Helen L. Reed

Supersonic laminar flow control on swept wings using distributed roughness

The present work addresses a new technology development that can lead to drag reduction on supersonic aircraft by means of passive laminar flow control (LFC). Recent developments in the understanding of stability and transition in swept-wing flows in low-disturbance environments have offered the promise of controlling transition without the use of complicated systems. The principal control problem with highly swept wings concerns the crossflow instability. It has been demonstrated in a series of low-speed experiments that distributed roughness near the attachment line can control the crossflow instability and can laminarize a boundary layer, provided an induced roughness wavelength is below a critical value. The present work extends this idea to supersonic flow over highly swept wings. The combined computational and experimental work gives design criteria and demonstrates LFC on airfoils swept beyond the characteristic Mach angle i.e. subsonic leading edges.

Passive control of transition in three-dimensional boundary layers, with emphasis on discrete roughness elements

A brief review of laminar flow control techniques is given and a strategy for achieving laminarization for transonic transport aircraft is discussed. A review of some flight-test results on swept-wing transition is presented. It is also shown that polished leading edges can create large regions of laminar flow because the flight environment is relatively turbulence free and the surface finish reduces the initial amplitude of the stationary crossflow vortex.

Direct Numerical Simulation of Discrete Roughness on a Swept-Wing Leading Edge

Direct numerical simulation is employed in order to describe the subsonic flow past an array of micron-sized discrete roughness elements, which were mounted near the leading edge of a 30-degree swept wing at a chord Reynolds number of 7.4 x 10 6 . The flow conditions correspond to flight receptivity experiments that were conducted to investigate the effects of roughness on crossflow instabilities. To make the computations tractable, the geometry is scaled by the radius of the wing leading edge, which magnifies the region of interest and enhances resolution. The leading-edge region is then approximated by the flow past an infinite parabolic cylinder. The numerical method is based upon a sixth-order-accurate time-implicit scheme to attain high fidelity and was used in conjunction with an eighth-order low-pass Pade-type nondispersive filter operator to maintain stability. A high-order overset-grid approach preserved spatial accuracy on a local mesh system representing the roughness elements, using domain decomposition to perform calculations on a parallel computing platform. The direct simulation for the flow about the roughness elements was used to capture crossflow vortices and served as input to the nonlinear parabolized stability equations, which were then solved in order to determine receptivity of the flow to the geometric perturbations. Three different geometric roughness elemental shapes were investigated in the study. For one shape, the effect of element height was examined. Features of the roughness-element flowfields are elucidated, and findings of the stability calculations are compared. Results are presented for receptivity of the crossflow instability to the size and shape of elements, as obtained by the direct numerical simulation and by two different stability approaches.

CROSSFLOW INSTABILITIES - THEORY & TECHNOLOGY

Over the years, the crossflow instability has been the primary challenge for Laminar Flow Control (LFC). Favorable pressure gradients used to stabilize streamwise instabilities destabilize crossflow. For years, it seemed as though the only solution to crossflow control was surface suction. The perceived complications with moving parts and additional maintenance were always discouraging factors toward laminarizing swept wings.

Laminar Flow Control on a Swept Wing with Distributed Roughness

The work cumulated in a series of laminar-turbulent transition flight-test experiments on a swept wing with the goal of validating the spanwise-periodic distributed roughness elements (DRE) technology in a Reynolds number range applicable to SensorCraft technology. Phase I of the program measured freestream turbulence levels that were nominally 0.05% to 0.06% of the freestream speed and thus established the suitability of the flight environment for the laminarization flights. Phase II of the program did the baseline transition measurements on the airfoil i.e. with and without DRE technology. The region of laminar flow was extended from 30% to 60% chord at a chord Reynolds number of Rec = 8 x10 6 and sweep angle, Λ = 30°. Establishing the origins of turbulent flow and transition from laminar to turbulent flow remains an important challenge of fluid mechanics. The common thread connecting aerodynamic applications is the fact that they deal with bounded shear flows (boundary layers) in open systems (with different upstream or initial amplitude conditions). It is well known that the stability, transition, and turbulent characteristics of bounded shear layers are fundamentally different from those of free shear layers. Likewise, open systems are fundamentally different from those of closed systems. The distinctions are trenchant and thus form separate areas of study. For the classic open system, no mathematical model exists that can predict the transition Reynolds number on a simple flat plate because the influences of freestream turbulence, sound, and surface roughness are incompletely understood. With the maturation of linear stability methods and the conclusions that breakdown mechanisms are initial-condition dependent, more emphasis is now placed on the understanding of the source of initial disturbances than on the details of the later stages of transition.

