Combustion Flow Field Research Articles

Effect of non-collimated light path on flow field's temperature reconstruction by moiré tomography

In this paper, the effect of non-collimated light path on flow field's parameter diagnosis by moiré tomography is discussed, and a candle combustion flow field is chosen for experiment. To obtain the non-collimated light path, a little lens which is used for beam expanding and collimation is equidistantly moved twice to the left (near the light source) and right (away from the light source) on the basis of the collimated position. Therefore, for a certain cross section, five sets of moiré fringes can be obtained from experiment for further analysis and discussion. And, three cross sections of the measured candle combustion are chosen for practical discussion. The results manifest that the moiré fringes are tilt to the right when the little lens moves to the left, and the temperature is larger than that of obtained by the collimated light path. Meanwhile, it is found that the moiré fringes are tilt to the left when the little lens moves to the right, and the temperature is smaller than that of obtained by the collimated light path. Finally, the possible factors which could affect the measurement accuracy of OCT methods for diagnosing flow fields are discussed.

Numerical Investigation on the Cold Flow Field of a Typical Cavity-based Scramjet Combustor with Double Ramp Entry

Abstract The effects of back pressure and cavity L/D ratio on the shock wave structure in the cold flow field of a typical cavity-based scramjet combustor with combined inlet and isolator is investigated numerically in the selected scramjet models. The scramjet with a throat ratio of TR 0.0 and cavity L/D 6.04 was analyzed. To perform such analysis, steady, 2-D RANS was used with SST k-ω. From the analysis, the value of static pressure along the cowl surface, contours of Mach number and pressure were obtained. The scramjet was modeled with different TR 0.1, 0.2, 0.25 and 0.3 with the same cavity L/D 6.04 and different cavity L/D 4.04, 9.04 and 12.04 with the same TR 0.25. All the models were analyzed with the same inlet conditions and the results were obtained. From the analysis, it was observed that the increase in back pressure moves the shock train towards the inlet of the isolator which leads to ‘engine unstart’ after the throat ratio of TR 0.1. Also, it is observed that there is an optimal L/D ratio of the cavity L/D 9.04 which restricts the propagation of high-pressure waves obtained in the combustor.

Numerical study on supersonic boundary-layer transition and wall skin friction reduction induced by fuel wall-jet combustion

In this study, the effect of wall-jet combustion on boundary layer transition and skin friction reduction was numerically investigated. To effectively capture the characteristics of boundary layer flow, the transition k−kl−w model was employed as the turbulence model and laminar finite-rate model was chosen as the combustion model. This numerical method was firstly validated by two sets of experimental results in open domain. After that, the research on wall-jet combustion was conducted and the numerical results showed that when injecting hydrogen in different directions, both the hydrogen self-ignition location and the skin friction on the wall were not altered significantly. When the injection angle to the airflow direction was increased to 30°, the intensity of combustion was insufficient and the skin-friction coefficient would be increased. Meanwhile, boundary layer transition occurred in a relatively smaller Reynolds number at this condition. The variation of the wall-jet height would have a greater impact on both boundary layer transition and the skin friction. More components can diffuse to the lower wall as the height of the wall-jet was enlarged, which can make the process of boundary layer transition be postponed and the skin friction be reduced as well. Furthermore, greater skin-friction reduction would be achieved downstream the self-ignition location when the wall was adiabatic, while the original Reynolds numbers for boundary layer transition is the smallest if the wall temperature is set to 600 K. Finally, in order to simulate the effect of the back pressure on the combustion flow field, another injector is added near the exit at the upper wall to produce air-throttling flow jet. The results showed that altering back pressure nearly has little influence on boundary layer transition and the skin friction decreases with more air-throttling flow rate.

Phase information extraction for moiré fringes based on multiresolution analysis.

