Combustion Flow Field Research Articles

An Investigation of Fuel Mixing and Reaction in a CH4/Syngas/Air Premixed Impinging Flame with Varied H2/CO Proportion

For industrial applications, we propose a concept of clean and efficient combustion through burning syngas on an impinging burner. We performed experimental measurements of particle image velocimetry, OH radical (OH*) chemiluminescence, flame temperature, and CO emission to examine the fuel mixing and reaction of premixed impinging flames of CH4/syngas/air with H2/CO in varied proportions. The velocity distribution of the combustion flow field showed that a deceleration area in the main flow formed through the mutual impingement of two jet flows, which enhanced the mixing of fuel and air because of an increased momentum transfer. The deceleration area expanded with an increased CO proportion, which indicated that the mixing of fuel and air also increased with the increased CO proportion. Our examination of the OH* chemiluminescence demonstrated that its intensity increased with increased CO proportion, which showed that the reaction between fuel and air accordingly increased. CO provided in the syngas hence participated readily in the reaction of the CH4/syngas/air premixed impinging flames when the syngas contained CO in a large proportion. Although the volume flow rate of the provided CO quadrupled, the CO emission increased by only 12% to 15%. The results of this work are useful to improve the feasibility of fuel-injection systems using syngas as an alternative fuel.

Study on atomization and combustion characteristics of LOX/methane pintle injectors

Influences of main structural parameters of the LOX/methane pintle injectors on atomization cone angles and combustion performances were studied by experiments and numerical simulation respectively. In addition, improvement was brought up to the structure of the pintle injectors and combustion flow fields of two different pintle engines were obtained. The results indicate that, with increase of the gas-liquid mass flow ratio, the atomization cone angle decreases. In the condition of the same gas-liquid mass flow ratio, as the thickness of the LOX-injection gap grows bigger, the atomization cone angle becomes smaller. In the opposite, when the half cone angle of the LOX-injection gap grows bigger, the atomization cone angle becomes bigger. Moreover, owing to the viscous effects of the pintle tip, with increase of the ‘skip distance’, the atomization cone angle gets larger. Two big recirculation zones in the combustor lead to combustion stability of the pintle engines. When the value of the non-dimensional ‘skip distance’ is near 1, the combustion efficiency of the pintle engines is the highest. Additionally, pintle engines with LOX injected in quadrangular slots can acquire better mixing efficiency of the propellants and higher combustion efficiency as the gas methane can pass through the adjacent slots. However, the annular-channel type of pintle injectors has an ‘enclosed’ area near the pintle tip which has a great negative influence on the combustion efficiency.

Effect of Fin-Guided Fuel Injection on Dual-Mode Scramjet Operation

Supersonic combustion experiments were conducted to determine the effect of a mixing enhancement fin on the combustion operation and reaction zone characteristics of a dual-mode scramjet. This fin-guided fuel injection approach was compared against a baseline that used no fin. A direct-connect vitiated-air facility was used to simulate the total enthalpy of a Mach 4.4 flight condition and generate a Mach 1.9 flow at the isolator inlet. Laboratory-scale tests were performed in a model combustor designed with a cavity flame holder and a 2 deg expansion. The experiments used hydrogen fuel at an overall equivalence ratio ranging between 0.05 and 0.25. The combustor flowfield was qualitatively analyzed using schlieren and chemiluminescence imaging, whereas quantitative assessments were made by comparing wall pressure distributions and exit temperature measurements. Low jet penetration with the baseline supplied the cavity with ample fuel, which concentrated the heat release and thermally choked the flow for an equivalence ratio higher than 0.10. Increased jet penetration with fin-guided injection displaced the reaction downstream of the cavity, enabling supersonic combustion up to an equivalence ratio of 0.15 but generating a flame 81–182% longer compared to the baseline. Practical considerations of using fin-guided fuel injection in a real system are discussed.

