Wake flows have been extensively investigated in the past mainly with the objective to reduce drag. The failure of Ariane 5 Flight 157 though revealed that the understanding has not been sufficient, which triggered a new series of base-flow investigations. Most of them scrutinized the dynamics of the flow at one specific point along the flight path either in the subsonic, transonic, or supersonic range. However, no coherent image of the flow topology development in terms of a parametric study is available so far. The current work addresses the question of a near-wake scaling of a generic space launcher configuration. This question is tackled experimentally by means of particle image velocimetry measurements in the vertical test section Cologne in the range of Mach 0.5, 0.6, 0.7, 0.8, and 0.9 for Reynolds numbers between Re D = 0.8 × 10 6 and 1.7 × 10 6. Results show the mean, the turbulence intensity, and the Reynolds shear stress distribution in the wake in comparison with the literature. In addition, data to the incoming boundary layer are provided and to the evolving shear layer. In that regime, the results indicate that the velocity and Reynolds stresses might be independent from the Mach and Reynolds number if scaled with the reattachment length. If this hypothesis holds true, it might be useful for the validation of numerical codes and for the initial determination of the flow field of space launchers in the design phase.

The Ariane 5 failure flight 157 made clear that the loads in the base region of space launcher configurations were underestimated and its near-wake dynamics required more attention. In the recent years, many studies have been published on buffet/ buffeting in the critical high subsonic flow regime. Nevertheless, not much experimental data are available on the interaction of the ambient flow with an exhaust jet over a wide subsonic Mach number range. Further, a preceding study without exhaust jet revealed questions regarding a similar distribution of the velocity and Reynolds stress in the near-wake if scaled with the reattachment length. Consequently, a generic space launcher configuration featuring a cold, supersonic, over-expanded jet is investigated experimentally in the vertical test section Cologne (VMK) by means of particle image velocimetry (PIV) for five subsonic Mach numbers ranging from 0.5 to 0.9 with corresponding Reynolds numbers between Re D = 0.8 × 10 6 to 1.6 × 10 6. The velocity and Reynolds stress distribution are provided for the near-wake flow and additionally for the incoming boundary layer. Just as in the preceding study, self-similar features are found in the flow field as long as the separated shear layer reattaches on the solid nozzle wall. Substantial changes are then measured for an alternating (hybrid) reattachment between the solid nozzle wall and supersonic exhaust jet as found for Mach 0.8, one of them being the increased axial turbulence in the recirculation bubble due to a 'dancing' large-scale, clockwise-rotating vortex.

The turbulent wake §ow of generic rocket con¦gurations is investigated experimentally and numerically at a freestream Mach number of 6.0 and a unit Reynolds number of 10 · 10 6 m −1 . The §ow condition is based on the trajectory of Ariane V-like launcher at an altitude of 50 km, which is used as the baseline to address the overarching tasks of wake §ows in the hypersonic regime like §uid-structural coupling, reverse hot jets and base heating. Experimental results using pressure transducers and the high-speed Schlieren measurement technique are shown to gain insight into the local pressure §uctuations on the base and the oscillations of the recompression shock. This experimental con¦guration features a wedgepro¦led strut orthogonally mounted to the main body. Additionally, the in §uence of cylindrical dummy nozzles attached to the base of the rocket is investigated, which is the link to the numerical investigations. Here, the axisymmetric model possesses a cylindrical sting support of the same diameter as the dummy nozzles. The sting support allows investigations for an undisturbed wake §ow. A time-accurate zonal Reynolds-Averaged NavierStokes / Large Eddy Simulation (RANS/LES) approach is applied to identify shocks, expansion waves, and the highly unsteady recompression region numerically. Subsequently, experimental and numerical results in the strut-averted region are compared with regard to the wall pressure and recompression shock frequency spectra. For the compared con¦gurations, experimental pressure spectra exhibit dominant Strouhal numbers at about Sr D = 0.03 and 0.27, and the recompression shock oscillates at 0.2. In general, the pressure and recompression shock §uctuations numerically calculated agree reasonably with the experimental results. The experiments with a blunt base reveal base-pressure spectra with dominant Strouhal numbers at 0.08 at the center position and 0. 145, 0.210.22, and 0.310.33 at the outskirts of the base.

The near-wake of a separated flow behind a generic space launcher model with different afterbody extensions is studied experimentally in the hypersonic flow regime (Mach number 6.0 and unit Reynolds number 16 × 10 6 m −1) with the main objective being to show unsteady flow phenomena. This objective is based on the desire to control the base flow of space launchers to reduce drag and dangerous vibrations. The investigations focus on the unsteadiness of the base-pressure fluctuations and the oscillations of the recompression shock. Base-pressure spectra and spectra of the vertical recompression shock oscillation are discussed for various pressure transducer locations, various nozzle lengths and nozzle diameters. The base-pressure spectra exhibits dominant Strouhal numbers centered at 0.08 at the center positions and 0.08, 0.2, 0.38 at the outskirts of the base, and shows a crucial change with a frequency shift to Strouhal number of 0.27 when a nozzle is applied. Additionally, the pressure spectra indicate an azimuthal mode. The spectra of the recompression shock feature a dominant Strouhal number of 0.2 independent of the geometry of the afterbody. Since the incoming boundary layer is suspected to have an influence on the near-wake flow, the state and the flow conditions are examined in introductory experiments by means of infrared images of the model surface and Pitot pressure at the body corner.

Rocket wake flows were under investigation within the Collaborative Research Centre SFB/TRR40 since the year 2009. The current paper summarizes the work conducted during its third and final funding period from 2017 to 2020. During that phase, focus was laid on establishing a new test environment at the German Aerospace Center (DLR) Cologne in order to improve the similarity of experimental rocket wake flow–jet interaction testing by utilizing hydrogen–oxygen combustion implemented into the wind tunnel model. The new facility was characterized during tests with the rocket combustor model HOC1 in static environment. The tests were conducted under relevant operating conditions to demonstrate the design’s suitability. During the first wind tunnel tests, interaction of subsonic ambient flow at Mach 0.8 with a hot exhaust jet of approx. 920 K was compared to previously investigated cold plume interaction tests using pressurized air at ambient temperature. The comparison revealed significant differences in the dynamic response of the wake flow field on the different types of exhaust plume simulation.