Flow characteristics of hypersonic inlets with different cowl-lip blunting methods

The unsteady effects in the shock tunnel flow establishment can aid hypersonic inlet starting. An unself-starting inlet may be started in the shock tunnel initial operation. To confirm the hypersonic inlet self-starting ability in shock tunnel, a new test method has been developed. It adopts a light obstacle to choke the inlet flow for unstarting the inlet at the initial stage of the shock tunnel operation. The obstacle can then be fast blown out to determine whether the inlet will be restarted under steady test flow conditions. Shutterlike fast response is required for the flow blockage generated by the obstacle because of the short steady flow duration of the shock tunnel. With the help of the shutter-like obstacle, the inlet could be ”pre-set” in either started or unstarted state prior to the steady test flow. Generic hypersonic inlets with different internal contraction ratios have been developed for inlet starting experiments. Various flow characteristics in the inlet starting process were achieved by applying a high speed schlieren system. The internal contraction ratio limits for inlets self-starting and unstarting were obtained and compared with literature. The effect of leading edge bluntness on the inlet starting characteristics were also examined. It was found that the leading edge bluntness play quite important roles in the inlet starting process.

Due to the limitations of size and low enthalpy of the general ground facilities, there are deviations between the flow structures and performance validated by ground tunnel experiments of a scaled inlet model and that of a full-scale model under practical flight conditions. To closely combine the results of ground and flight experiments and speed up the design of hypersonic vehicles, it’s important to investigate the influence of the test environment on the performance of the hypersonic inlet. The present study is motivated to elucidate the influence of incoming environment on the interference flow characteristics of a hypersonic inlet. Numerical simulations are conducted to investigate the effect of difference of flight and tunnel conditions on the performance of Ma7 inlet. The k-ω SST model is used for modeling turbulence and a turbulence model study is conducted to validate its ability. A grid sensitivity study is performed to examine the validity of simulation results. Details of flow characteristics of shock/shock interference (SSI) near the cowl lip and shock-wave/boundary-layer interference (SWBLI) in the isolator are compared and analyzed under the two conditions. The results indicate that SSI configurations near the cowl lip change greatly in different incoming environments. The flow structures at the lip directly determine the location and intensity of SWBLI downstream in the isolator. Compared with flight condition, the boundary layer separates earlier and the separation bubble extents larger under tunnel condition. Different flow characteristics lead to different distribution laws of surface aerodynamic parameters, and there does not exist simple conversion relationship. The nondimensionalized pressure and heat flux on the cowl lip are higher under the tunnel condition than those in the flight environment, whereas the surface aerodynamic parameters of the inlet are lower. When the hypersonic inlet operates at off-design point, the subsonic region generated by boundary layer separation moves closer to the cowl lip under the tunnel conditions than that in the flight environment. This indicates that the starting performance of the hypersonic inlet tested in the general tunnels tends to be a conservative estimation.

Present study simulates oscillatory supersonic impinging jet flows using an axisymmetric Navier -Stokes code. We focus on the staging behavior of the oscillation frequency of the impinging surface pressure associated with the variation of the nozzle -to -plate distance and nozzle pressure ratio. These two distinct staging behaviors of the oscillat ion frequency are found to be well correlated when the frequency and the distance are normalized by the length of the first shock cell. We show that the staging behavior is in line with the changing pattern of large coherent structures of the jet shear lay er . have clearly established that the self -sustained, highly unsteady behavior o f the jet and the resulting impinging tones is governed by a feedback mechanism. The energy of the feedback loop is provided by the instability waves in the shear layer of the jet. This instability wave usually generates large -scale vortical structures in the shear layer. The feedback loop is the coupling between the downstream -traveling instability waves (coherent structures) and the upstream -propagating acoustic waves. It has been reported that the frequencies of the dominant peak in both of sound and pre ssure signals of a supersonic impinging jet coincide. It is also well known that the frequencies of the impinging tone and the surface pressure oscillation exhibit staging behavior as the distance from the nozzle exit to the plate gradually increases. 1,4 This staging behavior for moderately underexpanded jets is typical of the class where disturbances from the jet interaction feed back to the nozzle, and so cause periodic self -induced oscillations. To predict the frequency and amplitude of the tones, many t heoretical modeling by using linear stability analysis have been attempted. However, the amplitude prediction and staging feature of the frequencies still remains elusive. 1-4

