Supersonic Flow Field Research Articles

Design of a high-resolution optical measuring ring for supersonic combustion flow field

Design of a high-resolution optical measuring ring for supersonic combustion flow field

Reacting Flow Simulations of a Dual Throat-Dual Fuel Thruster

Numerical simulation of the supersonic turbulent reacting flow field in a dual throat- dual fuel rocket thrust chamber is presented. Future single stage to orbit and high lift space transportation missions aspire a reliable, efficient, and cost-effective propulsion systems. The dual throat-dual fuel concept is a simple altitude compensating propulsion alternative with reusable possibilities. Turbulent reacting supersonic flow field emanating from independent thrust chambers needs to be resolved for a better understanding of the flow structures and design modifications for the performance improvement. The operation of a dual throat nozzle brings about a unique shock train in reacting supersonic flow. Two-dimensional axis-symmetric compressible reacting flow field has been solved using HLLC (Harten, Lax, van Leer, with Contact wave) scheme based finite volume Riemann solver with multi-step finite rate chemistry model for hydrocarbon/hydrogen-oxygen combustion. The computational procedure has been validated with experimental data for species distribution of a coaxial supersonic combustor. Chemical species distribution in the supersonic free shear layer is analyzed in detail to explore the nature of active reaction zones in the flow field.

On the formation and sustenance of the compressible vortex rings in starting axisymmetric jets: A phenomenological approach

The present work numerically investigates the formation and sustenance of a primary vortex ring (PVR) and counter-rotating vortex rings (CRVRs) in laminar, axisymmetric, continuously blowing compressible jets. The effect of jet Reynolds number (Rej), jet Mach number (Mj), and pressure-ratio (pjpa) on these compressible vortex rings and their flow dynamics is explored. The core bow shock is formed in the inviscid core of the transonic/supersonic jets for sustaining the supersonic flow field in the PVR. The embedded shock (ES) and vortex-induced shock (VIS) are also found to sustain the evolving PVR in the supersonic jets. The effect of Mj and pjpa on the circulation and pinch-off of the PVR is also investigated. For over-expanded and perfectly expanded supersonic jets, Kelvin–Helmholtz vortices (KHVs) entrained into the PVR interact with the ES to cause a shock–vortex interaction, which results in the appearance of multiple VISs. It is established that the different events of interaction of KHV with the ES inside the PVR after the pinch-off are a manifestation of two factors: (i) the constraint imposed on the PVR by quasi-time-invariance of circulation around a material curve enclosing the PVR and (ii) the requirement of spatially cyclic property distribution inside the PVR. The shock-shear layer–vortex interaction in under-expanded jets is also investigated using vorticity analysis and oblique shock theory. For under-expanded jets, apart from the Mach-reflection timescale, the CRVR pattern formation is also governed by the initial strength of the slip-stream, which, in turn, is regulated by all three parameters, i.e., Rej, Mj, and pjpa.

Inlet air and fuel flow pressure fluctuation effect on supersonic combustion

PurposeThe purpose of this paper is to analyze the effect of pressure fluctuations on the combustion efficiency of the hydrogen fuel injected into the supersonic oxidizing cross flow. The pressure fluctuations are imposed on inlet air flow and also on the fuel flow stream. Two different situations are considered: the combustion chamber once without and again with the inlet standing oblique shock wave.Design/methodology/approachThe pressure fluctuations are imposed on inlet air flow and also on the fuel flow stream. Two different situations are considered: the combustion chamber once without and again with the inlet standing oblique shock wave. The unsteady turbulent reacting flow solver is developed to simulate the supersonic flow field in the combustion chamber with detail chemical kinetics, to predict the time-variation of the combustion efficiency due to the imposed pressure fluctuations.FindingsThe results show that the response of the reacting flow field depends on both the frequency of fluctuations and the existence of the inlet shock wave. In addition, the inlet standing shock wave has some attenuating role, but the reacting flow shows an amplifying role on imposed oscillations which is also augmented by imposing anti-phase fluctuations on both inlet and fuel flow streams.Originality/valueThis study is performed to analyze the instabilities in the supersonic combustion which has not been considered before in this manner.