In this paper, the multiresolution analysis (MRA) method is used to preprocess moiré fringes, which can reduce the number of data points and increase computation speeds. To discuss the applicability of the method, a candle combustion flow field is chosen as an example for experiment by moiré deflectometry. First, moiré fringes are preprocessed by the MRA method. Then, phase information extraction and refractive index reconstruction are performed on the three-level low-frequency approximation components. Finally, the involved results prove that the calculation time required for phase information extraction and refractive index reconstruction is greatly reduced based on the moiré fringes preprocessed by MRA method. The relative error could be accepted if the suitable approximation level is applied.

Simulation of Combustion Flowfield in Porous Media with Lattice Boltzmann Method

In this paper, two-dimensional numerical modeling of a premixed porous combustion is presented by a Lattice Boltzmann approach. In this solution, the flow distribution is replaced by a pressure distribution equation using the Darcy and Forchheimer model as the multiple-relaxation-time code. The laminar propane-air mixture with Reynolds number of 100 and Prandtl number of 0.75 is put through a silicon carbide porous medium channel. The boundary conditions are included as the uniform inlet temperature and velocity, atmospheric output pressure, periodic wall temperature, and nonslip walls with initial zero gradient for products and reactants. For validation of this model, the presented data of temperature, molar concentration, and velocity comprised published data where the existence of acceptable agreements confirms the accuracy of this solution. The results show that, in the porous medium, the flame position is located at closer distance from the entrance of channel rather than the nonporous medium. Furthermore, the maximum temperature is reduced up to 5.8% when Darcy number is decreasing from 0.1 to 0.001, and with the reduction of Darcy number, the size of vortices is decreased. Moreover, when the porosity coefficient decreases, the size of high-temperature region decreases, and with increasing equivalence ratio from 0.6 to 1, the maximum temperature is significantly augmented with a value of 24.3%.

Numerical study of the influence of annular width on the rotating detonation wave in a non-premixed combustor

A numerical study of three-dimensional non-premixed rotating detonation combustor is performed to investigate the effects of combustor annular width (w) on the formation process of rotating detonation wave, characteristics of flow field, propagating stability and propulsive performance of combustor. Fixing the outer diameter and changing the inner diameter, three annular widths (i.e., 5 mm, 8 mm and 12 mm) are investigated. The results show that annular width plays a crucial role to affect the behavior of rotating detonation wave. Dual-wave mode is obtained in case of w=5 mm while single-wave mode is obtained in cases of w=8 mm and 12 mm. Additionally, the highest height of rotating detonation wave is observed in case of w=8 mm, which leads to the largest pressure peak. Besides, great fluctuation of detonation wave height is observed in case of w=12 mm. Furthermore, accumulated pressure wave appears in each case and its intensity is an important factor for the generation of new detonation wave front. Moreover, as the increase of annular width, the thrust and specific impulse both decrease. The largest thrust and specific impulse are both obtained in case of w=5 mm, which are 79.16N and 2692.52 s respectively.

Temperature measurement in combustion flow field with femtosecond laser-induced filament

Laser-based diagnostic techniques are critical nonintrusive methods of measuring the in-situ temperature in combustion flow fields. Developing temperature measurement techniques with high accuracy and precision is of great significance for studying the combustion. At present, nanosecond (ns) lasers are commonly used in these methods. However, the researches based on femtosecond (fs) lasers are relatively few. Here, we develop a thermometry technique for combustion fields based on fs laser-induced filament. When the fs laser propagates in an optical medium, a long uniformly distributed plasma channel (also named filament) will be generated. The clamped intensity inside the filament is high enough to generate excited atoms/molecules through fs laser-induced photochemical reactions. Subsequently, the excited atoms/molecules release fluorescence signals. The length of the filament can be measured by imaging the fluorescence signal with an ICCD camera, which is evaluated by the full width at half maximum (FWHM) of the spatial distribution of the filament emission signal. Based on theoretical analysis, the experimental data of the filament length are fitted with a power function, and the result is satisfactory compared with the <i>R</i>-squared measure of goodness (<i>R</i><sup>2</sup>) of 0.984. This indicates that the filament length is correlated well with the temperature of the combustion field. A monotonic quantitative relationship between the filament length and the temperature can be established by a calibration process, and then the temperature of the combustion field can be measured. When the temperature changes from 1630 to 2007 K, the length of the filament shortens by 38%. This indicates that the filament length is sensitive to the temperature of the flow field. When the temperature is 2007 K, the absolute uncertainty of the measurement is ±25 K, and the relative uncertainly is about 1.2%. The spatial resolution of the measurement system is 50 μm, which was determined by a USAF 1951 Target. Based on the spatial resolution, the measurement precision can arrive at 17 K. Although, at present, this temperature measurement technique based on femtosecond laser-induced filament is used only in laminar premixed flames, it has potential applications in temperature measurements ranging from room temperature to combustion temperatures.