It was observed in an earlier study [Lim, H. D., New, T. H., Mariani, R., & Cui, Y. D. (2019). Effects of bevelled nozzles on standoff shocks in supersonic impinging jets. Aerospace Science and Technology, 94, 105371.] that standoff shocks in M=1.45 supersonic impinging jets produced by non-bevelled and bevelled nozzles under small separation distances may undergo large and abrupt fluctuations in their positions above the impingement points. This occurs when the separation distance is h/d=1.5 and for nozzle-pressure-ratio (NPR) between NPR=4 and 5, where h is the distance between the nozzle exit and flat-wall and d is the jet diameter. It has to be highlighted that the observed behaviour is unlike the more minor positional fluctuations typically associated with supersonic impinging jets. To reveal further details on the unsteadiness of the standoff shock, color schlieren results will be presented here to shed more light on the flow phenomenon, while initial modelling efforts to reproduce these standoff shock behaviour will be elaborated. Present results reinforce the notion that the large standoff shock fluctuations are due to the standoff shock trying to achieve pressure equilibrium through abrupt displacements when variation in the NPR changes the location of the shock reflection point. Considering the small separation distance used, small NPR changes are able to trigger abrupt standoff shock unsteadiness.

Bluntness impact on performance of waverider

Hypersonic shock tunnel studies of Edney Type III and IV shock interactions

The pipeline with a closed side branch is a typical structure in nuclear systems; the shear layer flow and deep cavity in this structure can lead to a complex flow-acoustic coupling, which may result in severe vibration and noise. To clarify the shear layer flow characteristics in this typical structure filled with water, the numerical strategy is established by employing the localized dynamic k-equation large eddy simulation model. Then, the statistical and instantaneous characteristics of the fluid field are analyzed, respectively. The dynamic vortex evolution process is presented by time–frequency domain analysis. Also, the phase relationship of pressure in shear layer is adopted to investigate the formation process of the shear layer mode. The results verify that the pressure fluctuation near the downstream corner is the sound source as the vortex impacts the wall periodically. The convection effect and disturbance intensity effect are separated for the first time through the control of the kinematic viscosity. The results show that the convection velocity influences the vortex shedding frequency (VSF) by changing the vortex moving speed. Also, changes in the disturbance intensity and branch length rectify the VSF by changing the phase relationship of impinging shear layer mode. Finally, a suppressing method of shifting the VSF is put forward by modifying the disturbance only.

The low swirl flow is a novel method for stabilizing lean premixed combustion to achieve low emissions of nitrogen oxides. Understanding the characteristics of low swirl flows is of both practical and fundamental interest. In this paper, in order to gain better insight into low swirl stabilized combustion, large eddy simulation and dynamically thickened flame combustion modeling are used to characterize various features of non-reacting and reacting low swirl flows including vortex breakdown, shear layers’ instability, and coherent structures. Furthermore, four test cases with different equivalence ratios are studied to evaluate the effects of equivalence ratio on the flame and flow characteristics. A finite volume scheme on a Cartesian grid with a dynamic one equation eddy viscosity subgrid model is used for large eddy simulations. The obtained results show that the combustion heat release and increase in equivalence ratio toward the stoichiometric value decrease the local swirl number of the flow field, while increasing the flow spreading at the burner outlet. Results show that the flame becomes W shaped as the equivalence ratio increases. Moreover, the combination of the swirling motion and combustion heat release temporally imposes a vortex breakdown in the post-flame region, which leads to occurrence of a transient recirculation zone. The temporal recirculation zone disappears downstream of the burner outlet due to merging of the inner shear layer from all sides at the centerline. Also, various analyses of shear layers’ wavy and vortical structures show that combustion heat release has the effect of decreasing the instability amplitude and vortex shedding frequency.

