Showing papers on "Oblique shock published in 2009"

Large-eddy simulation of low-frequency unsteadiness in a turbulent shock-induced separation bubble

Abstract: The need for better understanding of the low-frequency unsteadiness observed in shock wave/turbulent boundary layer interactions has been driving research in this area for several decades. We present here a large-eddy simulation investigation of the interaction between an impinging oblique shock and a Mach 2.3 turbulent boundary layer. Contrary to past large-eddy simulation investigations on shock/turbulent boundary layer interactions, we have used an inflow technique which does not introduce any energetically significant low frequencies into the domain, hence avoiding possible interference with the shock/boundary layer interaction system. The large-eddy simulation has been run for much longer times than previous computational studies making a Fourier analysis of the low frequency possible. The broadband and energetic low-frequency component found in the interaction is in excellent agreement with the experimental findings. Furthermore, a linear stability analysis of the mean flow was performed and a stationary unstable global mode was found. The long-run large-eddy simulation data were analyzed and a phase change in the wall pressure fluctuations was related to the global-mode structure, leading to a possible driving mechanism for the observed low-frequency motions.

A simple model for low-frequency unsteadiness in shock-induced separation

Abstract: A model to explain the low-frequency unsteadiness found in shock-induced separation is proposed for cases in which the flow is reattaching downstream. It is based on the properties of fluid entrainment in the mixing layer generated downstream of the separation shock whose low-frequency motions are related to successive contractions and dilatations of the separated bubble. The main aerodynamic parameters on which the process depends are presented. This model is consistent with experimental observations obtained by particle image velocimetry (PIV) in a Mach 2.3 oblique shock wave/turbulent boundary layer interaction, as well as with several different configurations reported in the literature for Mach numbers ranging from 0 to 5.

Microramp Control of Supersonic Oblique Shock-Wave/Boundary-Layer Interactions

Abstract: The potential of microramp sub-boundary-layer vortex generators for flow control in supersonic engine inlets is investigated. In particular, the study focuses on the ability of these devices to beneficially affect oblique shock-wave/ boundary-layer interactions. Experiments have been conducted at Mach 2.5 to determine the nature of flow controlled by microramps and to investigate their ability to delay separation in a reflected shock interaction. Various ramp heights between 30 and 90% of the boundary-layer thickness were investigated. The details of the vortical flow generated by such devices were identified. The general flow features were found to scale with device height and it is suggested that smaller devices need to be placed closer to the expected adverse pressure gradients. When applied to a separated oblique shock-wave/boundary-layer interaction generated with a 7 degree wedge, microramps were not able to completely eliminate flow separation, although they were shown to break up separated regions. Other performance indicators across the shock-wave/boundary-layer interaction were also improved through the application of the devices.

Fundamentals of rotating detonations

Abstract: A rotating detonation propagating at nearly Chapman–Jouguet velocity is numerically stabilized on a two-dimensional simple chemistry flow model. Under purely axial injection of a combustible mixture from the head end of a toroidal section of coaxial cylinders, the rotating detonation is proven to give no average angular momentum at any cross section, giving an axial flow. The detonation wavelet connected with an oblique shock wave ensuing to the downstream has a feature of unconfined detonation, causing a deficit in its propagation velocity. Due to Kelvin–Helmholtz instability existing on the interface of an injected combustible, unburnt gas pockets are formed to enter the junction between the detonation and oblique shock waves, generating strong explosions propagating to both directions. Calculated specific impulse is as high as 4,700 s.

