Supersonic Diffuser Research Articles

Gas dynamics at starting and terminating phase of a supersonic exhaust diffuser with a conical nozzle

This research investigates the process of starting and breakdown of second throat diffuser during high-altitude test of conical nozzle with a high expansion ratio. The subscale experimental setup includes a conical nozzle with an expansion ratio of 53, plus a second throat diffuser with a contraction ratio of 1.85, using compressed air as the working fluid. Numerical simulation has been employed to identify the flow physics during the unsteady process of diffuser startup and breakdown. According to the results, the starting process is divided into three distinct stages. In the first stage, the exhausted flow from the nozzle enters the vacuum chamber, leading to an increase in vacuum chamber pressure. The second stage, corresponding to the period before full flow establishment in the conical nozzle, exhibits a relatively constant slope. In the final stage, associated with the transition from over-expanded to under-expanded states, the slope of the rate of evacuation development descends. Next, the vacuum degradation in termination process has been analyzed, and it has been found to include three stages: high slope, middle slope, and low slope. The flow physics during the start process, similar to the results observed in other conical nozzles, exhibits only a Mach reflection structure. However, during the termination process, the flow physics involve a combination of both structures, including Mach reflection and Cap Shock. The results indicate that during the start, an internal shock only interacts with the separation shock, and no special change occurs in the Mach reflection structure. In contrast, during the termination process, unlike what has been reported in previous studies on conical nozzles, the structure of cap shock waves and restricted shock separation patterns are also observed. Another distinction between the starting and termination processes is related to the pressure distribution in the diffuser wall. The wall pressure at the diffuser inlet during the termination process has been reported to be 90% higher than during the starting process.

Design and analysis of a supersonic diffuser for altitude testing of upper-stage liquid rocket engines

This study conducted a numerical simulation of high-altitude environments for upper-stage liquid rocket engines that use second-throat supersonic diffusers. To obtain the minimum second-throat area, a typical normal shock model was modified and validated using two experimental procedures in an existing high-altitude facility: a cold flow test with a conical nozzle, and a hot test of a 7-ton class liquid engine. The modified normal shock model, known as the passive diffuser model, was employed to design a new second-throat supersonic diffuser and to numerically test it under various conditions. In the numerical analysis, a velocity parameter, such as the Mach number, should be maintained constant with minimal v-velocity at the entrance of the diffuser contraction to realize a good agreement with the quasi-one-dimensional estimation. In the full-scale facility, geometrical parameters of the second-throat supersonic diffuser normally have a marginal effect on the performance of the liquid rocket engine under test, unless they affect the starting condition of the second-throat supersonic diffuser. It is possible to reduce the throttling of the new upper-stage liquid rocket engine to 63 % by utilizing the self-evacuation feature of the newly designed second-throat supersonic diffuser. We find that, as throttling decreases, the hot region shifts to the second throat, necessitating the use of a cooling system to protect the entire second-throat supersonic diffuser from thermal damage.

Numerical Simulations of Supersonic Exhaust Diffuser At Various Operating Conditions

A numerical study was carried out to determine starting characteristics of a straight cylindrical supersonic exhaust diffuser of a high altitude simulation of high expansion rocket nozzle. The starting conditions of the supersonic exhaust diffuser are numerically simulated for contour nozzle of exit-to-throat area ratio of 21.77, diffuser diameter-to-nozzle throat area ratio of 23.36, and diffuser length-to-diameter ratio of 15. The evacuation performances of supersonic exhaust diffuser at various operating conditions are investigated with help of the Mach contour plots. The Mach variations along the centre line and wall of the diffuser indicate the location of the shock and flow characteristics. Validation of numerical simulation is carried out with experimental data and shows a good agreement between them. Flowfield inside the straight cylinder supersonic exhaust diffuser is presented out for stagnation to ambient pressure ratio of 5.5 - 28. The nozzle is started at the stagnation to ambient pressure ratio of about 17. The straight cylindrical supersonic exhaust diffuser is started at the stagnation pressure to ambient pressure ratio of about 21. The numerical simulations of the flowfield are able to capture the flows filled in the diffuser and the shock system is fully established in the duct.

