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For aerodynamic profile tests on aircraft models, boundary-layer transition detection is of great interest. Under ambient flow conditions the infrared technique (IR) is a well-established image-based method for this purpose. In high Reynolds number tests that are conducted at cryogenic temperatures the IR technique is of only limited suitability. In contrast, the image-based Temperature-Sensitive Paint (TSP) technique is well suited for these conditions. Boundary layer transition detection by means of TSP generally requires an artificial temperature difference between model surface and flow. For wind tunnels operated under cryogenic conditions changing the liquid nitrogen injection rate of the working fluid causing a rapid change of the flow temperature can generate this temperature difference (“temperature-step method”). The drawback of this temperature-step method is that during the change of the flow temperature neither the Reynolds nor the Mach number can be kept constant. To overcome this problem, we have recently published an alternative approach where Carbon Nanotubes (CNT) are used to electrically heat the model surface and thus generate a well-defined temperature difference to visualize laminar-turbulent transition. The combination of CNT and TSP, which we call cntTSP, delivered excellent results in different wind tunnel tests from ambient temperatures down to 100 K. The cntTSP sensor visualizes the surface boundary-layer conditions as temperature distribution based on different heat transfer coefficients in the laminar and turbulent flow. Compared to the temperature-step method the cntTSP approach has the advantage and potential to measure boundary-layer transition in a time-resolved manner, e.g. on wind tunnel models when the wind tunnel is operated in continuous pitch traverse mode, which is desired in order to improve the data productivity of the wind tunnel tests. In this paper, the applicability of cntTSP to a continuous moving model (pitch-traverse or pitch-sweep test) is investigated in the pilot facility of the European Transonic Windtunnel (PETW). The laminar to turbulent boundary-layer transition appearing on a 2D model is visualized and compared between the pitch-traverse and pitch-pause tests.

In this work, laminar-turbulent boundary-layer transition is investigated on the suction side of Mach-scaled helicopter rotor blades in climb. The phenomenon is assessed by means of temperature-sensitive paint (TSP). Results are compared to a data sample acquired by infrared (IR) thermography and accompanied by integral thrust- and local surface pressure measurements at two radial blade sections. Spatially, high-resolved data allow for precise detection of boundary-layer transition along the outer 60% of the blade span. Results obtained via TSP and IR show remarkable agreement with minor deviations due to different surface qualities of the respective blades tested. TSP data are obtained at various collective pitch angles and three different rotating speeds corresponding to chord Reynolds and Mach numbers based on blade tip speed of $$Re_{\rm tip} = 4.8 - 9.3\times 10^5$$ and $$M_{\rm tip} = 0.29 - 0.57$$ , respectively. The transition position is detected with an accuracy of better than 1% chord and the findings show overall coherence as blade loading and tip chord Reynolds number are varied. Experimental findings are shown to be consistent with two-dimensional simulations using the $$e^N$$ -envelope method for transition prediction. Based on quantitative agreement between measured and calculated surface pressures, a comparison of the corresponding transition results suggests a critical amplification factor of $$N_{\rm cr.} = 5.5$$ best suited for transition prediction in the rotating test facility of the DLR Gottingen.

This paper deals with the application of thermochemical Lagrangian MDF (mass density function) methods for compressible sub- and supersonic RANS (Reynolds Averaged NavierStokes) simulations. A new approach to treat molecular transport is presented. This technique on the one hand ensures numerical stability of the particle solver in laminar regions of the flow field (e.g. in the viscous sublayer) and on the other hand takes differential diffusion into account. It is shown in a detailed analysis, that the new method correctly predicts first and second-order moments on the basis of conventional modeling approaches. Moreover, a number of challenges for MDF particle methods in high speed flows is discussed, e.g. high cell aspect ratio grids close to solid walls, wall heat transfer, shock resolution, and problems from statistical noise which may cause artificial shock systems in supersonic flows. A Mach 2 supersonic mixing channel with multiple shock reflection and a model rocket combustor simulation demonstrate the eligibility of this technique to practical applications. Both test cases are simulated successfully for the first time with a hybrid finite-volume (FV)/Lagrangian particle solver (PS).

Tests were conducted at the Cryogenic Ludwieg-Tube Gottingen to investigate the influence on onset of boundary-layer transition of thermal gradients on the model surface, whereby the temperature-sensitive paint technique has been used to detect transition. The laminar airfoil under investigation was tested at high Reynolds and at both low and high subsonic Mach numbers. The influence on boundary-layer transition of the temperature difference between the flow and the model surface due to the temperature drop occurring during a test run at the Cryogenic Ludwieg-Tube Gottingen was examined. Various combinations of flow conditions and angles of attack were employed to investigate different stability situations. Linear-stability computations, based on the experimental data, were also carried out. The unfavorable impact of surface temperatures larger than the adiabatic-wall temperature was identified to be dependent on the considered boundary-layer-stability situation. For certain stability situations, it was s...

