scholarly journals Shock-induced separated flows on the lee surface of delta wings

1987 ◽  
Vol 91 (903) ◽  
pp. 128-141 ◽  
Author(s):  
S. N. Seshadri ◽  
K. Y. Narayan
Keyword(s):  
Pressure Rise ◽  
Leading Edge ◽  
Experimental Results ◽  
Separated Flows ◽  
Delta Wings ◽  
Number Range ◽  
Pressure Distributions ◽  
Boundary Layer Interaction ◽  
Oil Flow ◽  
Attached Flow

Experiments were conducted to study shock-induced separated flows on the lee surface of delta wings with sharp leading edge at supersonic speeds. Two sets of delta wings of different thickness (10° and 25° normal angle), each with leading edge sweep angles varying from 45° to 70°, were tested. The measurements, carried out in a Mach number range from 1.4 to 3.0, included oil flow visualisations (on both sets of wings) and static pressure distributions (on the thicker wings only). Using the test results, some features of shock-induced separated flows, including in particular the boundary between this type of flow and fully attached flow, have been determined. The experimental results indicate that this boundary does not seem to show any significant dependence on wing thickness within the limit of thicknesses tested. It is shown that this boundary can be predicted for thin delta wings using a well known criterion for incipient separation in a glancing shock wave boundary layer interaction, namely that a pressure rise of 1.5 is required across the shock. Comparison of the predicted boundary with experimental results (from oil flow visualisations) shows good agreement.

The Pressure on Flat and Anhedral Delta Wings with Attached Shock Waves

1972 ◽  
Vol 23 (4) ◽  
pp. 253-262 ◽  
Author(s):  
J Pike
Keyword(s):  
Shock Wave ◽  
Pressure Distribution ◽  
Leading Edge ◽  
Surface Shape ◽  
Plane Shock ◽  
Delta Wings ◽  
Wave Condition ◽  
Pressure Distributions ◽  
Large Pressure ◽  
Attached Flow

SummaryAn expression is derived which relates the pressure on a wing in a supersonic free stream to the pressure on a thin wing with the same surface shape. The expression is used to find the pressure distribution for caret wings and flat delta wings with attached flow at their leading edges. The compression surface pressure distributions found are in good agreement with existing experimental and theoretical results, except when large pressure changes occur in the flow behind the attached shock wave. Some expansion surface results are also obtained for wings with an isentropic expansion at the leading edge. The effects of flow and geometry changes on the pressure distribution are investigated. It is found that a small improvement in the lift/drag ratio of a caret wing can be obtained by halving the anhedral required for the plane shock wave condition.

Steady longitudinal vortices in supersonic turbulent separated flows

2011 ◽  
Vol 672 ◽  
pp. 451-476 ◽  
Author(s):  
ERICH SCHÜLEIN ◽  
VICTOR M. TROFIMOV
Keyword(s):  
High Speed ◽  
Large Scale ◽  
Roughness Element ◽  
Vortex Generator ◽  
Vortex Pair ◽  
Leading Edge ◽  
Separated Flows ◽  
Vortex Generators ◽  
Longitudinal Vortices ◽  
Oil Flow

Large-scale longitudinal vortices in high-speed turbulent separated flows caused by relatively small irregularities at the model leading edges or at the model surfaces are investigated in this paper. Oil-flow visualization and infrared thermography techniques were applied in the wind tunnel tests at Mach numbers 3 and 5 to investigate the nominally 2-D ramp flow at deflection angles of 20°, 25° and 30°. The surface contour anomalies have been artificially simulated by very thin strips (vortex generators) of different shapes and thicknesses attached to the model surface. It is shown that the introduced streamwise vortical disturbances survive over very large downstream distances of the order of 104 vortex-generator heights in turbulent supersonic flows without pressure gradients. It is demonstrated that each vortex pair induced in the reattachment region of the ramp is definitely a child of a vortex pair, which was generated originally, for instance, by the small roughness element near the leading edge. The dependence of the spacing and intensity of the observed longitudinal vortices on the introduced disturbances (thickness and spanwise size of vortex generators) and on the flow parameters (Reynolds numbers, boundary-layer thickness, compression corner angles, etc.) has been shown experimentally.

