The flying hot wire and related instrumentation

A flying hot-wire technique is proposed for studies of separated turbulent flow in wind tunnels. The technique avoids the problem of signal rectification in regions of high turbulence level by moving the probe rapidly through the flow on the end of a rotating arm. New problems which arise include control of effects of torque variation on rotor speed, avoidance of interference from the wake of the moving arms, and synchronization of data acquisition with rotation. Solutions for these problems are described. The self-calibrating feature of the technique is illustrated by a sample X-array calibration

Structure of the Turbulent Separated Flow Around a Stalled Airfoil

Hot-wire measurements were made in the boundary layer, the separated region, and the near wake for flow past a NACA 4412 airfoil at maximum lift. The Reynolds number based on chord was 1,500,000. Special care was taken to achieve a two-dimensional mean flow. Data were obtained at several thousand locations in the flow field. These data include intermittency, two components of mean velocity, and mean values for three double, four triple, and five quadruple products of two velocity fluctuations. No information was obtained about the third (spanwise) velocity component. Smoothing and interpolating routines were used to determine intermittency, two components of mean velocity, and mean values of three double, four triple, and five quadruple products of two velocity fluctuations on a fine rectangular mesh aligned with the airfoil chord. The data are presented in contour plots, in three-dimensional plots, and in tabular form. The format used to store the experimental data in digital form is described and a computer program which illustrates how this data can be accessed is presented

Simple turbulence models and their application to boundary layer separation

Measurements in the boundary layer and wake of a stalled airfoil are presented in two coordinate systems, one aligned with the airfoil chord, the other being conventional boundary layer coordinates. The NACA 4412 airfoil is studied at a single angle of attack corresponding to maximum lift, the Reynolds number based on chord being 1.5 x 10 to the 6th power. Turbulent boundary layer separation occurred at the 85 percent chord position. The two-dimensionality of the flow was documented and the momentum integral equation studied to illustrate the importance of turbulence contributions as separation is approached. The assumptions of simple eddy-viscosity and mixing-length turbulence models are checked directly against experiment. Curvature effects are found to be important as separation is approached

Flying-Hot-Wire Study of 2-Dimensional Mean Flow Past an NACA 4412 Airfoil at Maximum Lift

Hot-wire measurements have been made in the boundary layer, the separated region, and the near wake for flow past an NACA 4412 airfoil at maximum lift. The Reynolds number based on chord was about 1,500,000. Special care was taken to achieve a two-dimensional mean flow. The main instrumentation was a hot-wire probe mounted on the end of a rotating arm. An unexpected effect of rotor interference was identified and brought under control. A digital computer was used to control synchronized sampling at closely spaced points along the probe arc. Ensembles of data were obtained at several thousand locations in the flow field. The data include intermittency, two components of mean velocity, and twelve mean values for double, triple, and quadruple products of two velocity fluctuations. The data are available on punched cards in raw form and also after use of smoothing and interpolation routines to obtain values on a fine rectangular grid aligned with the airfoil chord

Rotorcraft Downwash Flow Field Study to Understand the Aerodynamics of Helicopter Brownout

Rotorcraft brownout is caused by the entrainment of dust and sand particles in helicopter downwash, resulting in reduced pilot visibility during low, slow flight and landing. Recently, brownout has become a high-priority problem for military operations because of the risk to both pilot and equipment. Mitigation of this problem has focused on flight controls and landing maneuvers, but current knowledge and experimental data describing the aerodynamic contribution to brownout are limited. This paper focuses on downwash characteristics of a UH-60 Blackhawk as they pertain to particle entrainment and brownout. Results of a full-scale tuft test are presented and used to validate a high-fidelity Navier-Stokes computational fluid dynamics (CFD) calculation. CFD analysis for an EH-101 Merlin helicopter is also presented, and its flow field characteristics are compared with those of the UH-60

Separation delay via hydro-acoustic control of a NACA4412 airfoil in pre-stalled conditions

We have performed large-eddy simulations of turbulent separation control via impedance boundary conditions (IBCs) on a \nacafft airfoil in near-stalled conditions. The uncontrolled baseline flow is obtained for freestream Mach numbers of $M_\infty=0.3$, chord-Reynolds numbers $Re_c = 1.5\times10^6$ and angle of attack, $\alpha=14^{\circ{}}$. Flow control is applied via imposition of complex IBCs using the time-domain implementation developed by Scalo, Bodart, and Lele, \emph{Phys. Fluids} (2015). Separation is delayed due to the enhanced mixing associated with convectively amplified spanwise-oriented Kelvin-Helmholtz (KH) rollers, generated via hydro-acoustic instabilities. The latter are the result of the interaction of the wall-normal transpiration through the impedance panel and the overlying mean background shear. The result is an alteration of the coupled instability between the separating shear layer and the vortex shedding in the wake (already present in the uncontrolled baseline flow) yielding unique wake topologies associated with different intensities for the passively generated KH vortical structures. Specifically, enhancements up to +13\% in the lift coefficients have been obtained. Results show that tuning of the resonant cavities below the natural shedding frequency is required to generate KH rollers structures with a sufficiently large entrainment diameter to encompass the full extent of the separated region, thereby enhancing mixing and promoting reattachment. Overall, the results presented in this work show that the adoption of hydro-acoustically tuned resonant panels is a promising passive control technique for boundary layer separation control

Vortex Wake Geometry of a Model Tilt Rotor in Forward Flight

The vortex wake trajectory from one rotor of a 0.25-scale V-22 tiltrotor model was measured for four test conditions in the NASA Ames 40- by 80-Foot Wind Tunnel. Vortex wake images were acquired using a laser light sheet and video camera. Wake trajectories were constructed by extracting vortex positions from the video images. Wake trajectories were also calculated using the comprehensive analysis CAMRAD II. Measured and calculated wake geometries exhibit similar trends when advance ratio is varied at fixed thrust or when thrust is varied at fixed advance ratio