Numerical Investigations At High Altitudes For Nozzles Performance Monitoring

Large area ratio nozzles that are employed for space propulsion applications are not meant to be tested at the ground-level conditions because of the flow conditions that occurs in their divergent section due to their low exhaust pressures. Therefore, it is desired to perform the experimental evaluation at high-altitude to evaluate the performance of the said nozzles. Generally, a high-altitude test facility, consisting of a supersonic exhaust diffuser, is generally employed. In thispaper, a second-throat exhaust diffuser has been numerically investigated to predict its minimum starting pressure, and to understand the flow physics in the diffuser using Computational Fluid Dynamics (CFD). Numerical computation of the flow field in the diffuser system is done for the two cases: 1) when the nozzle and the diffuser system are initially evacuated to a low pressure, and 2) when the nozzle and the diffuser system are initially at ambient pressure (1 bar). Simulations have been carried out for cold flow situation (Ȗ=1.4) over a range of nozzle inlet stagnation pressures. Numerical results compare favorably with the theoretical and experimental results and provide adequate insight to the flow physics and internal shock structures formed in thediffuser in addition to minimum starting pressure, diffuser wall pressure and vacuum thrust.

Principles of Magnetohydrodynamical Control of Internal and External Supersonic Flows

This paper demonstrates the possibility of active magnetohydrodynamic (MHD) control of supersonic flows containing shock waves. The shock wave configurations that occur at the inlet to a supersonic diffuser and in front of a streamlined semicylindrical model are used for the purpose of investigation. The impact is carried out by organizing local gas discharge regions when applying a magnetic field transverse to gas discharge currents. It has been shown that by changing the local region of application, the intensity and the direction of the gas discharge currents, it is possible to change the intensity and direction of the ponderomotive force acting on the gas flow during MHD interaction. The ponderomotive force control allows for acting locally on the shape and position of shock waves, the speed and direction of the flow, and the increase or reduction of pressure near the surface of the streamlined body. The experiments were carried out on a gas dynamic setup based on a shock tube in a gas dynamic path, capable of creating supersonic flows in a wide range of Mach numbers at M = 4–7. There was a possibility of organizing the electric and pulsed magnetic fields with an intensity of up to 1.5 T. The given experimental Schlieren flow patterns and the analysis of the obtained data demonstrate the MHD effect on: the change in the angle of inclination of the attached shocks, both into increase and decrease; the bow shock wave approaching the body or the removal from it; and the change in the aerodynamic drag and lift force of the streamlined bodies.

Design and Parametric Analysis of a Mach 4 Wind Tunnel for SSLCs Studies

Abstract The use of shear sensitive liquid crystals (SSLCs) provides a nonintrusive technique for accurate visualization and quantification of shear stress on surfaces under investigation. However, investigations using liquid crystal involving supersonic and hypersonic flows poses specific requirements for the design of wind tunnel. These constraints are pertaining to time response, erosion, and repeated application of the layer of SSLCs on the model surface as well as its sensitivity to wall temperature under the influence of high-speed flows. In this research, the Mach 4 wind tunnel facility is designed using existing empirical, semi-empirical methods and detailed literature survey and is analyzed numerically in ansysfluent 19.0 environment using transient k–ω shear stress transport Reynolds-averaged Navier–Stokes turbulence model. The influence of geometric contouring of converging section of the supersonic nozzle on the reservoir pressure requirement to establish a M = 4 flow condition in test section is also analyzed. The iterative design method is adopted for the supersonic diffuser section of the wind tunnel where different geometric parameters are varied to achieve higher diffuser efficiency as well as minimum startup time requirement. The parametric study is carried out to ascertain the factors influencing test section wall shear stress during wind tunnel startup process and a novel approach is explored in which the Coanda effect is generated to keep the jet emanating from the supersonic nozzle away from the applied SSLC layer, thereby reducing its erosion effect. Furthermore, the influence of the location of a diaphragm for the partially ruptured case on the wind tunnel startup is also critically analyzed.