Simultaneous reconstruction of 3D refractive index, temperature, and intensity distribution of combustion flame by double computed tomography technologies based on spatial phase-shifting method

In this paper, a transient multi-parameter three-dimensional (3D) reconstruction method is proposed to diagnose and visualize a combustion flow field. Emission and transmission tomography based on spatial phase-shifted technology are combined to reconstruct, simultaneously, the various physical parameter distributions of a propane flame. Two cameras triggered by the internal trigger mode capture the projection information of the emission and moiré tomography, respectively. A two-step spatial phase-shifting method is applied to extract the phase distribution in the moiré fringes. By using the filtered back-projection algorithm, we reconstruct the 3D refractive-index distribution of the combustion flow field. Finally, the 3D temperature distribution of the flame is obtained from the refractive index distribution using the Gladstone–Dale equation. Meanwhile, the 3D intensity distribution is reconstructed based on the radiation projections from the emission tomography. Therefore, the structure and edge information of the propane flame are well visualized.

Numerical analysis of passive techniques for optimizing the performance of scramjet combustor

In this research paper numerical analysis of scramjet combustor has been carried out with different passive techniques to improve the mixing efficiency of supersonic airstream and hydrogen fuel. The actual mixing rate of air and hydrogen is very less because of the less resident time of supersonic air in the combustion chamber. The mixing of hydrogen fuel with supersonic airstream is greatly affected by the development of shock waves, streamline vortices and shear layer of mixing in the flow filed of the combustor. The experimental DLR scramjet model has been considered as a reference and standard model to compare and validate the numerical results. In this research article three different passive techniques are considered and defined as a uniform zigzag surface, small parabola shape cavities and bumps at lower wall of the combustion chamber. The computational scramjet domains are modeled as modified versions of DLR scramjet by the implementation of passive techniques. Two dimensional numerical analyses are carried out by solving Favre-averaged Navier–Stokes governing equations and SST k−ω turbulence model along with finite rate/eddy dissipation combustion chemistry turbulence model is also considered for accurate analysis of streamline vortices, shear layer development and shock waves. From the numerical results of various passive techniques it is found that the development or creation of gradient fluxes of flow property are increased. By incorporating these types of surfaces at combustor bottom wall it is identified that the growth rate of mixing layer has been increased along the flow field of combustor. Hence the more amount of hydrogen fuel being carried out by streamline vortices and velocity gradients or shock waves and this phenomenon improve the mixing efficiency of scramjet combustor and also it leads to increase in combustion efficiency. From the numerical analysis results and discussion it is to be concluded that highest mixing and combustion efficiency is identified with a uniform zigzag (wavy wall) surface combustor wall design.

Study on Line CARS for Temperature Measurement in Combustion Flow Field

Some laser beams meet at a single point by a convex len in normal coherent anti-Stokes Raman scattering (CARS), and the CARS signal with temperature information of the focal piont yields under phase matching. Normal CARS can only get the temperature of one spatial piont in one measurement, which can not meet the needs of deep research on combustion flow field. In order to get more information in one test and improve the measuring capacity of CARS, line CARS (L-CARS) was presented. In L-CARS, convex lens are replaced by cylindrical convex lens to get a focal line, and nearly all the points on the line meet the phase matching. So, the CARS signal of each point on the line yield in one test. Cylindrical convex lens are also used in subsequent beam path to focus the CARS signals into spectrometer, and ICCD camera transfers the signals to computer to acquire the temperature of each point. Then, the measuring capacity of CARS is advanced from piont measuring to line measuring. Experimental results based on plain flame furnace suggest that L-CARS can acquire about 200 points’ temperatures effectively in one test, and the length of the measuring line is about 3.6mm. The spatial resolution is about 18μm and the uncertainty is less than 7%, which is as the same as the ordinary CARS’s.

