Shock Oscillation Frequency Research Articles

Characteristics of self-excited shock wave oscillations in a supersonic isolator with exit duct configurations

The structure and dynamics of multiple shock waves inside a constant area isolator of a model scramjet engine are studied numerically using a finite volume solver. The flow is accelerated through a nozzle with an exit Mach number of 1.75 and continues in the constant area isolator where multiple shock waves are formed. To analyze the effect of duct acoustics on the dynamics of shock train oscillations, three ducts with a duct length of 0D, 50D, and 150D are attached to the exit of the isolator, where D is the isolator exit height. The wall static pressure variation along the bottom wall and four specific locations are monitored. The wall shear stress variation along the wall provides the location of separation and reattachment points in the flow field. The statistical analysis provided more insights into the oscillation mechanisms. The lower frequency components from the duct resonance are enhanced as the duct length is increased and it altered the dominant frequency of shock oscillation. The dependence of shock train length on the exit duct confirmation is analyzed by the pressure ratio method. The different frequency components in the flow field are analyzed using the power spectral density of wall static pressure fluctuations and are distinguished based on the frequency range. The cross-spectral analysis of the wall pressure fluctuations provided the correlation between pressure signals and their narrowband time delay. Thus, the origin and directions of the perturbations are obtained and discussed. The increase in exit duct length decreases the downstream disturbance velocity, indicating the duct resonance effect on shock train oscillations. • Simulation of the supersonic isolator flow field to understand the duct resonance effect on self-excited oscillations of shock train. • The duct resonance is varied by attaching ducts of different lengths to the exit of the isolator. • As the duct length is increased, the amplitude of shock oscillation is increased and the dominant frequency of shock oscillation is decreased. • The length of the shock train is increased with an increase in duct length. • The cross-spectral analysis of the wall pressure fluctuations provided the correlation between pressure signals and their narrowband time delay.

Separation Control for a Transonic Convex Corner Flow Using Ramp-Type Vortex Generators

A flap can be used to control wing camber and as a high-lift device. A convex corner is a simplified model of the upper surface of a flap. At transonic speeds, shock-induced boundary layer separation (SIBLS) occurs at greater freestream Mach numbers and deflection angles. This results in energy losses and a reduction in aerodynamic performance. This study installs ramp-type vortex generators (VGs) upstream of a convex corner, and the effect of the height of the VG on SIBLS is determined. As the height of the VG increases, the magnitude of the mean surface pressure upstream of the corner increases and downstream expansion decreases, which results in a reduction in lift. A reduction in peak surface pressure fluctuations, the separation length, and the frequency of shock oscillation is also determined. For flow control and lift enhancement, micro-VGs are more effective.

The Effect of Vortex Generators on Shock-Induced Boundary Layer Separation in a Transonic Convex-Corner Flow

Deflected control surfaces can be used as variable camber control in different flight conditions, and a convex corner resembles a simplified configuration for the upper surface. This experimental study determines the presence of passive vortex generators, VGs (counter-rotating vane type), on shock-induced boundary layer separation for transonic convex-corner flow. The mean surface pressure distributions in the presence of VGs for h/δ = 0.2 and 0.5 are similar to those for no flow control. If h/δ = 1.0 and 1.5, there is an increase in the amplitude of the mean surface pressure upstream of the corner’s apex, which corresponds to greater device drag and less downstream expansion. There is a decrease in peak pressure fluctuations as the value of h/δ increases, because there is a decrease in separation length and the frequency of shock oscillation. The effectiveness of VGs also depends on the freestream Mach number. For M = 0.89, there is an extension in the low-pressure region downstream of a convex corner, because there is greater convection and induced streamwise vorticity. VGs with h/δ ≤ 0.5 are preferred if deflected control surfaces are used to produce lift.

Strong interactions of incident shock wave with boundary layer along compression corner

The coherent structure and instability of the interaction of incident shock wave with boundary layer developing on a compression corner are experimentally studied. The experiments are carried out in a supersonic wind tunnel of Mach number 2. Particular attention is paid to shock patterns and unsteady shock motions induced by the separation bubble. The high-speed schlieren is used to visualize the flowfield evolution and to characterize the instability. The snapshot proper orthogonal decomposition of schlieren sequences is applied to investigate the primary coherent structure in the flowfield. Fast Fourier transform and continuous wavelet transformation are applied to characterize the instability. The results show that there are large-scale low-frequency oscillations of the shock waves and small-scale high-frequency pulsations in the separation region. The peak frequency of shock oscillation is mainly concentrated in the range of 100–1000 Hz. The pulsation of the small flow structure in the separation bubble is mainly concentrated above 12.5 kHz. Based on the results of experimental analysis, the preliminary mechanism of the large-scale instability of such interaction is obtained.