Study on Subsynchronous Vibration With Tilting Pad Journal Bearing Under Starved Lubrication

2018 ◽  
Author(s):  
Rimpei Kawashita ◽  
Tadasuke Nishioka ◽  
Shimpei Yokoyama ◽  
Makoto Iwasaki ◽  
Shuichi Isayama ◽  
...  
Keyword(s):  
Flow Rate ◽  
Journal Bearing ◽  
Damping Ratio ◽  
Leading Edge ◽  
Rotor System ◽  
Experimental Results ◽  
Oil Film ◽  
Tilting Pad ◽  
Oil Flow ◽  
Starved Lubrication

Industrial machines such as gas and steam turbines require high efficiency and reliability. Direct lubricated bearings have been developed and installed to reduce mechanical losses. In recent years, it has been reported in the literature that subsynchronous vibration can occur to rotor shafts with direct lubricated tilting pad journal bearings under reduced oil flow rate conditions. In this study, a test rig with a 200 mm diameter and 3.5 meter long rotor supported by a direct lubricated tilting two pad journal bearing was constructed. The primary critical speed is 2100rpm and rotational speed is 3600rpm. The oil-starved area, the non-oil film layer region at the leading edge of the bearing pads, was measured by observing oil film pressure in the bearing clearance with pressure transducers on the rotor surface. A sine sweep excitation test was carried out by using an inertial shaker installed on the bearing housing and the damping ratio of the rotor system was measured. Measured data showed that a larger starved area at the leading edge of the bearing pads due to reduced oil feeding results in a smaller damping ratio, and an increase in the natural frequency of the rotor. Experimental results of two types of oil feeding nozzles were compared with respect to the correlation between starved area and damping ratio of the rotor system, and a relationship between oil flow rate and starved area was discussed. A method for modeling bearing coefficients under starved lubrication has been proposed based on thermo-hydrodynamic lubrication (THL) analysis. A numerical analysis of a finite element-transfer matrix model of the test rotor with the bearing coefficients calculated by the proposed method is carried out, and it is found that the analytical results are in broad agreement with the experimental results.

Experimental Control of Vortex Breakdown by Pulsed Blowing Over a Delta Wing With Rounded Leading-Edge

2002 ◽  
Author(s):  
Renac Florent ◽  
Molton Pascal ◽  
Barberis Didier
Keyword(s):  
Vortex Breakdown ◽  
Delta Wing ◽  
Laser Sheet ◽  
Leading Edge ◽  
Control Device ◽  
Surface Flow ◽  
Sweep Angle ◽  
Laminar To Turbulent Transition ◽  
Pressure Distributions ◽  
Oil Flow

The purpose of this study is to construct and test an experimental device to control vortex on a delta wing. The model has a root chord of c = 690mm and a sweep angle of Λ = 60°. The control system is based on four rectangular slits 50 mm long and 0.2 mm wide running along the leading edge. This configuration produces jets normal to the leading edge. The mass flow rates and frequencies of injection can be varied independently. The results are shown in the form of surface flow visualizations, with the skin friction pattern exhibited by oil flow visualization, and the laminar-to-turbulent transition by acenaphthene. Mean and instantaneous surface pressure distributions were determined with Kulite™ sensors and the velocity field was determined by 3D laser Doppler velocimetry (LDV) measurements. Control device efficiencies were evaluated by laser sheet visualization.

Lift Force of Delta Wings

1990 ◽  
Vol 43 (9) ◽  
pp. 209-221 ◽  
Author(s):  
Mario Lee ◽  
Chih-Ming Ho
Keyword(s):  
Lift Force ◽  
Delta Wing ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Surface Flux ◽  
Experimental Results ◽  
Geometrical Parameters ◽  
Two Dimensional ◽  
Delta Wings ◽  
Vorticity Balance

On a delta wing, the separation vorticies can be stationary due to the balance of the vorticity surface flux and the axial convection along the swept leading edge. These stationary vortices keep the wing from losing lift. A highly swept delta wing reaches the maximum lift at an angle of attack of about 40°, which is more than twice as high as that of a two-dimensional airfoil. In this paper, the experimental results of lift forces for delta wings are reviewed from the perspective of fundamental vorticity balance. The effects of different operational and geometrical parameters on the performance of delta wings are surveyed.

