The near wall jet of a normally impinging, uniform, axisymmetric, supersonic jet

1974 ◽  
Vol 66 (1) ◽  
pp. 159-176 ◽  
Author(s):  
J. C. Carling ◽  
B. L. Hunt
Keyword(s):  
Mach Number ◽  
Shock Layer ◽  
Supersonic Jet ◽  
Surface Flow ◽  
Wall Jet ◽  
Viscous Effects ◽  
Number Range ◽  
Axisymmetric Jet ◽  
Pressure Distributions ◽  
Near Wall

The near wall jet produced by directing a uniform axisymmetric jet of air normally onto a large flat plate has been investigated experimentally and theoretically for four jets in the Mach number range 1·64–2·77. Detailed measurements of the surface pressure and shadowgraph and surface flow pictures are presented. The results show that the mechanism which mainly determines the supersonic near wall jet is the jet-edge expansion and its reflexions from the sonic line and the wall-jet boundaries. The near wall jet is found to consist of an alternating series of expansion and recompression regions whose strengths depend on the jet Mach number and decay with distance. At Mach numbers of 2·4 and above, shock waves are observed in the first recompression region and at a Mach number of 2·77 the boundary layer separates locally. Further out, viscous effects become increasingly important and a constant-pressure shear flow is established at a distance which increases with jet Mach number. The application of the method of characteristics in an approximate manner reproduces a number of the features of the near wall jet which are observed experimentally.Pressure distributions obtained in the shock layer show that a stagnation bubble can occur and that its occurrence depends on factors such as the flow upstream of the nozzle. The wall-jet region is found to be largely independent of whether or not a bubble occurs in the shock layer.

The Impingement of a Uniform, Axisymmetric, Supersonic Jet on a Perpendicular Flat Plate

1971 ◽  
Vol 22 (4) ◽  
pp. 403-420 ◽  
Author(s):  
J. H. Gummer ◽  
B. L. Hunt
Keyword(s):  
Mach Number ◽  
Flat Plate ◽  
Supersonic Jet ◽  
Theoretical Description ◽  
Alternative Methods ◽  
Number Range ◽  
Sonic Line ◽  
Pressure Distributions ◽  
Experimental Shock ◽  
Mach Numbers

SummaryThe impingement region produced by directing a uniform, axisymmetric, supersonic jet of air normally onto a large, flat plate has been investigated experimentally and theoretically for four jets in the Mach number range 1·64 to 2·77. A qualitative theoretical description of the flow in the neighbourhood of the sonic line is given. A single-strip version of the method of Polynomial Approximation and Integral Relations (PIR) is applied to the flow, using two alternative methods of determining the centre-line shock height. The PIR predictions are compared to experimental shock shapes and pressure distributions. It is found that a PIR method in which the sonic line is assumed to intersect the shock at the jet edge leads to very good agreement with experiment at the higher jet Mach numbers, but accuracy is much reduced at the lower Mach numbers, the shock height being in error by about 62 per cent at a jet Mach number of 1·64. A change in flow pattern at small nozzle-to-plate distances is reported.

Shock–boundary layer interaction and energetics in transonic flutter

2017 ◽  
Vol 832 ◽  
pp. 212-240 ◽  
Author(s):  
Pradeepa T. Karnick ◽  
Kartik Venkatraman
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Turbulence Model ◽  
Viscous Effects ◽  
Trailing Edge ◽  
Aeroelastic System ◽  
Number Range ◽  
Shock Reversal ◽  
Transonic Flutter ◽  
Flutter Boundary