Numerical study of supersonic turbulent boundary layer combustion with pressure gradient

Since boundary layer combustion is an effective method to reduce skin friction in supersonic flow-fields, effects of the combustion on typical pressure gradient flow fields are numerically investigated using Reynolds-averaged Navier-Stokes (RANS) simulations. RANS method is validated by comparing the numerical results with the conducted experimental data. Wall heat flux of combustion case is reduced rather than increased although the flame is restricted in boundary layer, especially in adverse pressure gradient state. Compared with the results of no-injection cases, great skin-friction reduction can be obtained by boundary layer combustion, and further reduction can be achieved as pressure gradient increase. Then, the reduction mechanism is analyzed in details. Due to low viscosity gas injection and decline of velocity gradient, the skin friction level is reduced. Moreover, the component of shear stress, Reynolds stress, is decreased due to low density caused by combustion. Additionally, wall velocity laws of combustion coupled with pressure gradient are explored. The results show that the corrected White and Christoph's law is appropriate for velocity description even in sub-layer because the parameters used in the law are obtained at wall and modified by combustion heat release. The research popularizes the application of boundary layer combustion to common pressure gradient situation, and gives advice on wall function selection of turbulent boundary layer combustion.

Visualization and measurement of supersonic jet flow field evolution issuing from a rectangular nozzle with aft-deck

An experimental study is reported of supersonic jet surface flow structure visualization and wall shear stress field measurement issuing from a rectangular nozzle with aft-deck. The near-field surface flow structures evolution from over-expansion to under-expansion with the increase of nozzle pressure ratio (NPR) are successfully captured by surface oil flow visualization and shear sensitive liquid crystal coating (SSLCC) technique. The quantitative measurement result of shear stress vector field obtained by SSLCC shows that shear stress directions change significantly across the shock wave and expansion fans, while the magnitudes of shear stress have no obvious changes. Surface streamlines calculated by SSLCC image keep great consistency with the streamlines visualized using oil flow technique, which demonstrates the accuracy and potential application of SSLCC in supersonic jet surface flow visualization.

Theory, Analysis, and Applications of the Entropic Lattice Boltzmann Model for Compressible Flows.

The entropic lattice Boltzmann method for the simulation of compressible flows is studied in detail and new opportunities for extending operating range are explored. We address limitations on the maximum Mach number and temperature range allowed for a given lattice. Solutions to both these problems are presented by modifying the original lattices without increasing the number of discrete velocities and without altering the numerical algorithm. In order to increase the Mach number, we employ shifted lattices while the magnitude of lattice speeds is increased in order to extend the temperature range. Accuracy and efficiency of the shifted lattices are demonstrated with simulations of the supersonic flow field around a diamond-shaped and NACA0012 airfoil, the subsonic, transonic, and supersonic flow field around the Busemann biplane, and the interaction of vortices with a planar shock wave. For the lattices with extended temperature range, the model is validated with the simulation of the Richtmyer–Meshkov instability. We also discuss some key ideas of how to reduce the number of discrete speeds in three-dimensional simulations by pruning of the higher-order lattices, and introduce a new construction of the corresponding guided equilibrium by entropy minimization.

Experimental research on the shock wave control based on one power supply driven plasma synthetic jet actuator array

The interaction between a ramp induced shock wave and a plasma synthetic jet actuator (PSJA) array was experimentally investigated using high-speed schlieren imaging technique. The PSJA array consisting of 10 actuators was driven by only one power supply. To enhance the actuation intensity, the actuators were located along the streamwise direction to form an accumulation effect. Experimental results showed that the plasma synthetic jet can penetrate into the supersonic flow field with a depth of more than 10 mm. Both the induced shock and the controlling gas bulbs generated by PSJA array can affect the shock wave. The separation shock wave bifurcates when interacting with the shock induced by the actuator. The foot portion of the separation shock wave disappeared when the controlling gas bulbs passed through the interaction region. What's more, as the actuators were located in streamwise direction, the controlling gas bulbs were completely emerged as a much larger gas bulb. As a result, the effective control time was longer than the single actuator and the control effect is also much better.

Numerical investigation of supersonic transverse jet interaction on CPU/GPU system

The main purpose of this paper is to develop a double-precision parallel algorithm implemented on graphics processing units (GPUs) for quick and accurate numerical simulations of large-scale supersonic transverse jet interaction problems. The finite volume method based on structured grid is considered; the AUSM + UP upwind scheme and the explicit multistage Runge–Kutta method are used for spatial discretization and time discretization, respectively. The turbulent solution is solved by K–ω SST two-equation model. Numerical investigation is performed for a supersonic missile body. Numerical results show that performing calculations on GPU can accurately capture the complex wave structures and vortex structures in the supersonic transverse jet flowfield. For single-GPU implementation, parallel computing can achieve an acceleration ratio of 90 times or more, and four GPU parallel computing can achieve an acceleration ratio of 106–245 times. Thus, GPU parallel computing can achieve a large-scale and efficient solution to supersonic transverse jet interaction problems.

