based fuel injector were considered as the testing models. The flow characteristics with and without combustion reactions of the two types combustor model were studied with a transient hydrogen/oxygen combustion model. The detailed results of aerodynamic data are obtained and discussed, moreover, the combustion properties of varying the equivalent ratio of hydrogen, including the concentration of reacting species, hydrogen and oxygen, and the reacting products, water, are demonstrated to study the combustion process and performance of the combustor. The comparisons of flow field structures, pressure on wall and velocity profiles between the experimental data and the solutions of the present algorithms, showed qualitatively as well as the quantitatively in good agreement, and validated the adequacy of the present simulation tool for hypersonic scramjet reacting flow analysis.

Numerical study of influence of biofuels on the combustion characteristics and performance of aircraft engine system

The atomization and combustion flowfield of the combustion chamber with swirl-nozzle were simulated using different biofuels; the thermodynamic cycle of the aircraft engine system were also analyzed, influences of biofuels on the combustion characteristics and performance of aircraft engine system were explored. Results show that viscosity and caloric value are key factors affecting the atomization and combustion characteristics of biofuels, and then dominate the distribution of the temperature and NO concentration. Due to the characteristic of low viscosity and low caloric value for biofuels adopted, the biofuels accumulate near the head of combustion chamber, and the corresponding NO emission is lower than that it has for conventional kerosene. When biofuels with low caloric value are used under the operation condition which is same as the condition for the conventional kerosene, lower turbine inlet temperature, lower thrust and higher specific fuel consumption would be achieved for the aircraft engine.

The unsteady turbulence flow of cold and combustion case in different trapped vortex combustor

The trapped vortex combustor (TVC) is a new design concept in which cavities are designed to trap a vortex flow structure. In order to improve combustion efficiency of trapped vortex combustor, a new structure with guide vane and blunt body is proposed. Numerical simulation of unsteady turbulence flow is carried out, and the flow field of different combustor with blunt body and without blunt body is compared to analyze the effect brought by blunt body. The results show that the setting of blunt body can increase the value and the distribution area of turbulent kinetic energy, and improve the heat and mass transfer of the combustion chamber. Double vortex structure, no matter size or location, basically remains unchanged since it is formed, setting blunt body will not affect the cavity flow. Furthermore, after setting blunt body, the combustion chamber exit temperature distribution is more uniform, combustion efficiency is higher.

50 N급 아산화질소/에탄올 추력기의 점화 과도 유동장 가시화

아산화질소/에탄올 추진제 조합을 사용하는 50 N급 추력기 분사기 인접부에서 shadowgraph 기법을 사용하여 연소 유동장을 가시화하였다. 설계 작동조건에서 폭발성 점화가 일어났으며, 급격히 팽창하는 연소가스에 의해 순간적으로 소염되는 현상이 관찰되었다. 최초 점화로부터 약 83 ms 뒤 추진제 분무는 재점화 되었으며, 그로부터 약 23 ms 후 화염이 안정화되는 것이 관찰되었다. 동일한 시퀀스에서 설계 조건보다 산화제 유량을 증가시킨 경우, 소염 현상 없이 완만한 연소 압력 과도 특성을 보였으며 최초 점화로부터 약 17 ms이내에 화염 안정화가 이루어졌다. 이는 점화 직전의 연소기 내부 미연 추진제 혼합 향상에 의한 것으로 생각되었다. The combustion flowfield at the near-injector region of a 50 N scale <TEX>$N_2O/C_2H_5OH$</TEX> thruster was visualized using shadowgraph technique. The explosive ignition was occurred at the design spray condition, and the expanding combustion gas quenched the flame immediately. Approximately after 83 ms from the initial ignition, the propellant spray was re-ignited, and the flame was stabilized after 23 ms elapsed. In the increased oxidizer flow rate condition, the transient pressure at the moment of ignition was smoother than explosive ignition, and the blow down phenomenon was not appeared in the same operating sequence. In addition, the flame was stabilized within 17 ms, and it is caused by improved propellants mixing before ignition.

