The present work focuses on the experimental investigation and characterization of vortex-flow phenomena of a generic triple-delta-wing configuration, like the DLR-F23 aircraft, under various transonic speeds and angles of attack. This experimental study is conducted at the Transonic Wind Tunnel in Gottingen, which allows for the exploration of multiple transonic flow regimes and several angles of attack using two advanced flow visualization techniques, namely stereoscopic particle-image velocimetry and fast-response, unsteady pressure-sensitive paint. From these flow visualizations, it is found that three distinct vortex systems develop over the surface of the DLR-F23 wing, namely an inboard, midboard, and outboard vortex, which respectively develop from the model's forebody, strake, and main wing. These vortical structures are strongly dependent on the flow conditions, whether this concerns low-to-high transonic conditions and/or the wing's angle of attack. In particular at higher transonic flows, significant compressible effects, such as shock fronts, develop that noticeably affect the vortices' characteristics and behavior, including - but not limited to - vortex-vortex and vortex-shock interactions, their breakdown and/or merging. Upon these observations, future work may be directed towards further exploring these complex vortex-dominated flows under trans- and even supersonic flow conditions, whether this being done through dedicated experimental efforts and/or by employing advances numerical means. All this would allow developing more efficient and high performant triple-delta-wing aircraft configurations that could operate robustly across a wide range of flight conditions (i.e., Mach numbers, pitch angles).

We measured and characterized the flow around a full-span, generic passenger aircraft model under buffet conditions at Mach numbers up to M = 0.9 and Reynolds numbers up to Re = 25 x 10(6) by adapting and applying the cryogenic Particle Image Velocimetry (cryoPIV) technique in the European Transonic Windtunnel (ETW). Three different scenarios, involving buffet on the main wing, interaction of the buffet-wake with the horizontal tail plane (HTP) and buffet-phenomena in the vicinity of the engine nacelle were addressed in three dedicated wind-tunnel entries incorporating a variety of fields of view and angles of attack between -4.5 degrees and +7 degrees. To enable these measurements and to improve performance, reliability and data quality under the harsh experimental conditions at ETW, we carried out a comprehensive extension and upgrade of our cryoPIV measurement system. As a result, three different stereo particle image velocimetry (PIV) systems, involving one state of the art time-resolved and two novel multi-pulse PIV systems, were installed at ETW and operated during the first two of three wind tunnel entries in order to measure buffet phenomena on the upper wing side and in the wing wake downstream to the HTP. In the third entry, a system of three scientific CMOS cameras was combined to two different stereo PIV systems in order to measure buffet phenomena at the lower wing in the vicinity of the nacelles. The captured flow fields reveal the large- and small-scale flow structures involved in the highly three-dimensional flow around the aircraft model. Beside position, shape and dynamics of the shock waves, the turbulent wake of the wing as well as its interaction with the horizontal tail plane is disclosed by the instantaneous, mean and fluctuating velocity fields. The partially overlapping measurement fields show reasonable agreement among different measurement positions and systems.

Hybrid RANS/LES simulations of the flow around the NASA Common Research Model aircraft configuration were carried out with the focus on understanding the interaction of the separated wake with the tailplane in the presence of massively separated flow on the main wing. Validation of the CFD data using PIV data obtained for the flow conditions at \(\alpha =16^{\circ }\), \(\alpha =18^{\circ }\) and \(\alpha =20^{\circ }\) was carried out, confirming the generally satisfactory performance of the DDES simulations observed in earlier publications. As a next step, the wake characteristics and tailplane forces were evaluated for three angles of attack in order to investigate the flow dynamics in low speed stall. The separation characteristics were found to vary over the span. The wake size and downwash direction varied significantly with higher values of \(\alpha \). The altered wing downwash influenced the tailplane inflow, with the load fluctuations on the latter being significantly affected by the amount of turbulent kinetic energy present in the wake.

The development of vortical flow strongly changes the properties of motion-induced aerodynamic loads. Tests were performed in the Transonic Wind Tunnel Gottingen with a half-wing model of a lambda wing. A hydraulic actuator imposed pitch oscillations with amplitudes and frequencies relevant for aeroelastic analyses. Unsteady surface-pressure and flowfield measurements were performed. With increasing angle of attack, a leading-edge vortex emerges and moves inboard. The unsteady loads show a strong dependency on the effects of the vortical flow. For the pitching cases, the effects are lagging the motion. In the rear outboard part of the wing, the unsteady pressures are small, but the amplitude of the motion is high. Hence, a significant amount of energy-force integrated over motion-is transferred from the fluid to the motion. For a free motion or deformation of the wing, this would cause a destabilization. A significant effect of the vortical flow is the discrepancy between the position where it emerges and is affected, and the position (e.g., further outboard) where the flow may generate unsteady, potentially destabilizing loads.

