An experimental investigation of the ELAC 1 configuration at supersonic speeds

Supersonic flight of aerospace planes is of marked interest since several flow regimes characterized by different local flow structures have to be flown through. This problem was investigated experimentally for the hypersonic research configuration ELAC 1. The aim of the study was to detect the influence of the rounded leading edge, of the thickness distribution prescribed, and of the Reynolds number, especially on the flow on the leeward side of the configuration. The experiments were carried out in the transonic wind tunnel of Aerodynamisches Institut of RWTH Aachen, at a freestream Mach number Ma _∞=2, a unit Reynolds number of Re _∞=13×10⁶, angles of attack between ?3°?α?10°, and in a wind tunnel of the Institute for Theoretical and Applied Mechanics of the Russian Academy of Sciences in Novosibirsk. The freestream Mach numbers covered in these experiments were varied between 2?Ma _∞?4, freestream Reynolds numbers per unit length between 25×10⁶?Re _∞?56×10⁶ and angles of attack between ?3°?α?10°. Flow visualization studies, measurements of surface pressure distributions and of aerodynamic forces were used to analyze the flow. The results, which will also be compared with numerical data, clearly indicate marked differences in the location of the separation and reattachment lines, and the formation of the primary, secondary and tertiary vortices, for the flow regimes investigated.     

Study of the aerodynamic stability of a blunt cone with a side flap in a hypersonic flow

The stationary and time-dependent aerodynamic coefficients of a slender blunt cone with a flap located near the base section of the model are experimentally investigated. The freestream parameters (M_∞ = 6, Re _L = 0.88 × 10⁷, and γ = 1.4) ensured a turbulent regime of flow over the conical surface and the flap. At high angles of attack (α ～ 10°) laminar-turbulent transition is observable in the separation zone on the leeward side of the body. Emphasis is placed on the determination of the trimming angles of attack for different positions of the center of rotation and the static and dynamic stability coefficients (the model oscillation damping coefficient).     

Aerodynamic optimization of the flat-plate leading edge for experimental studies of laminar and transitional boundary layers

This work is concerned with the design of a leading edge for a flat-plate model used to study laminar and transitional boundary layers. For this study, the flow over the complete boundary-layer model, including leading edge, flat section, and trailing-edge flap, is modeled. The effect of important geometrical features of the leading edge on the resulting pressure distribution, starting from the well-known symmetric modified super ellipse, is investigated. A minimal pressure gradient on the measurement side of the plate is achieved using an asymmetrical configuration of modified super ellipses, with a thickness ratio of 7/24. An aerodynamic shape optimization is performed to obtain a novel leading edge shape that greatly reduces the length of the non-zero pressure gradient region and the adverse pressure gradient region compared to geometries defined by ellipses. Wind tunnel testing is used to validate the numerical solutions.     

Experimental study of the boundary layer over an airfoil in plunging motion

This is an experimental study on the boundary layer over an airfoil under steady and unsteady conditions.It specifically deals with the effect of plunging oscillation on the laminar/turbulent characteristics of the boundary layer.The wind tunnel measurements involved surfacemounted hot-film sensors and boundary-layer rake.The experiments were conducted at Reynolds numbers of 0.42×10 6 to 0.84 × 10 6 and the reduced frequency was varied from 0.01 to 0.11.The results of the quasi-wall-shear stress as well as the boundary layer velocity profiles provided important information about the state of the boundary layer over the suction surface of the airfoil in both static and dynamic cases.For the static tests,boundary layer transition occurred through a laminar separation bubble.By increasing the angle of attack,disturbances and the transition location moved toward the leading edge.For the dynamic tests,earlier transition occurred with increasing rather than decreasing effective angle of attack.The mean angle of attack and the oscillating parameters significantly affected the state of the boundary layer.By increasing the reduced frequency,the boundary layer transition was promoted to the upstroke portion of the equivalent angle of attack,but the quasi skin friction coefficient was decreased.     

