A Computational Study of Thrust Vectoring Control Using Dual Throat Nozzle

Dual throat nozzle （DTN） is fast becoming a popular technique for thrust vectoring. The DTN is designed with two throats, an upstream minimum and a downstream minimum at the nozzle exit, with a cavity in between the upstream throat and exit. In the present study, a computational work has been carried out to analyze the performance of a dual throat nozzle at various mass flow rates of secondary flow and nozzle pressure ratios （NPR）. Two-dimensional, steady, compressible Navier-Stokes equations were solved using a fully implicit finite volume scheme. The present computational results were validated with available experimental data. Based on the present results, the control effectiveness of thrust-vectoring is discussed in terms of the thrust coefficient and the coefficient of discharge.     

Study on Transonic Tone in an Axisymmetric Supersonic Nozzle

This paper describes experimental and numerical works to investigate noise phenomenon in supersonic flow dis- charged from a convergent-divergent nozzle. The noise phenomenon of flow is generated by an emission of ＇transonic tones＇. The results obtained show that the frequency of a transonic tone, that differs from the frequency of a screech tone due to the shock-cell structures in a jet and originates in the shock wave in the nozzle, increases in proportion to the nozzle pressure ratio. The high-order transonic tone has the directivity in the direction of the flow. As for the transonic tone＇s frequency, the separated zone was calculated by using a simple flow model con- sidering the propagating perturbation. The results of the model corresponded to the results of this experiment well.     

Effect of plate position and size for self-induced flow oscillation of underexpanded supersonic jet impinging on perpendicular plate

When the underexpanded supersonic jet impinges on the obstacle, it is well known that the self-induced flow oscillation occurs at the specific condition of the pressure ratio in the flowfield, the position of an obstacle and so on. This oscillation is related with the noise problems of aeronautical and other industrial engineering so that the characteristic and the mechanism of self-induced flow oscillation have to be cleared to control the various noise problems. But, it seems that the characteristics of t...     

Experimental Study on Shock Wave Structures in Constant-area Passage of Cold Spray Nozzle

Cold spray is a technique to make a coating on a wide variety of mechanical or electric parts by spraying solidparticles accelerated through a high-speed gas flow in a converging-diverging nozzle.In this study,pseudo-shockwaves in a modeled cold spray nozzle as well as high-speed gas jets are visualized by schlieren technique.Theschlieren photographs reveals the supersonic flow with shock train in the nozzle,Static pressure along the barrelwall is also measured.The location of the head of pseudo-shock wave and its pressure distribution along the noz-zle wall are analytically explained by using a formula of pseudo-shock wave.The analytical results show that thesupersonic flow accompanying shock wave in the nozzle should be treated as pseudo-shock wave instead of nor-mal shock wave.     

The effect of diffuser angle on the discharge coefficient of a miniature critical nozzle

Many researches on critical nozzles have been performed to accurately measure the mass flow
rate of gas flow,and to standardize the performance as a flow meter.Recently,much interest is being
paid on the measurement of very small mass flow rate in industry fields such as MEMS
applications.However,the design and performance data of the critical nozzles obtained so far have
been applied mainly to the critical nozzles with comparatively large diameters,and the works available
on miniature critical nozzles are lacking.In the present study,a computational fluid dynamics method
has been applied to investigate the influence of the diffuser angle on discharge coefficient of the
miniature critical nozzles.In computations,the throat diameter of critical nozzle is varied from 0.2 mm to
5.0 mm and the diffuser angle is changed from 2 deg to 8 deg.The computational results are
validated with some experimental data available.The results show that the present computational
results predict appropriately the discharge coefficient of the gas flows through miniature critical
nozzles.It is known that the discharge coefficient is considerably influenced by the diffuser angle,as
the throat diameter of nozzle becomes small below a certain value.This implies that the miniature
critical nozzles should be carefully designed.     

Numerical study on carbon dioxide removal from the hydrogen-rich stream by supersonic Laval nozzle

A hydrogen purification system with a supersonic nozzle pretreatment process is proposed to improve the performance of traditional CO₂ removal processes from hydrogen-rich streams. A mathematical model of the H₂–CO₂ double-component condensation was established to investigate the feasibility of CO₂ capture in a hydrogen-rich stream using a supersonic nozzle. Compared to the single-phase model, this model is more similar to the objective flow facts and can effectively correct the deviation of the single-phase flow model by 21.4%. Furthermore, the parameters in the H₂–CO₂ double-components spontaneous condensation process in the nozzle were analyzed, and the microscopic mechanism of CO₂ spontaneous condensation was clarified. Finally, the effects of the inlet parameters on the carbon capture efficiency were analyzed. The results indicated that the nozzle is more suitable for purifying hydrogen-rich streams with low temperatures and high carbon content, confirming the possibility of using a supersonic nozzle as a carbon capture method.     