Application and theoretical analysis of the flamelet model for supersonic turbulent combustion flows in the scramjet engine

In the framework of Reynolds-averaged Navier–Stokes simulation, supersonic turbulent combustion flows at the German Aerospace Centre (DLR) combustor and Japan Aerospace Exploration Agency (JAXA) integrated scramjet engine are numerically simulated using the flamelet model. Based on the DLR combustor case, theoretical analysis and numerical experiments conclude that: the finite rate model only implicitly considers the large-scale turbulent effect and, due to the lack of the small-scale non-equilibrium effect, it would overshoot the peak temperature compared to the flamelet model in general. Furthermore, high-Mach-number compressibility affects the flamelet model mainly through two ways: the spatial pressure variation and the static enthalpy variation due to the kinetic energy. In the flamelet library, the mass fractions of the intermediate species, e.g. OH, are more sensible to the above two effects than the main species such as H2O. Additionally, in the combustion flowfield where the pressure is larger than the value adopted in the generation of the flamelet library or the conversion from the static enthalpy to the kinetic energy occurs, the temperature obtained by the flamelet model without taking compressibility effects into account would be undershot, and vice versa. The static enthalpy variation effect has only little influence on the temperature simulation of the flamelet model, while the effect of the spatial pressure variation may cause relatively large errors. From the JAXA case, it is found that the flamelet model cannot in general be used for an integrated scramjet engine. The existence of the inlet together with the transverse injection scheme could cause large spatial variations of pressure, so the pressure value adopted for the generation of a flamelet library should be fine-tuned according to a pre-simulation of pure mixing.

Absolute/Convective Instability Transition in a Backward Facing Step Combustor: Fundamental Mechanism and Influence of Density Gradient

Hydrodynamic instabilities of the flow field in lean premixed gas turbine combustors can generate velocity perturbations that wrinkle and distort the flame sheet over length scales that are smaller than the flame length. The resultant heat release oscillations can then potentially result in combustion instability. Thus, it is essential to understand the hydrodynamic instability characteristics of the combustor flow field in order to understand its overall influence on combustion instability characteristics. To this end, this paper elucidates the role of fluctuating vorticity production from a linear hydrodynamic stability analysis as the key mechanism promoting absolute/convective instability transitions in shear layers occurring in the flow behind a backward facing step. These results are obtained within the framework of an inviscid, incompressible, local temporal and spatio-temporal stability analysis. Vorticity fluctuations in this limit result from interaction between two competing mechanisms—(1) production from interaction between velocity perturbations and the base flow vorticity gradient and (2) baroclinic torque in the presence of base flow density gradients. This interaction has a significant effect on hydrodynamic instability characteristics when the base flow density and velocity gradients are colocated. Regions in the space of parameters characterizing the base flow velocity profile, i.e., shear layer thickness and ratio of forward to reverse flow velocity, corresponding to convective and absolute instability are identified. The implications of the present results on understanding prior experimental studies of combustion instability in backward facing step combustors and hydrodynamic instability in other flows such as heated jets and bluff body stabilized flames is discussed.

Experimental investigations on flow field and combustion characteristics of a model trapped vortex combustor

Experimental studies are carried out to investigate cold flow field and combustion performance of a workable liquid-fueled trapped vortex combustor (TVC). A remarkable design of the model is that all the air flows through the combustor is supplied from a single source, which makes it distinct from most of the models mentioned in open literatures. Planar particle image velocimetry (PIV) is used to measure the cold flow field, with particular emphasis placed on vortical flow pattern within cavities, turbulence intensity distribution and interaction between cavity stream and mainstream. The aim of the PIV measurement is to better understand the underlying flow physics in a TVC, as well as to provide a benchmark for validation and improvement of computational models. Combustion experiments are carried out with preheated air (473K) at atmospheric pressure using liquid kerosene. Flame images at different operating conditions are obtained and the effects of fuel to air ratio (FAR) and inlet Mach number on the physical appearance of flames are clearly revealed. The fuel to air ratio at lean blow out (LBO) is in the range from 0.0043 to 0.0056 as the inlet Mach number varies from 0.15 to 0.30, which indicates good flame stability. The combustion efficiency falls reasonably in the range of 92–100%. The improvement of the radial profile of outlet temperatures indicates the effectiveness of dilution holes.