Cellular automata ray tracing in two-dimensional aero-optical flow fields

<sec>For the supersonic flow field with large density fluctuation produced by the unsteady flow and turbulent large-scale structures, an effective method to obtain the beam path is to solve the ray equation. Then the optical path difference (OPD), Strehl ratio (SR), optical transmission function (OTF), etc. can be obtained to analyze the optical distortion, and the correction of aero optics effects can be realized to improve the optical system performance. Generally, when the refractive index distribution is arbitrary, the ray equation analytic solution is difficult to obtain.</sec> <sec>Cellular automata (CA) ray tracing algorithm is proposed in this paper for aero-optical calculation in the 2D discrete flow fields. Unlike numerically solving the ray equation (NSRE), the coordinate value and the offset angle are calculated according to the position and direction transformation rules in CA algorithm. The position transformation rule is used to obtain the end point of the beam vector and determine whether the offset angle needs calculating at each iteration, the direction transformation rule is to calculate the offset angles. Then the refractive index field is integrated along the beam path to obtain the optical path length (OPL). The OPD is calculated from OPL. In this paper, aero-optical calculation is based on two types of flow fields. The supersonic shear layer including supersonic mixing layer and boundary layer 2D density distribution is measured by the nano-tracer-based planar laser scattering (NPLS) technique. The supersonic flow field surrounding the optical dome is simulated based on detached-eddy simulation (DES).</sec><sec>The OPD<sub>rms</sub> and program running time quantitatively verify the calculation accuracy and high efficiency of CA. The results show that for the 2D supersonic NPLS flow field and the 2D supersonic flow field surrounding the optical dome, the calculation accuracy of CA is approximately equal to NSRE. Moreover, we find that the program running time of NSRE is about four times that of the CA algorithm, so the efficiency of ray tracing is effectively improved. Thus, the CA algorithm provides a new scheme for aero-optical calculation in the 2D supersonic flow field. Meanwhile, it also provides a guideline for the research on the ray tracing in 3D discrete aero-optical flow field.</sec>

A Numerical Study on the Characteristics of Air–Fuel Mixing Using a Fluidic Oscillator in Supersonic Flow Fields

In this study, numerical simulations were conducted to confirm the possibility of improved mixing performance by using a fluidic oscillator as a fuel injector. Three-dimensional URANS non-reacting simulations were conducted to examine air–fuel mixing in a supersonic flow field of Mach 3.38. The numerical methods were validated through simulations of the oscillating flow generated from the fluidic oscillator. The results show that the mass flow rate and momentum are reduced at the outlet because the total pressure loss increases inside the fluidic oscillator, which means that higher pressure needs to be applied to supply the same mass flow rate. The simulation showed that the flow structure varies over time as the injected flow is swept laterally. With lateral injection, the fuel distribution is long and narrow, and asymmetric vortexes are generated. However, with central injection, the fuel distribution is relatively similar to the case of using a simple injector. Compared to the simple injector, the penetration length, flammable area, and mixing efficiency were improved. However, the total pressure loss in the flow field increases as well. The results showed that the supersonic fluidic oscillator could be fully utilized as a means to enhance the mixing effect, however a method to reduce the total pressure loss is necessary for practical application.

Inversion and reconstruction of supersonic cascade passage flow field based on a model comprising transposed network and residual network

A data-driven model comprising a transposed network and a residual network is proposed to predict the flow field structure of the supersonic cascade passage by measuring the wall pressure of the cascade passage. This model is based on the close relationship between the wall pressure values of the supersonic cascade passage and the flow field structure. Wind tunnel experiments of the supersonic cascade passage were conducted to obtain real experimental data. The transposed network, which is the first part of the model, consists of transposed convolution, normal convolution, and fully connected layers. However, the mapping ability of the transposed network between the wall pressure values and the cascade passage flow field structures was insufficient. The maximum relative error by the transposed network in predicting the position of the leading edge of the shock train was 8.5%. Therefore, the residual network was introduced as the second part of the model to achieve end-to-end super-resolution and improve the mapping ability of the model. After training, the residual network reduced the maximum relative error of the position prediction of the leading edge of the shock train to 3.5%. In the testing set, the prediction results of the model indicated that the movement of the leading edge of the shock train can be distinguished well. Furthermore, a combination of the two networks revealed the mapping relationship between the wall pressure of the supersonic cascade passage and the flow field structure. Thus, the results of this study lay a good foundation for the subsequent determination of the flow field stability margin.