Combustion instability modeling using multi-mode flame transfer functions and a nonlinear Euler solver

Modern methods for predicting combustion dynamics in high-pressure combustors range from high-fidelity simulations of sub-scale model combustors, mostly for validation purposes or detailed investigations of physics, to linearized, acoustics-based analysis of full-scale practical combustors. Whereas the high-fidelity simulations presumably capture the detailed physics of mixing and heat addition, computational requirements preclude their application for practical design analysis. The linear models that are used during design typically use flame transfer functions that relate the unsteady heat addition [Formula: see text] to oscillations in velocity and pressure ([Formula: see text] and [Formula: see text]) that are obtained from the wave equation. These flame transfer functions can be empirically determined from measurements or derived from theory and analysis. This paper describes a hybrid approach that uses high-fidelity simulations to generate flame transfer functions along with nonlinear Euler CFD to predict the combustor flowfield. A model rocket combustor that presented a self-excited combustion instability with pressure oscillations on the order of 10% of mean pressure is used for demonstration. Spatially distributed flame transfer functions are extracted from a high-fidelity simulation of the combustor and then used in a nonlinear Euler CFD model of the combustor to verify the approach. It is shown that the reduced-fidelity model can reproduce the unsteady behavior of the single element combustor that was both measured in the experiment and predicted by a high-fidelity simulation reasonably well.

저선회 모델 연소기의 스월러 코어 막힘률에 따른 난류연소유동장에 관한 수치적 연구

A low swirl combustor is an alternative technology to sustain a stable flame and to meet the required emission restriction of industrial gas turbine engines. The flame is lifted above the combustor front wall due to a balance between the flame speed and the local flow velocity. The present study is on the numerical simulations of reacting flows of a low swirl model combustor. The flamelet-generated manifold (FGM) method, which is one of the combustion chemistry reduction techniques, is coupled with a large eddy simulation solver. For a premixed CH₄ and air inflow at an equivalence ratio of 0.65, the combustion flow field in the combustor is investigated for two kinds of core blockage configurations, using three-dimensional large eddy simulations based on FGM. Compared to conventional swirl combustors, the low swirl model combustor presents turbulent core flow through the center body of the swirler, which turns out to play an important role in producing stable lifted flames.

A numerical investigation of CO2 dilution on the thermochemical characteristics of a swirl stabilized diffusion flame

The turbulent combustion flow modeling is performed to study the effects of CO2 addition to the fuel and oxidizer streams on the thermochemical characteristics of a swirl stabilized diffusion flame. A flamelet approach along with three well-known turbulence models is utilized to model the turbulent combustion flow field. The k-ω shear stress transport (SST) model shows the best agreement with the experimental measurements compared with other models. Therefore, the k-ω SST model is used to study the effects of CO2 dilution on the flame structure and strength, temperature distribution, and CO concentration. To determine the chemical effects of CO2 dilution, a fictitious species is replaced with the regular CO2 in both the fuel stream and the oxidizer stream. The results indicate that the flame temperature decreases when CO2 is added to either the fuel or the oxidizer stream. The flame length reduction is observed at all levels of CO2 dilution. The H radical concentration indicating the flame strength decreases, following by the thermochemical effects of CO2 dilution processes. In comparison with the fictitious species dilution, the chemical effects of CO2 addition enhance the CO mass fraction. The numerical simulations show that when the dilution level is higher, the rate of the flame length reduction is more significant at low swirl numbers.

