W. Schröder

Nano-newton drag sensor based on flexible micro-pillars

A new simple and low-cost sensor concept to measure drag forces on fixed particles in the near-wall region with detectable forces down to pico-newton with a novel microoptomechanical system (MOMS) based on a flexible micro-pillar is presented. The cylindrical pillar with a diameter of a few microns is manufactured from an elastomer such that it is very flexible and easily deflected by the fluid forces or supplementary forces acting on flow obstacles attached to the tip. The pillar-tip bending is detected optically using a highly magnifying optical system. A feasibility study was carried out in a plate–cone rheometer with an air bubble of diameter 140 µm attached to the pillars tip to show that the micro-pillar sensor is capable of detecting net particle drag forces. The Reynolds number based on the bubble diameter Rebubble was varied in the range of 0.1–15. The calibration of the pure sensor structure shows linear behaviour of the pillar. The experimental results of the bubble drag in plane shear flow showed good agreement with theoretical predictions. The optics and the sensor used in this experiment allow a reliable detection of forces down to a minimum of about 5–10 nN. Using microscopic optics even smaller forces are detectable. The sensor geometry can be varied in a wide range, making it possible to measure forces of order down to nano-newtons on particles possessing a geometric length of a few µm to several hundred µm. Therefore, the method fits well into existing drag measurement concepts closing a gap between optical tweezers and recent atomic force cantilevers. Furthermore, the new sensor concept possesses very low intrusive interference. It has to be emphasized that unlike cantilever methods, the new micro-pillar sensor allows measurements of the two wall-parallel drag components. Using arrays of sensors, drag forces on multiple particles in a plane can be measured.

Coated Hot-Film Sensors for Transition Detection in Cruise Flight

In this paper results of wind-tunnel experiments and flight tests that demonstrate the effectiveness of coated multisensor hot-film arrays using constant current anemometry to determine boundary-layer transition under cruise conditions are presented. The study focuses on the analysis of several coating procedures, materials, and coating thicknesses to improve mechanical durability and signal quality. After several numerical simulations and environmental resistance tests, for example, in climate chambers and low-speed flight experiments, wind-tunnel experiments and flight tests at transonic speeds are carried out. The numerous experimental and numerical investigations and flight tests in cooperation with the Deutsche Airbus GmbH on a Beluga airplane (300-600ST) show a 1-μm parylene layer covering the extremely sensitive hot-film sensors and conducting paths to be sufficiently robust to stand the environmental conditions in cruise flight for several months. Furthermore, flight tests at transonic speeds on the Mystere Falcon 20 E5 at an altitude up to 30,000 ft and Mach numbers up to 0.79 are conducted in a joint project with DLR Oberpfaffenhofen. Modified wing models are mounted under both wings of the Falcon. The surfaces are instrumented with coated hot-film arrays using 128 sensors. The tests on the Falcon prove parylene coated hot-film sensors to still possess the sensitivity to locate transition under real flight conditions.

Source Term Evaluation of the APE-RF System

Acoustic perturbation equations for reacting flows (APE-RF) in conjunction with direct numerical simulations (DNS) are used to investigate in detail the thermo-acoustic effects resulting from turbulent premixed flames. The basic procedure is a two-step DNS/APE-RF method, where the flow is simulated by direct numerical simulations and the acoustic analysis is performed using the APE-RF system. Based on the DNS data, the source terms of the APE-RF system can be thoroughly evaluated, since the full chemical reaction is taken into account in the DNS. The acoustic impact of several source mechanisms are investigated, such as the effect of unsteady heat release, that of heat flux, viscous effects, the effect of non-isomolar combustion, and that of species diffusion. The study shows the unsteady heat release to be the dominant source. All source terms but the heat diffusion term possess a monopole-like structure in the low frequency range. At high frequencies a multipole-like pattern is also determined for the sources due to species diffusion and viscous effects. This deviation from the monopole structure is caused by the chemical reaction time scales. It is shown in this study that the radiated acoustic energy is in good agreement comparing the impact of the total time derivative of the density as major source term with the unsteady heat release rate.

Efficient Large-Eddy Simulation of Low Mach Number Flow

Turbulent flows at very low Mach numbers are numerically investigated using largeeddy simulation (LES). The numerical computations are carried out by solving the viscous conservation equations for compressible fluids. An implicit dual time stepping scheme combined with low Mach number preconditioning and a multigrid acceleration technique is developed for LES. The method is validated for turbulent channel flow at Re τ = 590 and flow across a cylinder at Re = 3900 and different low Mach numbers. The data are compared with numerical and experimental findings from the literature. The computations show an efficiency increase by a factor of up to 60.

