Mean and fluctuating pressure in boat-tail separated flows at transonic speeds

Experiments were carried out investigating the features of mean and unsteady surface pressure fluctuations in boat-tail separated flows relevant to launch vehicle configurations at transonic speeds. The tests were performed on a generic axisymmetric body in the Mach-number range of 0.7-1.2, and the important geometrical parameters, namely, the boat-tail angle and diameter ratio, were varied systematically. The measurements made included primarily the mean and unsteady surface-pressure fluctuations on nine different model configurations. Flow-visualization studies employing a surface oil flow, and schlieren techniques were carried out to infer features like boundary-layer separation, reattachment, and shock waves in the flow. The features of mean and fluctuating surface pressures are discussed in detail including aspects of similarity. It has been observed that, on a generic configuration employed in the present study, the maximum levels of surface-pressure fluctuations in the reattachment zone are appreciably lower than those found on launch vehicle configurations having a bulbous or hammerhead nose shape. A simple correlation is suggested for the maximum value of rms pressure fluctuations in the reattachment zone at different freestream Mach numbers.

A study of shock wave interaction with a rotating cylinder

Shock wave reflection over a rotating circular cylinder is numerically and experimentally investigated. It is shown that the transition from the regular reflection to the Mach reflection is promoted on the cylinder surface which rotates in the same direction of the incident shock motion, whereas it is retarded on the surface that rotates to the reverse direction. Numerical calculations solving the Navier-Stokes equations using extremely fine grids also reveal that the reflected shock transition from RRdouble right arrowMR is either advanced or retarded depending on whether or not the surface motion favors the incident shock wave. The interpretation of viscous effects on the reflected shock transition is given by the dimensional analysis and from the viewpoint of signal propagation.

Minimum error solutions of the Boltzmann equation for shock structure

‘Best’ solutions for the shock-structure problem are obtained by solving the Boltzmann equation for a rigid sphere gas by applying minimum error criteria on the Mott-Smith ansatz. The use of two such criteria minimizing respectively the local and total errors, as well as independent computations of the remaining error, establish the high accuracy of the solutions, although it is shown that the Mott-Smith distribution is not an exact solution of the Boltzmann equation even at infinite Mach number. The minimum local error method is found to be particularly simple and efficient. Adopting the present solutions as the standard of comparison, it is found that the widely used v2x-moment solutions can be as much as a third in error, but that results based on Rosen's method provide good approximations. Finally, it is shown that if the Maxwell mean free path on the hot side of the shock is chosen as the scaling length, the value of the density-slope shock thickness is relatively insensitive to the intermolecular potential. A comparison is made on this basis of present results with experiment, and very satisfactory quantitative agreement is obtained.

Linear stability, transient growth and the role of viscosity stratification in compressible Couette flow

Linear stability and the nonmodal transient energy growth in compressible plane Couette flow are investigated for two prototype mean flows: (a) the uniform shear flow with constant viscosity, and (b) the nonuniform shear flow with stratified viscosity. Both mean flows are linearly unstable for a range of supersonic Mach numbers (M). For a given M, the critical Reynolds number (Re) is significantly smaller for the uniform shear flow than its nonuniform shear counterpart; for a given Re, the dominant instability (over all streamwise wave numbers, α) of each mean flow belongs to different modes for a range of supersonic M. An analysis of perturbation energy reveals that the instability is primarily caused by an excess transfer of energy from mean flow to perturbations. It is shown that the energy transfer from mean flow occurs close to the moving top wall for “mode I” instability, whereas it occurs in the bulk of the flow domain for “mode II.” For the nonmodal transient growth analysis, it is shown that the maximum temporal amplification of perturbation energy, Gmax, and the corresponding time scale are significantly larger for the uniform shear case compared to those for its nonuniform counterpart. For α=0, the linear stability operator can be partitioned into L∼L̅ +Re2 Lp, and the Re-dependent operator Lp is shown to have a negligibly small contribution to perturbation energy which is responsible for the validity of the well-known quadratic-scaling law in uniform shear flow: G(t∕Re)∼Re2. In contrast, the dominance of Lp is responsible for the invalidity of this scaling law in nonuniform shear flow. An inviscid reduced model, based on Ellingsen-Palm-type solution, has been shown to capture all salient features of transient energy growth of full viscous problem. For both modal and nonmodal instability, it is shown that the viscosity stratification of the underlying mean flow would lead to a delayed transition in compressible Couette flow.

Transfer matrix modeling of hyperbolic and parabolic ducts with incompressible mean flow

A general differential equation for the propagation of sound in a variable area duct or nozzle carrying incompressible mean flow (of low Mach number) is derived and solved for hyperbolic and parabolic shapes. Expressions for the state variables of acoustic pressure and acoustic mass velocity of the shapes are derived. Self‐consistent expressions for the four‐pole parameters are developed. The conical, exponential, catenoidal, sine, and cosine ducts are shown to be special cases of hyperbolic ducts. Finally, it is shown that if the mean flow in computing the transmission loss of the mufflers involving hyperbolic and parabolic shapes was not neglected, little practical benefit would be derived.