Within the framework of the EU project ESWIRP, the particle image velocimetry (PIV) using high-speed camera and laser has been used to measure the turbulent flow in the wake of a stalled aircraft wing. The measurements took place on the common research model provided by NASA in the pressurized cryogenic European Transonic Wind tunnel. A specific cryoPIV system has been used and adapted for using high-speed PIV components under the cryogenic conditions of the wind tunnel facility. First results are presented comprising transonic and subsonic stall conditions at realistic flight Reynolds numbers of 11.6 and 30 million, respectively.

This paper examines the application of formation flight to micro air vehicles with regard to possible power savings. Results of an experimental investigation on echelon formations using low-aspect-ratio (AR = 2) flat plate rectangular wings at low Reynolds number (Re = 35,000) are presented. One-, two-, and three-wing configurations are tested in a low-speed wind tunnel. To quantify the power savings by lift enhancement and drag reduction, the aerodynamic loads acting on each wing are measured using specific balances while the trailing wings of the formation are being traversed laterally and vertically in fine steps. In addition, the flowfields of the wing wakes are measured using particle image velocimetry. The force and flowfield measurements show that the optimal positions for lift enhancement appear at slightly spanwise overlapping between the leading and trailing wings.

Transient pitch and plunge maneuvers of low aspect ratio wings in a special low Reynolds number wind tunnel have been investigated by means of high speed Particle Image Velocimetry to analyze the leading edge vortex dynamics quantitatively. A specific requirement of the measurement was to design and apply an optical arrangement for the laser light sheet which moves with the wing so as to keep the light sheet in fixed position on the upper side of the model. This became necessary because of the low pulse energy of high speed lasers and the large amplitudes of motion. This paper shows the principle and the implementation of this particular set-up. The evaluation procedure shows some specific steps concerning the preprocessing, for example, mask generation and disparity correction. The temporally-resolved data sets enable a tracking of vortices and a detailed analysis of the leading edge vortex dynamics in terms of circulation build-up and convection velocity.

Three-dimensional unsteady flow fields of a flapping, low-aspect-ratio wing have been investigated by means of highly resolved tomographic particle image velocimetry (Tomo-PIV) measurements and computational fluid dynamics (CFD). Furthermore, force measurements have been carried out. Tomo-PIV was applied to the flow above a flat plate wing during the downstroke. High spatial resolution and large volume thickness could be achieved by using sensitive sCMOS cameras and a traversing setup. The CFD calculations covered the complete period of motion. The analysis of the vortex-dominated flow fields provides a deeper understanding of vortex interaction and three-dimensionality of low Reynolds number (Re = 18; 000 and Re = 36; 000) flows. Two different Strouhal numbers (St = 0.06 and St = 0.13) are considered and their effects on the development of a leading edge and tip vortex are discussed. The PIV results show instantaneous flow fields after a leading edge separation that are dominated by small-scale turbulent vortex structures. The presented CFD approach is able to predict these vortices by using highly resolved meshes. Coarser grids compare well with the phase-averaged experimental flow fields, which feature multiple large-scale leading edge vortices developing during the downstroke. Turbulent effects decrease for the lower Reynolds number. Force and moment hystereses as well as large-scale leading edge vortice circulation, calculated from the PIV results, increase with increasing Strouhal number. Vortex breakdown of the wing tip vortex can be observed during the downstroke in the experimental data.

Stereo PIV measurements have been performed in multiple planes of the flow above the wing of a high-lift aircraft configuration in order to investigate the performance of strake vortex generators attached to the nacelle with regard to the evolution of flow separations on the main wing at high angles of attack. The measurements have been carried out in the cryogenic pressurized transonic wind tunnel facility of ETW that allows to perform the measurements at real flight Mach and Reynolds numbers by using total gas temperatures and pressures down to 110 K and up to 450 kPa, respectively. Because of this extreme wind tunnel conditions a specific cryo-PIV measurement system has been used. To provide insights into the streamwise development of the vortical flow a light-sheet, oriented perpendicular to the direction of free stream velocity, has been positioned to different chord stations of the wing. This paper focuses on a description of the employed PIV setup in the ETW showing first results for a standard configuration obtained at flight Reynolds numbers.

This paper presents results of high-resolution three-dimensional wing shape measurements performed on free-flying barn owls in flapping flight. The applied measurement technique is introduced together with a moving camera set-up, allowing for an investigation of the free flapping flight of birds with high spatial and temporal resolution. Based on the three-dimensional surface data, a methodology for parameterizing the wing profile along with wing kinematics during flapping flight has been developed. This allowed a description of the spanwise varying kinematics and aerodynamic parameters (e.g. effective angles of attack, camber, thickness) of the wing in dependence on the flapping phase. The results are discussed in detail using the data of a single flight, whereas a comparison of some kinematic parameters obtained from different flights is given too.