Engine Jet/Vortex Interaction in the Near Wake of an Airfoil

*† ‡ Experimental results concerning the engine jet/vortex wake interaction will be presented. A method is described that improves the evaluation of unsteady vortex measurements and allows to analyze the effect of unsteady vortex movement. The results indicate that the engine jet modifies the stability properties of the wake flow. Depending on the engine position and operation mode the mean wandering amplitudes and the velocity fluctuations in the vortex wake are different and the topology of the vortex is affected. The current results suggest that for high-lift configurations the preferable engine position is closer to the wingtip if the focus is on the stability properties of the vortex wake. I. Introduction IRCRAFT separation rules in international aviation are a growing handicap for the ongoing rise in passenger numbers. To reduce the distance between following aircraft the physics of trailing vortices needs to be investigated. It is still a matter of scientific research to understand the mechanisms that can lead to a rapid decay of the vorticity strength. Since three decades in international civil air transportation the separation distances between flying aircraft are prescribed by rules, which divide the air fleet into different weight categories. The ICAO rules, for example, are based on three categories, below 7 tons maximum take-off weight, between 7 and 136 tons take-off weight, and more than 136 tons. The definition of these rules reaches back to flight tests in the early seventies, when the problem of wake vortex hazard first became apparent with the introduction of the Boeing B747. Since then the passenger numbers are continuously rising and forecasts from IATA, Boeing, or Airbus predict that this rise will go on. Major airports are reaching their capacity limits and this causes delays for airlines and passengers. The economic drawback of these delays urges to raise airport capacities by increasing the number of take offs and landings per time. However, this necessitates a revision of the current vortex wake related separation standards. That is, new aircraft separation rules have to be derived that at least keep or even improve the current safety level of international air transportation. This corresponds to the question, which safety margins exist in the current spacing. The answer to this question needs to be based on detailed knowledge about the physics of aircraft wakes, since even under today’s conditions there are nearly daily vortex encounters 1

A Hybrid LES/CAA Method for Aeroacoustic Applications

This paper describes a hybrid LES/CAA approach for the numerical prediction of airframe and combustion noise. In the hybrid method first a Large-Eddy Simulation (LES) of the flow field containing the acoustic source region is carried out from which then the acoustic sources are extracted. These are then used in the second computational Aeroacoustics (CAA) step in which the acoustic field is determined by solving linear acoustic perturbation equations. For the application of the CAA method to a unconfined turbulent flame, an extension of the method to reacting flow fields is presented. The LES method is applied to a turbulent flow over an airfoil with a deflected flap at a Reynolds number of Re = 106. The comparison of the numerical results with the experimental data shows a good agreement which shows that the main characteristics of the flow field are well resolved by the LES. However, it is also shown that a zonal LES which concentrates of the trailing edge region on a refined local mesh leads to a further improvement of the accuracy. In the second part of the paper, the CAA method with the extension to reacting flows is explained by an application to a non-premixed turbulent flame. The monopole nature of the combustion noise is clearly verified, which demonstrates the capability of the hybrid LES/CAA method for noise prediction in reacting flows.

Large-eddy simulations for tundish and airfoil flows

Large-eddy simulations (LES) of two complex flow problems, a continuous tundish flow and the flow around multi-element airfoils are presented. The numerical computations are performed by solving the conservation equations for compressible fluids. An implicit dual time stepping scheme combined with low Mach number preconditioning and a multigrid accelerating technique is developed for LES computations. The method is validated by comparing data of turbulent pipe flow at Re τ=1280 and cylinder flow at Re=3900 at different Mach numbers with experimental findings from the literature. Finally, the characteristics of the flow in a one-strand tundish is analyzed and a solution for a flow around a two-element airfoil as well as a zonal solution for the trailing edge region are discussed.

Numerical simulation of combustion noise using acoustic perturbation equations

Combustion noise of unconfined turbulent flames has been investigated using a hybrid CFD/CAA Method. A large-eddy simulation (LES) of a turbulent non-premixed flame is used to determine the source terms for the computational aeroacoustics (CAA) simulation. The governing CAA equations, namely the Acoustic Perturbation Equations (APE), have been extended to take into account noise generated by reacting flow effects. The right-hand side of the pressure-density relation within the APE system shows that the major source term, the heat release per unit volume, is encoded in the density fluctuation. Therefore the total time derivative of the density is used as source term to simulate combustion noise.

Large-Eddy Simulation of Attached Airfoil Flow

A Large-eddy simulation version of the FLOWer code is introduced to compute attached flow around a quasi three-dimensional airfoil at a Mach number Ma = 0.088, Reynolds number Re = 8 × 105, and an angle of attack of 3.3°. For the treatment of the subgrid-scale stresses the MILES approach is chosen. The visualization of instantaneous flow fields shows the typical flow features such as the streaky structures in the near-wall region to be well resolved and the overall agreement of the computational results with experimental data to be satisfactory.