Shear flow instabilities in swirl-stabilized combustors and their impact on the amplitude dependent flame response: A linear stability analysis

The hydrodynamic instabilities in the flow field of a swirl-stabilized combustor are investigated theoretically. These instabilities give rise to large-scale flow structures that interact with the flame front causing unsteady heat release rate fluctuations. The streamwise growth of these coherent structures depends on the receptivity of the shear layer, which can be predicted numerically by means of linear stability analysis. This analysis is applied to the reacting flow field of a perfectly premixed swirl-stabilized combustor that is subjected to strong axial forcing mimicking thermoacoustic oscillations. The linear stability analysis reveals a clear correlation between the shear layer receptivity and the measured amplitude dependent flame transfer function. The stability analysis based on the natural flow predicts the distinctive frequency dependent flame response to low amplitude forcing. At these conditions, the flow reveals strong spatial amplification near the nozzle, causing the inlet perturbations to be significantly amplified before they reach the flame. At higher forcing amplitudes, the flow instabilities saturate, which manifests in a saturation of the flame response. The saturation of the shear layers predicted from the linear stability analysis is compared to phase-locked measurements of the forced flow field revealing good qualitative agreement. The analysis of the mean flow stability offers a powerful analytical tool to investigate the impact of shear flow instabilities on the flame describing function.

Ramp Injector Scale Effects on Supersonic Combustion

The combustion field downstream of a 10 deg compression ramp injector has been studied experimentally using wall static pressure measurement, OH planar laser-induced fluorescence, and 2 kHz intensified video filtered at 320 nm for OH emission. Nominal test section entrance conditions were Mach 2, 131 kPa static pressure, and 756 K stagnation temperature. The experiment was equipped with a variable-length inlet duct that facilitated varying the boundary-layer development length for nearly constant injector shock structure conditions. By varying the boundary-layer thickness from 40% of the ramp height to 150% of the ramp height, changes in the combustion flowfield downstream of the injector could be diagnosed. It was found that flame shape changed, the persistence of the vortex cores was reduced, and combustion efficiency rose as the incident boundary layer grew. Unsteadiness within the flameholding region also varied greatly over the range of conditions studied, and largest unsteadiness was observed at intermediate boundary-layer heights.

High-repetition-rate three-dimensional OH imaging using scanned planar laser-induced fluorescence system for multiphase combustion

Imaging dynamic multiphase combusting events is challenging. Conventional techniques can image only a single plane of an event, capturing limited details. Here, we report on a three-dimensional, time-resolved, OH planar laser-induced fluorescence (3D OH PLIF) technique that was developed to measure the relative OH concentration in multiphase combustion flow fields. To the best of our knowledge, this is the first time a 3D OH PLIF technique has been reported in the open literature. The technique involves rapidly scanning a laser sheet across a flow field of interest. The overall experimental system consists of a 5 kHz OH PLIF system, a high-speed detection system (image intensifier and CMOS camera), and a galvanometric scanning mirror. The scanning mirror was synchronized with a 500 Hz triangular sweep pattern generated using Labview. Images were acquired at 5 kHz corresponding to six images per mirror scan, and 1000 scans per second. The six images obtained in a scan were reconstructed into a volumetric representation. The resulting spatial resolution was 500×500×6 voxels mapped to a field of interest covering 30 mm×30 mm×8 mm. The novel 3D OH PLIF system was applied toward imaging droplet combustion of methanol gelled with hydroxypropyl cellulose (HPC) (3 wt. %, 6 wt. %), as well as solid propellant combustion, and impinging jet spray combustion. The resulting 3D dataset shows a comprehensive view of jetting events in gelled droplet combustion that was not observed with high-speed imaging or 2D OH PLIF. Although the scan is noninstantaneous, the temporal and spatial resolution was sufficient to view the dynamic events in the multiphase combustion flow fields of interest. The system is limited by the repetition rate of the pulsed laser and the step response time of the galvanometric mirror; however, the repetition rates are sufficient to resolve events in the order of 100 Hz. Future upgrade includes 40 kHz pulsed UV laser system, which can reduce the scan time to 125 μs, while keeping the high repetition rate of 1000 Hz.

Effect of Cross-flow Momentum on Opposing Jet Mixing

The flow of opposing jets was studied in computational fluid dynamics simulation. To clarify the flow field of the jet engine combustor, it is necessary to study air-jet dilution effects typical of opposing jets. The velocity distribution and turbulence intensity were obtained in large eddy simulation for air and an average Reynolds number of 2 × 10 4 corresponding to the jet diameter and velocity. Results obtained numerically generally agreed quantitatively with experimental results obtained in our previous study. Simulations were then carried out to clarify the effect of the momentum flux ratio J (4, 9, 16, and 64) on mixing. Unmixedness was found to be highest for J = 4 since the penetrations and thus collision of opposing jets were weak for J = 4. When J = 9, 16, 64, mixing was improved by jet collision. It is proposed that the mixing mechanisms are differed depending on J.