Experimental Investigation of Unstart in an Inlet/Isolator Model in Mach 5 Flow

Abstract: DOI: 10.2514/1.40966 The dynamics of unstart of a floor-mounted inlet/isolator model in a Mach 5 flow are investigated experimentally. The inlet section contains a 6-deg compression ramp, and the isolator is a rectangular straight duct that is 25.4-mm high by 50.8-mm wide by 242.3-mm long. Measurements made include 8-kHz schlieren imaging and simultaneous fast-response wall pressures along the length of the inlet/isolator. Unstart is initiated by deflecting a flap at the downstream end of the isolator. The shock system, induced by unstart, initially propagates upstream through the isolator at a velocity of about 35 m=s (in the laboratory frame of reference), then decelerates to about 20 m=s near the isolator entrance, and then accelerates to a velocity of about 74 m=s within the inlet. Throughout the isolator, unstart is seen to be strongly associated with boundary-layer separation. Once the inlet has unstarted, a highamplitudeoscillatory(periodic)unstarted flowensues,forwhichtheoscillationfrequencyisabout124Hz.However, under some conditions, an 84-Hz oscillatory unstarted flow mode, with lower pressure fluctuations, is observed. Under other conditions, a nonoscillatory unstarted flow, with much lower pressure fluctuations, is observed.

Direct numerical simulation of canonical shock/turbulence interaction

Abstract: A set of direct numerical simulations of isotropic turbulence passing through a nominally normal shock wave is presented. Upstream of the shock, the microscale Reynolds number is 40, the mean Mach number is 1.3–6.0, and the turbulence Mach number is 0.16–0.38. It is shown that the Kolmogorov scale decreases during the shock interaction, which implies that the grid resolution needed to resolve the viscous dissipation is finer than that used in previous studies. This leads to some qualitative differences with previous work, e.g., a rapid increase in the streamwise vorticity variance behind the shock and large anisotropy of the postshock Reynolds stresses. The instantaneous structure of the shock/turbulence interaction is examined using averages conditioned on the instantaneous shock strength. For locally strong compressions, the flow is characterized by overcompression, followed by an expansion. At points where the shock is locally weak, the profiles differ qualitatively depending on the strength of the inc...

Supersonic flow imaging via nanoparticles

Abstract: Due to influence of compressibility, shock wave, instabilities, and turbulence on supersonic flows, current flow visualization and imaging techniques encounter some problems in high spatiotemporal resolution and high signal-to-noise ratio (SNR) measurements. Therefore, nanoparticle based planar laser scattering method (NPLS) is developed here. The nanoparticles are used as tracer, and pulse planar laser is used as light source in NPLS; by recording images of particles in flow field with CCD, high spatiotemporal resolution supersonic flow imaging is realized. The flow-following ability of nanoparticles in supersonic flows is studied according to multiphase flow theory and calibrating experiment of oblique shock wave. The laser scattering characteristics of nanoparticles are analyzed with light scattering theory. The results of theoretical and experimental studies show that the dynamic behavior and light scattering characteristics of nanoparticles highly enhance the spatiotemporal resolution and SNR of NPLS, with which the flow field involving shock wave, expansion, Mach disk, boundary layer, sliding-line, and mixing layer can be imaged clearly at high spatiotemporal resolution.

Some physical aspects of shock wave/boundary layer interactions

Abstract: When the flow past a vehicle flying at high velocity becomes supersonic, shock waves form, caused either by a change in the slope of a surface, a downstream obstacle or a back pressure constraining the flow to become subsonic In modern aerodynamics, one can cite a large number of circumstances where shock waves are present The encounter of a shock wave with a boundary layer results in complex phenomena because of the rapid retardation of the boundary layer flow and the propagation of the shock in a multilayered structure The consequence of shock wave/boundary layer interaction (SWBLI) are multiple and often critical for the vehicle or machine performance The shock submits the boundary layer to an adverse pressure gradient which may strongly distort its velocity profile At the same time, in turbulent flows, turbulence production is enhanced which amplifies the viscous dissipation leading to aggravated performance losses In addition, shock-induced separation most often results in large unsteadiness which can damage the vehicle structure or, at least, severely limit its performance The article first presents basic and well-established results on the physics of SWBLI corresponding to a description in terms of an average two-dimensional steady flow Such a description allows apprehending the essential properties of SWBLIs and drawing the main features of the overall flow structure associated with SWBLI Then, some emphasis is placed on unsteadiness in SWBLI which constitutes a salient feature of this phenomenon In spite of their importance, fluctuations in SWBLI have been considered since a relatively recent date although they represent a domain which deserves a special attention because of its importance for a clear physical understanding of interactions and of its practical consequences as in aeroelasticity

