Yancheng You

On the flow physics of a low momentum flux ratio jet in a supersonic turbulent crossflow

A Detached Eddy Simulation (DES) study of a low momentum flux ratio jet, J=(ρu2)jet/(ρu2)∞=0.35, in the HyShot II scramjet system is carried out. The flow structure near the injector, shock pattern in the symmetry plane as well as the instantaneous coherent structures are presented and explained in the paper. Different from most previous studies, over-expanded and under-expanded flow states occur simultaneously at the exit of the current jet porthole. The shock system near the injector is therefore a combination of a detached normal shock and a small three-dimensional barrel shock terminated by a Mach disk. Further insight into the flow physics is conducted by visualizing instantaneous coherent structures. The formation of Ω-shaped vortices, which was observed in experiments previously, but never well-studied numerically, is discussed in detail. A new understanding of the key flow physics and mixing patterns in the low momentum flux ratio jet in supersonic crossflow is finally schematically provided.

Dual Waverider Concept for the Integration of Hypersonic Inward-Turning Inlet and Airframe Forebody

On the basis of the osculating waverider and intern al waverider inlet, a new integration concept for the hypersonic forebody and inlet called dual w averider design is presented in this paper. And a new internal convergent flowfield is proposed and n amed ICFC flowfield afterward. Inviscid calculation results show that ICFC flowfield could receive a conical impinging shock at the entrance and generate a relatively uniform outflow for its d ownstream component. The key point of the dual waverider concept is to organize an internal/planar /external shock wave, which could be accepted for both the internal and external flows. When ICFC flowfields strength of initial shock matches with the outer planar and external ones, such kind of combin ed internal/planar/external shock wave can be produced. With a sample shock wave structure in this category, a dual waverider system is traced out and will be validated in the next stage.

An Overview of the Advantages and Concerns of Hypersonic Inward Turning Inlets

Advantages and concerns of hypersonic inward turning inlets are discussed in this paper. Inward turning inlets have the following essential advantages: 1) high compression ratio, 2) short length in the axial direction, 3) small wetted area, 4) high mass capture rate, 5) good performance at off-design conditions, 6) low drag while operating. In addition, there are another three potential advantages proposed, which might further enhance the popularity of these inlets. That is, 1) suitability for working together with round components, 2) being designed by using inverse methods, i.e. the on-design performance is known in advance, 3) a good chance to be integrated with a waverider airframe in a three-dimensional way. Regarding the concerns, the integration of inward turning inlets with a planar forebody as well as the low Mach number starting ability of the inward turning inlets still requires more particular attention.

High Enthalpy Wind Tunnel Tests of Three- Dimensional Section Controllable Internal Waverider Hypersonic Inlet

The Mach 5 high enthalpy wind tunnel tests of a specific three-dimensional section controllable internal waverider hypersonic inlet are studied in this paper. The performance and static pressure distributions of the inlet at different back-pressure conditions (included free back-pressure) are obtained. The mass flow capture rate is 0.99 0.01 ± in all test statuses, which confirms the inlet’s high mass capture capacity. In the free back-pressure status, the mass-averaged total pressure recovery at the exit plane of the inlet is 0.609 and the average Mach number is 2.78, while the pressure rise reaches 13.98. The inlet can work in the maximum back-pressure status (51.4 times of upstream pressure) while its exit plane massaverage Mach number is close to 1. At that moment, the separation phenomenon could be identified near the ramp-side shoulder position. The inlet works unsteadily and becomes unstarted once the average Mach number is below 1 while the back-pressure keeps increasing. It is pointed out that the location where the reflection shock wave intersects with the ramp-side shoulder is the critical position and the shock wave/boundary layer interaction is the inducement of the inlet unstarting in back-pressure tests. Furthermore, for the elliptical exit inlet, the flow loss in the major axis direction is more remarkable, which slows the flow there into subsonic accompanying probable reversing flow. The three-dimension effect should be considered when one designs the elliptical exit inlet and its isolator.

Integration methodology for waverider-derived hypersonic inlet and vehicle forebody

The dual waverider method for the integration of hypersonic inlet and waverider forebody is presented. The method connects the hypersonic internal waverider inlet and the waverider forebody by application of the osculating waverider theory. As an inverse design method, one should construct a three-dimensional shock wave surface that has continuous local curvature centers. In particular, the centers have to locate internally and externally away from the shock to obtain a proper integration configuration. Unlike most previous integration techniques, this approach jointly creates waverider lower surface and inward turning inlet configuration. As a result, the integratedcomponents become essential elements of a hypersonic vehicle design. Furthermore, detailed numerical simulations of several integration schemes demonstrate the feasibility and diversity of the new approach.