Numerical Investigation of the Influence of Injection Scheme on the Ethylene Supersonic Combustion

Ethylene supersonic combustion flow field in different injection schemes was studied numerically in the flight Mach 4. The results show that injection pressure has significant influence on the location of the separation zone and the heat release region, but the starting point of the separation region was mostly influenced by the heat release rather than by the injection pressure; the combustion efficiency of the injection schemes including two injection points is higher than that of three injection points, while the total pressure recovery coefficient of the former injection schemes is lower than the latter; excessive ethylene injected in upstream will lead to the change of free-stream flow conditions, which behaves as the inlet unstart in practical application; more ethylene could be injected in downstream to avoid the problem; on the condition of avoiding thermal choke in isolator, it is more advantageous that injection points were arranged more closely to the starting point of separation zone in upside and to the front of the cavity in downside.

Quasi-One-Dimensional Numerical Method for Solid Fuel Scramjet Combustor Analysis and Design

AbstractTo research the performance of a solid fuel scramjet combustor, a solid fuel regression rate model was coupled into a quasi-one-dimensional flow equation, and so the flow parameters and fuel regression rate could be calculated together. According to the regression rate and the combustor diameter in the previous moment, the combustor diameter of the next moment could be obtained. The unsteady combustion and flow matter in the combustor was simplified into a steady calculation of every moment by solving the boundary condition at different moments independently. The numerical results were compared to actual experimental data from the literature, and they agreed well with the experiment. Taking the fuel/air ratio and Mach number as the optimization conditions, the initial size of the combustor could be obtained. On the basis of an optimized combustor, the relevant parameters’ variation rules were calculated and analyzed. It was found that during the working process, the combustor flow field could be k...

The Influence of Bluff-Body Structural Parameters and Incoming Velocities on Advanced Vortex Combustor

The influence of bluff-body structural parameters and incoming velocity on the cold flow field of advanced vortex combustor (AVC) has been studied, and the results show that the vortex in cavity can be the most stable and the total pressure loss can be the least when structural parameter is B2/B1=0.7, L/B1=0.8. The incoming velocity should not be too high or too low, it should be combined with specific structural parameters of the combustor, the combustion flame speed, fuel type, and the stoichiometric ratio and other factors. If AVC design is reasonable, the cavity can not only stabilize the flame, but also not be sensitive to the speed changes.

Comparison of Flow and Heat Transfer Distributions in a Can Combustor for Radial and Axial Swirlers Under Cold Flow Conditions

A comparison study between axial and radial swirler performance in a gas turbine can combustor was conducted by investigating the correlation between combustor flow field geometry and convective heat transfer at cold flow conditions for Reynolds numbers of 50,000 and 80,000. Flow velocities were measured using particle image velocimetry (PIV) along the center axial plane and radial cross sections of the flow. It was observed that both swirlers produced a strong rotating flow with a reverse flow core. The axial swirler induced larger recirculation zones at both the backside wall and the central area as the flow exits the swirler, and created a much more uniform rotational velocity distribution. The radial swirler however, produced greater rotational velocity as well as a thicker and higher velocity reverse flow core. Wall heat transfer and temperature measurements were also taken. Peak heat transfer regions directly correspond to the location of the flow as it exits each swirler and impinges on the combustor liner wall.

Ethylene Fueled Scramjet Combustion Experiments

Effect of fueling scheme on ethylene-fueled scramjet combustion performance is investigated based on pulse wind-tunnel direct connected experiment platform. Both numerical computing and optical diagnosing are employed to reveal the physical features of the scramjet model. The structure of combustor flow-field is uncovered by high speed schlieren experiments. Also the structure is product by numerical computing. The good agreement between experiment results and computing results validated the reliability of numerical computing. The first and second injector’s lean limits and rich limits for combustion are obtained through injecting characteristics experiments. The time evolution of flame with ethylene injected at different single injector are demonstrated by high speed videos. The flame stabilization mode for the first injector is “fuel injection jet-wake stabilized mode”, which for the second injector is “cavity stabilized mode”.