Газодинамическое проектирование и численное исследование сверхзвукового контура аэродинамической трубы

For an advanced trisonic wind tunnel of a straight-flow type with a test section size of 1.2 × 1.2, intended for ground tests of rocket-space and aviation aircraft models, we implemented a gas-dynamic design of the circuit and did numerical simulation of the flow for the main supersonic regimes (M = 2, M = 4). The gas-dynamic design of the wind tunnel circuit was carried out on the basis of techniques developed at TsAGI and operating experience of existing facilities. The study considers both traditional configurations of the duct with the bending of the walls of all elements, i.e., nozzle, test section and diffuser in the XY plane, and alternative design developments with the bending of the diffuser walls in the XZ plane. When carrying out numerical studies in all areas of the wind tunnel, the ANSYS Fluent software package was used to solve the Navier --- Stokes equations for viscous and heat-conducting air using the turbulence model, i.e., Spalart --- Allmaras, SST. The paper investigates the effect of the wall opening angle, compensating the increasing thickness of boundary-layer longwise displacement, on the flow characteristics; shows the possibilities of obtaining a sufficiently uniform flow with the Mach number accuracy ΔM = ± 0.005 in the area of the model, and analyzes the influence of geometric parameters and boundary conditions on the efficiency of the supersonic diffuser

Passive Control of Base Pressure: A Review

In the present world, passive control finds application in various areas like flow over blunt projectiles, missiles, supersonic parallel diffusers (for cruise correction), the engine of jets, static testbeds of rockets, the ports of internal combustion engines, vernier rockets, and single expansion ramp nozzle (SERN) rockets. In this review, various passive control techniques to control the base pressure and regulate the drag force are discussed. In the study, papers ranging from subsonic, sonic, and supersonic flow are discussed. Different types of passive control management techniques like cavity, ribs, dimple, static cylinder, spikes, etc., are discussed in this review article. This study found that the passive control device can control the base pressure, resulting in an enhancement in the base pressure and reducing the base drag. Also, passive control is very efficient whenever there is a favorable pressure gradient at the nozzle exit.

Numerical Study on the Flow and Heat Transfer Characteristics of a Second Throat Exhaust Diffuser According to Variations in Operating Pressure and Geometric Shape

A numerical study was conducted to investigate the flow and heat transfer characteristics of a supersonic second throat exhaust diffuser for high-altitude simulations. The numerical results were satisfactorily validated by the experimental results. A subscale diffuser using nitrogen was utilized to investigate starting pressure and pressure variation in the diffuser wall. Based on the validated numerical method, the flow and heat transfer characteristics of the diffuser using burnt gas were evaluated by changing operating pressure and geometric shape. During normal diffuser operation without cooling, high-temperature regions of over 3000 K appeared, particularly near the wall and in the diffuser diverging section. After cooling, the flow and pressure distribution characteristics did not differ significantly from those of the adiabatic condition, but the temperature in the subsonic flow section decreased by more than 1000 K. Furthermore, the tendency of the heat flux from the diffuser internal flow to the wall was similar to that of the pressure variations, and it increased with operating pressure. It was confirmed that the heat fluxes of the supersonic and subsonic flows in the diffuser were proportional to the operating pressure to the 0.8 and −1.7 power, respectively. In addition, in the second throat region after separation, the heat flux could be scaled to the Mach number ratio before and after the largest oblique shock wave because the largest shock train affected the heat flux of the diffuser wall.