We study the influence of shock and boundary layer interactions in transonic flutter of an aeroelastic system using a Reynolds-averaged Navier–Stokes (RANS) solver together with the Spalart–Allmaras turbulence model. We show that the transonic flutter boundary computed using a viscous flow solver can be divided into three distinct regimes: a low transonic Mach number range wherein viscosity mimics increasing airfoil thickness thereby mildly influencing the flutter boundary; an intermediate region of drastic change in the flutter boundary due to shock-induced separation; and a high transonic Mach number zone of no viscous effects when the shock moves close to the trailing edge. Inviscid transonic flutter simulations are a very good approximation of the aeroelastic system in predicting flutter in the first and third regions: that is when the shock is not strong enough to cause separation, and in regions where the shock-induced separated region is confined to a small region near the trailing edge of the airfoil. However, in the second interval of intermediate transonic Mach numbers, the power distribution on the airfoil surface is significantly influenced by shock-induced flow separation on the upper and lower surfaces leading to oscillations about a new equilibrium position. Though power contribution by viscous forces are three orders of magnitude less than the power due to pressure forces, these viscous effects manipulate the flow by influencing the strength and location of the shock such that the power contribution by pressure forces change significantly. Multiple flutter points that were part of the inviscid solution in this regime are now eliminated by viscous effects. Shock motion on the airfoil, shock reversal due to separation, and separation and reattachment of flow on the airfoil upper surface, also lead to multiple aerodynamic forcing frequencies. These flow features make the flutter boundary quantitatively sensitive to the turbulence model and numerical method adopted, but qualitatively they capture the essence of the physical phenomena.

A numerical study on the single pulsed energy addition based unsteady wave drag reduction at varied hypersonic flow regimes

2020 ◽  
pp. 095441002097313
Author(s):  
Dathi SNV Rajasekhar Rao ◽  
Bibin John
Keyword(s):  
Steady State ◽  
Mach Number ◽  
Drag Reduction ◽  
Shock Layer ◽  
Blunt Body ◽  
Wave Drag ◽  
Flow Features ◽  
Energy Pulse ◽  
Energy Addition ◽  
Wave Drag Reduction

In this study, unsteady wave drag reduction in hypersonic flowfield using pulsed energy addition is numerically investigated. A single energy pulse is considered to analyze the time-averaged drag reduction/pulse. The blast wave creation, translation and its interaction with shock layer are studied. As the wave drag depends only on the inviscid aspects of the flowfield, Euler part of a well-established compressible flow Navier-Stokes solver USHAS (Unstructured Solver for Hypersonic Aerothermodynamics) is employed for the present study. To explore the feasibility of pulsed energy addition in reducing the wave drag at different flight conditions, flight Mach numbers of 5.75, 6.9 and 8.0 are chosen for the study. An [Formula: see text] apex angle blunt cone model is considered to be placed in such hypersonic streams, and steady-state drag and unsteady drag reductions are computed. The simulation results indicate that drag of the blunt-body can be reduced below the steady-state drag for a significant period of energy bubble-shock layer interaction, and the corresponding propulsive energy savings can be up to 9%. For energy pulse of magnitude 100mJ deposited to a spherical region of 2 mm radius, located 50 mm upstream of the blunt-body offered a maximum percentage of wave drag reduction in the case of Mach 8.0 flowfield. Two different flow features are found to be responsible for the drag reduction, one is the low-density core of the blast wave and the second one is the baroclinic vortex created due to the plasma energy bubble-shock layer interaction. For the same freestream stagnation conditions, these two flow features are noted to be very predominant in the case of high Mach number flow in comparison to Mach 5.75 and 6.9 cases. However, the ratio of energy saved to the energy consumed is noted as a maximum for the lower Mach number case.

Direct numerical simulations of Rayleigh–Bénard convection in water with non-Oberbeck–Boussinesq effects

2019 ◽  
Vol 881 ◽  
pp. 1073-1096 ◽  
Author(s):  
Andreas D. Demou ◽  
Dimokratis G. E. Grigoriadis
Keyword(s):  
Numerical Simulations ◽  
Two Dimensions ◽  
Direct Numerical Simulations ◽  
Wall Region ◽  
Number Range ◽  
Bénard Convection ◽  
Benard Convection ◽  
Rayleigh Benard Convection ◽  
Rayleigh Benard ◽  
Near Wall

Rayleigh–Bénard convection in water is studied by means of direct numerical simulations, taking into account the variation of properties. The simulations considered a three-dimensional (3-D) cavity with a square cross-section and its two-dimensional (2-D) equivalent, covering a Rayleigh number range of $10^{6}\leqslant Ra\leqslant 10^{9}$ and using temperature differences up to 60 K. The main objectives of this study are (i) to investigate and report differences obtained by 2-D and 3-D simulations and (ii) to provide a first appreciation of the non-Oberbeck–Boussinesq (NOB) effects on the near-wall time-averaged and root-mean-squared (r.m.s.) temperature fields. The Nusselt number and the thermal boundary layer thickness exhibit the most pronounced differences when calculated in two dimensions and three dimensions, even though the $Ra$ scaling exponents are similar. These differences are closely related to the modification of the large-scale circulation pattern and become less pronounced when the NOB values are normalised with the respective Oberbeck–Boussinesq (OB) values. It is also demonstrated that NOB effects modify the near-wall temperature statistics, promoting the breaking of the top–bottom symmetry which characterises the OB approximation. The most prominent NOB effect in the near-wall region is the modification of the maximum r.m.s. values of temperature, which are found to increase at the top and decrease at the bottom of the cavity.