The jet plume effects on sonic boom for hypersonic slender bodies are studied systematically for the first time by simulating nose cone and nose double-cone geometries. The jet plume effects on hypersonic sonic boom are found to be different from the effects on low-supersonic sonic boom because the significantly smaller leading shock angle makes the shock directly interact with the underexpanded jet plume. When sufficiently underexpanded, the jet-induced shocks and shear layers significantly amplify the peak overpressure by a factor of 2–3 at hypersonic flight conditions. A detailed analysis of the dependence of the jet plume effects on the static-pressure ratio, jet Mach number, and vehicle shape has been performed. The static-pressure ratio and jet Mach number are found to have a significant impact on the jet plume effects. The jet plume effects also show sensitivity to the vehicle shape when the static-pressure ratio is small. To quantify the jet plume effects, a nondimensional parameter, Cj, representing the fraction of jet plume energy transferred to the sonic boom signature, is proposed and used. Finally, the amplification of peak overpressure by the jet plume is found to increase when the signature is propagated downward.

Measurement and analysis of the shear layer instabilities in supersonic impinging jets

Centrally staged combustion technique is often used in the military high-temperature-rise combustor. The pilot-stage structure affects the flow characteristics in the centrally staged combustor, which further affects the performance of ignition, combustion, and emission of military aero-engines. In order to increase the flow capacity of the swirler, the swirler with a non-rotating channel structure was designed. In this work, the influences of the pilot-stage structure on the flow characteristics in the centrally staged high-temperature-rise combustor are investigated. The flow fields of combustors with different pilot-stage swirl numbers (0.44, 0.60, and 0.71) are analyzed by large eddy simulation (LES). The results demonstrate that the primary recirculation zone (PRZ) becomes gradually longer and wider as the pilot-stage swirl number increases. In the combustors with three different pilot-stage structures, the precessing vortex core (PVC) was formed near the shear layer at the outlet of the pilot stage. The PVC frequency decreased from 1670 Hz to 1425 Hz and 1400 Hz with the increase of the pilot-stage swirl number from 0.44 to 0.60 and 0.71, respectively, and the breakdown position of the PVC shifted forward. The proper orthogonal decomposition (POD) and dynamic mode decomposition (DMD) methods are used to analyze the dynamic flow fields. It was observed that the corresponding frequency of the main pulsation structure decreased, and the flow instability was aggravated with the increase of the pilot-stage swirl number. The results deepen the understanding of the influences of the pilot-stage structure on the flow characteristics in the centrally staged high-temperature-rise combustor.

Flow and flow-induced acoustic fields of a supersonic jet impinging on an inclined flat plate is computationally studied. Effects of nozzle-plate distance and jet temperature are discussed as these are important parameters for the estimation of the acoustic wave strength from the engineering viewpoint such as a rocket plume. Three-dimensional compressible Navier-Stokes equations are solved with the modified weighted compact nonlinear scheme for the numerical simulations. Simulation results are analyzed from the viewpoints of acoustic emission and propagation mechanism, and acoustic field characteristics such as directivity, their spectra, and acoustic wave source positions are investigated. Time-averaged pressure contour on the plate shows two pressure peaks that are commonly observed in the jet-plate interaction; one by the jet impingement and the other by the plate shock wave. The acoustic fields indicate that there are at least three kinds of acoustic waves in all the cases considered in the study: (i) Mach waves generated from the shear layer of the main jet, (ii) acoustic waves generated from the impingement region, and (iii) Mach waves generated from the shear layer of the supersonic flow downstream of the jet impingement. Indication of the second one (ii) is important because the commonly-used empirical method for the estimation of the acoustic waves from a rocket plume does not consider such acoustic waves. Directivity and sound pressure level of the acoustic waves (ii) and the Mach waves (iii) change strongly with the temperature increase. Directivity and sound pressure level of the acoustic waves (ii) and the Mach waves (iii) change slightly with the nozzle-plate distance increase. The results show that whether the impinging point is inside or outside of potential core does not strongly affect the characteristics of the induced acoustic waves (ii).