Interaction mechanisms of fuel momentum with flashback limits in lean-premixed combustion of hydrogen

The impact of fuel momentum on the combustor flow field is studied experimentally in a swirl-stabilized, technically premixed hydrogen flame. The volumetric heating value of hydrogen is about 3.5 times lower compared to natural gas, which leads to significantly higher volumetric fuel flow rates at the same power level. This additional fuel momentum significantly alters the combustor flow field. Therefore, the fuel momentum also affects the combustor stability limits. Previous studies were mostly conducted at perfectly premixed conditions, where the fuel momentum does not alter the combustor flow field. In the current study, non-reacting and reacting combustor flow fields of a technically premixed model combustor injecting fuel in axial direction are recorded. Results reveal a strong impact of fuel momentum on axial velocity distribution at the mixing tube outlet and, thus, on the stability limits. Additionally, OH-PLIF recordings for different flow rates, air preheat temperatures, and equivalence ratios show that the axial location of the upstream flame front, xf, poses a telling estimator for flashback resistance. No flashback was observed when the upstream flame front was located downstream of the mixing tube. However, the flame tip always located upstream of the mixing tube outlet prior to flashback. A high value of xf was, thus, identified as a sufficient condition for flashback resistance. At the investigated conditions, xf is shifted downstream with increasing equivalence ratio due to the added momentum of the fuel flow, thereby superseding any parallel augmentation in the turbulent flame speed. This has been identified as the driving mechanism affecting the combustor stability limit.

Ignition Transition of Coaxial Kerosene/Gaseous Oxygen Jet

ABSTRACTThe pressure and temperature during liquid rocket engine ignition drastically increase because of energy generation from the severe combustion reaction of pure oxidizer and fuel. Studies on the ignition of oxygen/kerosene combustion are scarce. This research observes ignition transition of a gaseous oxygen/kerosene spray by directly visualizing the combustion flow field using a windowed combustor and high-speed shadowgraph imaging technique. The hydrodynamic characteristics of a propellant feed system are investigated before the ignition experiments by analyzing the time responses of propellant injection pressures and a high-speed movie of developing spray during injection. The experiments are performed with a 22.5 ms fuel pre-injection sequence. The high-speed shadowgraph imaging and dynamic pressure measurement results are analyzed. The effects of ignition timing on the ignition transition are focused on. The early ignition timing results in the smoothest ignition and longest ignition delay time from the propellant injection related to an ignitable transient spray condition. The combustion pressure rapidly increases with ignition timing delay. A peak pressure caused by propellant mixture accumulation is also observed. The ignition delay is less than 5 ms, and slightly decreases as the ignition timing is further delayed.

Numerical study of hollow rotating detonation engine with different fuel injection area ratios

Three-dimensional numerical simulations are performed to study rotating detonation engines (RDEs) with a hollow combustor. The computations are based on the reactive Euler equations. The reactions of stoichiometric hydrogen–air mixtures are calculated by the one-step Arrhenius chemistry model. The hollow RDE combustors with different fuel injection area ratios on the head end wall are considered. The fuel injection area ratios range from 55.6% to 88.9%. The number of rotating detonation waves in the flow field of the hollow combustor is found to be related to the fuel injection area ratio. In the hollow combustors with fixed dimensions, the number of detonation waves increases when the fuel injection area ratio is higher. In addition, a comparative analysis of the hollow RDE model and the conventional annular RDE model is performed with respect to the flow field and fuel injection conditions. The propulsive performance of the hollow RDE with different fuel injection area ratios is also evaluated and compared with experimental data and other simulation results. It is demonstrated from several aspects that the removal of the inner wall does not sacrifice the propulsive performance of the RDE.