Numerical investigation of the effects of ITD length on low pressure nozzle

The advantage of high efficiency,low SFC (Specific Fuel Consumption),ultra-high bypass ratio turbofan engine attracts more and more attention in modem commercial engine.The intermediate turbine duct (ITD),which connects high pressure turbine (HPT) with low pressure turbine (LPT),has a critical impact on the overall performance of turbine by guiding flow coming from HPT to LPT without too much loss.Therefore,it becomes more and more urgent to master the technique of designing aggressive,even super-aggressive ITD.Much more concerns have been concentrated on the duct.However,in order to further improve turbine,LPT nozzle is arranged into ITD to shorten low pressure axle.With such design concept,it is obvious that LPT nozzle flow field is easily influenced by upstream duct structure,but receives much less interests on the contrary.In this paper,numerical method is used to investigate the effects of length of ITD with upstream swirl blades on LPT nozzle.Nine models with the same swirl and nozzle blades,while the length of ITD is the only parameter to be changed,will be discussed.Finally,several conclusions and advices for designers are summarized.After changing axial length of ducts,inlet and outlet flow field of nozzle differs,correspondingly.On the other hand,the shearing stress on nozzle blade (suction and pressure) surface presents individual feature under various inlet flow.In addition to that,"Clocking-like effect" is described in this paper,which will contribute much to the pressure loss and should be paid enough attention.     

Characteristic and mechanism of pressure fluctuation caused by self-induced oscillation of supersonic impinging jet

When the underexpanded supersonic jet impinges on the obstacle, it is well known that the self-induced flow oscillation occurs. This oscillation depends on the pressure ratio in the flowfield, the position of an obstacle and is related with the noise problems of aeronautical and other industrial engineering. The characteristic and the mechanism of self-induced flow oscillation, have to be clarified to control various noise problems. But, it seems that the characteristics of the oscillated flowfield and the mechanism of an oscillation have to be more cleared to control the oscillation. This paper aims to clarify the effect of the pressure ratio and the obstacle position and the mechanism of self-induced flow oscillation by numerical analysis and experiment, when the underexpanded supersonic jet impinges on the cylindrical body. From the result of this study, it is clear that occurrence of the self-induced flow oscillation depends on the pressure balance in the flowfield.     

Numerical Simulation of the Supersonic Flows in the Second Throat Ejector —Diffuser Systems

INTRODUCTIONThe ejector system is a device which employs ahigh-velocity primary motive fluid to enirain and accelerate a slower moving secondary fluid. The resulting kinetic energy of the mixture is subsequently usedfor self-compression to a higher pressure, thus performing the function of a compressor. The ejectorsystem has lOng been applied to jet pumps, vacuumpumps, high-altitude simulators, V/STOLs, etc[lrv4],because of the major advantages of its structural simplicity and reliabili…     

Navier—Stokes Computations of the Supersonic Ejector—Diffuser System with a Second Throat

The supersonic ejector-diffuser system with a second throat was simulated using CFD.An explicit finite volume scheme was applied to solve two-dimensional Navier-Stokes equations with standard κ-εturbulence model.The vacuum performance of the supersonic ejector-diffuser system was investigated by changing the ejector throat area ration and the operating pressure ratio.Two convergent-divergent nozzles with design Mach number of 2.11 and 3.41 were selected to give the supersonic operation of the ejector-diffuser system.The presence of a second throat strongly affected the shock wave structure inside the mixing tube as well as the spreading of the under-expanded jet discharging from the primary nozzle.There were optimum values of the operating pressure ratio and ejector throat area ratio for the vacuum performance of the system to maximize.     

Application of Rainbow Schlieren Deflectometry for Axisymmetric Supersonic Jets （Comparison of Experiments with Numerical Analysis）

This paper analyzes the correctly-expanded supersonic jet from a convergent-divergent
axisymmetric nozzle by using numerical simulation of turbulent flow.And the calculated density
distributions in this flow are compared with the present experimental data using rainbow schlieren
deflectometry.The value of the density from the experimental data agrees well with the results
calculated by this simulation.Therefore,the present method of the measurement using rainbow
schlieren deflectometry is useful for the measurement of the density of the correctly-expanded
supersonic jet.     

Effect of shock waves on supersonic film cooling with a slotted wall

A slotted wall with a cavity which reduces the effect of the shock wave on the film cooling was developed through understanding of the mechanism by which the shock wave affects the supersonic film cooling. Numerical results show that the supersonic film cooling effectiveness with the slotted wall is improved after the shock wave incidence, even better than that without the shock wave effect. The cooling stream flows into the cavity upstream of the slotted wall and flows out downstream, which bypasses more cooling gas to protect the surface downstream after the shock wave incidence, which weakens the effect of the shock wave on the film cooling. Upstream of the shock wave incidence, the slotted wall reduces the mixing between the mainstream and the cooling stream and the coolant boundary layer thickness, which reduces the film cooling effectiveness for both structures than without the slotted wall, with an effectiveness which nearly the same as or even a little better than without the slotted wall for another structure.     

