Experimental studies on the “Phantom Yaw Effect” at maneuvering slender bodies

Asymmetric vortices can occur unexpectedly on slender bodies at high angles of attack. These vortices separating from the nose or/and shoulder region induce a side force and also a corresponding yawing moment often referred to as "phantom yaw". In the last decades, there have been many experimental and also numerical studies on this phenomenon. The aim was to understand this effect and to find the in fluencing parameters. There have also been investigations on using the asymmetric vortices for control purposes in addition to the fins. Despite this, another target of research has been the suppression of the vortex inducing side forces and yawing moments in order to increase the stability of e.g. a missile in a maneuver. Most of the wind tunnel tests have been done without model motion at several but fixed angles of attack. Since slender bodies as missiles achieve these high angles of attack via very rapid pitching maneuvers, the model motion is supposed to have some impact on the test results. One reason for the lack of dynamic test data at high Reynolds numbers are high inertial and aerodynamical forces acting on the test model and it's support. They result in contradicting design issues. On the one hand, the support needs to be stiff to withstand all forces and moments and on the other hand, the aerodynamic behaviour of the model shall not be changed by the support. Nonetheless, a maneuver simulator has been built at the DLR Goettingen. By means of this device, wind tunnel tests in a transonic wind tunnel at high Mach and Reynolds numbers, pitching rates of up to omega = 700°/s and pitching maneuvers from alpha = 0 ... 45° have been done. We compared the "phantom yaw"' emergence at a clean configuration with the ones at a configuration housing a pair of symmetric longitudinal slot nozzles which were fed by natural ventilation. The results showed a yawing moment for the clean configuration at angles of attack higher than alpha = 38°. They also showed that the jet flow through the slot nozzles successfully suppressed the yawing moment by causing a fixed separation. Differences between static and dynamic tests could be seen as well

Wall-Normal Focused Laser Differential Interferometry

A new approach for measuring boundary-layer disturbances with focused laser differential interferometry (FLDI) for planar models is presented. By integrating a glass window into a flat plate, the optical axis was aligned normal to the model surface, and the focal plane was set inside the boundary layer. By determining the extent of the sensitive volume along the optical axis and calculating the analytical transfer function of the setup, the implications of the FLDI properties on the measured data are analyzed. Measurements performed on a flat plate at Mach 6 are used to demonstrate the effects of the laminar–turbulent transition on the spectral distribution of the power density and to explicitly verify the detectability of the expected second-mode instabilities. Advantages and disadvantages of the proposed setup compared to the conventional one are discussed

Interaction of a moving shock wave with a turbulent boundary layer

In the present study, the influence of a uniformly moving impinging shock on the resulting shock wave–turbulent boundary layer interaction is numerically investigated. The relative Mach number of the shock front travelling above a flat-plate model varied between 0 and 2.3, while the quasi-steady inflow conditions remained constant with a Mach number of 3. To quantitatively evaluate the effect of shock travelling speed, a well-known scaling method for interaction length in quasi-steady flows was applied as a reference after significant improvements in modelling the effects of Reynolds number and wall temperature using new and existing data. Moreover, previously obtained experimental results for a limited range of travelling speeds were employed to validate the obtained numerical results. Three ranges of shock travelling speeds with distinctly different properties were extracted and quantitatively described using a developed correlation-based approach built on the extended quasi-stationary scaling law. In the first range, the scaled interaction length reaches its maximum for the given interaction strength and can be directly described by the scaling law obtained for quasi-stationary interactions. In the second travelling-speed range, the dependence of the interaction length on the interaction strength is explicitly influenced by the shock movement. With increasing shock travelling speed, the scaled interaction length here decreases significantly faster than in the quasi-stationary reference case. The end of this speed range is reached when the absolute shock front speed has caught up with the maximum speed of sound in the interaction zone, and thus the interaction length has fallen to zero. This travelling-speed limit signifies the transition to the third range, where upstream influence is no longer possible

Experimental analysis of the log-law at adverse pressure gradient

The experimental data for the mean velocity are analysed in the inner layer for a turbulent boundary layer at significant adverse pressure gradient and Reynolds numbers up to Re-theta=57000. The aim is to determine the resilience of the log law for the mean velocity, the possible change of the von Karman constant kappa and the appearance of a square-root law above the log law at significant adverse pressure gradients. In the wind-tunnel experiment, the adverse pressure gradient is imposed by an S-shaped deflection of the contour model which is mounted on a wind-tunnel sidewall. A large-scale particle imaging velocimetry method is applied to measure the streamwise evolution of the flow over a streamwise distance of 15 boundary layer thicknesses. In the adverse pressure gradient region, microscopic and three-dimensional Lagrangian particle tracking velocimetry are used to measure the mean velocity and the Reynolds stresses down to the viscous sublayer. Oil-film interferometry is used to determine the wall shear stress. The log law in the mean-velocity profile is found to be a robust feature at adverse pressure gradient, but its region is thinner than its zero pressure gradient counterpart, and its slope is altered. A square-root law emerges above the log law, extending to the wall distance the log law typically occupies at zero pressure gradient. Lower values for kappa are found than for zero pressure gradient turbulent boundary layers, but the reduction is within the uncertainty of the measurement

Combined free-stream disturbance measurements and receptivity studies in hypersonic wind tunnels by means of a slender wedge probe and direct numerical simulation

