Pseudo-shock waves and their interactions in high-speed intakes

In an air-breathing engine the flow deceleration from supersonic to subsonic conditions takes places inside the isolator through a gradual compression consisting of a series of shock waves. The wave system, referred to as a pseudo-shock wave or shock train, establishes the combustion chamber entrance conditions, and therefore influences the performance of the entire propulsion system. The characteristics of the pseudo-shock depend on a number of variables which make this flow phenomenon particularly challenging to be analysed. Difficulties in experimentally obtaining accurate flow quantities at high speeds and discrepancies of numerical approaches with measured data have been readily reported. Understanding the flow physics in the presence of the interaction of numerous shock waves with the boundary layer in internal flows is essential to developing methods and control strategies. To counteract the negative effects of shock wave/boundary layer interactions, which are responsible for the engine unstart process, multiple flow control methodologies have been proposed. Improved analytical models, advanced experimental methodologies and numerical simulations have allowed a more in-depth analysis of the flow physics. The present paper aims to bring together the main results, on the shock train structure and its associated phenomena inside isolators, studied using the aforementioned tools. Several promising flow control techniques that have more recently been applied to manipulate the shock wave/boundary layer interaction are also examined in this review

Numerical modeling and simulation of supersonic flows in propulsion systems by open-source solvers

Two open-source solvers, Eilmer and hyFoam, are here considered for their performance in simulating high-speed flows in different flow conditions and geometric configurations typical of propulsive systems at supersonic speeds. The goal is to identify the open-source platform providing the best compromise between accuracy, flexibility and computational cost to eventually simulate the flow fields inside ramjet and scramjet engines. The differences in terms of discretization and solution methods of the selected solvers are discussed in terms of their impact on solution accuracy and computational efficiency and in view of the aerothermodynamic analysis and design of future trans-atmospheric propulsive systems. In this work steady state problems are considered. Numerical results of two scramjet type engines demonstrated a similar predictive capability of both codes in non-reacting conditions. These results highlight their potential to be considered for further characterization of overall engine performance

Analytical and Numerical Study of the Non-uniformity induced Type II Asymmetric Cap Shock Mach Reflection in Over-expanded Supersonic Jets

A combined analytical and numerical study is conducted to investigate the asymmetric cap-shock non-uniform Mach Reflection (csMR) phenomenon outside of an over-expanded supersonic jet for the first time. Prior analytical works have only considered the wedge-induced steady symmetric and asymmetric Mach reflection configurations. However, there is another structure occurring in nozzle flow fields known as a non-uniformity induced cap-shock pattern. We derive a new analytical model to predict the wave structure of the asymmetric csMR in the absence of internal shocks by extending on a prior symmetric Mach reflection (MR) model. Different from the wedge flow case, flow non-uniformity is incorporated by assuming different upstream Mach numbers in both upper and lower domains, where the asymmetry is predicted through averaged flowfields and slipstream inclination angles. The numerical approach utilises an Euler solver for comparisons to the developed theory. It is found that the model adequately predicts the shock structure obtained from numerical simulations, and can be utilised for various sets of parameters to capture the direct Mach reflection (DiMR-DiMR) configuration. The von Neumann criterion is also predicted by the new analytic model, along with the Mach stem profile and shock curvatures. Based on the analytical and numerical observations, a hypothesis is also made regarding the stability of MR structures within an over-expanded jet.Comment: 33 pages, 18 figure

Numerical Study of Two-Dimensional Secondary Injection Into a Mach 3.5 Freestream

Upstream interaction within a dual-mode scramjet engine is investigated numerically. The upstream interaction is investigated by increasing the freestream-to-injector pressure ratio. The constant area duct and sudden expansion geometries are used to achieve these results. Analysis using the symmetry boundary condition is compared to the full boundary condition for the constant area duct. Numerical analysis of the Mach 3.5 freestream is conducted using normal sonic injection of nitrogen gas to create the upstream interaction. Comparisons, where applicable, are made to experimental results. Due to the high speed of the flowfield, oblique shockwaves are present causing numerical convergence difficult to achieve. The flowfield symmetry assumptions break down due to the high shear stresses present in the boundary layer separation region. As the freestream-to-injector pressure ratio is increased this separation region begins to move upstream within the isolator region. This study proves that the upstream interaction is not related to reacting flows, but rather to the high turbulent shear stresses that are present in the boundary layer separation regions