The influence of annular lobe-injector on the fuel mixing of hydrogen jet at supersonic combustion chamber

The performance of the engine highly depends on the fuel mixing process as a significant process to achieve efficient supersonic flight. Current article has attempted to release the effects of different annular lobe-injectors on fuel mixing when Ma>1. Three various annular jet nozzles are expansively investigated for injection of the sonic hydrogen jet at supersonic air crossflow with Mach-4. Comprehensive comparison of the jet structure of these models are performed through the evaluation of Mach and fuel concentration downstream of these lobe-injectors. Comparison of mixing efficiency also indicates that the nozzle with 3-lobe configuration has 25% more fuel mixing performance than other configurations. Our findings also show that mixing performance of annular lobe-injector is about 15% more than simple one for cases with 2-lobe and 4-lobe injectors.     

Uncertainty-quantification analysis of the effects of residual impurities on hydrogen–oxygen ignition in shock tubes

This study addresses the influences of residual radical impurities on the computation and experimental determination of ignition times in H₂/O₂ mixtures. Particular emphasis is made on the often-times encountered problem of the presence of H-atoms in the initial composition of H₂/O₂ mixtures in shock tubes. Two methods are proposed for quantifying experimentally H-residual impurities in shock tubes, namely, an a priori method that consists of detecting OH traces upon shocking unfueled mixtures, and a posteriori method in which the amount of impurities is inferred by comparing fueled experimental autoignition data with calculations. A stochastic Arrhenius model that describes the amount of H-radical impurities in shock tubes is proposed on the basis of experimental measurements as a function of the test temperature. It is suggested that this statistical model yields a probability density function for the residual concentration of hydrogen radicals in standard shock tubes. Theoretical quantifications of the uncertainties induced by the impurities on autoignition times are provided by using the 5-step short chemistry of Del Álamo et al. [1]. The analysis shows that the relative effects of H-impurities on delay times above crossover become more important as the dilution increases and as the temperature and pressure decrease. Below crossover, the effects of H-impurities on the ignition delay vanish rapidly, and are negligible compared to the departures produced by the non-ideal pressure rise that is seen in some shock-tube experiments at such low temperatures. The influences of kinetic uncertainties on the ignition time are typically negligible compared to the effects of the uncertainties induced by H-impurities when the short mechanism is used, except for air at high temperatures for which kinetic uncertainties dominate. Furthermore, calculations performed with the short mechanism show that correlations between the uncertainties in the rates of branching and termination steps have only some small influences on the ignition-time variabilities near crossover, where a global sensitivity analysis shows an increasing importance of the recombining kinetics. Computational quantifications of uncertainties are carried out by using numerical simulations of homogeneous ignition subject to Monte-Carlo sampling of the concentration of impurities. For the conditions analyzed, these computations show that the variabilities produced in ignition delays by the uncertainties in H-impurities are comparable to the experimental data scatter and to the effects of typical uncertainties of the test temperature when the Stanford chemical mechanism [2] is used. The calculations also unveil that the utilization of two other different chemical mechanisms, namely San Diego [3] and GRI v3.0 [4], yields variations in the ignition delays which are within the range of the uncertainties induced by the H-impurities. Finally, the effects of residual impurities in kinetic-isolation experiments and in supersonic-combustion ramjets are briefly discussed.     

Numerical investigation of the auto-ignition of transient hydrogen injection in supersonic airflow

The auto-ignition process in high Mach number airflow is performed by Large Eddy Simulation. The predictions are in good agreement with experiment. The temporal and spatial analysis of the auto-ignition process are carried out. The results show that the auto-ignition process can be divided into five stages in terms of time sequence, and there are three regions in space at the steady state. It is also found that the maximum mass fraction of HO₂ is one of the best indicators to see whether the scramjet working on high Mach number reaches steady state, and the flame stabilization is dominated by auto-ignition. The heat release rate of different combustion modes is calculated. It reveals that the supersonic combustion mode dominates in all stages and regions, but premixed combustion mode dominates only in transient auto-ignition stage and the auto-ignition region.     

Effect of Mach number on thermoelectric performance of SiC ceramics nose-tip for supersonic vehicles

This paper focus on the effects of Mach number on thermoelectric energy conversion for the limitation of aero-heating and the feasibility of energy harvesting on supersonic vehicles. A model of nose-tip structure constructed with SiC ceramics is developed to numerically study the thermoelectric performance in a supersonic flow field by employing the computational fluid dynamics and the thermal conduction theory. Results are given in the cases of different Mach numbers. Moreover, the thermoelectric performance in each case is predicted with and without Thomson heat, respectively. Due to the increase of Mach number, both the temperature difference and the conductive heat flux between the hot side and the cold side of nose tip are increased. This results in the growth of the thermoelectric power generated and the energy conversion efficiency. With respect to the Thomson effect, over 50% of total power generated converts to Thomson heat, which greatly reduces the thermoelectric power and efficiency. However, whether the Thomson effect is considered or not, with the Mach number increasing from 2.5 to 4.5, the thermoelectric performance can be effectively improved.     