Combined free-stream disturbance measurements and receptivity studies in hypersonic wind tunnels were conducted by means of a slender wedge probe and direct numerical simulation. The study comprises comparative tunnel noise measurements at Mach 3, 6 and 7.4 in two Ludwieg tube facilities and a shock tunnel. Surface pressure fluctuations were measured over a wide range of frequencies and test conditions including harsh test environments not accessible to measurement techniques such as Pitot probes and hot-wire anemometry. A good agreement was found between normalized Pitot pressure fluctuations converted into normalized static pressure fluctuations and the wedge probe readings. Quantitative results of the tunnel noise are provided in frequency ranges relevant for hypersonic boundary-layer transition. Complementary numerical simulations of the leading-edge receptivity to fast and slow acoustic waves were performed for the applied wedge probe at conditions corresponding to the experimental free-stream conditions. The receptivity to fast acoustic waves was found to be characterized by an early amplification of the induced fast mode. For slow acoustic waves an initial decay was found close to the leading edge. At all Mach numbers, and for all considered frequencies, the leading-edge receptivity to fast acoustic waves was found to be higher than the receptivity to slow acoustic waves. Further, the effect of inclination angles of the acoustic wave with respect to the flow direction was investigated. An inclination angle was found to increase the response on the wave-facing surface of the probe and decrease the response on the opposite surface for fast acoustic waves. A frequency-dependent response was found for slow acoustic waves. The combined numerical and experimental approach in the present study confirmed the previous suggestion that the slow acoustic wave is the dominant acoustic mode in noisy hypersonic wind tunnels.The present study was supported by an ESA funded Technology Research Project (ESA-Contract number $4200022793/09/NL/CP$).Published versio

Modification of the SSG/LRR-omega RSM for adverse pressure gradients using turbulent boundary layer experiments at high Re

A modification of the SSG/LRR-omega model for turbulent boundary layers in adverse pressure gradient is presented. The modification is based on a new wall law for the mean velocity at adverse pressure gradient. The wall law is found from two new joint DLR/UniBw experiments and from the analysis of a data base from the literature. The mean velocity profile in the inner layer is found to consist of a log-law region, which is thinner than its zero pressure gradient counterpart, and a half-power law region above the log law. An empirical correlation for the wall-distance of the transition from the log-law to the half-power law is presented. Then a modification of the omega-equation to account for a half-power law behaviour of the mean velocity is described. The modified SSG/LRR-omega model is then applied to the two joint DLR/UniBw experiments. The modification leads to a reduction of the mean velocity in the inner part of the boundary layer and makes the model more susceptible for flow separation, which is in good agreement with the experimental data

Modification of turbulence models for pressure-induced separation on smooth surfaces using the DLR VicToria experiment

A new experiment of a turbulent boundary layer flow at a large adverse pressure gradient at a high Reynolds number is presented. The strong pressure gradient leads to pressure-induced separation on the smooth surface of the geometry model with a thin separation bubble. The experiment was performed within the DLR internal project VicToria. First, the design of the test case, the set-up in the wind tunnel, and the measurement technique using both large-scale and high-magnification particle imaging and Lagrangian particle tracking are described. Then the experimental results for the mean velocity are described as the flow evolves downstream from the zero-pressure gradient region into the adverse pressure gradient region. From the measurement data a wall law for the mean velocity with a thin log-law region and a half-power law region above the log-law is observed in the adverse pressure gradient region. Then the differential Reynolds stress transport model SSG/LRR-omega is considered. Based on the observation that the length-scale equation is not consistent with the assumed wall laws at adverse-pressure gradient, a modification of the equation for the dissipation rate omega in the model is proposed, so that the modified model can predict the observed wall law at adverse-pressure gradient. Finally, the numerical results using the modified SSG/LRR-omega model are shown. The modification causes a reduction of the mean velocity in the inner part of the boundary layer at adverse-pressure gradients, making the modified model more susceptible for flow separation. The numerical predictions of the modified model are found to be in good agreement with the experimental data

Simplified model for flow-heating effect on wave drag and its validation

A simplified prediction method for determination of flow-heating effects on the wave drag of bodies based on the combined analytical–empirical model of the thermal-spike phenomenon is presented. The existing model conceptions addressing certain aspects of this phenomenon were complemented and refined to develop a method suitable for parameter studies. Some reliable experimental and numerical results for the Mach 3 supersonic flow over conically nosed bodies were used as training data to estimate empirically the model parameters. Finally, the method was cross validated by the different available results for blunt bodies with hemispherical, flat, and conical front faces. By this opportunity, the ability to predict the influence of some crucial parameters has successfully been demonstrated for the heat input ratio at steady and periodic heating, the normalized heated-wake/filament diameter, the Mach number, and the specific heat capacity. Read More: http://arc.aiaa.org/doi/abs/10.2514/1.J05460

About the Assessment of Heat Flux and Skin Friction of the DLR TAU-code for Turbulent Supersonic Flows

The correct prediction of skin friction and surface heat flux by CFD tools is an important prerequisite for the design of hypersonic flight vehicles. Results of a test campaign of a flow past a flat plate configuration with an impinging oblique shock carried out in the Ludwieg Tube Facility of DLR Göttingen were used to assess the applicability and accuracy of the DLR TAU-code at respective flow conditions. A shock-wave/turbulent boundary layer interaction (SWTBLI) on the flat plate model was created with a planar shock generator at different incidence angles. This two dimensional test case involving the flat plate flow and the interaction of an oblique shock with the turbulent boundary layer is well documented with various measurements like wall pressure, boundary layer velocity profiles, skin friction, and heat flux, and offers a good basis for the assessment of the two latter variables for the DLR TAU-code. The assessment includes grid studies, the influence of various turbulence models as well as the external turbulence intensity set in the calculations