Laminar supersonic sphere wake unstable bifurcations

The laminar sphere unstable bifurcations are sought at a Mach number of M∞ = 1.2. Global stability performed on steady axisymmetric base flows determines the regular bifurcation critical Reynolds number at Rereg cr = 650, identifying a steady planar-symmetric mode to cause the loss of the wake axisymmetry. When global stability is performed on steady planar-symmetric base flows, a Hopf bifurcation is found at ReHopf cr = 875 and an oscillatory planar-symmetric mode is temporally amplified. Despite some differences due to highly compressible effects, the supersonic unstable bifurcations present remarkably similar characteristics to their incompressible counterparts, indicating a robust laminar wake behavior over a large range of flow speeds. A new bifurcation for steady planar-symmetric base flow solutions is found above Re > 1000, caused by an anti-symmetric mode consisting of a 90○ rotation of the dominant mode. To investigate this reflectional symmetry breaking bifurcation in the nonlinear framework, unsteady nonlinear calculations are carried out up to Re = 1300 and dynamic mode decomposition (DMD) based on the combination of input data low-dimensionalization and compressive sensing is used. While the DMD analysis confirms dominance and correspondence in terms of modal spatial distribution with respect to the global stability mode responsible for the Hopf bifurcation, no reflectional symmetry breaking DMD modes were found, asserting that the reflectional symmetry breaking instability is not observable in the nonlinear dynamics. The increased complexity of the wake dynamics at Re = 1300 can be instead explained by nonlinear interactions that suggest the low-frequency unsteadiness to be linked to the destabilization of the hairpin vortex shedding limit cycle.CNE

Wp-2 basic investigation of transition effect

An important goal of the TFAST project was to study the effect of the location of transition in relation to the shock wave on the separation size, shock structure and unsteadiness of the interaction area. Boundary layer tripping (by wire or roughness) and flow control devices (Vortex Generators and cold plasma) were used for boundary layer transition induction. As flow control devices were used here in the laminar boundary layer for the first time, their effectiveness in transition induction was an important outcome. It was intended to determine in what way the application of these techniques induces transition. These methods should have a significantly different effect on boundary layer receptivity, i.e. the transition location. Apart from an improved understanding of operation control methods, the main objective was to localize the transition as far downstream as possible while ensuring a turbulent character of interaction. The final objective, involving all the partners, was to build a physical model of transition control devices. Establishing of such model would simplify the numerical approach to flow cases using such devices. This undertaking has strong support from the industry, which wants to include these control devices in the design process. Unfortunately only one method of streamwise vortices was developed and investigated in the presented study

Active debris removal mission from LEO with electric and hybrid propulsion

Debris proliferation in space environment is an impending problem for preserving future missions. Although prevention measures (IADC guidelines, NASA Handbook for Limiting Orbit Debris) with the intent to regulate and limit the space trac are adopted, the necessity to reduce the increasing quantity of space debris leads to the will to conceive and test the effectiveness of active debris removal missions (ADR). ADR missions are devised for controlling the number of large objects, such as launch vehicle orbital stages or derelict spacecraft which no longer serve a useful purpose [1], orbiting in densely populated and commercially interesting areas (Sun-Synchronous Orbit, SSO)

Forced response of a laminar shock-induced separation bubble

The source of unsteadiness in shock-wave/boundary-layer interactions is currently disputed. This paper considers a two-dimensional separation bubble induced by an oblique shock wave interacting with a laminar boundary layer at a free-stream Mach number of 1.5. The global response of the separated region to white noise forcing is analyzed for different interaction strengths, which generate small and large separation bubbles. Forcing location and amplitude effects have been examined. For both interaction strengths and for forcing both upstream and inside the bubble, thewall-pressure spectra downstream of the separation show a high-frequency peak that is demonstrated to be a Kelvin-Helmholtz instability. A low-frequency response atthe separation point is also found when the separation bubble is only forced internally, therefore with a disturbance-free upstream boundary layer. For low-amplitude internal forcing, the low-frequency response at the separation point and downstream of the bubble is linear. However, when forced upstream the low-frequency unsteadinessof the large separation bubble is found to be driven by nonlinearities coming from the downstream shedding. The same nonlinear behavior is found when the separationbubble is internally forced over a narrowband around the shedding frequency, without low-frequency disturbances. This analysis for a laminar interaction is used to interpretthe low-frequency unsteadiness found at the foot of the shock of turbulent interactions. Here, the low-frequency unsteadiness occurs in the absence of upstream disturbancesand a linear relationship is found between the internal forcing and the response near the separation point. When low-frequencies are not present in the forcing they aregenerated from weak nonlinearities of the shear-layer instability modes