Underexpanded Supersonic Plume Surface Interactions: Applications for Spacecraft Landings on Planetary Bodies

Numerical and experimental investigations of both far-field and near-field supersonic steady jet interactions with a flat surface at various atmospheric pressures are presented in this paper. These studies were done in assessing the landing hazards of both the NASA Mars Science Laboratory and Phoenix Mars spacecrafts. Temporal and spatial ground pressure measurements in conjunction with numerical solutions at altitudes of approx.35 nozzle exit diameters and jet expansion ratios (e) between 0.02 and 100 are used. Data from steady nitrogen jets are compared to both pulsed jets and rocket exhaust plumes at Mach approx.5. Due to engine cycling, overpressures and the plate shock dynamics are different between pulsed and steady supersonic impinging jets. In contrast to highly over-expanded (e 5 (lunar atmospheric regime), the ground pressure is minimal due to the development of a highly expansive shock structure. We show this is dependent on the stability of the plate shock, the length of the supersonic core and plume decay due to shear layer instability which are all a function of the jet expansion ratio. Asymmetry and large gradients in the spatial ground pressure profile and large transient overpressures are predominantly linked to the dynamics of the plate shock. More importantly, this study shows that thruster plumes exhausting into martian environments possess the largest surface pressure loads and can occur at high spacecraft altitudes in contrast to the jet interactions at terrestrial and lunar atmospheres. Theoretical and analytical results also show that subscale supersonic cold gas jets adequately simulate the flow field and loads due to rocket plume impingement provided important scaling parameters are in agreement. These studies indicate the critical importance of testing and modeling plume-surface interactions for descent and ascent of spacecraft and launch vehicles

Impact of Vibrational Nonequilibrium on the Simulation and Modeling of Dual-Mode Scramjets

The practical realization of supersonic flight relies on the development of a robust propulsion system. These air-breathing scramjet engines process fuel and high-speed air to generate propulsive thrust. Unlike conventional jet engines, scramjets achieve efficient thrust by compressing air through a system of shocks. As a result, the reliability of the engine is highly sensitive to the stability of these shock structures. Physically, these shocks are located in an engine component called the isolator. The shock structures are spatially distributed leading to a region of pressure increase, which is termed the pseudoshock. As vehicle operating conditions change, the length of the pseudoshock will change, reflecting changes to inflow conditions and operation of downstream combustor component. The overall objective of this thesis is to understand the complex flow inside these isolators. Of particular focus is the role of molecular processes in the development of the shocks. At high enthalpy conditions, the internal motions of the molecules are moved out of equilibrium due to compression shocks, which affects not only the thermophysical properties of air, but more critically the fuel-air mixing and chemical reactions. While there exists a vast body of literature on scramjet isolators, almost all of these works focus on low enthalpy conditions due to laboratory experimental limitations, or simply rely on equilibrium thermodynamics. In this work, the effect of nonequilibrium on isolator and scramjet combustors at high-altitude high-enthalpy flight conditions was studied using high-fidelity numerical simulations. Detailed models for the description of molecular nonequilibrium, in the form of multi-temperature model was used. Computational chemistry derived reaction rates were used to describe the combustion processes. These studies revealed the following key features: a) nonequilibrium of vibrational states greatly increases pseudoshock length, b) contrary to external hypersonics, nonequilibrium accelerates chemical reactions in the combustor, reducing the distance from fuel injection to flame ignition and stabilization, c) while multi-temperature models are adequate to express such nonequilibrium effects, more detailed state-specific representations clearly demonstrate that molecular populations do not follow the Boltzmann relation even at subsonic but compressible flow conditions. In a related study but using equilibrium thermal conditions, it was shown that the isolator shock structure can develop a resonance to inflow perturbations that can vastly increase the pseudoshock spatial oscillations. These results verify that isolator flow is a complex nonlinear process and clearly demonstrate that the design of scramjets needs to include the effect of thermal nonequilibrium. To begin addressing this process, reduced-order models in the form of a flux-conserved one-dimensional formulation for estimating pseudoshock length was developed for thermal equilibrium conditions.PHDAerospace EngineeringUniversity of Michigan, Horace H. Rackham School of Graduate Studieshttps://deepblue.lib.umich.edu/bitstream/2027.42/143909/1/rfievet_2.pdfhttps://deepblue.lib.umich.edu/bitstream/2027.42/143909/2/rfievet_1.pd