The use of driver inserts to reduce non-ideal pressure variations behind reflected shock waves

Abstract: Non-ideal shock tube facility effects, such as incident shock wave attenuation, can cause variations in the pressure histories seen in reflected shock wave experiments. These variations can be reduced, and in some cases eliminated, by the use of driver inserts. Driver inserts, when designed properly, act as sources of expansion waves which can counteract or compensate for gradual increases in reflected shock pressure profiles. An algorithm for the design of these inserts is provided, and example pressure measurements are presented that demonstrate the success of this approach. When these driver inserts are employed, near- ideal, constant-volume performance in reflected shock wave experiments can be achieved, even at long test times. This near-ideal behavior simplifies the interpretation of shock tube chemical kinetics experiments, particularly in experiments which are highly sensitive to temperature and pressure changes, such as measurements of ignition delay time of exothermic reactions.

Electron shock surfing acceleration in multidimensions: two-dimensional particle-in-cell simulation of collisionless perpendicular shock

Abstract: Electron acceleration mechanisms in high-Mach-number collisionless shocks propagating in a weakly magnetized medium is investigated using a self-consistent two-dimensional particle-in-cell simulation. Simulation results show that strong electrostatic waves are excited via the electron-ion electrostatic two-stream instability at the leading edge of the shock transition region as in the case of earlier one-dimensional simulations. We observe strong electron acceleration that is associated with the turbulent electrostatic waves in the shock transition region. The electron energy spectrum in the shock transition region exhibits a clear power-law distribution with spectral index of 2.0-2.5. By analyzing the trajectories of accelerated electrons, we find that the acceleration mechanism is very similar to shock-surfing acceleration of ions. In contrast to the ion shock surfing, however, the energetic electrons are reflected by electron-scale electrostatic fluctuations in the shock transition region and not by the ion-scale cross-shock electrostatic potential. The reflected electrons are then accelerated by the convective electric field in front of the shock. We conclude that the multidimensional effects as well as the self-consistent shock structure are essential for the strong electron acceleration at high-Mach-number shocks.

Unsteady aspects of an incident shock wave/turbulent boundary layer interaction

Abstract: An incident shock wave/turbulent boundary layer interaction at Mach 2.1 is investigated using particle image velocimetry in combination with data processing using the proper orthogonal decomposition, to obtain an instantaneous and statistical description of the unsteady flow organization. The global structure of the interaction is observed to vary considerably in time. Although reversed flow is often measured instantaneously, on average no reversed flow is observed. On an instantaneous basis, the interaction exhibits a multi-layered structure, characterized by a relatively highvelocity outer region and low-velocity inner region. Discrete vortical structures are prevalent along their interface, which create an intermittent fluid exchange as they propagate downstream. A statistical analysis suggests that the instantaneous fullness of the incoming boundary layer velocity profile is (weakly) correlated with the size of the separation bubble and position of the reflected shock wave. The eigenmodes show an energetic association between velocity fluctuations within the incoming boundary layer, separated flow region and across the reflected shock wave, and portray subspace features that represent the phenomenology observed within the instantaneous realizations.

Experimental study on a plane shock wave accelerating a gas bubble

Abstract: A detailed experimental study of the interaction between a planar shock wave and an isolated spherical gas inhomogeneity is presented here. Different configurations have been considered: a shock wave moving from one gas into another, of similar density, lower density and one of higher density. Sequences of shadowgraph pictures obtained during the same run provided useful insights into several mechanisms such as shock wave reflection, refraction and focusing, distortion of the bubble interface, and vortex formation. Based on these sequences, the changes with time in the characteristic bubble sizes were plotted and the results showed that the influence of the shock wave Mach number is significantly greater in the case of light gas bubbles. The displacement of the inhomogeneity relative to the surrounding gas was determined and compared to Rudinger and Somers’ model. In all the cases studied, although the measurements were found to agree well with the theoretical predictions, in the initial acceleration phas...