Cross Section Controllable Hypersonic Inlet Design Using Streamline-Tracing and Osculating Axisymmetric Concepts

In the passed decades, conical waverider design and osculating axisymmetric flows waverider theory have been extensively reported for the design of airframe geometrics. These two concepts make it quite easy to determine complicated three dimensional waverider configurations for constant strength shocks assuming a piecewise conical flowfield. And in hypersonic inlet design region, the up-to-date focus has been concentrated into the development of a detailed design methodology for high performance three dimensional hypersonic inlets. Streamline tracing technology has been proved to be an effective way for the job [Smart, 1998 #58] [Billig, 2003 #57]. In this paper, an innovative cross section controllable three dimensional hypersonic inlet (also named internal waverider hypersonic inlet) are successfully designed by the integration of streamline tracing and osculating axisymmetric concepts respectively.

Application of the Re-theta Transition Model in High Speed Flows

The prediction of boundary layer transition in high-speed flows plays a crucial role in the design of hypersonic vehicles. The paper contains a detailed description of a correlation-based transition model using local variables, namely the Langtry-Menter transition model. The model was implemented into DLR-TAU code and tested for hypersonic test cases obtained by other partners. Discussions of the physical background as well as the fundamental understanding of the main variables, correlations and functions inside the model are provided. Test results show that direct applications of the model for high-speed flows can capture hypersonic transition. For the blunted double ramp case, the present model can accurately predict the transition process. However, for the shock impingement cases with assigned sharp leading edges, the present model tends to over-estimate the heat flux by over 30% than the experimental data. This suggests a further detailed study of both numerical and experimental settings.

Numerical Study of the Subsonic Base Flow with a Side Support

The base flow of a generic rocket configuration is investigated numerically with different levels of turbulence modeling. At the nominal flow conditions, the comparison of numerical results with the experiments shows significant deviations in the vertical plane where a side support stands. A simulation of the open test section indicates two necessities of correction. On the one hand, an C p increase of 0.015 is necessary to correlate the measured plenum pressure with the inflow location of the numerical simulation. On the other hand, a − 0.32° angle of attack modification should be accounted for a justified comparison between the DES results and the experiments. A strong sensitivity towards such small angles of attack has been observed later in the experiments but not in respective RANS solutions. The DES results agree well with the experiment based on the above-mentioned corrections.

Evaluation of Turbulence Models in Predicting Hypersonic and Subsonic Base Flows Using Grid Adaptation Techniques

The flows behind the base of a generic rocket, at both hypersonic and subsonic flow conditions, are numerically studied. The main concerns are addressed to the evaluation of turbulence models and the using of grid adaptation techniques. The investigation focuses on two configurations, related to hypersonic and subsonic experiments. The applicability tests of different turbulence models are conducted on the level of two-equation models calculating the steady state solution of the Reynolds-averaged Navier-Stokes(RANS) equations. All used models, the original Wilcox k-ω, the Menter shear-stress transport (SST) and the explicit algebraic Reynolds stress model(EARSM) formulation, predict an asymmetric base flow in both cases caused by the support of the models. A comparison with preliminary experimental results indicates a preference for the SST and EARSM results over the results from the older k-ω model. Sensitivity studies show no significant influence of the grid topology or the location of the laminar to turbulent transition on the base flow field, but a strong influence of even small angles of attack is reported from the related experiments.

A Non-linear Eddy-Viscosity View of Shock Wave/Boundary Layer Interaction Flow Simulation

Shock wave/boundary-layer interaction (SWBLI) is a common but important flow phenomenon, within various engineering design areas such as engine inlets, compressors and turbines. In past decades, researchers have made great efforts towards better understanding and modeling of SWBLI flows. The reviews by Knight and Degrez[1], Zheltovodov[2], Dolling[3] examine the capability of Reynoldsaveraged Navier-Stokes (RANS) turbulence models in the prediction of SWBLI. The common conclusion is that most RANS models based on the linear formulation of the Boussinesq assumption are difficult to accurately predict details of flow separation, i.e. the distributions of pressure loads, heat transfer and skin friction. In consequence, a lot of efforts are put in deriving non-linear RANS turbulence models, either in an explicit algebraic form or through transport equations for the Reynolds stress components. The ongoing research emphasis is to get a physically reliable understanding of SWBLI and to reach a point where a unique non-linear formulation could be used for the modeling of Reynolds stress in a large range of flow configurations.