Roughness receptivity in swept-wing boundary layers – experiments

The receptivity of boundary layer stability to micron-sized, spanwise-periodic discrete roughness elements (DREs) was studied. The DREs were applied to the leading edge of a 30-degree swept-wing. The test article was attached vertically to the port wing of a Cessna O-2A aircraft and operated at a chord Reynolds number of 6.5 to 7.5 million. Critically spaced DREs were applied at the leading edge to excite the crossflow instability and move transition forward. In this case, calibrated, multi-element hotfilm sensors were used to measure disturbance wall shear stress. The roughness height was varied from 0 to 50 microns both in the positive (bumps) and negative (dimples) sense. Thus, the disturbance amplitude variations were determined as a function of modulated DRE heights.

Computational Fluid Dynamics Validation Issues in Transition Modeling

Laminar-turbulent transition is highly initial- and operating-condition dependent. Finding careful, archival experiments for comparison is the main validation issue for computational fluid dynamics (CFD) modeling. The CFD formulations validated to date demonstrate that if the environment and operating conditions can be modeled and input correctly, the computations (nonlinear parabolized stability equations and direct numerical simulations) agree quantitatively with the experiments. Future challenges for validation include successful CFD simulations of other available complete databases, CFD leadership in the identification and modeling of the effects of freestream disturbances, CFD leadership in the determination of relevant validation experiments for supersonic and hypersonic flows, careful validation experiments and CFD solutions for complex three-dimensional geometries, and simulations and validations for the high Reynolds numbers of flight.

JoKHeR: NPSE Simulations of Hypersonic Crossflow Instability

Transition prediction and control are critical enablers for hypersonic ight and there is a need for physics-based approaches, especially for three-dimensional (3-D) boundary layers. A nonlinear parabolized stability equation (NPSE) capability for quasi 3-D geometries, called JoKHeR, has been developed. The NPSE approach for modeling stability and transition is particularly attractive since it includes curvature, nonparallel, and nonlinear e ects at a fraction of the resources required for a direct numerical simulation (DNS). Results for the cross ow instability on the much studied 7 degree half angle cone yawed to 6 degrees at Mach 6 are presented. Steady basic state calculations are obtained with Pointwise and GASP. New methods for determining cross ow vortex trajectories as well as wave number variations along the trajectory are proposed. Veri cation with prior results including a DNS is shown.

Roughness receptivity in swept-wing boundary layers - computations

The crossflow instability responsible for transition over a swept wing has been found to be ultra-sensitive to micron-sized roughness at the leading edge. Transition-predictive tools are limited because of the lack of models connecting physical roughness characteristics with initial and boundary conditions needed by the computational codes. The Texas A&M Flight Research Lab (FRL) is currently conducting flight tests of a laminar flow 30° swept wing model (SWIFT – swept wing in flight tests) mounted vertically below the port wing hard-point of a Cessna O-2A Skymaster and operated at chord Reynolds numbers on the order of 7.5 million. Various roughness configurations are correlated with local skin-friction measurements downstream. As a companion to the flight experiments, the group has engaged in a computational study aimed at relating roughness features to the resulting initial amplitude of the instability. This will provide a critical connection between stability analysis design tools and transition locat...

Boundary-layer instability and transition on a flared cone in a Mach 6 quiet wind tunnel

Measurements of boundary-layer transition location and boundary layer profiles on a sharp-tipped 5o-half-angle flared cone were made in a low-disturbance Mach 6 wind tunnel. Uncalibrated boundary-layer profiles of mean and fluctuating voltage representative of mass flux are obtained using constant temperature hot-wire anemometry at several axial locations, and are notionally compared with preliminary simulations. Spectral energy content is observed between 250 and 310 kHz – the first measurements of frequencies typical of the second mode instability at Texas A&M. Growth of this high-frequency content is compared with N-factor results from linear parabolised stability equation (LPSE) computations. Possible sources of disagreement between the experimental and computed frequencies for second-mode growth are discussed, as are future improvements to the hotwire anemometry technique. Nevertheless, the successful measurement of high-frequency content highlighted here constitutes an important step toward acquisit...