Modeling of the unsteady force for shock-particle interaction

Abstract: The interaction between a particle and a shock wave leads to unsteady forces that can be an order of magnitude larger than the quasi-steady force in the flow field behind the shock wave. Simple models for the unsteady force have so far not been proposed because of the complicated flow field during the interaction. Here, a simple model is presented based on the work of Parmar et al. (Phil Trans R Soc A 366:2161–2175, 2008). Comparisons with experimental and computational data for both stationary spheres and spheres set in motion by shock waves show good agreement in terms of the magnitude of the peak and the duration of the unsteady force.

Analytical linear theory for the interaction of a planar shock wave with an isotropic turbulent vorticity field.

Abstract: An exact analytical model for the interaction between an isolated shock wave and an isotropic turbulent vorticity field is presented. The interaction with a single-mode two-dimensional (2D) divergence-free vorticity field is analyzed in detail, giving the time and space evolutions of the perturbed quantities downstream. The results are generalized to study the interaction of a planar shock wave with an isotropic three-dimensional (3D) or 2D preshock vorticity field. This field is decomposed into Fourier modes, and each mode is assumed to interact independently with the shock front. Averages of the downstream quantities are made by integrating over the angles that define the orientation of the upstream velocity field. The ratio of downstream/upstream kinetic energies is in good agreement with existing numerical and experimental results for both 3D and 2D preshock vorticity fields. The generation of sound and the sonic energy flux radiated downstream from the shock front is also discussed in detail, as well as the amplification of transverse vorticity across the shock front. The anisotropy is calculated for the far downstream fields of both velocity and vorticity. All the quantities characteristic of the shock-turbulence interaction are reduced to closed-form exact analytical expressions. They are presented as explicit functions of the two parameters that govern the dynamics of the interaction: the adiabatic exponent $\ensuremath{\gamma}$ and the shock Mach number ${M}_{1}$. These formulas are further reduced to simpler exact asymptotic expressions in the limits of weak and strong shock waves $({M}_{1}\ensuremath{-}1⪡1,\text{ }{M}_{1}⪢1)$ and high shock compressibility of the gas $(\ensuremath{\gamma}\ensuremath{\rightarrow}1)$.

Comparison of three large-eddy simulations of shock-induced turbulent separation bubbles

Abstract: Three different large-eddy simulation investigations of the interaction between an impinging oblique shock and a supersonic turbulent boundary layer are presented. All simulations made use of the same inflow technique, specifically aimed at avoiding possible low-frequency interferences with the shock/boundary-layer interaction system. All simulations were run on relatively wide computational domains and integrated over times greater than twenty five times the period of the most commonly reported low-frequency shock-oscillation, making comparisons at both time-averaged and low-frequency-dynamic levels possible. The results confirm previous experimental results which suggested a simple linear relation between the interaction length and the oblique-shock strength if scaled using the boundary-layer thickness and wall-shear stress. All the tested cases show evidences of significant low-frequency shock motions. At the wall, energetic low-frequency pressure fluctuations are observed, mainly in the initial part of interaction.

On Transonic Shocks in a Nozzle with Variable End Pressures

Abstract: In the book, Courant and Friedrichs (Supersonic Flow and Shock Waves. New York: Interscience Publishers, 1948) described the following transonic shock phenomena in a de Laval nozzle: Given the appropriately large receiver pressure p r , if the upstream flow is still supersonic behind the throat of the nozzle, then at a certain place in the diverging part of the nozzle a shock front intervenes and the gas is compressed and slowed down to subsonic speed. The position and the strength of the shock front are automatically adjusted so that the end pressure at the exit becomes p r . When the end pressure p r varies and lies in an appropriate scope, in general, it is expected that a curved transonic shock is still formed in a nozzle. In this paper, we solve this problem for the two-dimensional steady Euler system with a variable exit pressure in a nozzle whose divergent part is an angular sector. Both existence and uniqueness are established.