Possible types of flow on lee-surface of delta wings at supersonic speeds

1988 ◽  
Vol 92 (915) ◽  
pp. 185-199 ◽  
Author(s):  
S. N. Seshadri ◽  
K. Y. Narayan
Keyword(s):  
Shock Wave ◽  
Static Pressure ◽  
Delta Wing ◽  
Separation Bubble ◽  
Separated Flow ◽  
Leading Edge ◽  
Delta Wings ◽  
Oil Flow ◽  
Normal Angle ◽  
Secondary Separation

SummaryExperiments were conducted to study the types of flow that occur on the lee surface of delta wings at supersonic speeds. Two sets of flat topped delta wings of different thickness (wedges with 10° and 25° normal angle respectively), each with leading edge sweep angles of 45°, 50°, 60° and 70°, were tested. The measurements, carried out at Mach numbers of 1·4, 1·6, 1·8, 2·0, 2·5 and 3·0, included oil flow visualisations (on both sets of wings) and static pressure distributions (on the thicker wing only). In addition, a 60° sweptback delta wing with a normal angle of 40° was also tested. The tests on this wing included both oil flow visualisations and static pressure measurements. From these and other existing measurements, the leeside flows have been classified into nine distinct types, namely (i) leading edge separated flow with secondary separation, (ii) leading edge separated flow with secondary and tertiary separation, (iii) leading edge separated flow with a shock wave beneath the primary vortex, (iv) leading edge separated flow with shock-induced secondary separation, (v) fully attached flow, (vi) flow attached at the leading edge with inboard shock-induced separation, (vii) mixed type of flow, (viii) flow with a leading edge separation bubble and (ix) leading edge separated flow with a shock wave lying on the lee surface in between the leading edge vortices. These types of flow have been displayed in a plane of Mach number and angle of attack normal to the leading edge. The experimental results indicate that increasing wing thickness has no qualitative effect on the types of flow observed but does shift the boundaries between some of the types of flow.

scholarly journals Analysis of Transonic Flow Fields Inside a High Pressure Ratio Centrifugal Compressor at Design and Off Design Conditions

1999 ◽  
Author(s):  
Chunill Hah ◽  
Hartmut Krain
Keyword(s):  
Experimental Data ◽  
Boundary Layer ◽  
Centrifugal Compressor ◽  
Flow Separation ◽  
Pressure Rise ◽  
Leading Edge ◽  
Layer Interaction ◽  
Shock Boundary Layer Interaction ◽  
Blade Thickness ◽  
Boundary Layer Interaction

This paper reports on an experimental and numerical study of detailed flow structures in a transonic centrifugal compressor impeller at various operating conditions. Experimental data were obtained from conventional and laser-two-focus measurements inside the impeller. Numerical results were obtained from steady, three-dimensional Reynolds-averaged Navier-Stokes calculations. Both the experimental data and the numerical solutions at the design condition indicate that the flow incidence is high near the hub and flow separation exists near the leading edge. Due to the flow separation, low momentum fluid migrates rapidly to the tip area resulting in further loss generation through increased shock/boundary layer interaction. At a higher flow rate, a second passage shock develops near the leading edge of the splitter blade, further increasing shock/boundary layer interaction. Numerical studies were performed to explore possible design modifications for better efficiency and higher pressure rise. First, the blade camber near the leading edge was modified to improve the incidence. Second, the blade thickness was reduced by 50 percent. The incidence modification eliminates the flow separation near the leading edge and makes a more uniform flow split between the two channels, resulting in better flow distribution at the impeller exit. The simulated blade thickness reduction, along with the modified incidence, improves the efficiency by about 5 percent and increases the impeller pressure rise from 6.1:1 to 7.1:1.