The Effect of Transverse Injection Upstream of an Axisymmetric Aft Wall Angled Cavity in a Supersonic Flow Field

Experiments are performed in a supersonic non-reacting flow facility to investigate the performance of an axisymmetric aft wall angled cavity with upstream fuel injection in a Mach 1.8 flow field. The supersonic combustor has a circular cross sectional duct in which cavities are introduced at a distance of 20 mm from the inlet. The aft wall of the cavity is tapered towards flow downstream and inclined with two step consecutive angles. Flush wall mounted injector is mounted at the upstream of the cavity. The tests are conducted at three fuel injection pressures to simulate the flow field in the present study. The mixing performance of the aft wall angled cavities are analysed based on the momentum flux distribution at the exit of the combustor and the stagnation pressure loss across the combustor and compared with the rectangular cavity. Transverse upstream injection of aft wall angled cavities enhances mixing than rectangular cavities, deliberated with less stagnation pressure loss from the former. Increase in injection pressures resulted in more uniform mixing across the flow direction of the combustor irrespective of the cavity configuration, concurrently induces more stagnation pressure loss due to increase in jet penetration depth into the main stream.

Characteristics of Flow Field for Supersonic Oxygen Multijets with Various Laval Nozzle Structures

The Laval nozzle is used widely in steelmaking processes to increase the velocity of an oxygen jet up to approximately 2 Mach numbers. The present research is aimed at determining the effect of the Laval nozzle structure on a supersonic oxygen jet flow field. Both one-dimensional isoentropic flow theory and a characteristic-line method were used to design five-hole oxygen lances of two types and, applicable to oxygen multijets, to analyze their flow field characteristics by a series of numerical simulations and experimental studies. When compared to the conventional oxygen lance designed according to the one-dimensional isoentropic flow theory, the new oxygen lance designed with employment of the characteristic-line method is more effective in both suppressing the shock wave generation and improving the initial axial velocity of oxygen jets under the same test conditions. As a result, a larger impaction cavity was generated by the new oxygen lance and the mass transfer within oxygen, liquid slag, and molten steel was enhanced.

Study on the Initiation Mechanism of Non‐Premixed Shock Induced Combustion in Supersonic Propellant Gas Jet

AbstractThe reflection of a supersonic propellant gas jet always happens during the firing of the weapon system of aircraft. The reflection induces complex non‐premixed shock induced combustion. It is of great importance to investigate the initiation mechanism of the combustion in the supersonic jet of propellant gas. The mechanism is vital for the design of any suppression method of the secondary combustion. To investigate the combustion in the reflected supersonic jet, a solid propellant combustion model and a numerical model based on compressible multispecies Navier‐Stokes equations are established. A detailed reaction mechanism consisting of 9 species and 12 reactions is utilized for the combustion of the propellant gas. Based on the models, the supersonic precursor flow field and the supersonic propellant gas jet flow are obtained and discussed. A secondary jet is observed which has not been reported before. A sandwich structure is proposed to reveal the initiation of the combustion. The oxygen supply in the secondary jet is found to be dominated by a vortex. An inert gas atmosphere can be adopted to suppress the mass transfer at the contact surface. Special devices could be designed to suppress the formation of the vortex in the secondary propellant jet.

Numerical investigation of scramjet strut injector cooling for different fuel mass fluxes and strut material properties at mach 8 flight conditions

The thermal load on a lobed scramjet (supersonic combustion ramjet) strut injector is investigated numerically for Mach 8 flight conditions. For this purpose, coupled 3D RANS (Reynolds Averaged Navier-Stokes) simulations of the supersonic external flow field around the strut injector and the subsonic strut internal hydrogen flow are performed. The temperature distribution in the solid strut material is simulated by a separate code, which exchanges data with the fluid solvers. Due to the predominantly low Mach numbers inside the strut, an all-Mach number preconditioning is required to enable convergence of the compressible flow solver. The air flow Mach number in the combustion chamber is three. Different strut materials (copper, silicone carbide, molybdenum, and tungsten) with different heat capacities and thermal conductivities are compared. Moreover, the impact of the height of the blunt tip of the strut on the heat load is investigated. Due to a small detached normal shock wave upstream of the strut, the temperatures at the tip become very high; this part is notably difficult to cool. In order to keep the strut temperatures acceptable, it is cooled actively by the hydrogen that is injected for combustion. A sophisticated internal strut structure is needed to direct the hydrogen towards regions of highest heat load. Different strut geometries, fuel mass fluxes (equivalence ratios), and strut materials are investigated.