Investigation on supersonic high-speed internal flows and the tools to study their interactions

Air-breathing vehicles are characterised by a high level of integration between the propulsion system and the vehicle frame. Since the peculiarity of this type of aircraft is the absence of moving parts, before the air flow arrives at the combustion chamber, it must be slowed down to lower supersonic speeds, in scramjets, or to subsonic speeds, in ramjets, by means of a shock wave structure, called a shock train. The prediction and control of the shock train is important for the evaluation of the engine performance. This work aims to improve the understanding of the flow mechanisms occurring in the shock train as a consequence of the interaction of shock waves with the boundary layer in long and narrow ducts. A full pressure sensitive paint system was developed. Polymer-based and ruthenium-based compounds were identified as suitable for the investigation of the shock train in the wind tunnel. Before being able to collect experimental data, the design and manufacture of an indraft supersonic wind tunnel able to operate at mach numbers M= 2 and M= 4 was accomplished. The air at ambient conditions is drawn into the tunnel and then discharged into a vacuum tank with a volume of 34 m^3. Preliminary attempts to run the wind tunnel have identified the presence of leakages between the vacuum tank and the wind tunnel that prevented the establishment of the pressure difference required to obtain a supersonic flow in the test section. In support of the experimental approach, different flow configurations are numerically studied using the RANS equations. The k-w Wilcox model provided the most accurate results for such a complex flow field. Sensitivity studies are carried out since the characteristics of the shock train depend on several variables, including the duct geometry and the back pressure. The numerical findings revealed that the location of the shock train strongly varies with the grid size. Transient simulation is used to reproduce the shock train oscillation due to the pressure fluctuations that occur in the combustion chamber of an air-breathing aircraft. Under a sinusoidal forcing, the shock train executes a motion around its mean position that deviates from a perfect sinusoidal profile depending on the oscillation amplitude, frequency, and whether the pressure is first increased or decreased. With large oscillation amplitudes the shock train is greatly influenced by a pressure increase rather than a pressure drop, but the opposite is observed at small oscillation amplitudes. With varying forcing frequency, the shock displacement around its mean position decreases as the forcing frequency increases

EVALUATION OF GEOMETRIC SCALE EFFECTS FOR SCRAMJET ISOLATORS

A numerical analysis was conducted to study the effects of geometrically scaling scramjet inlet-combustor isolators. Three-dimensional fully viscous numerical simulation of the flow inside constant area rectangular ducts, with a downstream back pressure condition, was analyzed using the SolidWorks Flow Simulation software. The baseline, or 1X, isolator configuration has a 1” x 2.67” cross section and 20” length. This baseline configuration was scaled up based on the 1X configuration mass flow to 10X and 100X configurations, with ten and one hundred times the mass flow rate, respectively. The isolator aspect ratio of 2.67 was held constant for all configurations. To provide for code validation, the Flow Simulation program was first used to analyze a converging-diverging channel and a wind tunnel nozzle. The channel case was compared with analytical theory and showed good agreement. The nozzle case was compared with AFRL experimental data and showed good agreement with the entrance and exit conditions (Pi0= 40 psia, Ti0= 530ºR, Pe= 18.86 psia, Te= 456ºR, respectively). While the boundary layer thickness remained constant, the boundary layer thickness with respect to the isolator height decreased as the scale increased. For all the isolator simulations, a shock train was expected to form inside the duct. However, the flow simulation failed to generate this flow pattern, due to improper sizing of the isolator and combustor for a 3-D model or having a low pressure ratio of 2.38. Instead, a single normal shock wave was established at the same relative location within the length of each duct, approximately 80% of the duct length from the isolator entrance. The shape of the shock changed as the scale increased from a normal shock wave, to a bifurcated shock wave, and to a normal shock train, respectively for the 1X, 10X, and 100X models