Transonic buffet instability: From two-dimensional airfoils to three-dimensional swept wings

The objective of the present study is to explain the evolution of the transonic buffet phenomenon from two-dimensional airfoils to three-dimensional swept wings by a global stability analysis. With respect to two-dimensional buffet, shock oscillation frequency increases by a factor of 4 to 7 in the case of a swept 30° wing and three-dimensional patterns in the detached boundary layer are convected outboard. Crouch et al. [J. Comput. Phys. 224, 924 (2007)] explained the two-dimensional transonic buffet phenomenon by the appearance of a real positive complex eigenvalue of the linearized Jacobian matrix. In the case of an infinite unswept wing, the present study shows that two unstable modes actually exist: The two-dimensional transonic buffet mode already identified by Crouch et al. [J. Fluid Mech. 628, 357 (2009)] and a strongly amplified three-dimensional zero-frequency mode. The latter exhibits regular patterns in the separated boundary layer, which relates to the so-called buffet cells as named by Iovnovich et al. [AIAA J. 53, 449 (2015)]. The nonzero sweep angle generates a spanwise velocity component on the wing which convects the cells outboard. This impacts both modes identified in the unswept case: The two-dimensional mode is weakly damped by the sweep while the three-dimensional buffet cells mode, even if weakly damped, remains strongly unstable and now exhibits a nonzero frequency which increases with the sweep angle. The frequency and wavelength of the most unstable three-dimensional mode for a sweep angle of 30° agree well with numerical and experimental values of the three-dimensional transonic buffet on wings. The analysis of the wavemaker of the three-dimensional modes indicates that the core of the instability is nearly solely located in the separated region, with a maximum along the separation line. In contrast, the wavemaker of the two-dimensional buffet mode exhibits stronger values all along the

Characterization of a Griffith-Type Transonic, Laminar-Flow Airfoil

An experimental test campaign was conducted on a transonic Griffith-type airfoil to evaluate the effectiveness of its laminar flow qualities and boundary-layer flow control characteristics in the transonic regime. From the pressure distributions it was observed that the flow control application had a beneficial influence, allowing for a more aggressive pressure recovery downstream of the suction slot and higher recovery pressure values at the trailing edge. At and , a net profile drag reduction of 10.70% and an increase in the ratio of 14.68% was observed. PIV experiments were performed to characterize the wake velocity deficit produced by the airfoil, which showed good agreement with the drag performance across every Mach number condition. Surface-oil flow visualization experiments were also performed at the design Mach number of 0.7 to find that the airfoil experienced extensive laminar flow runs regardless of suction application, which were followed by a shock-induced laminar separation bubble before the slot location. In addition, Schlieren imaging experiments were performed at to characterize the unsteady shock behavior. A characteristic shock oscillation frequency of 22.38 Hz was identified and observed to significantly stabilize with the application of suction.

Large-eddy simulation of laminar transonic buffet

A large-eddy simulation of laminar transonic buffet on an airfoil at a Mach number $M=0.735$, an angle of attack $\unicode[STIX]{x1D6FC}=4^{\circ }$, a Reynolds number $Re_{c}=3\times 10^{6}$ has been carried out. The boundary layer is laminar up to the shock foot and laminar/turbulent transition occurs in the separation bubble at the shock foot. Contrary to the turbulent case for which wall pressure spectra are characterised by well-marked peaks at low frequencies ($St=f\cdot c/U_{\infty }\simeq 0.06{-}0.07$, where $St$ is the Strouhal number, $f$ the shock oscillation frequency, $c$ the chord length and $U_{\infty }$ the free-stream velocity), in the laminar case, there are also well-marked peaks but at a much higher frequency ($St=1.2$). The shock oscillation amplitude is also lower: 6 % of chord and limited to the shock foot area in the laminar case instead of 20 % with a whole shock oscillation and intermittent boundary layer separation and reattachment in the turbulent case. The analysis of the phase-averaged fields allowed linking of the frequency of the laminar transonic buffet to a separation bubble breathing phenomenon associated with a vortex shedding mechanism. These vortices are convected at $U_{c}/U_{\infty }\simeq 0.4$ (where $U_{c}$ is the convection velocity). The main finding of the present paper is that the higher frequency of the shock oscillation in the laminar regime is due to a different mechanism than in the turbulent one: laminar transonic buffet is due to a separation bubble breathing phenomenon occurring at the shock foot.