An experimental investigation of unsteady turbulent-wake/boundary-layer interaction

An experimental investigation of unsteady-wake/boundary-layer interaction, similar to that occurring in turbomachinery, has been conducted in a specially modified wind tunnel. Unsteadiness in a turbomachine is periodic in nature, due to the relative motion of rotor and stator blades, resulting in travelling-wave disturbances that affect the blade boundary layers. In the experimental rig, travelling-wave disturbances were generated by a moving airfoil apparatus installed upstream of a flat plate to provide a two-dimensional model of a turbomachine stage. The boundary layer on the flat plate was tripped near the leading edge to generate a turbulent flow prior to interaction with the wakes, and measurements of velocity throughout the boundary layer were taken with a hot-wire probe. The Reynolds number, based on distance along the plate, ranged from 0.144×10⁵ to 1.44×10⁵, and all data were reduced through a process of ensemble averaging. Due to the nonlinear interactions with the boundary layer, the travelling discrete frequency wakes were found to decrease the shape factor of the velocity profile and to increase the level of turbulent fluctuations. Unlike the phase advance found with stationary-wave external disturbances, velocity profiles subject to the travelling wake fluctuations exhibited increasingly negative phase shifts from the free-stream towards the wall.     

Computation of two‐dimensional flows past ram‐air parachutes

Computational results for flow past a two‐dimensional model of a ram‐air parachute with leading edge cut are presented. Both laminar (Re=10⁴) and turbulent (Re=10⁶) flows are computed. A well‐proven stabilized finite element method (FEM), which has been applied to various flow problems earlier, is utilized to solve the incompressible Navier–Stokes equations in the primitive variables formulation. The Baldwin–Lomax model is employed for turbulence closure. Turbulent flow computations past a Clarck‐Y airfoil without a leading edge cut, for α=7.5°, result in an attached flow. The leading edge cut causes the flow to become unsteady and leads to a significant loss in lift and an increase in drag. The flow inside the parafoil cell remains almost stagnant, resulting in a high value of pressure, which is responsible for giving the parafoil its shape. The value of the lift‐to‐drag ratio obtained with the present computations is in good agreement with those reported in the literature. The effect of the size and location of the leading edge cut is studied. It is found that the flow on the upper surface of the parafoil is fairly insensitive to the configuration of the cut. However, the flow quality on the lower surface improves as the leading edge cut becomes smaller. The lift‐to‐drag ratio for various configurations of the leading edge cut varies between 3.4 and 5.8. It is observed that even though the time histories of the aerodynamic coefficients from the laminar and turbulent flow computations are quite different, their time‐averaged values are quite similar. Copyright © 2001 John Wiley & Sons, Ltd.     

Non-Darcy Free Convection Along a Horizontal Heated Surface

In this paper we analyse how the presence of inertia (Forchheimerform-drag) affects the steady free convective boundary layer flow over anupward-facing horizontal surface embedded in a porous medium. The surfacetemperature is assumed to display a power-law variation,x ⁿ with distance from the leading edge, x. It is shown thatthere are three distinct cases to consider: n<0.5, n=0.5 and0.5相似文献   

Turbulent free convection in a vertical converging channel

This paper presents the results of experimental and numerical investigations of the problem of turbulent natural convection in a converging-plate vertical channel. The channel has two isothermally heated inclined walls and two adiabatic vertical side walls. The parameters involved in this study are the channel geometry represented by the channel width at exit, the inclination of the heated walls and the temperature difference between the heated walls and the ambient. The investigation covered modified Rayleigh numbers up to 10⁸ in the computational study and up to 9.3 × 10⁶ in the experimental work. The experimental measurements focused on the velocity field and were carried out using a PIV system and included measurements of the mean velocity profiles as well as the root-mean-square velocity and shear stress profiles. The experiments were conducted for an inclination angle of 30°, a gap width of 10 mm and two temperature differences (∆T=25.4°C and 49.8°C). The velocity profiles in the lower part of the channel indicated the presence of two distinct layers. The first layer is adjacent to the heated plate and driven by buoyancy forces while the second layer extends from the point of maximum velocity to the channel center plane and driven mainly by shear forces. The velocity profile at the upper portion of the channel has shown the merging of the two boundary layers growing over the two heated walls. The measured values of the Reynolds shear stress and root mean square of the horizontal and vertical velocity fluctuation components have reached their maximum near the wall while having smaller values in the core region. The computational results have shown that the average Nusselt number increases approximately linearly with the increase of the modified Rayleigh number when plotted on log–log scale. The variation of the local Nusselt number indicated infinite values at the channel inlet (leading edge effect) and high values at the channel exit (trailing edge effect). For a fixed value of the top channel opening, the increase of the inclination angle tended to reduce flow velocity at the inlet section while changing the flow structure near the heated plates in such a way to create boundary-layer type flow. The maximum value of the average Nusselt number occurs when θ = 0 and decreases with the increase of the inclination angle. On the other hand, the increase of the channel width at exit for the same inclination angle caused a monotonic increase in the flow velocity at the channel inlet.     

