Robert Konrath

Static and Dynamic SACCON PIV Tests, Part II: Aft Flow Field

The flow above delta wing shaped UCAV (Unmanned Combat Air Vehicle) configurations is often dominated by strong vortices, especially at moderate and high angle of attack. On delta wings with sharp leading edges, the flow separates at the leading edges already at small angles of attack. The shear layer rolls up and two vortices are formed on each side of the wing starting close to the wing apex and increasing in size towards the wing aft. The vortices produce strong suction peaks on the upper wing surface, so that additional lift is generated which is of benefit by performing low speed manoeuvres like take-off and landing. UCAV configurations often consist of rounded leading edges to improve their performance under cruise conditions. In this case the flow topology becomes more complicated because the flow separations at the leading edges are delayed to higher angles of attack and depend strongly on the Reynolds and Mach number. In regions of attached flow around blunt leading edges, other flow effects may produce additional vortices as observed on the VFE2 delta wing where vortices develop inboard the leading edge vortices caused by a kind of instability within the boundary layer. The onset of the leading edge vortices, their interactions and accompanying phenomena like vortex breakdown can cause adverse loading effects. To avoid such non-linear aerodynamic characteristics detailed flow investigations are of great importance during the design chain by improving the understanding of the aerodynamics. This all the more so if the dynamic behaviour of the vehicle has to be considered as the flow effects become even further complicated. Within the scope of the Task Group of AVT-161 (Chapter 1), Stereoscopic Particle Image Velocimetry (PIV) has been applied for flow field investigations on a SACCON (Stability And Control CONfiguration) wind tunnel model.

Analysis of PIV Results Obtained for the VFE-2 65° Delta Wing Configuration at Sub- and Transonic Speeds

The objectives of the International Vortex Flow Experiment 2 (VFE-2) are to perform new wind-tunnel tests on a delta wing by using modern measurement techniques and to compare these data with results of numerical state-of-the-art codes. In the present paper results of the first part of the VFE-2 experiments carried out at DLR Gottingen on a NASA wind tunnel model are described. These investigations comprise pressure distribution measurements by means of the Pressure Sensitive Paint (PSP) technique which captures the complete model surface giving much more insights in details of the flow topology than would be possible by discrete pressure taps. The delta wing of 65° sweep angle was equipped with sharp as well as with rounded leading edges. The pressure distributions are analyzed for two Mach numbers M = 0.4 and 0.8 and two Reynolds numbers 2 and 3 million. Dependent on the angle of attack (10o -25o) the aerodynamics are described with emphasis to secondary vortex formation and vortex breakdown. For the rounded leading edge the aerodynamics are also described with regard to attached flow, onset of outer primary vortex and particularly the formation of an inner primary vortex.

Going for Experimental and Numerical Unsteady Wake Analyses Combined with Wall Interference Assessment by Using the NASA CRM Model in ETW

Detailed experimental and accompanying numerical studies on the development of unsteady wakes past an aircraft under stall conditions are currently prepared by a consortium of research institutions and universities. The experiments will be performed in the ETW cryogenic wind tunnel on the NASA Common Research Model. Besides wake surveys using time-resolved cryo PIV measuring technique, wall interference measurements are planned. The tests scheduled for July 2013 are funded by the European Commission in the 7th framework program. In this paper results of preparatory CFD studies and wake analyses of the CRM model, the TR-PIV measuring technique and the ETW facility are presented along with the wind tunnel model and the planned test program.

Advanced Measurement Techniques for High Reynolds Number Testing in Cryogenic Wind Tunnels

The present paper addresses the development, qualification trials and application of some non-intrusive measurement techniques suitable for operation in industry-scale, pressurised cryogenic wind tunnels. The application of cryogenic Temperature-Sensitive Paint (cryoTSP) as a tool for transition detection is described as well as the implementation of the Image Pattern Correlation Technique (IPCT) and the Backward Oriented Schlieren method (BOS) in the European Transonic Windtunnel (ETW). Progress on the development of cryogenic Pressure-Sensitive Paint (cryoPSP) is shown, and considerations for the establishment of a Particle Image Velocimetry system suited for low temperatures (cryoPIV) are presented. Furthermore, the state of adaptation of the microphone array technique