Experiment and numerical investigation of flow control on a supersonic inlet diffuser

Design of a two-dimensional supersonic inlet was developed. Numerical simulation and wind tunnel experiment were carried out to investigate the aerodynamic performance and variable geometric rules of the inlet. The result indicates that the variable geometry scheme adopted solves the contradiction between starting performance at low Mach number and aerodynamic performance at high Mach number. Inlet works normally and stably over a wide speed range. Compared with the diffuser without suction, the performance of the diffuser with suction is higher. Total pressure recovery coefficient σe improves by 2%-3% while outlet Mach number decreases by about 0.03 under different inflow Mach numbers. During the process of shock train moving upstream, the propagation gradually transforms from up and down mode to left and right mode. Flow separation caused by the shock train is mainly concentrated on the sidewall in the upstream of the diffuser. Sidewall suction case is more effective to eliminates the low-energy flow and stabilize the terminal shock wave. As backpressure increases, shock train changes from weak λ-type to strong λ-type and finally to X-type near the suction slots.

Numerical investigation on the flow structures of the multi-strut mixing enhancement ejector

A new type of multiple jets ejector called the multi-strut mixing-enhancement ejector (MSE) has been developed in previous studies. The excellent suction and supercharging abilities of this ejector have been demonstrated through a series of cold flow tests. However, the current knowledge on MSE is limited. Numerical investigations on MSE have not been conducted in previous studies. Therefore, the present study aims to conduct a series of numerical investigation on the MSE to analyse the transform regularities of the flow structures at various secondary flow conditions and explore the evolution processes of multiple 2D jet interactions in a confined space. Results show that the flow structures of the MSE are extremely complicated and different from the traditional single jet supersonic ejector. When the secondary flow is zero, the collisions of multiple supersonic primary flows generate multiple abreast oblique shocks, and the intersections and reflections of these oblique shock waves form a shock diamond matrix in the mixer and convergent section. In addition, a novel head-shape pseudo shock is observed in the secondary throat due to the coupling interactions of multiple oblique and adverse pressure gradients, which contain three parallel X-type shocks and an extremely intricate ‘butterfly-type’ shock behind. Under various secondary flow mass rate conditions, the secondary flows limit the expansions of the primary flows in the supersonic diffuser, thereby decreasing the number and intensity of oblique shocks in the mixer and convergent section, followed by the changes in the head shape of the pseudo shock in the secondary throat. Besides, the evolution processes of multiple 2D jets interactions demonstrate a complicated 3D effect. The wave loss, friction loss and mixing loss are the main factors that affect the evolution processes of multiple jets in the mixer and convergent section. The waves cause the jets to expand and compress along the flow direction, the friction loss results in vertical separation at the interface between the primary jets and the walls, and the mixing of flows causes the horizontal recess of the primary jets. These factors also have different effects on the jet’s evolution under various secondary flow conditions. The shock and friction losses are the main factors that affect the jet’s evolution when ER is small (ER ≤ 0.037). However, with the increase in ER, the effect of both losses weakens, and mixing loss becomes the dominant factor.

A study on the performance characteristics of two-stream supersonic diffusers

This paper investigates the mixing of two supersonic streams in a supersonic exhaust diffuser using theoretical, numerical, and experimental approaches. It has focused attention on the startup and stationary operation of the two-stream supersonic diffusers. The diffuser geometry consists of a constant area throat which connects two ducts of convergent and divergent cross-sections. The effects of diffuser contraction ratio (CR) and convergence angle (θ) on the starting process, pressure recovery, and flow features. Experiments have been performed at five different secondary Mach number (Ms) values in the range of 1.8 <Ms < 2.6, at a fixed primary Mach number (Mp) value of 2. Synchronized pressure measurements are used to study the diffuser performance characteristics. Numerical simulations were also carried out to obtain a better insight into the flow field and computed results have been compared with the experimental results. It was found that the starting behavior has a hysteresis. As a result, the tunnel can be operated at lower pressure ratio than the starting pressure ratio. It is observed that the settling chamber pressure required to start a Constant-Area Diffuser (CAD) is decreased with increase in the diffuser length up to an L/D ratio of 9; subsequently, starting pressure remains almost constant with the further increase in the length due to frictional and pressure recovery losses. The present study provides valuable insights into the multi-stream supersonic mixing and diffusion problem.

