Shock-boundary layer interaction and transonic flutter

The transonic flutter dip of an aeroelastic system is primarily caused by compressibility of the flowing fluid. Viscous effects are not dominant in the pre-transonic dip region. In fact, an Euler solver can predict this flutter boundary with considerable accuracy. However with an increase in Mach number the shock moves towards the trailing edge causing shock induced separation. This shock-boundary layer interaction changes the flutter boundary in the transonic and post-transonic dip region significantly. We discuss the effect of viscosity in changing the flutter boundary in the post-transonic dip region using a RANS solver coupled to a two-degree of freedom model of the structural dynamics of a wing.

Detonation-driven-shock wave interactions with perforated plates

The study of detonations and their interactions is vital for the understanding of the high-speed flow physics involved and the ultimate goal of controlling their detrimental effects. However, producing safe and repeatable detonations within the laboratory can be quite challenging, leading to the use of computational studies which ultimately require experimental data for their validation. The objective of this study is to examine the induced flow field from the interaction of a shock front and accompanying products of combustion, produced from the detonation taking place within a non-electrical tube lined with explosive material, with porous plates with varying porosities, 0.7-9.7%. State of the art high-speed schlieren photography alongside high-resolution pressure measurements is used to visualise the induced flow field and examine the attenuation effects which occur at different porosities. The detonation tube is placed at different distances from the plates' surface, 0-30 mm, and the pressure at the rear of the plate is recorded and compared. The results indicate that depending on the level of porosity and the Mach number of the precursor shock front secondary reflected and transmitted shock waves are formed through the coalescence of compression waves. With reduced porosity, the plates act almost as a solid surface, therefore the shock propagates faster along its surface.

Shock tunnel experiments on control of shock induced large separation bubble using boundary layer bleed

Shock-Boundary Layer Interaction (SBLI) often occurs in supersonic/hypersonic flow fields. Especially when accompanied by separation (termed strong interaction), the SBLI phenomena largely affect the performance of the systems where they occur, such as scramjet intakes, thus often demanding the control of the interaction. Experiments on the strong interaction between impinging shock wave and boundary layer on a flat plate at Mach 5.96 are carried out in IISc hypersonic shock tunnel HST-2. The experiments are performed at moderate flow total enthalpy of 1.3 MJ/kg and freestream Reynolds number of 4 million/m. The strong shock generated by a wedge (or shock generator) of large angle 30.96 degrees to the freestream is made to impinge on the flat plate at 95 mm (inviscid estimate) from the leading edge, due to which a large separation bubble of length (75 mm) comparable to the distance of shock impingement from the leading edge is generated. The experimental simulation of such large separation bubble with separation occurring close to the leading edge, and its control using boundary layer bleed (suction and tangential blowing) at the location of separation, are demonstrated within the short test time of the shock tunnel (similar to 600 mu s) from time resolved schlieren flow visualizations and surface pressure measurements. By means of suction - with mass flow rate one order less than the mass flow defect in boundary layer - a reduction in separation length by 13.33% was observed. By the injection of an array of (nearly) tangential jets in the direction of mainstream (from the bottom of the plate) at the location of separation - with momentum flow rate one order less than the boundary layer momentum flow defect - 20% reduction in separation length was observed, although the flow field was apparently unsteady. (C) 2014 Elsevier Masson SAS. All rights reserved.

Novel supersonic nozzles for mixing enhancement in supersonic ejectors

Two novel supersonic nozzles Tip Ring Supersonic Nozzle and Elliptic Sharp Tipped Shallow (ESTS) Lobed Nozzle have been developed to enhance mixing at high speeds which is beneficial to supersonic ejectors. A circular ring protruding at the exit of a conical nozzle forms the tip ring nozzle. The innovative ESTS lobed nozzle comprising of four elliptic lobes with sharp tips that do not protrude deep into the core supersonic flow is produced by a novel yet simple methodology. A comparative experimental study is conducted between a conical nozzle, an ESTS lobed nozzle and a tip ring nozzle with exit Mach number of 2.3. For the first time, the three dimensional flow structure from ESTS lobed nozzle and tip ring nozzle is revealed from laser scattering flow visualization experiments on the free jet. A doubling of jet spreading rate is observed in the ESTS lobed nozzle. When applied to a supersonic ejector, both nozzles achieve a 30% increase in entrainment of secondary flow. The loss of compression ratio is 15% for the ESTS lobed nozzle while it is 50% for the tip ring nozzle. Further, the behavior of wall static pressure profile corroborates mixing enhancement. (C) 2014 Elsevier Ltd. All rights reserved.