Application of Stereo PIV to the VFE-2 65° Delta Wing Configuration at Sub- and Transonic Speeds

The present paper describes the application of Particle Image Velocimetry (PIV) to the flow around a delta wing in a pressurized transonic wind tunnel. These investigations are the second part of International Vortex Flow Experiment 2 (VFE-2) measurements carried out at DLR Gottingen on a NASA wind tunnel model. The first part comprised surface pressure measurements by means of Pressure Sensitive Paint. These results were used to select the test cases for PIV. Flow fields for the Mach numbers of 0.4 and 0.8 and for the Reynolds numbers of 2 and 3 million at four incidences of 11, 13, 20 and 25 degree were captured within different planes perpendicular to the model axis. Delta wings with sharp as well as with rounded leading edges were investigated. Details of the stereoscopic PIV system such as the overall arrangement, the image evaluation and the flow seeding technique will be described. Techniques to overcome problems caused by density changes within the wind tunnel are described too, such as deflections of the laser light beam and camera viewing. Laser light flare on the model surface is reduced by a specially developed fluorescent paint which allows for the detection of small flow structures very close to the model surface.

High-Speed PIV Applied to Wake of NASA CRM Model in ETW Under High Re-Number Stall Conditions for Sub- and Transonic Speeds

Particle Image Velocimetry (PIV) using high-speed camera and laser has been applied to turbulent flow regions in the wake of a stalled aircraft wing. The measurements took place on the Common Research Model (CRM) of NASA in the pressurized cryogenic European Transsonic Wind tunnel (ETW) within the framework of the EU project ESWIRP. The employed high-speed PIV setup is described and preliminary results are presented comprising sub- (M = 0.17) and transsonic (M = 0.85) stall conditions at flight Reynolds numbers of 7 and 30 million, respectively.

Accepting a Challenge -The Development of PIV for Application in pressurized cryogenic Wind Tunnels

The development and application of a Stereo-PIV system in the cryogenic transonic wind tunnel ETW enabling detailed flow field analyses on scaled models of transport aircraft at flight Mach- and Reynolds numbers is described. The specific conditions in such wind tunnels as cryogenic gas temperatures down to 110 K and moderate gas pressures of up to 450kPa as well as a large plenum encapsulating the test section made specific developments necessary. This comprises the generation of suitable flow tracers for cryogenic flows considering the specific requirements for an application in the ETW as well as the supply of laser light of high pulse energy and optical components to be placed within the cryogenic environment. Optical effects due to gas density changes within the wind tunnel plenum and light beam deflections or shifts are considered in the optical design of the CRYO-PIV system. Following trials in the cryogenic research tunnel PETW the applicability of the developed system is successfully demonstrated in the ETW during two test campaigns for a full scale and a half- model of a transport aircraft. The employed Stereo-PIV system and first results of the measurements are discussed. Scheduled next steps for improvements based on the gained experiences are outlined.

Analysis of Flow Field Measurements Obtained in a Large Tow Tank Regarding the Decay of Wake Vortices in the Far- Field for Two- and Four-Vortex Systems

Stereo PIV measurements on a 2- and 4-vortex system have been performed in a large tow tank of SVA in Potsdam. The considered 4-vortex system consists of two counter-rotating vortex pairs generated by wings of a span width ratio of b 2 /b 1 = 0.3 and a circulation ratio of Γ 2 /Γ 1 = -0.3. As reference the corresponding 2-vortex system is also considered. A submersible Stereo PIV system attached to a vertical traversing system was employed. Cross-plane velocity fields are obtained of the descending wake vortex system to access the vortex dynamics well into the wake far field (x/b > 500, τ > 11). The spatio-temporal evolutions of the vortex systems are analyzed with respect to the vortex trajectories, circulations and core radii as well as the total kinetic energy contained in the measured flow field.

Flow Field Measurements by PIV at High Reynolds Numbers

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.

Tracking the Nacelle Vortex Above Aircraft Wing in the ETW at Real Mach- and Reynolds Numbers by Means of PIV

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.