Supersonic flow separation in planar nozzles

Abstract: We present experimental results on separation of supersonic flow inside a convergent–divergent (CD) nozzle. The study is motivated by the occurrence of mixing enhancement outside CD nozzles operated at low pressure ratio. A novel apparatus allows investigation of many nozzle geometries with large optical access and measurement of wall and centerline pressures. The nozzle area ratio ranged from 1.0 to 1.6 and the pressure ratio ranged from 1.2 to 1.8. At the low end of these ranges, the shock is nearly straight. As the area ratio and pressure ratio increase, the shock acquires two lambda feet. Towards the high end of the ranges, one lambda foot is consistently larger than the other and flow separation occurs asymmetrically. Downstream of the shock, flow accelerates to supersonic speed and then recompresses. The shock is unsteady, however, there is no evidence of resonant tones. The separation shear layer on the side of the large lambda foot exhibits intense instability that grows into large eddies near the nozzle exit. Time-resolved wall pressure measurements indicate that the shock oscillates in a piston-like manner and most of the energy of the oscillations is at low frequency.

Shock geometry and spectral breaks in large sep events

Abstract: Solar energetic particle (SEP) events are traditionally classified as "impulsive" or "gradual." It is now widely accepted that in gradual SEP events, particles are accelerated at coronal mass ejection-driven (CME-driven) shocks. In many of these large SEP events, particle spectra exhibit double power law or exponential rollover features, with the break energy or rollover energy ordered as (Q/A)^α, with Q being the ion charge in e and A the ion mass in units of proton mass m_p . This Q/A dependence of the spectral breaks provides an opportunity to study the underlying acceleration mechanism. In this paper, we examine how the Q/A dependence may depend on shock geometry. Using the nonlinear guiding center theory, we show that α ~ 1/5 for a quasi-perpendicular shock. Such a weak Q/A dependence is in contrast to the quasi-parallel shock case where α can reach 2. This difference in α reflects the difference of the underlying parallel and perpendicular diffusion coefficients κ_(||) and κ ⊥. We also examine the Q/A dependence of the break energy for the most general oblique shock case. Our analysis offers a possible way to remotely examine the geometry of a CME-driven shock when it is close to the Sun, where the acceleration of particle to high energies occurs.

Unstable combustion induced by oblique shock waves at the non-attaching condition of the oblique detonation wave

Abstract: The instability of oblique shock wave (OSW) induced combustion is examined for a wedge with a flow turning angle greater than the maximum attach angle of the oblique detonation wave (ODW), where archival results rarely exist for this case in previous literatures. Numerical simulations were carried out for wedges of different length scales to account for the ratio of the chemical and fluid dynamic time scales. The results reveal three different regimes of combustion. (1) No ignition or decoupled combustion was observed if a fluid dynamic time is shorter than a chemical time behind an OSW. (2) Oscillatory combustion was observed behind an OSW if a fluid dynamic time is longer than a chemical time behind an OSW and the fluid dynamic time is shorter than the chemical time behind a normal shock wave (NSW) at the same Mach number. (3) Detached bow shock-induced combustion (or detached overdriven detonation wave) was observed if a fluid dynamic time is longer than a chemical time behind a NSW. Since no ignition or decoupled combustion occurs as a very slow reaction and the detached wave occurs as an infinitely fast reaction, the finite rate chemistry is considered to be the key for the oscillating combustion induced by an OSW over a wedge of a finite length with a flow turning angle greater than the maximum attach angle for an ODW. Since this case has not been previously reported, grid independency was tested intensively to account for the interaction between the shock and reaction waves and to determine the critical time scale where the oscillating combustion can be observed.

Shock wave boundary layer interaction

Abstract: Experiments based on laser Doppler anemometry (LDA) were performed to investigate shock/boundary layer interaction in the presence of separation. Applicability and limits of this experimental method are briefly discussed. Experiments are concentrated on flow conditions given by M = 2.5/Re = 29⋅10 m, interaction is generated by an unswept, 2-d flat plate/24 degree ramp configuration. Time averaged locations of separation, reattachment and the mean shock position are determined by measurements in closest proximity to the surfaces. Results are compared with data from pressure measurements given in literature.