Study of buzz phenomenon using visualization of external shock structure

An experimental study was carried out on an axisymmetric supersonic inlet with external compression in order to investigate the buzz phenomenon at different angles of attack and mass flow rates. The model was equipped with accurate and high-frequency pressure sensors, and the tests were conducted at Mach numbers varying from 1.8 to 2.5, for various angles of attack. Shadowgraph visualization technique, together with a high-speed camera, was used to provide the visual description of the shock structure in front of the inlet and to study the characteristics of buzz. Furthermore, pressure distribution over the spike surface was measured using several pressure sensors. Frequency of the buzz and shock displacement were measured by inspection of visualization pictures in each test. The obtained data from shadowgraphs were compared with those obtained from pressure measurements, and good agreement was found between them. The results revealed that for a moderate value of mass flow rate, the frequency of shock oscillation decreases as Mach number increases. Further, by increasing angle of attack, the shock displacement of oscillation will increase. At non-zero angles of attack, the displacement and frequency of shock motion show different behaviors on the leeward and windward sides of the body.

Back-pressure effects on unsteadiness of separation shock in a rectangular duct at Mach 3

To evaluate back-pressure effects on unsteadiness of separation shock in a rectangular duct, some test cases were designed and conducted in a supersonic air-breathing wind tunnel. High-speed Schlieren technique and high-frequency pressure measurements were utilized for data acquisition. Different back-pressure levels were applied by changing downstream throttling ratio. Some parameters for estimating canonical separation flows, such as zero-crossing frequency and intermittency, have been introduced to analyze the duct separation flow. The experimental results show that the separation shock oscillation frequencies increase and the intermittent region lengths decrease with the increasing back-pressure level. A comparison between the unsteadiness of shock/boundary layer interactions for the canonical flow and the duct internal flow were made. It is found that the normalized Strouhal number for the duct internal separation flow ranges from 0.01 to 0.03, which share the similar range with canonical separation flows. That is to say, the physics of separation shock unsteadiness is similar over a wide range of interaction flows, which may share the inherent flow mechanism.

Effect of plate inclination on the noise of impinging jets

This article explains the flow and acoustic variation due to the jet impinging on inclined plates. There is a misconception that the acoustic radiation reduces by changing the angle of impingement. However, in the present study it is found that the reduction or enhancement of the acoustic radiation depends on the location of measurement and nozzle pressure ratio. A detailed directivity study of jet impinging on the inclined plate is carried out. Further, high speed flow visualization is carried out to compare shock oscillation and tonal frequencies.

Impingement Tones and Associated Shock Instabilities in Supersonic Plug Nozzle Flows

An investigation into a supersonic impinging jet emitted from a truncated plug nozzle is presented for a nozzle pressure ratio of 3.4. High-resolution schlieren and acoustic measurements verify that the jet behavior close to the nozzle is significantly different from round jets. Impingement tones are not observed for plate distances less than 3.2 nozzle diameters from the throat. Sufficiently far downstream from the nozzle, the jet shock structure is similar to that of a round jet. Ultrahigh-speed schlieren is used to track the motion of the standoff shock located above the impingement plate. The shock-oscillation frequency matches the impingement-tone frequency within statistical uncertainty. The average shock standoff distance and oscillation amplitude are related to the relative position of the impingement plate with respect to the shock structure in the equivalent freejet. Shock splitting is directly observed for certain plate distances.

Shock-Tracking Procedure for Studying Screech-Induced Oscillations

A SHOCK-CONTAINING supersonic jet radiates two so-called shock-associated noise components in addition to the classical mixing noise. These two components are the broadband shockassociated noise and the screech. The latter was studied in the early 1950s by Powell [1]. Screech is a tonal noise component. Its generation mechanism has been explained with some success as a feedback loop involving sound production through shock– turbulence interaction. The broadband shock-associated noise component has been investigated at least since Martlew [2]. HarperBourne and Fisher [3] have adapted Powell’s stationary source array model to derive some observed properties of this noise component. In his review about high-speed jet aeroacoustics, Seiner [4] points out the necessity of studying shockmotion to confirmPowell’s [1]model of stationary sources. Furthermore, fluid disturbances having a relative motion to shocks may be responsible for broadband shock noise [4]. A certain insight into shock behavior within a jet plume is thus essential to accurately understand shock-associated noise generation. The screech phenomenon in a round jet involves different modes. They have been studied extensively in the past, for example, by Davies and Oldfield [5], Merle [6], or Powell et al. [7]. Mode switching appears as a sudden change of jet plume structure and screech frequency with operating conditions. Five modes are traditionally quoted: A1 and A2 are axisymmetric, B is sinuous or flapping, C is helical, andD is again sinuous. The reader is referred to Raman [8,9] for a review of screech properties. The coincidence between shock oscillation frequency and screech frequency has been first observed by Lassiter and Hubbard [10] from a shadowgraph technique. Sherman et al. [11] havemade use of highspeed schlieren recording to investigate the shock distortion during screech. They have also found the coincidence between oscillation and screech frequencies and estimated oscillation amplitudes for the third shock. Using an optical shock detection technique based on laser light scattering by a shock, Panda [12,13] has significantly contributed to the characterization of shock motion during screech. He has reported that every shock oscillates at the screech frequency andmotion amplitudes have been assessed. In the latter reference, an analyticalmodel for shock oscillation is also proposed and is found to be in good agreement with measurements. To the authors’ knowledge, Sherman et al.’s [11] and Panda’s [12,13] contributions are the only occurrences of shock oscillation measurements in the screech literature. In the present Note, oscillation frequency and amplitude are examined by means of high-speed schlieren images and simultaneous near-field acoustic recordings. A procedure for shock tracking is developed and applied on two different screech modes. The experimental setup is presented in Sec. II. The procedure is explained in Sec. III, and some results are discussed. Conclusions are finally given in Sec. IV.