Purpose The purpose of this paper is to carry out an extensive numerical study in order to understand the flow structures and the resulting noise generated by a supersonic impinging jet on a flat plate. One of the parameters varied in this study is the distance between the jet exit plane and the flat plate. Design/methodology/approach Because of the unsteady nature of the problem a time-dependent computation is carried out using the detached eddy simulation turbulence model. The OVERFLOW 2 CFD code was used with a highly resolved grid and small time steps. Findings The authors found that as the separation distance increases, the dominant frequencies in the noise spectrum decrease. In addition, the relative strength of the various frequencies to each other changes with changing distance, indicating the changing modes of the jet. The CFD results indicate a strong interaction between the acoustic waves emanating from the impingement plate and the jet plume. This feedback mechanism is responsible for destabilizing the jet shear layer leading to the jet changing modes. The computed near field spectra, convection velocities of the jet vortical structures and mean jet centerline velocity profile are in good agreement with experimental measurements. The results also show very high sound pressure levels all over the impingement plate but especially near the impingement point. These levels, if sustained, are detrimental to both human operators as well as the surrounding structures. Research limitations/implications Given the large-scale nature of the computations carried out, it is very costly to run the computations long enough to collect a good, statistically steady time sample to achieve a low frequency bandwidth resolution. Such a long time sample could actually improve the results in terms of frequency resolution and obtained an even better agreement with experiments. Off course there is always the issue of grid resolution as well, but given the good agreement with experiments that the authors obtained, the authors are confident in their results. Practical implications The practical implications of the results the authors obtained are significant in that, the authors now know that hybrid RANS-large eddy simulation methods can be used for this complex, unsteady engineering problems. In addition, the results also show the high noise level both on the impingement surface and in the surroundings of the jet. This could have a negative impact on the structural integrity of the flat surface. Social implications Noisy environments are never desirable anywhere especially in places where human operations take place. Therefore, given the high noise levels obtained in the simulations and confirmed by experiments, any human presence around the jet will be harmful to hearing and precautions need to be taken. Originality/value This is a physics-based study; i.e. understanding the physical phenomena involved in supersonic jet impingement. Of particular interest is the interaction of the jet shear layer with the acoustic waves emanating from the impingement area. This feedback loop is found to be responsible for intensifying the instability of the jet shear layer.

At hypersonic speeds, the external wall temperatures of an aerospace vehicle vary significantly. As a result, there is a considerable heat transfer variation between the boundary layer and the wall of the hypersonic vehicle. In this article, numerical computations are performed to investigate the effect of wall temperature on the separation bubble length in laminar hypersonic shock-wave/boundary-layer interaction flows over double-cone configuration at the Mach number of 12.2. The flow field is described in detail in terms of different shocks, expansion fans, shear layer and separation bubble. The variation of the Prandtl number has a negligible effect on the flow field and wall data. A specific heat ratio of less than 1.4 results in the better prediction of wall pressure and heat flux in the shock/boundary-layer interaction region. It is observed that as the wall temperature is increased, the separation bubble size and hence the separation shock length increases. The high firmness of the laminar boundary-layer at a high Mach number shows that the wall temperature in the shock/boundary-layer interaction region has little effect. The peak wall pressure and heat flux decrease with an increase in wall temperature. An estimation is developed between separation bubble length and wall temperature based on the computed results.

This study employs numerical simulation methods to investigate the flow characteristics of hypersonic inlets with different aspect ratio configurations under unstart conditions. Through comprehensive analysis of shock wave structure evolution, wall pressure distribution characteristics, and key performance parameters including mass flow ratio and total pressure recovery coefficient, the influence mechanism of aspect ratio on inlet unstart modes is revealed. The results demonstrate three distinct unstart modes: small separation on body (SSB), large separation on body (LSB), and dual-separation on both body and lip (DSBL). Furthermore, as the aspect ratio increases, the broadwise spillage effect weakens, and when exceeding a critical value, the DSBL transitions to LSB flow mode. Increasing the aspect ratio reduces the self-start performance of inlet while simultaneously improving key performance indicators such as pressure ratio, mass flow ratio and total pressure recovery coefficient, which provides important theoretical support for the design and optimization of hypersonic inlets.