Shock oscillation in highly underexpanded jets

The oscillatory motions of shocks in highly underexpanded jets with nozzle pressure ratios of 5.60, 7.47, 9.34, and11.21 are quantitatively studied by using large eddy simulation. Two types of shock oscillations are observed： one is the Mach disk oscillation in the streamwise direction and the other is the shock oscillation in the radial direction. It is found that the Mach disk moves quickly in the middle of the oscillatory region but slowly at the top or bottom boundaries. The oscillation cycles of Mach disk are the same for different cases, and are all dominated by an axisymmetric mode of 5.298 k Hz. For the oscillation in the radial direction, the shocks oscillate more toward the jet centerline but less in the jet shear layer, and the oscillation magnitude is an increasing function of screech amplitude. The cycles of the radial shock oscillation switch randomly between the two screech frequencies for the first two cases. However, the oscillation periodicity is more complex for the jets with high nozzle pressure ratios of 9.34 and 11.21 than for the jets with the low nozzle pressure ratios of 5.6 and 7.47. In addition, the shock oscillation characteristics are also captured by coarse mesh and Smagorinsky model,but the coarse mesh tends to predict a slower and weaker shock oscillation.

A CALCULATING METHOD OF SHOCK WAVE OSCILLATING FREQUENCY DUE TO TURBULENT SHEAR LAYER FLUCTUATIONS IN SUPERSONIC FLOW

One of the more severe fluctuating pressure environments encountered in supersonic or hypersonic flows is the shock wave oscillation driven by interaction of a shock wave with boundary layer. The high intensity oscillating shock wave may induce structure resonance of a high speed vehicle. The research for the shock oscillation used to adopt empirical or semiempirical methods because the phenomenon is very complex. In this paper a theoretical solution on shock oscillating frequency due to turbulent shear layer fluctuations has been obtained from basic conservation equations. Moreover, we have attained the regularity of the frequency of oscillating shock varying with incoming flow Much numbers M and turning angle . The calculating results indicate excellent agreement with measurements. This paper has supplied a valuable analytical method to study aeroelastic problems produced by shock wave oscillation.

Experimental and Numerical Investigations of Shock-Wave Boundary Layer Interactions in a Highly Loaded Transonic Compressor Cascade

Experimental and numerical investigations were conducted to investigate the variations of shock-wave boundary layer interaction(SBLI) phenomena in a highly loaded transonic compressor cascade with Mach numbers.The schlieren technique was used to observe the shock structure in the cascade and the pressure tap method to measure the pressure distribution on the blade surface.The unsteady pressure distribution on blade surface was measured with the fast-response pressure-sensitive paint(PSP) technique to obtain the unsteady pressure distribution on the whole blade surface and to capture the shock oscillation characteristics caused by SBLI.In addition,the Reynolds Averaged Navier Stokes simulations were used to compute the three-dimensional steady flow field in the transonic cascade.It was found that the shock wave patterns and behaviors are affected evidently with the increase in incoming Mach number at the design flow angle,especially with the presence of the separation bubble caused by SBLI.The time-averaged pressure distribution on the blade surface measured by PSP technique showed a symmetric pressure filed at Mach numbers of 0.85,while the pressure field on the blade surface was an asymmetric one at Mach numbers of 0.90 and 0.95.The oscillation of the shock wave was closely with the flow separation bubble on the blade surface and could transverse over nearly one interval of the pressure taps.The oscillation of the shock wave may smear the pressure jump phenomenon measured by the pressure taps.

Experimental Investigation of the Transonic Shock Oscillation Characteristics in a Heavy-Duty Gas Turbine Compressor Cascade

This paper presents an experimental study of the self-sustained transonic shock oscillating behaviors in a heavy-duty gas turbine compressor cascade under the inlet Mach number of 0.85,0.90 and 0.95.The transonic shock patterns and the surface flow structures are captured by schlieren imaging and oil flow visualization.The time-averaged and instantaneous transonic shock oscillating behaviors at the near choke point and the near stall point are investigated by the Anodized Aluminum Pressure-Sensitive Paint(AA-PSP)surface pressure measurement.The normal passage shock dominant pattern and the detached bow shock dominant pattern at the near choke point and the near stall point are experimental characterized,respectively.The passage shock oscillation behaviors at the near choke point have been observed to undergo periodic pressure perturbations of the shock shift between the upstreamλshock feet mode and the downstreamλshock feet mode.The detached bow shock oscillation behaviors at the near stall point have been observed to undergo the pressure perturbations of the shock cycle movement between the upstream detached bow shock mode and the downstream detached bow shock mode.The differences between the shock shift mode and the shock cycle movement mode lead to the different streamwise oscillation travel ranges and different shock intensity variations under the same inlet Mach number.