Wp-2 basic investigation of transition effect

An important goal of the TFAST project was to study the effect of the location of transition in relation to the shock wave on the separation size, shock structure and unsteadiness of the interaction area. Boundary layer tripping (by wire or roughness) and flow control devices (Vortex Generators and cold plasma) were used for boundary layer transition induction. As flow control devices were used here in the laminar boundary layer for the first time, their effectiveness in transition induction was an important outcome. It was intended to determine in what way the application of these techniques induces transition. These methods should have a significantly different effect on boundary layer receptivity, i.e. the transition location. Apart from an improved understanding of operation control methods, the main objective was to localize the transition as far downstream as possible while ensuring a turbulent character of interaction. The final objective, involving all the partners, was to build a physical model of transition control devices. Establishing of such model would simplify the numerical approach to flow cases using such devices. This undertaking has strong support from the industry, which wants to include these control devices in the design process. Unfortunately only one method of streamwise vortices was developed and investigated in the presented study

Wp-1 reference cases of laminar and turbulent interactions

In order to be able to judge the effectiveness of transition induction in WP-2, reference flow cases were planned in WP-1. There are two obvious reference cases—a fully laminar interaction and a fully turbulent interaction. Here it should be explained that the terms “laminar” and “turbulent” interaction refer to the boundary layer state at the beginning of interaction only. There are two basic configurations of shock wave boundary layer interaction and these are a part of the TFAST project. One is the normal shock wave, which typically appears at the transonic wing and on the turbine cascade. The characteristic incipient separation Mach number range is about M = 1.2 in the case of a laminar boundary layer and about M = 1.32 in the case of turbulent boundary layer. The second typical flow case is the oblique shock wave reflection. The most characteristic case in European research is connected to the 6th FP IP HISAC project concerning a supersonic business jet. The design speed of this airplane is M = 1.6. Therefore the TFAST consortium decided to use this Mach number as the basic case. Pressure disturbance at this Mach number is not very high and can be compared to the disturbance of the normal shock at the incipient separation Mach number mentioned earlier. As mentioned earlier, shock reflection at M = 1.6 may be related to incipient separation. Therefore two additional test cases were planned with different Mach numbers. ITAM conducted an M = 1.5 test case, and TUD an M = 1.7 test case. These partners have also previously made very specialized and successful contributions to the UFAST project

Wp-1 reference cases of laminar and turbulent interactions

In order to be able to judge the effectiveness of transition induction in WP-2, reference flow cases were planned in WP-1. There are two obvious reference cases—a fully laminar interaction and a fully turbulent interaction. Here it should be explained that the terms “laminar” and “turbulent” interaction refer to the boundary layer state at the beginning of interaction only. There are two basic configurations of shock wave boundary layer interaction and these are a part of the TFAST project. One is the normal shock wave, which typically appears at the transonic wing and on the turbine cascade. The characteristic incipient separation Mach number range is about M = 1.2 in the case of a laminar boundary layer and about M = 1.32 in the case of turbulent boundary layer. The second typical flow case is the oblique shock wave reflection. The most characteristic case in European research is connected to the 6th FP IP HISAC project concerning a supersonic business jet. The design speed of this airplane is M = 1.6. Therefore the TFAST consortium decided to use this Mach number as the basic case. Pressure disturbance at this Mach number is not very high and can be compared to the disturbance of the normal shock at the incipient separation Mach number mentioned earlier. As mentioned earlier, shock reflection at M = 1.6 may be related to incipient separation. Therefore two additional test cases were planned with different Mach numbers. ITAM conducted an M = 1.5 test case, and TUD an M = 1.7 test case. These partners have also previously made very specialized and successful contributions to the UFAST project