Transonic Shock Problem for the Euler System in a Nozzle

Abstract: In this paper, we study the well-posedness problem on transonic shocks for steady ideal compressible flows through a two-dimensional slowly varying nozzle with an appropriately given pressure at the exit of the nozzle. This is motivated by the following transonic phenomena in a de Laval nozzle. Given an appropriately large receiver pressure P r , if the upstream flow remains supersonic behind the throat of the nozzle, then at a certain place in the diverging part of the nozzle, a shock front intervenes and the flow is compressed and slowed down to subsonic speed, and the position and the strength of the shock front are automatically adjusted so that the end pressure at exit becomes P r , as clearly stated by Courant and Friedrichs [Supersonic flow and shock waves, Interscience Publishers, New York, 1948 (see section 143 and 147)]. The transonic shock front is a free boundary dividing two regions of C2,α flow in the nozzle. The full Euler system is hyperbolic upstream where the flow is supersonic, and coupled hyperbolic-elliptic in the downstream region Ω+ of the nozzle where the flow is subsonic. Based on Bernoulli’s law, we can reformulate the problem by decomposing the 3 × 3 Euler system into a weakly coupled second order elliptic equation for the density ρ with mixed boundary conditions, a 2 × 2 first order system on u2 with a value given at a point, and an algebraic equation on (ρ, u1, u2) along a streamline. In terms of this reformulation, we can show the uniqueness of such a transonic shock solution if it exists and the shock front goes through a fixed point. Furthermore, we prove that there is no such transonic shock solution for a class of nozzles with some large pressure given at the exit.

Separation control by vortex generator devices in a transonic channel flow

Abstract: An experimental study was conducted in a transonic channel to control by mechanical vortex generator devices the strong interaction between a shock wave and a separated turbulent boundary layer. Control devices—co-rotating and counter-rotating vane-type vortex generators—were implemented upstream of the shock foot region and tested both on a steady shock wave and on a forced shock oscillation configurations. The spanwise spacing of vortex generator devices along the channel appeared to be an important parameter to control the flow separation region. When the distance between each device is decreased, the vortices merging is more efficient to reduce the separation. Their placement upstream of the shock wave is determinant to ensure that vortices have mixed momentum all spanwise long before they reach the separation line, so as to avoid separation cells. Then, vortex generators slightly reduced the amplitude of the forced shock wave oscillation by delaying the upstream displacement of the leading shock.

Time-Asymptotic Behavior of Wave Propagation Around a Viscous Shock Profile

Abstract: We study the nonlinear stability of shock waves for viscous conservation laws. Our approach is based on a new construction of a fundamental solution for a linearized system around a shock profile. We obtain, for the first time, the pointwise estimates of nonlinear wave interactions across a shock wave. Our results apply to all ranges of weak shock waves and small perturbations. In particular, our results reduce to the time-asymptotic behavior of constant state perturbation, uniformly as the strength of the shock wave tends to zero.

Wall shocks in high-energy-density shock tube experiments

Abstract: The radiative precursor of a sufficiently fast shock has been observed to drive the vaporization of shock tube material ahead of the shock. The resulting expansion drives a converging blast wave into the gas volume of the tube. The effects of this wall shock may be observed and correlated with primary shock parameters. We demonstrate this process in experiments performed on the Omega Laser Facility, launching shocks propagating through xenon with speeds above 100 km/s driven by ablation pressures of approximately 50 Mbars. Wall shocks in laser experiments, in which the principal shock waves themselves should not be radiative, are also reported—in which the wall shocks have been launched by some other early energy source.