Spray combustion characteristics of kerosene/bio-oil part II: Numerical study

A numerical method was employed to simulate kerosene with a bio-oil additive derived through a fast pyrolysis process in a model for spray combustion flow field. The influences of various oxidizer velocities and proportions of the bio-oil additive on kerosene spray combustion were investigated. The results indicate that for the kerosene with the bio-oil additive, an increase in the oxidizer velocity decreased and intensified the spray combustion regime mainly because an increase in the spray angle reduced the unit area of the droplets, and the turbulence increased the mixing of fuel and oxidizer. The kerosene spray combustion regimes with the bio-oil additive were reduced when the proportion of the additive was increased, mainly because the bio-oil was derived from the biomass in a fast pyrolysis process; specifically, the condense temperature influenced the volatility of the bio-oil components. Compared with pure kerosene, the volatile bio-oil underwent early vaporization into fuel vapor; this resulted in an early reaction after mixing with the oxidizer and an indirect early kerosene reaction, thereby contributing to the decrease in the kerosene spray combustion regime with the bio-oil additive. Because the heating value of the bio-oil was low, the addition of excessive bio-oil reduced the combustion efficiency.

Advanced Identification of Coherent Structures in Swirl-Stabilized Combustors

We present an application of a newly introduced method to analyze the time-resolved experimental data from the flow field of a swirl-stabilized combustor. This method is based on the classic proper orthogonal decomposition (POD) extended by a temporal constraint. The filter operation embedded in this method allows for continuous fading from the classic POD to the Fourier mode decomposition. This new method—called spectral proper orthogonal decomposition (SPOD)—allows for a clearer separation of the dominant mechanisms due to a clean spectral separation of phenomena. In this paper, the fundamentals of SPOD are shortly introduced. The actual focus is put on the application to a combustor flow. We analyze high-speed particle image velocimetry (PIV) measurements from flow fields in a combustor at different operation conditions. In these measurements, we consider externally actuated, as well as natural dynamics and reveal how the natural and actuated modes interact with each other. As shown in the paper, SPOD provides detailed insight into coherent structures in the swirl flames. Two distinct PVC structures are found that are very differently affected by acoustic actuation. The coherent structures are related to the heat release fluctuations, which are derived from simultaneously acquired OH* chemiluminescence measurements. Besides the actuated modes, a low frequency mode was found that significantly contribute to the global heat release fluctuations.

Numerical simulation of swirling flow effect on the first stage vane film cooling distribution

In advanced gas turbine technology, lean premixed combustion is an effective strategy to reduce peak temperature and thus, NO[Formula: see text] emissions. The swirler is adopted to establish recirculation flow zone, enhancing mixing and stabilizing the flame. Therefore, the swirling flow is dominant in the combustor flow field and has impact on the vane. This paper mainly investigates the swirling flow effect on the turbine first stage vane cooling system by conducting a group of numerical simulations. Firstly, the numerical methods of turbulence modeling using RANS and LES are compared. The computational model of one single swirl flow field is considered. Both the RANS and LES results give reasonable recirculation zone shape. When comparing the velocity distribution, the RANS results generally match the experimental data but fail to at some local area. The LES modeling gives better results and more detailed unsteady flow field. In the second step, the RANS modeling is incorporated to investigate the vane film cooling performance under the swirling inflow boundary condition. According to the numerical results, the leading edge film cooling is largely altered by the swirling flow, especially for the swirl core-leading edge aligned case. Compared to the pressure side, the suction side film cooling is more sensitive to the swirling flow. Locally, the film cooling jet is lifted and turned by the strong swirling flow.

Passive control of thermoacoustic instabilities in swirl-stabilized combustion at elevated pressures

In this study, a porous insert is placed at the dump plane of a swirl-stabilized lean premixed combustor to passively suppress thermoacoustic instabilities. The diffuser-shaped annular ring of porous inert material influences the turbulent flow field directly, including recirculation zones and vortical and/or shear layer structures to passively control the acoustic performance of the combustor. The porous inert material is made of silicon carbide–hafnium carbide coated, high-strength, high-temperature-resistant open-cell foam materials. In this study, the porous insert concept is investigated at above-ambient operating pressures to demonstrate its suitability for practical combustion applications. Experiments are conducted in quartz and metal combustors, without and with the porous insert while varying operating pressure, equivalence ratio, and reactant flow rate. Measurements show that the porous insert, and consequent changes in the combustor flow field, decrease the sound pressure levels at the frequency of combustion instability at all operating conditions investigated in this study. The porous insert also decreases the broadband combustion noise, i.e. the measured sound pressure levels over a wide frequency range.