Oblique shock train motion based on schlieren image processing

In this paper,shock train motion in a Mach number 2.7 duct is studied experimentally,and large numbers of schlieren images are obtained by a high-speed camera.An image processing method based on Maximum Correlation Detection(MCD)is proposed to detect shock train motion from the schlieren images,based on which the key structures,e.g.,separation positions and separation shock angles on the top and bottom walls,can be analysed in detail.The oscillations of the shock train are generated by rhombus and ellipse shafts at various excitation frequencies.According to the analysis of MCD results,the distributions of the frequency components of shock train oscillation generated by the two shafts are distinctly different,in which the motion generated by the ellipse shaft is much smoother;shock train motion is mainly characterized by the oscillation of separation position while the separation shock strength is not so sensitive to downstream disturbance;there is a hysteresis loop relation between the downstream pressure and separation position.

Effect of Non-Equilibrium Condensation of Moist Air on Unsteady Behaviour of Shock Waves around a Circular Arc Blade

The unsteady phenomena in the transonic flow around airfoils are observed in the flow field of fan, compressor blades and butterfly valves, and this often causes serious problems such as the aeroacoustic noise, the vibration. In the transonic or supersonic flow where vapour is contained in the main flow, the rapid expansion of the flow may give rise to a non-equilibrium condensation. However, the effect of non-equilibrium condensation on the transonic internal flows around the airfoil has not yet been clarified satisfactorily. In the present study, the effect of non-equilibrium condensation of moist air on the self-excited shock wave oscillation on a circular arc blade was investigated numerically. The results showed that in the case with non-equilibrium condensation, frequencies of the flow oscillation became smaller than those without the non-equilibrium condensation.

Unsteadiness of Shock Wave/Boundary Layer Interaction in Supersonic Cascade

An investigation of the passive control of shock wave/boundary laper interaction for reducing the amplitude of the shock oscillation was conducted on the circular arc-wedge (CW) profile cascade in a 220×290 mm transonic compressor cascade wind tunnel. A perforated surface with a cavity beneath it was positioned on the suction surface of the blade at the location of shock impingement. The Schlieren and high-speed photographs for flow over perforated blade are presented and compared with the results for solid blades. With the perforated surface, the high-speed photographs indicated an significant suppression of shock oscillation.

Matter Effects on Neutrino Oscillations in Different Supernova Models

In recent years,with the development of simulations about supernova explosion,we have a better understanding about the density profiles and the shock waves in supernovae than before.There might be a reverse shock wave,another sudden change of density except the forward shock wave,or even no shock wave,emerging in the supernova.Instead of using the expression of the crossing probability at the high resonance,PH,we have studied the matter effects on neutrino oscillations in different supernova models.In detail,we have calculated the survival probability of νe(P_s)and the conversion probability of ν_x(P_c) in the Schrodinger equation within a simplified two-flavor framework for a certain case,in which the neutrino transfers through the supernova matter from an initial flavor eigenstate located at the core of the supernova.Our calculations was based on the data of density in three different supernova models obtained from simulations.In our work,we do not steepen the density gradient around the border of the shock wave,which differs to what was done in most of the other simulations.It is found that the mass and the density distribution of the supernova do make a difference on the behavior of P_s and P_c.With the results of P_s and P_c,we can estimate the number of νe(and ν_x) remained in the beam after they go through the matter in the supernova.

Unsteady flow characteristic analysis of turbine based combined cycle(TBCC)inlet mode transition10.1016/j.jppr.2015.07.006

A turbine based combined cycle(TBCC)propulsion system uses a turbine-based engine to accelerate the vehicle from takeoff to the mode transition flight condition,at which point,the propulsion system performs a“mode transition”from the turbine to ramjet engine.Smooth inlet mode transition is accomplished when flow is diverted from one flowpath to the other,without experiencing unstart or buzz.The smooth inlet mode transition is a complex unsteady process and it is one of the enabling technologies for combined cycle engine to become a functional reality.In order to unveil the unsteady process of inlet mode transition,the research of over/under TBCC inlet mode transition was conducted through a numerical simulation.It shows that during the mode transition the terminal shock oscillates in the inlet.During the process of inlet mode transition mass flow rate and Mach number of turbojet flowpath reduce with oscillation.While in ramjet flowpath the flow field is non-uniform at the beginning of inlet mode transition.The speed of mode transition and the operation states of the turbojet and ramjet engines will affect the motion of terminal shock.The result obtained in present paper can help us realize the unsteady flow characteristic during the mode transition and provide some suggestions for TBCC inlet mode transition based on the smooth transition of thrust.