Study the effect of electrical and mechanical shock loading on liberation and milling characteristics of mineral materials

Milling is an energy intensive process and it is considered as one of the most energy inefficient processes. Electrical and mechanical shock loading can be used to develop a pre-treatment methodology to enhance energy efficiency of comminution and liberation of minerals. Coal and Banded Hematite Jasper (BHJ) Iron ores samples were taken for the study to know the effect of shock loading. These samples were exposed to 5 electric shocks of 300 kV using an electric shock loading device. A diaphragmless shock tube was used to produce 3 and 6 compressed air shocks of Mach number 2.12 to treat the coal and Iron ore samples. Microscopic, comminution and liberation studies were carried out to compare the effectiveness of these approaches. It was found that electric shock loading can comminute the coal samples more effectively and increases the yield of carbon by 40% at 1.6 gm/cc density over the untreated coal samples. Mechanical shock loading showed improved milling performance for both the materials and 12.90% and 8.1% reduction in the D-80 of the particles was observed during grinding for treated samples of coal and iron, respectively. Liberation of minerals in BHJ Iron ore was found unaffected due to low intensity of the mechanical shock waves and non conductivity of minerals. Compressed air based shock loading is easier to operate than electrical shock loading and it needs to be explored further to improve the energy efficacy of comminution. (C) 2014 Elsevier Ltd. All rights reserved.

Large carbon cluster thin film gauges for measuring aerodynamic heat transfer rates in hypersonic shock tunnels

Different types of Large Carbon Cluster (LCC) layers are synthesized by a single-step pyrolysis technique at various ratios of precursor mixture. The aim is to develop a fast responsive and stable thermal gauge based on a LCC layer which has relatively good electrical conduction in order to use it in the hypersonic flow field. The thermoelectric property of the LCC layer has been studied. It is found that these carbon clusters are sensitive to temperature changes. Therefore suitable thermal gauges were developed for blunt cone bodies and were tested in hypersonic shock tunnels at a flow Mach number of 6.8 to measure aerodynamic heating. The LCC layer of this thermal gauge encounters high shear forces and a hostile environment for test duration in the range of a millisecond. The results are favorable to use large carbon clusters as a better sensor than a conventional platinum thin film gauge in view of fast responsiveness and stability.

Quantitative estimation of density variation in high-speed flows through inversion of the measured wavefront distortion

A simple method employing an optical probe is presented to measure density variations in a hypersonic flow obstructed by a test model in a typical shock tunnel. The probe has a plane light wave trans-illuminating the flow and casting a shadow of a random dot pattern. Local slopes of the distorted wavefront are obtained from shifts of the dots in the pattern. Local shifts in the dots are accurately measured by cross-correlating local shifted shadows with the corresponding unshifted originals. The measured slopes are suitably unwrapped by using a discrete cosine transform based phase unwrapping procedure and also through iterative procedures. The unwrapped phase information is used in an iterative scheme for a full quantitative recovery of density distribution in the shock around the model through refraction tomographic inversion. Hypersonic flow field parameters around a missile shaped body at a free-stream Mach number of 5.8 measured using this technique are compared with the numerically estimated values. (C) 2014 Society of Photo-Optical Instrumentation Engineers (SPIE)

Behaviour of turbulent Prandtl/Schmidt number in compressible mixing layer

The behaviour of turbulent Prandtl/Schmidt number is explored through the model-free simulation results. It has been observed that compressibility affects the Reynolds scalar flux vectors. Reduced peak values are also observed for compressible convective Mach number mixing layer as compared with the incompressible convective Mach number counterpart, indicating a reduction in the mixing of enthalpy and species. Pr-t and Sc-t variations also indicate a reduction in mixing. It is observed that unlike the incompressible case, it is difficult to assign a constant value to these numbers due to their continuous variation in space. Modelling of Pr-t and Sc-t would be necessary to cater for this continuous spatial variation. However, the turbulent Lewis number is evaluated to be near unity for the compressible case, making it necessary to model only one of the Pr-t and Sc-t..