An Inverse Design Method for Non-uniform Flow Inlet with a Given Shock Wave

For the hypersonic inlet and fore-body integrated design, the non-uniform incoming flow generated by the fore-body will bring a relatively big challenge to the inward-turning inlet design. To make the inlet match the non-uniform incoming flow, this paper, based on previous studies, develops a cross-stream marching plus (CSMP) method, by which an aerodynamic surface used to generate a given shock shape can be acquired. The method can correct such solution points as may give rise to grid distortions or flow-field abnormity and overcome the shortcoming of the insufficient stability of previous methods. Numerical simulation results of the conical supersonic flowfield show that the error obtained from the proposed CSMP method drops with the reduction of the grid dimension and the marching step, being less than 0.05% for reducing the marching step to 10%; that with this method the maximum relative error of the pressure on the profile is less than 0.23%. In the design process of the inward-turning inlets that match the fuselage fore-body, it’s found that in comparison with the results of the inviscid CFD results, the aerodynamic surface designed with the CSMP method can fully generate the given shock wave shape. Thus, the CSMP method provides a new direction for the inlet/fore-body integrated design.

A comprehensive analysis on the structure of groove-induced shock waves in a linear turbine

A shock wave mostly occurs around a compression corner in supersonic flow and, when the corner is followed with another, two shock waves normally generate and, between them, an intersection is possible because of the increased shock angle of the downstream shock. A grooved-surface can generate multiple-shockwave structure in a local supersonic flow field as well. Different from the shockwave structure resulted from the compression corners, its structure may be approximately parallel or divergent when an additional expansion process occurs between two adjacent shock waves. This paper focuses exclusively on the shockwave structures based on both experimental and numerical methods. The physical mechanism for the generation of the approximately parallel and divergent shockwave structures is understood and then the approximately parallel shock wave structure is proved to belong to the same family with the divergent one. In addition, the relationship of the total oblique shock loss with the groove number is analyzed theoretically and numerically in entropy.

Flow control in supersonic flow field based on micro jets

The flow field around supersonic aircraft is usually accompanied by complex flow phenomena, such as shock wave and shock wave/boundary layer interaction, which cause some adverse effects on aircraft performance. Seeking effective flow control methods has been a hot topic for many researchers. As an important method to improve the flow characteristics in supersonic flows, micro jet technology and its control mechanism have been paid much attention. In this article, we used compression corner calculation model and conducted detailed numerical investigations in the supersonic flow field with different injection pressure ratios, various actuation positions, and different nozzle types. The interaction between the micro jets and supersonic upstream flows generates complex flow structures, which contain bow shocks, barrel shocks, Mach disk, counter-rotating vortex pairs, and so on. The flow characteristics with micro jet schemes are superior to those in the no-control case. The controlling performance of micro jet is mainly determined by the following aspects. First, the downwash effect of counter-rotating vortex pairs can bring high-energy fluid into the bottom of the boundary layer to activate low-energy fluid and then strengthen the ability of resisting the flow separations. Second, the bow shock, which is generated upstream of the micro jet, significantly decelerates the downstream flows. Thus, the shock intensity at the corner is weakened and the characteristic of shock wave/boundary layer interaction is improved. In addition, the effective function range of MJ, that is, the distance between the counter-rotating vortex pair and the wall surface, is also an important factor. When both the counter-rotating vortex pairs and the bow shock are further from the wall, the flow characteristics around the corner in a larger area can be improved. Research shows that the micro jet scheme with Laval nozzle gives better controlling effect on shock wave/boundary layer interaction when the injection pressure radio is set to be 0.6, with the actuation location being 20 times the jet outlet diameter upstream of the corner.

Skin Friction Measurements Using Oil Film Interferometry in a 3-D Supersonic Flowfield

The application of oil film interferometry (OFI) for the measurement of global skin friction on the surface beneath a complex three-dimensional supersonic flow is explored in this study. The flowfield is produced by a fin-generated oblique shock interacting with a compressible turbulent boundary layer. Measurements were obtained by modifying a comprehensive acquisition and processing software, previously developed for two-dimensional, low-speed flows, to measure skin friction distributions underneath such flows. This approach required the local flow direction on the surface, and therefore traditional oil flow visualization was coupled with OFI measurements. A cross-correlation-based algorithm was incorporated to extract flow direction from sequences of oil flow visualization images. Photogrammetry was used with both surface oil flow and OFI to ensure proper spatial overlap for the two measurements. Results from this study demonstrate the capability of the approach in a complex, supersonic flow field that contains regions of both separation and reattachment concomitant with a wide range of skin friction values. The skin friction results show excellent homogeneity, standard deviation of 5.1% of the mean, in the undisturbed boundary layer upstream of the interaction and capture well-resolved high shear gradients in the interaction region. A comparison to legacy data from a similar interaction shows matching trends within the developed flow.