Andrew L. Carpenter

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.

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.

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...

Swept-Wing Laminar Flow Control Studies Using Cessna O-2A Test Aircraft

This paper presents the results of a series of flight tests using a test wing section, mounted on a pylon of a Cessna O-2A aircraft, to study the effects of micron-sized spanwise-periodic distributed roughness elements (DRE) on a 30° swept-wing model. Laminar to turbulent transition data were acquired at a chord Reynolds number of 6.5 to 8 million at angles of attack between ± 2°. The swept-wing test article was mounted vertically from the hard points located on the wing. The modification to the original aircraft design resulted in a campaign of flying quality tests and flutter clearance tests up to 170 KIAS and sideslip angles of ± 7 degrees. The flight test results are presented in three categories: flying qualities with the test article, freestream turbulence measurements, and laminar to turbulent transition studies on the test article. Flying qualities with the test article showed changes in static stability, but very little change in dynamic stability of the aircraft. Freestream turbulence measurements demonstrated turbulence intensity values suitable for boundary-layer stability experiments in the location of the test article. Transition results showed substantial amounts of natural laminar flow beyond the pressure minimum with a polished leading edge. An infrared camera was used as the laminar to turbulent transition diagnostics tool. This unobtrusive, qualitative imaging technique has proven valuable in a flight test environment, where weight and simplicity were a driving factor. Nomenclature ΛLE = leading edge sweep Cp = pressure coefficient � = aircraft angle of attack (model sweep angle) β = aircraft sideslip angle (model angle of attack, AoA) c = chord U∞ = freestream velocity Rec = chord Reynolds Number Rek = Reynolds Number based on the height of a roughness element, k Rex = Reynolds Number based on chord location, x, with respect to the leading edge k = roughness element height CG = aircraft center of gravity KIAS = indicated airspeed in knots Vne = never exceed speed Vno = top speed in turbulent conditions

Roughness receptivity studies in a 3-D boundary layer – Flight tests and computations

The receptivity of 3-D boundary layers 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 at the wavelength of the most unstable disturbance. 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. Thus, the disturbance-shear-stress amplitude variations were determined as a function of modulated DRE heights. The computational work was conducted parallel to the flight experiments. The complete viscous flowfield over the O-2 aircraft with the SWIFT model mounted on the port wing store pylon was successfully modeled and validated with the flight data. This highly accurate basic-state solution was incorporated into linear stability calculations and the wave growth associated with the crossflow instability was calculated.