Hypersonic wave drag reduction in re-entry capsules using concentrated energy deposition

Experiments are carried out in a shock tunnel at a nominal Mach number of 5.75 in order to study the effect of concentrated energy deposition on the drag force experienced by a 120° blunt cone. Electrical energy was deposited along the stagnation streamline of the model using a high voltage DC discharge circuit (1.5 – 3.5KW) and the drag force was measured by a single component accelerometer balance. Numerical simulations were also carried complimenting the experiments. These simulations showed a substantial drag reduction (20% ~ 65%) whereas the experiments show no appreciable reduction in drag

Force measurements on hypersonic waveriders in the IISc hypersonic shock tunnel HST2

The drag and lift coefficients for a viscous optimized Mach 6 conical waverider has been measured using an accelerometer force balance system in the IISc hypersonic shock tunnel. A rubber bush placed in between the waverider model and the steel sting ensures unrestrained motion to the model during shock tunnel testing (500 mu s). Two accelerometers mounted on the model are used to measure the model accelerations in the axial and normal directions. The measured value of lift to drag ratio at zero angle of incidence for the IISc conical waverider with viscous optimized leading edge is 2.149, which compares well with the value reported in the open literature (Anderson et al 1991) for similar class of waveriders designed for a flight Mach number of 6. The details of the experimental study along with illustrative numerical results are discussed in this paper.

A new 6-component accelerometer force balance for short duration ground testing facilities

A new six-component accelerometer force balance is developed and used in the HST2 shock tunnel of Indian Institute of Science. Aerodynamic forces and moments for a hypersonic slender body measured using this balance system at a free stream Mach number of 5.75 and Reynolds number of 1.5 million and stagnation enthalpy of 1.5 and 2 MJ/kg are presented. These measured values compare well with the theoretical values estimated using modified Newtonian theory.

Surface pressure and heat transfer measurements in the unsteady separated hypersonic flow field over double cones

Reliable bench mark experimental database in the separated hypersonic flow regime is necessary to validate high resolution CFD codes. In this paper we report the surface pressure and heat transfer measurements carried out on double cones (first cone semi-apex angle = 15, 25 deg.; second cone semi-apex angle= 35, 68 deg.) at hypersonic speeds that will be useful for CFD code validation studies. The surface pressure measurements are carried out at nominal Mach number of 8.35 in the IISc hypersonic wind tunnel. On the other hand the surface heat transfer measurements are carried out at a nominal Mach number of 5.75 in the IISc hypersonic shock tunnel. The flow separation point on the first cone, flow reattachment on the second cone and the wild fluctuation of the transmitted shock on the second cone surface (25/68 deg. double cone) in the presence of severe adverse pressure gradient are some of the flow features captured in the measurements. The results from the CFD studies indicate good agreement with experiments in the attached flow regime while considerable differences are noticeable in the separated flow regime.

A novel free floating accelerometer force balance system for shock tunnel applications

In order to overcome the interference of the model mounting system with the external aerodynamics of the body during shock tunnel testing, a new free floating internally mountable balance system that ensures unrestrained model motion during testing has been designed, fabricated and tested. Minimal friction ball bearings are used for ensuring the free floating condition of the model during tunnel testing. The drag force acting on a blunt leading edge flat plate at hypersonic Mach number has been measured using the new balance system. Finite element model (FEM) and CFD are exhaustively used in the design as well as for calibrating the new balance system. The experimentally measured drag force on the blunt leading edge flat plate at stagnation enthalpy of 0.7 and 1.2 MJ/kg and nominal Mach number of 5.75 matches well with FEM results. The concept can also be extended for measuring all the three fundamental aerodynamic forces in short duration test facilities like free piston driven shock tunnels.

Two-dimensional flow past circular cylinders using finite volume lattice Boltzmann formulation

In this article, an extension to the total variation diminishing finite volume formulation of the lattice Boltzmann equation method on unstructured meshes was presented. The quadratic least squares procedure is used for the estimation of first-order and second-order spatial gradients of the particle distribution functions. The distribution functions were extrapolated quadratically to the virtual upwind node. The time integration was performed using the fourth-order RungeKutta procedure. A grid convergence study was performed in order to demonstrate the order of accuracy of the present scheme. The formulation was validated for the benchmark two-dimensional, laminar, and unsteady flow past a single circular cylinder. These computations were then investigated for the low Mach number simulations. Further validation was performed for flow past two circular cylinders arranged in tandem and side-by-side. Results of these simulations were extensively compared with the previous numerical data. Copyright (C) 2011 John Wiley & Sons, Ltd.