Lagrangian Analysis of Flame Dynamics in the Flow Field of a Bluff Body-Stabilized Combustor

Abstract Experiments are performed in a partially premixed bluff body-stabilized turbulent combustor by varying the mean flow velocity. Simultaneous measurements obtained for unsteady pressure, velocity, and heat release rate are used to investigate the dynamic regimes of intermittency (10.1 m/s) and thermoacoustic instability (12.3 m/s). Using wavelet analysis, we show that during intermittency, modulation of heat release rate occurring at the acoustic frequency fa by the heat release rate occurring at the hydrodynamic frequency fh results in epochs of heat release rate fluctuations where the heat release rate is phase locked with the acoustic pressure. We also show that the flame position during intermittency and thermoacoustic instability are essentially dictated by saddle point dynamics in the dump plane and the leading edge of the bluff body.

Simultaneous deflection tomography and PIV measurements of non-premixed combustion

The multi-parameter distributions of non-premixed combustion were investigated in this work by using deflection tomography and particle image velocimetry (PIV). An experimental apparatus composed of a burning system, a deflection tomography system, and a PIV system is initially designed and constructed. Two charge-coupled diode cameras are controlled by a synchroniser to capture Moiré fringes and tracer particles simultaneously, and the velocity contour and streamline map are obtained from the tracer particle distributions and flame images. The temperature distribution is reconstructed from the Moiré fringes based on the wave-front recovery method and Abel-inversion algorithm, and the velocity and temperature distributions are matched in the combustion field by using the pixel registration method. The matching results show that the high-temperature and high-speed zones of the combustion flow field do not coincide under the experimental conditions. Finally, the coupling imaging effect and measurement errors in the multi-parameter measurement are discussed.

Effect of location of a transverse sonic jet on shock augmented mixing in a SCRAMJET engine

Transverse sonic injection into supersonic cross flow is a well-established technique to inject fuel into a SCRAMJET combustor. A SCRAMJET combustor with a backward facing step acting as a flame-holder has been used for this study. The jet is placed at various locations downstream of the step, where each location represents a distinct flow region. Three-dimensional simulations have been performed using Menter's SST model in our in-house parallel 3-D RANS unstructured grid CFD solver. In such a SCRAMJET configuration, mixing between air and fuel is augmented by shocks generated by the under-expanded jet injected into the supersonic cross flow, hence the jet location is expected to be critical. The performance and mixing of the combustor has been quantified for each of the distinct configurations. The length of the combustor required for complete mixing has also been estimated for the different cases. It is observed that the mixing and performance are strongly affected by the location of the jet in the combustor flow-field. From the results presented in this paper, the optimal location for the jet is somewhat before the end of the recirculation region behind the backward facing step.

Numerical investigation of wavy wall strut fuel injector for hydrogen fueled scramjet combustor

Mixing and combustion of a fuel with supersonic airstream in a scramjet combustor is a complex phenomenon because of very less resident time of the air in the combustion chamber. Mixing of fuel and air at supersonic speed and the subsequent combustion are greatly affected by the disturbance of the flow field in the form of shock waves, vortices and recirculation regions. In this research paper, the same concept has been considered by introducing an innovative strut fuel injector for the development of more shock waves and streamline vortices. The basic or standard computational domain of the scramjet combustor is considered from the reference of DLR experimental scramjet. The basic scramjet model consists of the wedge-shaped strut fuel injector. In this research, the strut injector has been re-designed such a way that to generate more oblique shock waves. Numerical analysis of the scramjet internal flow field has been performed with basic and innovative strut by solving the Reynolds-averaged Navier-Stokes equations with the help of computational fluid dynamics tool defined as ANSYS-FLUENT 16.0. The internal flow field of scramjet combustor with basic and innovative strut fuel injectors has been visualized from the analysis of pressure, temperature and velocity along with the analysis of flow structure, shock waves, and streamlines vortices. From the analysis of numerical results, it is identified that multiple numbers of oblique shock waves are being generated from the leading curved edge of the newly introduced strut. Both the pressure and temperature of airstream at the entrance of the combustion chamber are higher in the case of the wavy wall strut and it reduces the ignition delay time as compared to the basic strut model.