Numerical study on spontaneous ignition of pressurized hydrogen release through a length of tube

The issue of spontaneous ignition of highly pressurized hydrogen release is of important safety concern, e.g. in the assessment of risk and design of safety measures. This paper reports on recent numerical investigation of this phenomenon through releases via a length of tube. This mimics a potential accidental scenario involving release through instrument line. The implicit large eddy simulation (ILES) approach was used with the 5th-order weighted essentially non-oscillatory (WENO) scheme. A mixture-averaged multi-component approach was used for accurate calculation of molecular transport. The thin flame was resolved with fine grid resolution and the autoignition and combustion chemistry were accounted for using a 21-step kinetic scheme.The numerical study revealed that the finite rupture process of the initial pressure boundary plays an important role in the spontaneous ignition. The rupture process induces significant turbulent mixing at the contact region via shock reflections and interactions. The predicted leading shock velocity inside the tube increases during the early stages of the release and then stabilizes at a nearly constant value which is higher than that predicted by one-dimensional analysis. The air behind the leading shock is shock-heated and mixes with the released hydrogen in the contact region. Ignition is firstly initiated inside the tube and then a partially premixed flame is developed. Significant amount of shock-heated air and well developed partially premixed flames are two major factors providing potential energy to overcome the strong under-expansion and flow divergence following spouting from the tube.Parametric studies were also conducted to investigate the effect of rupture time, release pressure, tube length and diameter on the likelihood of spontaneous ignition. It was found that a slower rupture time and a lower release pressure will lead to increases in ignition delay time and hence reduces the likelihood of spontaneous ignition. If the tube length is smaller than a certain value, even though ignition could take place inside the tube, the flame is unlikely to be sufficiently strong to overcome under-expansion and flow divergence after spouting from the tube and hence is likely to be quenched.     

Investigation of the effect of chemistry models on the numerical predictions of the supersonic combustion of hydrogen

In this numerical study, the influence of chemistry models on the predictions of supersonic combustion in a model combustor is investigated. To this end, 3D, compressible, turbulent, reacting flow calculations with a detailed chemistry model (with 37 reactions and 9 species) and the Spalart-Allmaras turbulence model have been carried out. These results are compared with earlier results obtained using single step chemistry. Hydrogen is used as the fuel and three fuel injection schemes, namely, strut, staged (i.e., strut and wall) and wall injection, are considered to evaluate the impact of the chemistry models on the flow field predictions. Predictions of the mass fractions of major species, minor species, dimensionless stagnation temperature, dimensionless static pressure rise and thrust percentage along the combustor length are presented and discussed. Overall performance metrics such as mixing efficiency and combustion efficiency are used to draw inferences on the nature (whether mixing- or kinetic-controlled) and the completeness of the combustion process. The predicted values of the dimensionless wall static pressure are compared with experimental data reported in the literature. The calculations show that multi step chemistry predicts higher and more wide spread heat release than what is predicted by single step chemistry. In addition, it is also shown that multi step chemistry predicts intricate details of the combustion process such as the ignition distance and induction distance.     

Numerical study on supersonic combustion of hydrogen and its mixture with Ethylene and methane with strut injection

In this paper, supersonic combustion and flow field of hydrogen and its mixture with ethylene and methane from strut injections in a Mach 2 supersonic flow are studied numerically. The fuel mixture of hydrogen, methane and ethylene represents the major products of pyrolysis of hydrocarbon fuels with large molecules such as kerosene as it acts as coolant and flows through cooling channels and absorbs heat. Detached Eddy Simulation with a reduced kinetic mechanism and steady flamelet model are applied to simulate turbulent combustion. The calculated temperature profiles of hydrogen are compared to the experimental results of DLR supersonic combustor for validation of the present numerical method. The supersonic combustion flows associated with shock waves, turbulent vortices and flame structures are studied. With addition of methane and ethylene, the flame zone moves further downstream of the strut and the maximum flow temperature at chamber exit decreases by 200 K. With analysis of total temperature ratios, it is found that combustion efficiency for hydrogen combustion is 0.91 and it decreases to 0.78 for the fuel mixture. The calculation of ignition delay time and flame speed reveals that fuel mixture of hydrogen and hydrocarbons has considerably larger delay time and smaller flame speed, that contributes to the weakened flame zone and lower combustion efficiency.