Design and analysis of a cooling system for a supersonic exhaust diffuser

High-altitude test facilities are usually used to evaluate the performance of space mission engines. The supersonic exhaust diffuser, a main part of high-altitude test facility, provides the required test cell vacuum conditions by self-pumping the nozzle exhaust gases to the atmosphere. However, the plume temperature is often much higher than the temperature the diffuser structure is able to withstand, usually above 2500 K. In this study, an efficient cooling system is designed and analyzed to resolve the thermal problem. A water spray cooling technique is preferred among various existing techniques. Here, a new algorithm is developed for a spray cooling system for a supersonic exhaust diffuser. This algorithm uses a series of experimental and geometrical relationships to resize the governing parameters and remove the required heat flux from the diffuser surface. The efficiency of the newly designed cooling system is evaluated via numerical simulations. The utilized numerical technique is based on the discrete-phase method. Various computational studies are accomplished to enhance the accuracy of numerical prediction and validation. The present numerical study is validated using experimental results. The results show that the realizable k-ɛ method is superior compared to other Reynolds-averaged Navier–Stokes models.

Aerothermodynamic analyses and redesign of GHIBLI Plasma Wind Tunnel hypersonic diffuser

This article deals with the aerothermodynamic analyses of the hypersonic diffuser of GHIBLI Plasma Wind Tunnel at CIRA (Centro Italiano Ricerche Aerospaziali). This diffuser presented an unexpected behavior not ensuring nozzle under-expanded flow in test chamber, even in several free jet conditions. This experimental evidence was investigated following two approaches. First, a literature review pointed out that the design rules, generally recommended in supersonic diffusers design, were not observed in the case of GHIBLI diffuser. Second, a simplified numerical model was introduced and validated with experimental results, performing Computational Fluid Dynamics analyses of GHIBLI diffuser. Finally, in order to enhance the pressure recovery and to obtain the desired under-expanded conditions, two new candidate designs compliant with the literature recommended values have been proposed and analyzed via CFD.

Numerical and experimental investigation of an adjustable dual-channel supersonic diffuser

The conventional diffuser performance is constrained by the optimal contraction ratio and the start-up contraction ratio. In order to balance the start-up performance and operating performance, the diffuser can be designed to be adjustable. Existing adjustable diffuser configurations have different drawbacks and most of them lack experimental verification. Therefore, a new parallel dual-channel diffuser scheme is proposed. The diffuser operates at Ma 4 and uses simple transmission devices to achieve an adjustment in the contraction ratio. The performances of the dual-channel mode, single-channel mode and the mode conversion are tested experimentally. The diffuser exhibits good starting and anti-backpressure capabilities. The unsteady numerical results show that the increase in the anti-backpressure capability of the diffuser is mainly due to the increase in the number of oblique shock waves in the main channel and the backward movement of the leading shock.

Operation matching model and analysis between an air inlet and a compressor in an Air Turbo Rocket