We report the first observation of the Mach cones excited by a larger microparticle (projectile) moving through a cloud of smaller microparticles (dust) in a complex plasma with neon as a buffer gas under microgravity conditions. A collective motion of the dust particles occurs as propagation of the contact discontinuity. The corresponding speed of sound was measured by a special method of the Mach cone visualization. The measurement results are incompatible with the theory of ion acoustic waves. The estimate for the pressure in a strongly coupled Coulomb system and a scaling law for the complex plasma make it possible to derive an evaluation for the speed of sound, which is in a reasonable agreement with the experiments in complex plasmas., 5 pages, 2 figures, 1 table

The near-wake flow of a generic space launcher geometry with jet is investigated experimentally in the hypersonic flow regime at Mach 6 and an unit Reynolds number of 16 · 106 by means of unsteady pressure measurements. This corresponds to an altitude of 50 km for an Ariane V-like launcher with an underexpanded nozzle flow. The investigations with a truncated ideal contoured wind tunnel model nozzle designed for Mach number of 2.65 is used to simulate the nozzle flow. The objective of this study is the characterization of the plume-induced loads on the main structural components in the near wake region during flight. A parametric study with the reservoir pressures of 11, 14, 17 and 20 bar is presented. The transducers capture the pressure signal in the radial and azimuthal direction on the base and on the surface of the nozzle. The signals are spectrally analyzed with the objective to characterize and identify unsteady flow effects in order to reduce base drag and vibrations of future space launchers. The base-pressure spectra exhibits global modes at Strouhal numbers of 0.05 to 0.06, 0.09 to 0.1 and 0.2 to 0.25, which are attributed in literature to flapping and shedding. Local modes seem to occur Locally restricted modes can be detected for the tonal disturbances in the high frequency range for transducers on the base closer to the nozzle. The increase of the reservoir pressure shows a consistent increase of the pressure fluctuation level in the base region.

The framework for a semi-empirical model for the prediction of aerodynamically generated trailing edge sound from subsonic low Mach number 2D airfoils with mildly separated turbulent boundary layers is presented. Aerodynamic and acoustic measurements were performed and results indicate that the separated trailing edge flow noise can be modeled in a similar fashion to trailing edge noise generated by attached turbulent boundary layers. Comparisons with aeroacoustic measurements show agreement between prediction and results.

The influence of streamwise pressure gradient and a non-adiabatic wing surface on boundary layer transition was experimentally investigated at the DNW-KRG blow-down wind tunnel facility in Gottingen, Germany. Boundary layer transition was detected non-intrusively by means of the Temperature-Sensitive Paint technique. A new wind tunnel model was designed with the aim of systematically investigating the influence of streamwise pressure gradient and a non-adiabatic wing surface, including Reynolds number effects and Mach number effects, on boundary layer transition. The model was tested at high Reynolds numbers and at a high subsonic Mach number. Favorable, neutral, and unfavorable streamwise pressure gradients were considered and various temperature differences between flow and model surface were implemented. More pronounced negative streamwise pressure gradients and surface temperatures closer to the adiabatic wall temperature were shown to stabilize the boundary layer and allowed larger transition Reynolds numbers to be achieved. The resulting effect of the coupling of streamwise pressure gradient and a non-adiabatic wing surface was found to be strongly dependent on the considered stability situation. The favorable effect on boundary layer transition of surface temperatures closer to the adiabatic wall temperature was shown to be more pronounced for stability situations characterized by a markedly negative pressure gradient.

A significant part of the cabin noise of airplanes comes from the engines. The noise field of the engines is considerable changed due to the fuselage’s boundary layer before it hits the fuselage skin. In this work we study the influence of the mean boundary layer as well as of the instationary turbulent fluctuating boundary layer on the sound field on the fuselage using a point noise source. The flow field is calculated with the Reynolds-averaged Navier-Stokes (RANS) equation and the turbulent fluctuations are reconstructed stochastically by means of a Random-Particle-Mesh method (RPM). The sound propagation is calculated with a computational aeroacoustic (CAA) approach solving the linearized Euler equations. At high Mach numbers we find a strong refraction effect due to the mean boundary layer and a substantial scattering effect of the turbulent vortices. The surface pressure completely differs from the free acoustic field: The frequency is increased, the pressure field is spatially incoherent and a considerable part of the original signal is reflected by the boundary layer. According to our knowledge, the combined effect of refraction and scattering of high Mach numbers boundary layers was not studied before. We think that the results are very important for the assessment of the structural response of the fuselage skin to the acoustic surface pressures.

This study deals with experimental and numerical investigations of the time response of a mantle thermocouple. Using two different experiments, a clear influence of a glue covering is indicated. The numerical analyses are done using a stand-alone structure and a coupled flow-thermal calculation. Both computational results agree with the experimental data and established a detailed flow investigation of the VxG-experiment. The experimental and numerical investigatons confirm that a thin glue layer of about 1 mm is responsible for a time delay of few seconds.

Zapfluft und Wellenleistung wird den Triebwerken entnommen, um die Energie für beispielsweise die Kraftstoffpumpen, das Inflight Entertainment oder die Flügelvorderkantenenteisung zu erzeugen. Diese Energiegenerierung, hat einen Anstieg des Kraftstoffverbrauches zur Folge. Es hat sich herausgestellt, dass die Stelle der Zapfluftentnahme einen starken Einfluss auf den Gradienten des Brennstoffverbrauches hat. Das Projekt beschäftigt sich mit zwei- und drei-welligen Turbofantriebwerken und untersucht an ihnen, die Effekte der Leistungsnahmen. Als Simulationssoftware wurde GasTurb 8.0 eingesetzt und auf die integrierten Triebwerkskonfigurationen zurückgegriffen. Ziel der Arbeit ist die Ermittlung einer mathematischen Beziehung zur Berechnung des zusätzlichen Kraftstoffmassenstromes infolge einer Zapfluft- oder Wellenleistungsentnahme. So stellt sich die Frage, welche Triebwerksparameter dafür berücksichtigt werden müssen. Eine Wellenleistungsentnahme verursacht beispielsweise einen linearen Anstieg des spezifischen Kraftstoffverbrauches. Ist diese Zunahme, identisch mit der einer Zapfluftentnahme? Am Ende der Kapitel werden die Ergebnisse mit Literaturwerten verglichen und versucht Tendenzen zu erkennen bzw. bestehende zu erhärten.