Auto and Cross Correlation Measurements in a Turbine Cascade Using a Digital Correlator

1982 ◽  
Author(s):  
P. J. Bryanston-Cross ◽  
J. J. Camus
Keyword(s):  
Mach Number ◽  
High Speed ◽  
Cross Correlation ◽  
Turbine Blades ◽  
Image Plane ◽  
Strongly Correlated ◽  
Trailing Edge ◽  
Number Range ◽  
Cross Correlations ◽  
Digital Correlator

A simple technique has been developed which samples the dynamic image plane information of a schlieren system using a digital correlator. Measurements have been made in the passages and in the wakes of transonic turbine blades in a linear cascade. The wind tunnel runs continuously and has independently variable Reynolds and Mach number. As expected, strongly correlated vortices were found in the wake and trailing edge region at 50 KHz. Although these are strongly coherent we show that there is only limited cross-correlation from wake to wake over a Mach no. range M = 0.5 to 1.25 and variation of Reynolds number from 3 × 105 to 106. The trailing edge fluctuation cross correlations were extended both upstream and downstream and preliminary measurements indicate that this technique can be used to obtain information on wake velocity. The vortex frequency has also been measured over the same Mach number range for two different cascades. The results have been compared with high speed schlieren photographs.

Compressible Flowfield Characteristics of Butterfly Valves

1989 ◽  
Vol 111 (4) ◽  
pp. 400-407 ◽  
Author(s):  
M. J. Morris ◽  
J. C. Dutton
Keyword(s):  
Flow Separation ◽  
Pressure Ratio ◽  
Disk Surface ◽  
Surface Flow ◽  
Operating Conditions ◽  
Local Geometry ◽  
Operating Pressure ◽  
Butterfly Valve ◽  
Pressure Distributions ◽  
Flow Separation And Reattachment

The results of an experimental investigation into the flowfield characteristics of butterfly valves under compressible flow operating conditions are reported. The experimental results include Schlieren and surface flow visualizations and flowfield static pressure distributions. Two valve disk shapes have been studied in a planar, two-dimensional test section: a generic biconvex circular arc profile and the midplane cross-section of a prototype butterfly valve. The valve disk angle and operating pressure ratio have also been varied in these experiments. The results demonstrate that under certain conditions of operation the butterfly valve flowfield can be extremely complex with oblique shock waves, expansion fans, and regions of flow separation and reattachment. In addition, the sensitivity of the valve disk surface pressure distributions to the local geometry near the leading and trailing edges and the relation of the aerodynamic torque to flow separation and reattachment on the disk are shown.

Spin-up from rest of a compressible fluid in a rapidly rotating cylinder

1992 ◽  
Vol 237 ◽  
pp. 413-434 ◽  
Author(s):  
Jae Min Hyun ◽  
Jun Sang Park
Keyword(s):  
Mach Number ◽  
Compressible Fluid ◽  
Numerical Solutions ◽  
Stokes Equations ◽  
Axial Direction ◽  
Ekman Layer ◽  
Infinite Cylinder ◽  
Full Time ◽  
Viscous Effects ◽  
Initial State

Spin-up flows of a compressible gas in a finite, closed cylinder from an initial state of rest are studied, The flow is characterized by small reference Ekman numbers, and the peripheral Mach number is O(1). Comprehensive numerical solutions have been obtained for the full, time-dependent compressible Navier-Stokes equations. The details of the flow, temperature, and density evolution are described. In the early phase of spin-up, owing to the thermoacoustic disturbances caused by the compressible Rayleigh effect, the flows are oscillatory, and this oscillatory behaviour is pronounced at higher Mach numbers. The principal dynamical role of the Ekman layer is dominant over moderate times of orders of the homogeneous spin-up timescales. Owing to the density stratification in the radial direction, the Ekman layer is thicker in the central region of the interior. The interior azimuthal flows are mainly uniform in the axial direction. As the Mach number increases, the rate of spin-up in the interior becomes slower, and the propagating shear front is more diffusive. Explicit comparisons with the results for an infinite cylinder are made to ascertain the contributions of the endwall disks. In contrast to the usual incompressible spin-up from rest, the viscous effects are relatively more important for the case of a compressible fluid.