Aerodynamic Design of a Supersonic Inlet with a Parametric Bump

Abstract: Numerical investigations were performed with an external-compression inlet with a three-dimensional bump at Mach 2 to scrutinize the geometrical effects of the bump in controlling the interaction of a shock wave with a boundary layer. The inlet was designed for two oblique shock waves and a terminal normal shock wave followed by a subsonic diffuser, with a circular cross section throughout. The bump-type inlet that replaced the aft ramp of the conventional ramp-type inlet was optimized with respect to the inlet performance parameters as well as compared with the conventional ramp-type inlet. The current numerical simulations showed that a bump-type inlet can provide an improvement in the total pressure recovery downstream of the shock wave/boundary layer interaction over a conventional ramp-type inlet.

Shock structure in separated nozzle flows

Abstract: In the case of high overexpansion, the exhaust jet of the supersonic nozzle of rocket engines separates from nozzle wall because of the large adverse pressure gradient. Correspondingly, to match the pressure of the separated flow region, an oblique shock is generated which evolves through the supersonic jet starting approximately at the separation point. This shock reflects on the nozzle axis with a Mach reflection. Thus, a peculiar Mach reflection takes place whose features depend on the upstream flow conditions, which are usually not uniform. The expected features of Mach reflection may become much difficult to predict, depending on the nozzle shape and the position of the separation point along the divergent section of the nozzle.

Supersonic Boundary Layer Interactions with Various Micro-Vortex Generator Geometries

Abstract: Various types of micro-vortex generators (μVGs) are investigated for control of a supersonic turbulent boundary layer subject to an oblique shock impingement, which causes flow separation. The micro-vortex generators are embedded in the boundary layer to avoid excessive wave drag while still creating strong streamwise vortices to energize the boundary layer. Several different types of μVGs were considered including micro-ramps and microvanes. These were investigated computationally in a supersonic boundary layer at Mach 3 using Monotone Integrated Large Eddy Simulations (MILES). The results showed that vortices generated from μVGs can partially eliminate shock induced flow separation and can continue to entrain high momentum flux for boundary layer recovery downstream. The micro-ramps resulted in thinner downstream displacement thickness in comparison to the micro-vanes. However, the strength of the streamwise vorticity for the micro-ramps decayed faster due to dissipation especially after the shock interaction. In addition, the close spanwise distance between each vortex for the ramp geometry causes the vortex cores to move upwards from the wall due to induced upwash effects. Micro-vanes, on the other hand, yielded an increased spanwise spacing of the streamwise vortices at the point of formation. This resulted in streamwise vortices staying closer to the floor with less circulation decay, and the reduction in overall flow separation is attributed to these effects. Two hybrid concepts, named “thick-vane” and “split-ramp”, were also studied where the former is a vane with side supports and the latter has a uniform spacing along the centerline of the baseline ramp. These geometries behaved similar to the micro-vanes in terms of the streamwise vorticity and the ability to reduce flow separation, but are more physically more robust than the thin vanes.

Study of Restricted Shock Separation Phenomena in a Thrust Optimized Parabolic Nozzle

Abstract: An experimental investigation has been carried out on a subscale thrust optimized parabolic nozzle (area ratio of 30) to study the flow characteristics prevalent during a partially formed restricted shock separation and a fully formed restricted shock separation condition, each of which are observed to be discrete in nature Whereas the wall pressure signal near the nozzle exit randomly alternates between its value in the backflow region and that above ambient (ie, flow randomly alternates between a free shock separation and restricted shock separation transition condition and vice versa) as a function of time for a partially formed restricted shock separation condition, the wall pressure in a fully formed restricted shock separation condition fluctuates in values above ambient in the region of flow reattachment Further, the transient conditions of free shock separation to partially formed restricted shock separation and end-effect regime are studied in detail The preceding transitions and retransitions suggest a variation in the relative axial positions of normal and separation shocks that favors a fully formed restricted shock separation to occur during shutdown A second separation bubble is also observed in the restricted shock separation condition, the formation and opening of which is seen to contribute toward side-load peaks Results also indicate that the separation shock translates back and forth (flapping motion) and experiences spanwise perturbations (rippling motion) The results are based on simultaneously acquired real-time wall pressure measurements, surface oil visualization technique, high-speed schlieren images, and signals from strain gauges installed on the nozzle bending tube