Dirk Otter

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.

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.

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.

Flow Field Investigations on a UCAV with a Ducted Engine Air Intake at Transonic Conditions

Stereoscopic Particle Image Velocimetry (SPIV) measurements have been performed on an unmanned combat air vehicle (UCAV) wind tunnel model (DLR F17E) in the DNW-TWG transonic wind tunnel. The DLR F17E model is equipped with ducted engine air intake. The model has a leading edge sweep angle of 53 deg. At various streamwise locations SPIV results have been obtained for the investigation of the flow field phenomena induced by the ducted engine air intake. In addition, flow features in the exhaust region are also obtained. The SPIV measurements are part of a project dedicated to the validation of computational codes.