Numerical Simulations of Static Tested Ramjet Dump Combustor

The flow field of a Liquid Fuel Ram Jet engine side dump combustor with kerosene fuel is numerically simulated using commercial CFD code CFX-11. Reynolds Averaged 3-D Navier–Stokes equations are solved alongwith SST turbulence model. Single step infinitely fast reaction is assumed for kerosene combustion. The combustion efficiency is evaluated in terms of the unburnt kerosene vapour leaving the combustor. The comparison of measured pressures with computed values show that the computation underpredicts (~5 %) pressures for non reacting cases but overpredicts (9–7 %) for reacting cases.

Numerical exploration of mixing and combustion in a dual-mode combustor with backward-facing steps

Dual-mode scramjet combustor is the crucial component for the combined cycle engine of space mission vehicles. The Reynolds Averaged Navier-Stokes (RANS) equations, Spalart-Allmaras turbulence model and the finite-rate reaction model have been utilized to investigate the mixing and combustion in a dual-mode combustor with backward-facing steps, and the species distributions and the Mach number profile of a turbulent diffusion combustion problem have been employed to validate the numerical approach. Moreover, the influences of the injection strategy and the fuel equivalence ratio arrangement on its mixing and combustion flow fields have been explored. The obtained results show that the vertical injection is beneficial for mode transition, and an obvious high pressure region is generated with the vertical injection strategy. Accordingly, its mass-weighted average Mach number drops more sharply. The lower total equivalence ratio is beneficial for the mixing augmentation, as well the smaller fuel equivalence ratio discrepancy. However, the larger total equivalence ratio is beneficial for the mode transition.

5 kHz thermometry in a swirl-stabilized gas turbine model combustor using chirped probe pulse femtosecond CARS. Part 1: Temporally resolved swirl-flame thermometry

Single-laser-shot temperature measurements at 5 kHz were performed in a gas turbine model combustor using femtosecond (fs) coherent anti-Stokes Raman scattering (CARS). The combustor was operated at two conditions; one exhibiting a low level of thermoacoustic instability and the other a high level of instability. Measurements were performed at 73 locations within each flame in order to resolve the spatial flame structure and compare to previously published studies. The measurement procedures, including the procedure for calibrating the laser system parameters, are discussed in detail. Despite the high turbulence levels in the combustor, signals were obtained on virtually every laser shot, and these signals were strong enough for spectral fitting analysis for determination of flames temperatures. The spatial resolution of the single-laser shot temperature measurements was approximately 600 µm, the precision was approximately ±2%, and the estimated accuracy was approximately ±3%. The dynamic range was sufficient for temperature measurements ranging from 300 K to 2200 K, although some detector saturation was observed for low temperature spectra. These results demonstrate the usefulness of fs-CARS for the investigation of highly turbulent combustion phenomena. In a companion paper, the time-resolved fs CARS data are analyzed to provide insight into the temporal dynamics of the gas turbine model combustor flow field.

The hybrid RANS/LES of partially premixed supersonic combustion using G/Z flamelet model

In order to describe partially premixed supersonic combustion numerically, G/Z flamelet model is developed and compared with finite rate model in hybrid RANS/LES simulation to study the strut-injection supersonic combustion flow field designed by the German Aerospace Center. A new temperature calculation method based on time-splitting method of total energy is introduced in G/Z flamelet model. Simulation results show that temperature predictions in partially premixed zone by G/Z flamelet model are more consistent with experiment than finite rate model. It is worth mentioning that low temperature reaction zone behind the strut is well reproduced. Other quantities such as average velocity and average velocity fluctuation obtained by developed G/Z flamelet model are also in good agreement with experiment. Besides, simulation results by G/Z flamelet also reveal the mechanism of partially premixed supersonic combustion by the analyses of the interaction between turbulent burning velocity and flow field.