Control of NOx and other emissions in micro gas turbine combustors fuelled with mixtures of methane and ammonia

Methane–ammonia mixtures have potentials as low-carbon fuels for gas turbines, however significantly high fuel NOx production in their flames present challenges to their application. This study aims to provide deep insight into the physical and chemical processes involved in the formation and control of emissions from the combustion of CH4NH3–air with up to 30% ammonia by heat fraction in gas turbine combustors. Hence, laser diagnostics techniques such as Particle Image Velocimetry (PIV), and Planar Laser Induced Fluorescence (PLIF) imaging, in addition to Fourier Transform Infrared (FTIR) gas analysis were employed to study the flow field, flame structure and emissions characteristics of a micro gas turbine swirl combustor fuelled with CH4NH3–air mixtures. The control of emissions from the flames was further studied using Large Eddy Simulation (LES) of a model swirl combustor. The results show that NOx emissions from premixed CH4NH3–air in single-stage combustion were more than 5000 ppmv at equivalence ratios, Φ = 0.8–1.1, which is about twice more than the values already reported for NH3–air. Trends in NOx emissions correspond with the trends in OH radicals concentration in the combustor owing to the relevance of OH radicals in fuel NOx production. Emissions control leading to significantly low emissions such as 49 ppmv of NOx, 2 ppmv of CO and approximately zero N2O, HCN and NH3 emissions with a 99.8% combustion efficiency was achieved using rich-lean combustion. An optimum Φ of the primary combustion zone for low NOx emission was identified, which varied from 1.30 to 1.35 depending on the ammonia fraction. For Φ richer (leaner) than the optimum Φ, NOx emission increased due to an increase in NOx production in the secondary (primary) combustion zone. Rich-lean combustion of CH4NH3–air emitted less NOx than that of NH3-air because the higher flame speed of CH4NH3–air mixtures ensured lower NOx production in the secondary combustion zone.

Femtosecond laser-induced plasma spectroscopy for combustion diagnostics in premixed ammonia/air flames.

Ammonia (NH3) is a sustainable fuel with excellent emission characteristics. Hence, it is important to develop diagnostic techniques for NH3 combustion. In this paper, femtosecond laser-induced plasma spectroscopy (FLIPS) was performed in premixed NH3/air flames. The plasma emission spectra induced by the femtosecond laser in the flame and the chemiluminescence spectra of the flame itself were both measured. Through calibration, we found that the spectral intensity ratio of NH (336 nm)/N2 (337nm) could be used for equivalence ratio measurements in NH3 combustion flow fields. This work is the first attempt at using a femtosecond laser-based technique for NH3 combustion diagnostics.

Flow field and injector heat characteristics of hybrid rocket motor with annular-gap injector

A throttleable annular-gap axial injector, which could automatically adjust the injection gap area, is proposed to decrease injection pressure drop variation amplitude during wide range thrust regulation process in hybrid rocket motor. With the annular-gap injector adopted, liquid oxidizer is unhomogeneously injected into combustion chamber through a narrow annular injection gap. In conventional hybrid rocket motor, the oxidizer is usually homogeneously injected through full-inlet injection method. The injection area covers the whole fuel grain port. Difference of the injection methods greatly influences the oxidizer flow characteristics. Consequently, combustion and heat transfer characteristics of the motor are significantly changed. This paper is aimed to analyze the two-phase combustion flow field and coupled injector heat transfer characteristics of a lab-scale hybrid rocket motor with annular-gap injector through two-dimensional axisymmetric steady numerical simulations. The motor adopts 98% hydrogen peroxide and polyethylene as the propellants. Numerical analysis reveals that position of the injection gap influences the regression rate distribution in the first half of solid fuel grain but has little effect on that in the second half. The regression rate is relatively high when the injection gap is close to the fuel inner surface. In addition, the flowing liquid hydrogen peroxide in the injector could cool the chamber head. Sharp turns that produces vortex in oxidizer flow channel decreases the cooling effect and increases the overheating risk. Smooth bend could improve the injector heat transfer characteristics and eliminate the risk.