A career in vortices and edge forces

2012 ◽  
Vol 116 (1176) ◽  
pp. 101-152 ◽  
Author(s):  
J. Lamar
Keyword(s):  
Leading Edge ◽  
Time Frame ◽  
Vortical Flow ◽  
Delta Wings ◽  
Analysis And Design ◽  
Research Aircraft ◽  
Attached Flow ◽  
Computer Codes ◽  
Langley Research Center ◽  
Analysis Mode

Abstract This lecture recognises the background and distinguished work of Frederick William Lanchester, and notes that my background has a few similarities with his. These include a shared interest in wings, lift and vortices. My career at the NASA Langley Research Center spans the time-frame from America’s Super Sonic Transport through 2009. An early emphasis involved wind-tunnel testing of research aircraft models and the development of computer codes for subsonic aerodynamics of wing planforms. These attached-flow codes were applied to various configurations, including those with variable-sweep, dihedral, and more than one planform in both the analysis- and design-modes. These codes were used to provide a connection between leading-edge-forces and the associated additional lift on delta-wings with shed-vortex systems through the leading-edge suction analogy of Edward C. Polhamus. Subsequently, I extended the suction analogy to configurations with side-edges to predict the vortical-flow aerodynamics on complex configurations, including wing-strake combinations. These analysis codes could also be used in a design-by-analysis mode for configurations with leading-edge shed vortices. Later, I was involved in vortical-flow flight research with the F-106B and the F-16XL aircraft at cruise and maneuver conditions. Associated CFD predictions, generated by me and other members of the RTO/AVT-113 task group, have increased our understanding of the flight flow-physics measured on the F-16XL aircraft.

Separated Flow Past a Slender Delta Wing at Incidence

1973 ◽  
Vol 24 (2) ◽  
pp. 120-128 ◽  
Author(s):  
J E Barsby
Keyword(s):  
Delta Wing ◽  
Separated Flow ◽  
Leading Edge ◽  
Vortex Sheet ◽  
Simultaneous Equations ◽  
Delta Wings ◽  
Nested Iteration ◽  
Finite Difference Technique ◽  
Flow Past ◽  
Vortex Systems

SummarySolutions to the problem of separated flow past slender delta wings for moderate values of a suitably defined incidence parameter have been calculated by Smith, using a vortex sheet model. By increasing the accuracy of the finite-difference technique, and by replacing Smith’s original nested iteration procedure, to solve the non-linear simultaneous equations that arise, by a Newton’s method, it is possible to extend the range of the incidence parameter over which solutions can be obtained. Furthermore for sufficiently small values of the incidence parameter, new and unexpected results in the form of vortex systems that originate inboard from the leading edge have been discovered. These new solutions are the only solutions, to the author’s knowledge, of a vortex sheet leaving a smooth surface.Interest has centred upon the shape of the finite vortex sheet, the position of the isolated vortex, and the lift, and variations of these quantities are shown as functions of the incidence parameter. Although no experimental evidence is available, comparisons are made with the simpler Brown and Michael model in which all the vorticity is assumed to be concentrated onto an isolated line vortex. Agreement between these two models becomes very close as the value of the incidence parameter is reduced.

Performance and Boundary Layer Development of a High Turning Compressor Cascade at Sub- and Supercritical Flow Conditions

2012 ◽  
Author(s):  
Christoph Bode ◽  
Dragan Kožulović ◽  
Udo Stark ◽  
Heinz Hoheisel
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Density Ratio ◽  
Boundary Layer Separation ◽  
Experimental Results ◽  
Numerical Code ◽  
Compressor Cascade ◽  
Boundary Layer Development ◽  
Boundary Layer Interaction ◽  
Inlet Angle

Based on current numerical investigations, the present paper reports on new Q2D midspan-calculations and results for the well known high turning (Δβ = 50°) supercritical (Ma1 = 0.85) compressor cascade V2. A Q2D treatment of the problem was chosen in order to avoid the difficult modelling of the porous endwalls in a corresponding 3D approach. All simulations were done with the RANS solver TRACE of the DLR Cologne in combination with modified versions of the Wilcox turbulence model and Langtry/Menter transition model. Existing experimental Q2D midspan-results for the V2 compressor cascade were used to demonstrate the improved ability of the numerical code to determine performance characteristics, blade pressure and Mach number distributions as well as boundary layer parameter and velocity distributions. The loss characteristics show minimum loss regions when plotted against inlet angle or axial velocity density ratio. Within these regions, increasing with decreasing Mach number, the experimental results were adequately predicted. Outside these regions it turned out difficult to reproduce the experimental results due to increasing boundary layer separation. Furthermore, the prediction quality was very good for subsonic conditions (Ma1 = 0.60) and still reasonable for supercritical conditions (Ma1 = 0.85), where shock/boundary layer interaction made the prediction more difficult.

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