Effect of leading edge cut on the aerodynamics of ram‐air parachutes

The effect of the configuration of leading edge cut on the aerodynamic performance of ram‐air parachutes is studied via two‐dimensional flow simulations. The incompressible Reynolds‐averaged Navier–Stokes equations, in primitive variables, are solved using a stabilized finite‐element formulation. The Baldwin–Lomax model is employed for turbulence closure. Flow past an LS(1) 0417 airfoil is investigated for various configurations of the leading edge cut and results are compared with those from a Clark‐Y airfoil section. It is found that the configuration of the leading edge cut affects the lift‐to‐drag ratio (L/D) of the parachute very significantly. The L/D value has strong implications on the flight performance of the parachute. One particular configuration results in a L/D value that is in excess of 25 at 7.5° angle of attack. Results are presented for other angles of attack for this configuration. Copyright © 2004 John Wiley & Sons, Ltd.     

Turbulent separated flow over and downstream of a two-element airfoil

Flow characteristics in the vicinity of the flap of a single-slotted airfoil are presented and analysed. The flow remained attached over the model surfaces except in the vicinity of the flap trailing edge where a small region of boundary-layer separation extended over the aft 7% of flap chord. The airfoil configuration was tested at a Mach number of 0.09 and a chord Reynolds number of 1.8 × 10⁶ in the NASA Ames Research Center 7- by 10-Foot Wind Tunnel. The flow was complicated by the presence of a strong, initially inviscid, jet, emanating from the slot between airfoil and flap, and a gradual merging of the main airfoil wake and flap suction-side boundary layer.Research Engineer, NRC Research AssociateAerospace Engineer     

Effect of incidence angle with wake passing on a film cooled leading edge: A numerical study

This work presents the numerical study of a film‐cooled blade under the influence of wake passing at different incidence angles. The film cooling technology has been proven to be effective to increase the blade life of first turbine stages. However, the leading edge is affected by an high heat transfer rate and cooling this region is difficult. Moreover, separated regions downstream the coolant injection increases the local heat transfer coefficient and can have a detrimental effect in terms of airfoil life. This work analyses how the flow field is affected by the wake passing at different incidence angles (?5, 0, 5) and the impact on heat transfer coefficient. The test case is a linear cascade with two rows of cylindrical holes at the leading edge. Two different holes arrangements are compared in terms of film cooling structures, namely AGTB‐B1 and AGTB‐B2 with 0 and 45^° spanwise inclination. The numerical results show a good agreement with the experiments. A deeper investigation is carried out on AGTB‐B1. The results obtained show that the wake passing and the incidence angle have a strong effect on coolant jets. In particular, there is a significative impact on coolant redistribution near the leading edge. The wake passing has a stronger effect on pressure side, mainly at negative incidence. The predictive approach is based on an U‐RANS in‐house CFD solver using a conventional two‐equations closure. In order to avoid extra turbulence production, critical in the leading edge region, the turbulence model incorporates an extra algebraic equation that enforces a realizability constraint. The unsteady formulation is based on a dual time stepping approach with a sliding plane between the moving bars and the cascade. Copyright © 2009 John Wiley & Sons, Ltd.     

Leading edge bluntness effect on the flow in a model air-inlet

Flow in an idealized air-inlet with plane walls and a rectangular cross-section is experimentally investigated. The air-inlet is mounted on a plate, at a distance well removed from its leading edge. The experiments were conducted in a Ludwig tube at M_∞ = 5 and Re_∞L = 23×10⁶ and 13×10⁶. A panoramic (optical) technique of measuring the heat transfer coefficient is for the first time applied to study the internal flow in an air-inlet. The data on the effect of the bluntness of the leading edges of the plate and the air-inlet cowl on the heat transfer coefficient distribution and the flow structure within the air-inlet are obtained. It is shown that in an air-inlet with large channel constriction an even small bluntness of the plate or the cowl can lead to global changes in the flow structure.     