Using Mobile Health Technology to Improve Well-Being in Caregivers of Patients with Bipolar Disorder: Poster Abstract Submission

A piloted, partially premixed, liquid-fueled swirl burner is operated at high pressure (1 MPa). High-speed (6 KHz) stereoscopic PIV is used to investigate the characteristics of the stagnation line separating the pilot jet and the central recirculation zone (CRZ) with varying pilot-main ratio and global equivalence ratio. The mean curvature of the stagnation line displayed a large spatial scale pattern that was present for all operating conditions. All three components of velocity, in-plane shear, and swirling strength are conditioned upon the instantaneous stagnation line. Mean distributions of the velocity normal to the stagnation line show that velocity is oriented towards the CRZ when the stagnation line is found nearer the centerline of the combustor. The conditioned out-of-plane velocity (w) shows a distinct concentration of large mean and fluctuation RMS values towards the center of the measurement domain. Varying fuel flow does not significantly change this spatial structure, only the magnitudes of the w statistics. The in-plane shear stress was the largest for the pilot biased condition as a stronger shear layer develops. For the leanest flame, large fluctuation RMS values of shear stress were confined to a region where the pilot jet begins to interact more heavily with the main jet. Operating with less pilot fuel flow enhanced the mean conditional swirling strength indicating that the pilot shear layer was shedding more intense eddies. Disregarding spatial relations, a scatter plot of w, shear stress, and swirling strength displayed trends between the variables. The largest swirling strength values coincide with highest magnitude shear stresses and the widest range of w. These conditioned statistics highlight how certain aspects of the combustor flow field are invariant with fuel distribution. This is desirable for aeropropulsive combustors that must maintain stable ignition from a range of conditions from landing/take-off to cruise.

Numerical investigation of inlet guiding ring effect on combustor flow field with multiple swirler structures

ABSTRACT Lean-premixed combustion technology has been adopted widely for heavy-duty gas turbine application. The multiple-swirl structure might shorten the flame brush length and reduce emissions. However, the inlet flow cannot be set equally for each swirl burner. The inlet guiding ring structure should be carefully adjusted and validated to reach the design goal. This paper introduces a typical geometry of heavy-duty combustor, which is designed with multiple swirler structures. The inlet guiding ring is a semicircle type. The parametric analysis is trying to investigate the axial location of the inlet guiding ring effect on the downstream multiple-swirl flow structure. Computational results show that both the Counter-Swirl and Co-Swirl configurations possess quite well aerodynamics which can support the recirculation establishment and flame stabilization. As the axial distance between the inlet guiding ring and swirl nozzle increases from 16 mm to 37 mm, the width of the recirculation will increase, which intensify the turbulence mixing for the pilot and main stage. Meanwhile, the length of the recirculation also increases, and the turbulence intensity keeps relatively strong. When the gap distance increases from 37 mm to 51 mm, the flow phenomenon would return to the condition like the case with an axial distance equal to 16 mm. Therefore, the condition with the axial gap distance 37 mm is quite critical. Thus, the flow field is quite sensitive to the inlet guiding ring location.

Study on ignition process of air/liquid oxygen/ethanol tripropellant coaxial jets in a subscale air heater

Up to the present, no studies have been published to investigate the ignition process in air heater. In this paper, spray combustion flow field during ignition process in air heater was first visualized by high-speed shadowgraph technique. High-speed shadowgraph movies and pressure data indicated that ignition-quenching–reignition phenomenon occurred during one test run. The experimental results demonstrated that combustion of the accumulated propellant mixture during a very short period led to a peak pressure at the moment of last ignition. This process caused reverse flow of flame and ethanol spray into air injector annulus, which was one of the reasons leading to local ablation of injector. In order to obtain reliable ignition, effective improvements were necessary.