Pressure-sensitive paint on a truncated cone in hypersonic flow at incidences

The flow over a truncated cone is a classical and fundamental problem for aerodynamic research due to its three-dimensional and complicated characteristics. The flow is made more complex when examining high angles of incidence. Recently these types of flows have drawn more attention for the purposes of drag reduction in supersonic/hypersonic flows. In the present study the flow over a truncated cone at various incidences was experimentally investigated in a Mach 5 flow with a unit Reynolds number of 13.5�10 6m -1. The cone semi-apex angle is 15° and the truncation ratio (truncated length/cone length) is 0.5. The incidence of the model varied from -12° to 12° with 3° intervals relative to the freestream direction. The external flow around the truncated cone was visualised by colour Schlieren photography, while the surface flow pattern was revealed using the oil flow method. The surface pressure distribution was measured using the anodized aluminium pressure-sensitive paint (AA-PSP) technique. Both top and sideviews of the pressure distribution on the model surface were acquired at various incidences. AA-PSP showed high pressure sensitivity and captured the complicated flow structures which correlated well with the colour Schlieren and oil flow visualisation results. © 2012 Elsevier Inc.

Investigations on supersonic combustion in a combustion-driven shock tunnel

This article reports experimental results on supersonic combustion in a new facility. The facility is a combustion-driven shock tunnel, which is cheaper to build than the facilities in which such experiments are carried out conventionally. The observation region is a zone between two parallel flat plates with a 33 degrees wedge attached to the upstream end of the bottom plate. Gaseous hydrogen is injected at an angle of 45 degrees into an oncoming supersonic flow of Mach 2 (approximate) from a port on the bottom plate. The resulting flow field is visualized by a high speed camera in a dark background. Three different test gases, namely nitrogen, air, and oxygen-rich air are used, and the results are compared. A distinct luminosity due to combustion for oxygen-containing test gases is observed. Heat-transfer rates on a probe placed at the downstream end of the observation region and midway between the parallel plates are measured and compared for the three cases. Wall static pressure at 28 mm downstream of the injection port on the bottom plate is also presented.

Directional Diffusion Regulator (DDR) for some numerical solvers of hyperbolic conservation laws

A computational tool called ``Directional Diffusion Regulator (DDR)'' is proposed to bring forth real multidimensional physics into the upwind discretization in some numerical schemes of hyperbolic conservation laws. The direction based regulator when used with dimension splitting solvers, is set to moderate the excess multidimensional diffusion and hence cause genuine multidimensional upwinding like effect. The basic idea of this regulator driven method is to retain a full upwind scheme across local discontinuities, with the upwind bias decreasing smoothly to a minimum in the farthest direction. The discontinuous solutions are quantified as gradients and the regulator parameter across a typical finite volume interface or a finite difference interpolation point is formulated based on fractional local maximum gradient in any of the weak solution flow variables (say density, pressure, temperature, Mach number or even wave velocity etc.). DDR is applied to both the non-convective as well as whole unsplit dissipative flux terms of some numerical schemes, mainly of Local Lax-Friedrichs, to solve some benchmark problems describing inviscid compressible flow, shallow water dynamics and magneto-hydrodynamics. The first order solutions consistently improved depending on the extent of grid non-alignment to discontinuities, with the major influence due to regulation of non-convective diffusion. The application is also experimented on schemes such as Roe, Jameson-Schmidt-Turkel and some second order accurate methods. The consistent improvement in accuracy either at moderate or marked levels, for a variety of problems and with increasing grid size, reasonably indicate a scope for DDR as a regular tool to impart genuine multidimensional upwinding effect in a simpler framework. (C) 2012 Elsevier Inc. All rights reserved.

Theoretical-experimental study of shock wave-assisted metal forming process using a diaphragmless shock tube

The use of high-velocity sheet-forming techniques where the strain rates are in excess of 10(2)/s can help us solve many problems that are difficult to overcome with traditional metal-forming techniques. In this investigation, thin metallic plates/foils were subjected to shock wave loading in the newly developed diaphragmless shock tube. The conventional shock tube used in the aerodynamic applications uses a metal diaphragm for generating shock waves. This method of operation has its own disadvantages including the problems associated with repeatable and reliable generation of shock waves. Moreover, in industrial scenario, changing metal diaphragms after every shot is not desirable. Hence, a diaphragmless shock tube is calibrated and used in this study. Shock Mach numbers up to 3 can be generated with a high degree of repeatability (+/- 4 per cent) for the pressure jumps across the primary shock wave. The shock Mach number scatter is within +/- 1.5 per cent. Copper, brass, and aluminium plates of diameter 60 mm and thickness varying from 0.1 to 1 mm are used. The plate peak over-pressures ranging from 1 to 10 bar are used. The midpoint deflection, circumferential, radial, and thickness strains are measured and using these, the Von Mises strain is also calculated. The experimental results are compared with the numerical values obtained using finite element analysis. The experimental results match well with the numerical values. The plastic hinge effect was also observed in the finite element simulations. Analysis of the failed specimens shows that aluminium plates had mode I failure, whereas copper plates had mode II failure.