Numerical analysis of combustion characteristics of hybrid rocket motor with multi-section swirl injection

This paper is aimed to analyse the combustion characteristics of hybrid rocket motor with multi-section swirl injection by simulating the combustion flow field. Numerical combustion flow field and combustion performance parameters are obtained through three-dimensional numerical simulations based on a steady numerical model proposed in this paper. The hybrid rocket motor adopts 98% hydrogen peroxide and polyethylene as the propellants. Multiple injection sections are set along the axis of the solid fuel grain, and the oxidizer enters the combustion chamber by means of tangential injection via the injector ports in the injection sections. Simulation results indicate that the combustion flow field structure of the hybrid rocket motor could be improved by multi-section swirl injection method. The transformation of the combustion flow field can greatly increase the fuel regression rate and the combustion efficiency. The average fuel regression rate of the motor with multi-section swirl injection is improved by 8.37times compared with that of the motor with conventional head-end irrotational injection. The combustion efficiency is increased to 95.73%. Besides, the simulation results also indicate that (1) the additional injection sections can increase the fuel regression rate and the combustion efficiency; (2) the upstream offset of the injection sections reduces the combustion efficiency; and (3) the fuel regression rate and the combustion efficiency decrease with the reduction of the number of injector ports in each injection section.

Effects of Wall-Injection Length on Spray and Combustion in a Coaxial Porous Injector

The mixing performance of an injector determines the combustion performance of a bipropellant rocket engine. In the previous study, a novel coaxial porous injector concept was developed to improve the mixing performance of a shear coaxial injector. It demonstrated a higher characteristic velocity efficiency than that of a typical shear coaxial injector. To clarify the efficiency-improving mechanism of a coaxial porous injector, the combustion flowfields of both injectors were visualized using a high-speed shadowgraph technique. The effect of geometry factors on the spray core length and the combustion performance as well as the differences in the reactive spray behavior between the coaxial porous injector and the shear coaxial injector were observed. The coaxial porous injector with a different wall-injection length showed higher characteristic velocity efficiencies, shorter spray core lengths, and thicker phase-change patterns at the spray core boundary. Changes in wall-injection length had a nonmonotonic effect on both the spray core length and the characteristic velocity efficiency. As opposed to the expected results, the coaxial porous injector with the highest combustion efficiency showed the longest spray core length. The phase-change pattern width, caused by the evaporation of liquid fuel droplets, was proportional to the efficiency.

Experimental investigation on ignition schemes of partially covered cavities in a supersonic flow

In this study, ignition schemes of the partially covered cavity in a scramjet combustor were investigated under inflow conditions of Ma=2.1 with stagnation pressure P0=0.7Mpa and stagnation temperature T0=947K. It reveals that the ignition scheme of the partially covered cavity has a great impact on the ignition and flame stabilization process. There always exists an optimized global equivalence ratio of a fixed ignition scheme, and the optimized global equivalence ratio of ignition in the partially covered cavity is lower than that of the uncovered cavity. For tandem dual-cavities, ignition in the partially covered cavity could be enhanced with the optimization of global equivalence ratio. However, ignition in the partially covered cavity would be exacerbated with further increasing the global equivalence ratio. The global equivalence ratio and the jet penetration height have a strong coupling with the combustion flow-field. For multi-cavities, it is assured that fuel injection on the opposite side could hardly be ignited after ignition in the partially covered cavity even with the optimized global equivalence ratio. It is possible to realize ignition enhancement in the partially covered cavity with the optimization of global equivalence ratio, but it is not beneficial for thrust increment during the steady combustion process.