Reynolds averaged Navier-Stokes simulations of compressible mixing layers of similar and dissimilar gases: Performance of k- turbulence model

The issue of growth rate reduction of high speed mixing layer with convective Mach number is examined for similar and dissimilar gases using Reynolds averaged Navier-Stokes (RANS) methodology with k- turbulence model. It is observed that the growth rate predicted using RANS simulations closely matches with that predicted using model free simulations. Velocity profiles do not depend on the modelled value of Pr-t and Sc-t; while the temperature and species mass fraction distributions depend heavily on them. Although basic k- turbulence model could not capture the reduced growth rate for the mixing layer formed between similar gases, it predicts very well the reduced growth rate for the mixing layer for the dissimilar gases. It appears that density ratio changes caused by temperature changes for the dissimilar gases have profound effect on the growth rate reduction.

EXPERIMENTAL INVESTIGATION OF UNSTEADY FLOW IN A TRANSONIC UNI-STAGE AXIAL COMPRESSOR

Detailed steady and unsteady experimental measurements and analysis were performed on a Single stage Transonic Axial Compressor with asymmetric rotor tip clearance for studying the compressor stall phenomena. The installed compressor had asymmetric tip clearance around the rotor casing varying from about 0.65mm to 1.25mm. A calibrated 5-hole aerodynamic probe was traversed radially at exit of rotor and showed the characteristics of increased flow angle at lower mass flow rates for all the speeds. Mach number distribution and boundary layer effects were also clearly captured. Unsteady measurements for velocity were carried out to study the stall cell behavior using a single component calibrated hotwire probe oriented in axial and tangential directions for choke/free flow and near stall conditions. The hotwire probe was traversed radially across the annulus at inlet to the compressor and showed that the velocity fluctuations were dissimilar when probe was aligned axial and tangential to the flow. Averaged velocities across the annulus showed the reduction in velocity as stall was approached. Axial mean flow velocity decreased across the annulus for all the speeds investigated. Tangential velocity at free flow condition was higher at the tip region due to larger radius. At stall condition, the tangential velocity showed decreased velocities at the tip and slightly increased velocities at the hub section indicating that the flow has breakdown at the tip region of the blade and fluid is accelerated below the blockage zone. The averaged turbulent intensity in axial and tangential flow directions increased from free flow to stall condition for all compressor rated speeds. Fast Fourier Transform (FFT) of the raw signals at stall flow condition showed stall cell and its corresponding frequency of occurrence. The stalling frequency of about half of rotational speed of the rotor along with large tip clearance suggests that modal type stall inception was occurring.

Shock tunnel measurements of surface pressures in shock induced separated flow field using MEMS sensor array

Characterized not just by high Mach numbers, but also high flow total enthalpies-often accompanied by dissociation and ionization of flowing gas itself-the experimental simulation of hypersonic flows requires impulse facilities like shock tunnels. However, shock tunnel simulation imposes challenges and restrictions on the flow diagnostics, not just because of the possible extreme flow conditions, but also the short run times-typically around 1 ms. The development, calibration and application of fast response MEMS sensors for surface pressure measurements in IISc hypersonic shock tunnel HST-2, with a typical test time of 600 mu s, for the complex flow field of strong (impinging) shock boundary layer interaction with separation close to the leading edge, is delineated in this paper. For Mach numbers 5.96 (total enthalpy 1.3 MJ kg(-1)) and 8.67 (total enthalpy 1.6 MJ kg(-1)), surface pressures ranging from around 200 Pa to 50 000 Pa, in various regions of the flow field, are measured using the MEMS sensors. The measurements are found to compare well with the measurements using commercial sensors. It was possible to resolve important regions of the flow field involving significant spatial gradients of pressure, with a resolution of 5 data points within 12 mm in each MEMS array, which cannot be achieved with the other commercial sensors. In particular, MEMS sensors enabled the measurement of separation pressure (at Mach 8.67) near the leading edge and the sharply varying pressure in the reattachment zone.