Unsteady supercritical/critical dual flowpath inlet flow and its control methods

The characteristics of unsteady flow in a dual-flowpath inlet, which was designed for a Turbine Based Combined Cycle（TBCC） propulsion system, and the control methods of unsteady flow were investigated experimentally and numerically. It was characterized by large-amplitude pressure oscillations and traveling shock waves. As the inlet operated in supercritical condition,namely the terminal shock located in the throat, the shock oscillated, and the period of oscillation was about 50 ms, while the amplitude was 6 mm. The shock oscillation was caused by separation in the diffuser. This shock oscillation can be controlled by extending the length of diffuser which reduces pressure gradient along the flowpath. As the inlet operated in critical condition, namely the terminal shock located at the shoulder of the third compression ramp, the shock oscillated,and the period of oscillation was about 7.5 ms, while the amplitude was 12 mm. At this condition,the shock oscillation was caused by an incompatible backpressure in the bleed region. It can be controlled by increasing the backpressure of the bleed region.

Transonic Moist Air Flow around a Circular Arc Blade with Bump

The unsteady phenomena in the transonic flow around airfoils are observed in the flow field of fan, compressor blades and butterfly valves, and this causes often serious problems such as aeroacoustic noise and the vibration. In recent years, the effect of bump wall on the flow field around an airfoil has been investigated experimentally and as a result, it was observed that the bump wall is effective for the control of shock wave on the airfoil. In the transonic or supersonic flow field, a rapid expansion of moist air or steam gives rise to non-equilibrium condensation. In the present study, the effect of non-equilibrium condensation of moist air on the self-excited shock wave oscillation around a circular arc blade with or without a bump on the blade was investigated numerically. The results showed that the non-equilibrium condensation significantly reduced the flow field unsteadiness such as root mean of pressure oscillation and frequency compared to the case without the non-equilibrium condensation.

Compressible flow characteristics around a biconvex arc airfoil in a channel

Shock wave-boundary layer interactions(SWBLI)are observed in several practical high-speed internal flows,such as compressor blades,turbine cascades,nozzles and so on.Shock induced oscillations(SIO),aerodynamic instabilities so-called buffet flows,flutter,aeroacoustic noise and vibration are the detrimental consequences of this unsteady shockboundary layer interactions.In the present study,a numerical computation has been performed to investigate the compressible flow characteristics around a 12%thick biconvex circular arc airfoil in a two dimensional channel.Reynolds averaged Navier-Stokes equations with two equation k-ωshear stress transport(SST)turbulence model have been applied for the computational analysis.The flow field characteristics has been studied from pressure ratio(ratio of back pressure,pb to inlet total pressure,p01)of 0.75 to 0.65.The present computational results have been compared and validated with the available experimental data.The results showed that the internal flow field characteristics such as shock wave structure,its behavior(steady or unsteady)and the corresponding boundary layer interaction are varied with pressure ratio.Self-excited shock oscillation was observed at certain flow conditions.Moreover,the mode of unsteady shock oscillation and its frequency are varied significantly with change of pressure ratio.

Unsteady Transonic Flow Control around an Airfoil in a Channel

Transonic internal flow around an airfoil is associated with self-excited unsteady shock wave oscillation. This unsteady phenomenon generates buffet, high speed impulsive noise, non-synchronous vibration, high cycle fatigue failure and so on. Present study investigates the effectiveness of perforated cavity to control this unsteady flow field. The cavity has been incorporated on the airfoil surface. The degree of perforation of the cavity is kept constant as 30%. However, the number of openings(perforation) at the cavity upper wall has been varied. Results showed that this passive control reduces the strength of shock wave compared to that of baseline airfoil. As a result, the intensity of shock wave/boundary layer interaction and the root mean square(RMS) of pressure oscillation around the airfoil have been reduced with the control method.