Measurements of Endwall Flows in Transonic Linear Turbine Cascades: Part II—High Flow Turning

2010 ◽  
Author(s):  
F. Taremi ◽  
S. A. Sjolander ◽  
T. J. Praisner
Keyword(s):  
Mach Number ◽  
Kinetic Energy ◽  
Flow Visualization ◽  
Flow Separation ◽  
Surface Flow ◽  
Streamwise Vorticity ◽  
Vortical Structures ◽  
Turbine Cascades ◽  
Mach Numbers ◽  
Surface Flow Visualization

An experimental investigation of two low-turning (90°) transonic linear turbine cascades was presented in Part I of the paper. Part II examines two high-turning (112°) turbine cascades. The experimental results include total pressure losses, streamwise vorticity and secondary kinetic energy distributions. The measurements were made using a seven-hole pressure probe downstream of the cascades. In addition to the measurements, surface flow visualization was conducted to assist in the interpretation of the flow physics. The turbine cascades in Part II, referred to as SL1F and SL2F, have the same inlet and outlet design flow angles, but different aerodynamic loading levels: SL2F is more highly loaded than SL1F. The surface flow visualization results show evidence of small flow separation on the suction side of both airfoils. At the design conditions (outlet Mach number ≈ 0.8), SL2F exhibits stronger vortical structures and larger secondary velocities than SL1F. The two cascades, however, produce similar row losses based on the measurements at 40% axial chord lengths downstream of the trailing edge. Additional data were collected at off-design outlet Mach numbers of 0.65 and 0.91. As the Mach number is raised, the cascades become more aft-loaded. The absolute blade loadings increase, but the Zweifel coefficients decrease due to higher outlet dynamic pressures. Both profile and secondary losses decrease at higher Mach numbers; the main vortical structures and the corresponding peak losses migrate towards the endwall, and there are reductions in secondary kinetic energy and exit flow angle variations. The streamwise vorticity distributions show smaller peak vorticities associated with the passage and the counter vortices at higher exit Mach numbers. The corner vortex, on the other hand, becomes more intensified, resulting in reduction of flow overturning near the endwall. The results for SL1F and SL2F are compared and contrasted with the results for the lower turning cascades presented in Part I. The possible effects of suction-surface flow separation on profile and secondary losses are discussed in this context. The current research project is part of a larger study concerning the effects of endwall contouring on secondary losses, which will be presented in the near future.

CFD Aided Design of Heat Transfer Plates for Gasketed Plate Heat Exchangers

2014 ◽  
Author(s):  
Ece Özkaya ◽  
Selin Aradag ◽  
Sadik Kakac
Keyword(s):  
Heat Transfer ◽  
Pressure Drop ◽  
Heat Exchangers ◽  
Three Dimensional ◽  
Separate Flow ◽  
Working Fluid ◽  
Number Range ◽  
Plate Heat ◽  
Pressure Distributions ◽  
Cfd Analyses

In this study, three-dimensional computational fluid dynamics (CFD) analyses are performed to assess the thermal-hydraulic characteristics of a commercial Gasketed Plate Heat Exchangers (GPHEx) with 30 degrees of chevron angle (Plate1). The results of CFD analyses are compared with a computer program (ETU HEX) previously developed based on experimental results. Heat transfer plate is scanned using photogrammetric scan method to model GPHEx. CFD model is created as two separate flow zones, one for each of hot and cold domains with a virtual plate. Mass flow inlet and pressure outlet boundary conditions are applied. The working fluid is water. Temperature and pressure distributions are obtained for a Reynolds number range of 700–3400 and total temperature difference and pressure drop values are compared with ETU HEX. A new plate (Plate2) with corrugation pattern using smaller amplitude is designed and analyzed. The thermal properties are in good agreement with experimental data for the commercial plate. For the new plate, the decrease of the amplitude leads to a smaller enlargement factor which causes a low heat transfer rate while the pressure drop remains almost constant.

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