Characterization of noise amplifiers with global singular modes: the case of the leading-edge flat-plate boundary layer

This article deals with the linear dynamics of a transitional boundary layer subject to two-dimensional Tollmien–Schlichting instabilities. We consider a flat plate including the leading edge, characterized by a Reynolds number based on the length of the plate equal to Re = 6 × 10⁵, inducing a displacement thickness-based Reynolds number of 1,332 at the end of the plate. The global linearized Navier–Stokes equations only display stable eigenvalues, and the associated eigen-vectors are known to poorly represent the dynamics of such open flows. Therefore, we resort to an input–output approach by considering the singular value decomposition of the global resolvent. We then obtain a series of singular values, an associated orthonormal basis representing the forcing (the so-called optimal forcings) as well as an orthonormal basis representing the response (the so-called optimal responses). The objective of this paper is to analyze these spatial structures and to closely relate their spatial downstream evolution to the Orr and Tollmien–Schlichting mechanisms. Analysis of the spatio-frequential diagrams shows that the optimal forcings and responses are respectively localized, for all frequencies, near the upstream neutral point (branch I) and the downstream neutral point (branch II). It is also shown that the spatial growth of the dominant optimal response favorably compares with the e ^N method in regions where the dominant optimal forcing is small. Moreover, thanks to an energetic input–output approach, it is shown that this spatial growth is solely due to intrinsic amplifying mechanisms related to the Orr and Tollmien–Schlichting mechanisms, while the spatial growth due to the externally supplied power by the dominant optimal forcing is negligible even in regions where the dominant optimal forcing is strong. The energetic input–output approach also yields a general method to assess the strength of the instability in amplifier flows. It is based on a ratio comparing two quantities of same physical dimension, the mean-fluctuating kinetic energy flux of the dominant optimal response across some boundary and the supplied mean external power by the dominant optimal forcing. For the present boundary-layer flow, we have computed this amplification parameter for each frequency and discussed the results with respect to the Orr and Tollmien–Schlichting mechanisms.     

Multiple laminar-turbulent transition cycles around a swept leading edge

Certain interesting flow features involving multiple transition/relaminarization cycles on the leading edge of a swept wing at low speeds are reported here. The wing geometry tested had a circular nose and a leading edge sweep of 60°. Tests were made at a chord Reynolds number of 1.3 × 10⁶ with model incidence α varied in the range of 3°?18° in discrete steps. Measurements made included wing chord-wise surface pressure distributions and wall shear stress fluctuations (using hot-film gages) within about 10 % of the chord in the leading edge zone. Results at α = 16° and 18° showed that several (often incomplete) transition cycles between laminar-like and turbulent-like flows occurred. These rather surprising results are attributable chiefly to the fact that the Launder acceleration parameter K (appropriately modified for swept wings) can exceed a critical range more than once along the contour of the airfoil in the leading edge region. Each such crossing results in a relaminarization followed by direct retransition to turbulence as K drops to sufficiently low values. It is further shown that the extent of each observed transition zone (of either type) is consistent with earlier data acquired in more detailed studies of direct transition and relaminarization. Swept leading edge boundary layers therefore pose strong challenges to numerical modelling.     

Experimental study of evolution of disturbances in a supersonic boundary layer on a swept-wing model under controlled conditions

Experimental data on stability of a three-dimensional supersonic boundary layer on a swept wing are presented. Evolution of artificial wave trains was studied. The experiments were conducted for Mach numberM=2.0 and unit Reynolds numberRe ₁=6.6·10⁶m⁻¹ on a swept-wing model with a lenticular profile and a40° sweep angle of the leading edge at zero incidence. Excitation of high-frequency disturbances caused by secondary-flow instability at a high initial amplitude was observed. It is shown that the evolution of disturbances at frequencies of10, 20, and30 kHz is similar to the development of travelling waves for the case of subsonic velocities. Institute of Theoretical and Applied Mechanics, Siberian Division, Russian Academy of Sciences, Novosibirsk 630090. Translated from Prikladnaya Mekhanika i Tekhnicheskaya Fizika, Vol. 41, No. 1, pp. 50–56, January–February, 2000.     