Experimental and numerical study of forced shock-wave oscillations in a transonic channel

The aim of the present investigation is to analyse the response of a transonic channel flow when the shock-wave is subjected to a periodic motion at a well defined and controlled frequency. The shock periodic motion was produced by means of a second throat whose section was varied thanks to a rotating elliptical or rectangular shaft located in the middle of the channel outlet section. The induced shock oscillations had amplitudes of about 10 times the incoming boundary layer thickness and frequencies several orders lower than turbulence frequencies. Schlieren visualizations and unsteady pressure measurements, coupled with a spectral analysis, made it possible to highlight the rotating shaft shape influence on the shock oscillation in the channel. Conditional two component laser Doppler velocimeter probing was carried out for a shock oscillation frequency of 40 Hz produced by the elliptical shaft only. Phase-averaged field results have allowed a precise description of the unsteady flow field showing both the wave propagation in the core flow and the response of the boundary layer subjected to an oscillating shock-wave. In the shock oscillation region, no phase lag was observed between velocities in the core flow and in the boundary layers whereas a significant one has appeared in the downstream subsonic region. Numerical simulations of the channel flow were made by means of a two-dimensional unsteady Reynolds averaged Navier–Stokes solver using a dual time-stepping method and several two-equation transport turbulence models. The shock oscillation frequency of 40 Hz was only produced by the rotating elliptical shaft. Apart from some discrepancies in the interaction regions, the cycle of the shock oscillation is well reproduced by the computations. The good agreement between the evolutions of the measured and computed fluctuating quantities has allowed to confirm the assumption about the quasi instantaneous adaptation of the turbulence to the mean flow variations in such channel flows.

Study of the shock motion in a hypersonic shock system/turbulent boundary layer interaction

Wall pressure fluctuations and surface heat transfer signals have been measured in the hypersonic turbulent boundary layer over a number of compression-corner models. The distributions of the separation shock oscillation frequencies and periods have been calculated using a conditional sampling algorithm. In all cases the oscillation frequency distributions are of broad band, but the most probable frequencies are low. The VITA method is used for deducing large scale disturbances at the wall in the incoming boundary layer and the separated flow region. The results at present showed the existence of coherent structures in the two regions. The zero-cross frequencies of the large scale structures in the two regions are of the same order as that of the separation shock oscillation. The average amplitude of the large scale structures in the separated region is much higher than that in the incoming boundary layer. The length scale of the separation shock motion region is found to increase with the disturbance strength. The results show that the shock oscillation is of inherent nature in the shock wave/turbulent boundary layer interaction with separation. The shock oscillation is considered to be the consequence of the coherent structures in the separated region.

Control-nonlinear-nonstationary structural response and radiation near a supersonic jet

This paper is on the control of nonlinear-nonstationary vibration of a frame-stringer structure resulting from high levels of excitaation from a nearby supersonic jet exhaust. The structure exhibits periodic, chaotic, or random behaviors when forced by high-intensity sound from a supersonic jet exhaust with shock loading superimposed on the broadband response. The time history of the pressure, showing the rotation and flapping of the shock structure in the jet column due to large-scale instabilities, indicates that the response is not only nonlinear but also nonstationary. The acoustic pressure radiated by the structure also contains shocks and the formation of harmonics with distance. Control of the structural response is achieved by actively forcing the structure with an actuator at the shock oscillation frequency whose amplitude is locked into a self-control cycle. Results show that the peak power level is reduced by a factor of 63, or 18 dB. As a result, new broadband components emerge with at least four harmonics. At accelerating and decelerating supersonic speeds, the exhaust from the jet induces higher transient loading on the nearby flexible structure due to the occurence of multiple shocks from the jet.