Operation matching between an air inlet and a compressor in an Air Turbo Rocket (ATR) is crucial to the full development of the advantages of engines. In this study, operation characteristic models of an air inlet and a compressor were established and verified through numerical simulation to gain the matched operation characteristics of the air inlet and the compressor in an ATR. The air inlet–compressor matching model was constructed on the basis of coupling relations, such as “flow equilibrium,” “interface parameter equilibrium,” and “pressure ratio equilibrium,” between the air inlet and the compressor in the ATR. Finally, the integrated 3D numerical simulation of the air inlet–compressor was verified. Results demonstrate that (1) the operation characteristic model of the air inlet can accurately calculate the effects of backpressure in the air inlet (Pb) on the total pressure recovery coefficients of the throat, the supersonic diffuser, and the whole air inlet in the ATR. Model prediction results are highly accurate as confirmed by 3D numerical simulation results in the following ranges: 2.75 < Ma < 3.75 and 0.5 atm <Pb < 3 atm. The maximum absolute error is less than 0.05. (2) The centrifugal compressor model can indicate the airflow and performance parameters of different characteristic sectors of impeller entrance, compressor exit, and axial diffuser entrance and exit. The 3D numerical simulation reports that the maximum relative errors of static pressure, total pressure, and total temperature predictions are 13.25%, 11.30%, and 5.65%, respectively. (3) The constructed air inlet–compressor matching model can rapidly predict the airflow parameters and performance parameters of their characteristic sections under specific operation conditions. The “most stable operation curve” introduced in the pressure ratio characteristic curve of the compressor can quickly and accurately accomplish the initialization of matching iterative computations, thereby solving aerodynamic coupling problems on an air inlet–compressor interface. (4) The integrated 3D numerical simulation of the air inlet/compressor reveals that the relative errors of 17 in 18 airflow parameters of two components are lower than 8%, and the relative error of the other parameters is 10.91%. The model exhibits high accuracy and can satisfy the requirements for the performance prediction of air inlet systems in ATR.

Invariants of velocity gradient tensor in supersonic turbulent pipe, nozzle, and diffuser flows

Velocity gradient tensor (VGT) analysis of high-order accurate direct numerical simulation data of supersonic pipe, nozzle, and diffuser flows at computationally moderate Reynolds numbers is performed. Joint probability density functions of second and third invariants of the VGT conditioned on positive and negative dilatation levels are presented in the near-wall viscous layer, buffer layer, log layer, and core region of these flows. For flow regions with positive dilatation, there is a preference for unstable flow topologies, while regions with negative dilatation show a preference for stable flow topologies.

Large-eddy simulation of shock-turbulence interaction in supersonic diffuser flows

ABSTRACTLES based on explicit filtering is used to study the shock train phenomenon in turbulent supersonic diffuser flows with circular cross-section and isothermal wall with an incoming pipe flow at friction Reynolds number 245 and centerline Mach number 1.7. Alternate regions of compression and expansion are found in the shock train which is followed by a shock-free ‘mixing’ region as observed in experiments and simulations in the literature. Turbulence amplification and local peaks in pressure-dilatation correlation are observed in the vicinity of the shocks. Low-frequency oscillations of the shock train are also observed.

Numerical Characterisation of Supersonic Exhaust Diffusers

&lt;p&gt;Rocket motors that are designed to operate at high altitudes need a nozzle with a large expansion ratio to maximize the thrust at much lower atmospheric pressure than that of the sea level pressure. Accurate performance of these nozzles cannot be obtained when static tested on ground. Computational fluid dynamics (CFD) analyses have been performed to characterise the supersonic exhaust diffuser (SED). The results obtained from the CFD analyses have been found to be in good agreement with experimental and numerical values reported in the published literature. Started and un-started regions of the SED have been identified with the CFD results.&lt;/p&gt;

LES of supersonic turbulent flow in nozzles and diffusers

Large-eddy simulations of supersonic nozzle and diffuser flows with circular cross-sections using high-order compact schemes and an explicit filtering version of the approximate deconvolution method are presented in this paper. Two flow cases each for nozzle and diffuser having different outlet to inlet area ratios are presented. The effect of the geometry variations on the Reynolds stresses as well as on the production and pressure-strain terms in their transport equations is demonstrated. A Green’s function analysis of the Poisson equation for pressure fluctuations using LES data is presented and the results show similar trends as found in previous analyses using DNS data. The effects of geometry changes on the rapid and slow parts of pressure-strain correlations is also demonstrated.