A visual study in the near field of turbulent jets and implications for estimating accidental discharges

A series of dye flow visualization experiments are carried out in water to study the visible flow features in the near field of turbulent jets and to assess their usefulness in estimating the discharge rate of a turbulent jet in a homogeneous medium. The jet Reynolds numbers are 0.3–2.2 × 10⁵. The large eddies at the core of the flow and the smaller eddies at the edge show disparate, independent length scales. Their convection speeds are more than an order of magnitude apart. Discharge rate estimates based on large-scale core features are useful. However, their reliability depends on a priori knowledge of the state of the bulk flow upstream of the discharge location. A useful method for estimating discharge rates based on the small-scale outer edge features is not obvious.     

Sail aerodynamics: Understanding pressure distributions on upwind sails

The pressure distributions on upwind sails is discussed and related to the flow field around the headsail and the mast/mainsail. Pressures measured on several horizontal sections of model-scale and full-scale sails are used to provide examples. On the leeward side of the sails, leading edge separation and turbulent reattachment occurs, sometimes followed by trailing edge separation. On the windward side, leading edge separation occurs on the mast/mainsail and, at low angles of attack, it can also occur on the headsail. Differences were found between the leading edge bubbles on the two sails. Pressure trends for different angles of attack are presented, and these can be explained in terms of standard aerodynamic theory, particularly in terms of short and long leading edge separation bubble types. It was found that the pressure distributions measured on mainsails at full- and model-scale showed good agreement on both the windward and leeward sides.     

Acoustic receptivity of Mach 4.5 boundary layer with leading-edge bluntness

Boundary layer receptivity to two-dimensional slow and fast acoustic waves is investigated by solving Navier–Stokes equations for Mach 4.5 flow over a flat plate with a finite-thickness leading edge. Higher order spatial and temporal schemes are employed to obtain the solution whereby the flat-plate leading edge region is resolved by providing a sufficiently refined grid. The results show that the instability waves are generated in the leading edge region and that the boundary-layer is much more receptive to slow acoustic waves (by almost a factor of 20) as compared to the fast waves. Hence, this leading-edge receptivity mechanism is expected to be more relevant in the transition process for high Mach number flows where second mode instability is dominant. Computations are performed to investigate the effect of leading-edge thickness and it is found that bluntness tends to stabilize the boundary layer. Furthermore, the relative significance of fast acoustic waves is enhanced in the presence of bluntness. The effect of acoustic wave incidence angle is also studied and it is found that the receptivity of the boundary layer on the ‘windward’ side (with respect to the acoustic forcing) decreases by more than a factor of four when the incidence angle is increased from 0° to 45°. However, the receptivity coefficient for the ‘leeward’ side is found to vary relatively weakly with the incidence angle.      

Laminar-turbulent transition in the boundary layer on cones in a hypersonic flow at high Reynolds numbers per meter

The laminar-turbulent transition is experimentally studied in boundary-layer flows on cones with a rectangular axisymmetric step in the base part of the cone and without the step. The experiments are performed in an A-1 two-step piston-driven gas-dynamic facility with adiabatic compression of the working gas with Mach numbers at the nozzle exit M _∞ = 12–14 and pressures in the settling chamber P₀ = 60–600 MPa. These values of parameters allow obtaining Reynolds numbers per meter near the cone surface equal to Re _1e = (53–200) · 10⁶ m ⁻¹. The transition occurs at Reynolds numbers Re _tr = (2.3–5.7) · 10⁶. __________ Translated from Prikladnaya Mekhanika i Tekhnicheskaya Fizika, Vol. 48, No. 3, pp. 76–83, May–June, 2007.     

Non-symmetric bi-stable flow around the Ahmed body

The flow around the Ahmed body at varying Reynolds numbers under yawing conditions is investigated experimentally. The body geometry belongs to a regime subject to spanwise flow instability identified in symmetric flow by Cadot and co-workers (Grandemange et al., 2013b). Our experiments cover the two slant angles 25° and 35° and Reynolds numbers up to 2.784 × 10⁶. Special emphasis lies on the aerodynamics under side wind influence. For the 35° slant angle, forces and moments change significantly with the yawing angle in the range 10° ≤ |β| ≤ 15°. The lift and the pitching moment exhibit strong fluctuations due to bi-stable flow around a critical angle β of ±12.5°, where the pitching moment changes sign. Time series of the forces and moments are studied and explained by PIV measurements in the flow field near the rear of the body.