The underlying physics of the flight of birds is still not fully understood, in part due to the very complex wing movement and a lack of appropriate measurement techniques. Modern measurement technologies for fluid flows are often not suited for experiments with living animals, since these techniques may potentially harm the animal (e.g. lasers). The objective of this project was to modify the established technology Particle Image Velocimetry (PIV) to measure the flow around trained barn owls with the least possible hazard to the health of the animals. Furthermore, the shape of the wings was measured simultaneously using Projected Pattern Correlation Technique (PROPAC) to correlate the flow with the shape and movement of the wings. Results of the measurements of gliding and flapping flight are presented in this chapter.

Aircraft wake vortex evolution in ground proximity is investigated experimentally in a water towing tank, as well as numerically with wall-resolved large eddy simulation (LES). With these complementary instruments the enhancement of wake vortex decay by obstacles, introduced at the ground surface, is analyzed. The experimental methods include time-resolved stereo particle image velocimetry and vortex core visualization. For comparison with the experiment, the LES considers the turbulent wake of the strut, holding the towed aircraft model. Wake vortex trajectories and circulation decay are compared at different distances from the obstacle. Tracers are employed to visualize the obstacle’s effects on the vortex core, in LES and experiment. The experimentally obtained trajectories and decay characteristics are reproduced qualitatively by simulations, whereas the agreement is degraded at later times. Beyond that, the vortex dynamics, deduced from the LES results, help to understand the experimental observations. The obstacles trigger helical secondary vortex structures, propagating along the primary vortices. The observed propagation speed of the helical disturbance is fairly well predicted by the suggested simple model. It is shown that the obstacles significantly modify the vortex interaction with the ground and substantially accelerate vortex decay. Two neighboring obstacles lead to colliding disturbances that further enhance vortex decay rates.

Three-dimensional, unsteady flow fields of a flapping low aspect ratio wing are investigated by means of tomographic PIV (Tomo-PIV) measurements and computational fluid dynamics (CFD). Furthermore force measurements have been done. Tomo-PIV was applied to the flow above the flat-plate-wing during the down-stroke. A high spatial resolution and a large volume thickness could be achieved by using sensitive sCMOS cameras and a traversable set-up. The CFD calculations cover the complete period of motion. Analyzing the vortex dominated flow fields provides a deeper understanding of vortex interaction and three-dimensionality of low Reynolds number (Re = 18,000 and Re = 36,000) flows. Two different Strouhal numbers (St = 0.06 and St = 0.13) are considered and their effects on the development of a leading edge and tip vortex are discussed.

The wake flow of a high-lift aircraft model has been measured in the cryogenic transonic wind tunnel ETW using PIV. The wind tunnel, providing a 2.4 m wide and 2 m high test section, allows for measurements at the design point of the landing configuration, i.e. at a Reynolds number of 16.7 million and a Mach number of 0.2. To capture the wake flow along the complete wing span of the half model (half span width is ca. 1.5 m), a specific for the ETW developed cryo PIV system has been extended to use two Stereo-PIV setups in parallel. In the paper the employed cryo PIV setup is described with emphasis on the seeding technique using ice particles, the optical modules for an operation under cryogenic conditions and the automatic data acquisition system. Finally, the obtained results of the vortical wing wake flow are discussed for different angles of attack.

Nowadays, advanced optical measurement methods for wind tunnel or laboratory applications become more and more sophisticated and relatively easy to employ, even in complex and severe experimental environments. Hence, the European research project AIM (Advanced In-flight Measurement Techniques) intended to transfer the present experiences of optical deformation and flow visualisation tools to flight tests in order to measure at real aircraft and flight Reynolds numbers. Within three flight tests conducted in September 2009, PIV (Particle Image Velocimetry) recordings have been acquired with the help of a commonly used and well approved PIV system. A highly turbulent and intermittent flow field was recorded which was mainly influenced by the propeller slipstream, the fuselage boundary layer and the flap downwash. The outcome of this campaign proved the applicability of this optical flow measurement tool for flight testing. In this regard, the main advantages of the PIV technique are its non-intrusive nature as well as the capability of providing velocity information with high spatial resolution in a flow field. This ability ensures the determination of valuable instantaneous velocity vector field data during flight tests, whereas on the other side external probes or intrusive sensors could only obtain discrete single point values of the desired parameters. Nevertheless, a number of significant challenges for the application of in-flight PIV to future flight tests still remain related to the aircraft operation and certification. This paper presents and discusses the preparation, realisation and results of the in-flight PIV measurement campaign as well as future developments and improvements which have to be pursuit in order to make PIV a cost efficient flight test technique.

Stereo PIV measurements have been applied to the wake flow on a high-lift aircraft configuration in the cryogenic pressurized transonic wind tunnel ETW at realistic Mach and Reynolds numbers. A cryo PIV system, developed specially for the ETW to enable flow field measurements at gas temperatures down to 110 K and gas pressures of up to 450 kPa, has been adapted to a large measurement field extending over the complete span width of the half wing model. The specific seeding technique using ice particles and the optical systems are described as well as the automation the PIV measurements linked with the wind tunnel control system. First results of the vortical wing wake flow obtained at a Reynolds number of 17 million will be presented.