Showing papers on "Supersonic speed published in 2018"

Supersonic Combustion in Air-Breathing Propulsion Systems for Hypersonic Flight

Abstract: Great efforts have been dedicated during the last decades to the research and development of hypersonic aircrafts that can fly at several times the speed of sound. These aerospace vehicles have revolutionary applications in national security as advanced hypersonic weapons, in space exploration as reusable stages for access to low Earth orbit, and in commercial aviation as fast long-range methods for air transportation of passengers around the globe. This review addresses the topic of supersonic combustion, which represents the central physical process that enables scramjet hypersonic propulsion systems to accelerate aircrafts to ultra-high speeds. The description focuses on recent experimental flights and ground-based research programs and highlights associated fundamental flow physics, subgrid-scale model development, and full-system numerical simulations.

Alterations of Cowl Lip for the Improvement of Supersonic-Intake Performance

Abstract: This paper discusses the performance enhancement of supersonic air intake model through the implementation of blunted leading edge to the cowl lip section of the model. A supersonic air intake model with sharp cowl leading edge is initially considered to numerically investigate its performance. Mach 3, supersonic intake flow through the base model has been simulated using commercial CFD package Ansys Fluent-15. Comparison of numerical predictions and experimental measurements is presented to demonstrate the correctness and accuracy of numerical frame work followed in the present study. Higher order spatial accuracy of the solver along with suitably refined mesh helped in accurate capturing of the flow field. Modification to the cowl lip is proposed as an effective method to improve the performance of the supersonic air intake. Two different blunted cowl leading edge geometries were investigated to identify the possible enhancement in performance parameters. Improvement in mass capture and combustion stability attained through the use of forward shifted blunt cowl leading edge is presented. It is also revealed through the present study that the blunt cowl leading edge can reduce the intensity of shock wave boundary layer interaction occurring at the isolator entry section. Deviation in total pressure recovery and flow distortion observed with different supersonic air intake models are also discussed with reasons for the same. This study demonstrates the scope of overall improvement in scramjet engine performance through the use of suitably positioned blunt cowl leading edge.

Letter: Görtler-like vortices in an impinging shock wave/turbulent boundary layer interaction flow

Abstract: Gortler-like vortices (GLVs) are ubiquitous in supersonic flows and have a significant influence on the local flow. However, the existence of these streamwise vortices in impinging shock wave/turbulent boundary layer interaction (ISTBLI) has been a topic in dispute for decades. By the ice-cluster based planar laser scattering technique and a simplified aerodynamic model, we obtain and analyze high-resolution images acquired at two orthogonal planes. The results support the presence of GLVs in the ISTBLI flow. A possible mechanism of their generation is provided.

Characterization of a Supersonic Turbine Downstream of a Rotating Detonation Combustor

Abstract: Rotating detonation combustors (RDCs) offer theoretically a significant total pressure increase, which may result in enhanced cycle efficiency. The fluctuating exhaust of RDC, however, induces low supersonic flow and large flow angle fluctuations at several kHz, which affects the performance of the downstream turbine. In this paper, a numerical methodology is proposed to characterize a supersonic turbine exposed to fluctuations from RDC without any dilution. The inlet conditions of the turbine were extracted from a three-dimensional (3D) unsteady Reynolds-averaged Navier–Stokes simulation of a nozzle attached to a rotating detonation combustor, optimized for minimum flow fluctuations and a mass-flow averaged Mach number of 2 at the nozzle outlet. In a first step, a supersonic turbine able to handle steady Mach 2 inflow was designed based on a method of characteristics solver and total pressure loss was assessed. Afterward, unsteady simulations of eight stator passages exposed to periodic oblique shocks were performed. Total pressure loss was evaluated for several oblique shock frequencies and amplitudes. The unsteady stator outlet profile was extracted and used as inlet condition for the unsteady rotor simulations. Finally, a full stage unsteady simulation was performed to characterize the flow field across the entire turbine stage. Power extraction, airfoil base pressure, and total pressure losses were assessed, which enabled the estimation of the loss mechanisms in supersonic turbine exposed to large unsteady inlet conditions.

Formation of surface trailing counter-rotating vortex pairs downstream of a sonic jet in a supersonic cross-flow

Abstract: Direct numerical simulations were conducted to uncover physical aspects of a transverse sonic jet injected into a supersonic cross-flow at a Mach number of 2.7. Simulations were carried out for two different jet-to-cross-flow momentum flux ratios () of 2.3 and 5.5. It is identified that collision shock waves behind the jet induce a herringbone separation bubble in the near-wall jet wake and a reattachment valley is formed and embayed by the herringbone recirculation zone. The recirculating flow in the jet leeward separation bubble forms a primary trailing counter-rotating vortex pair (TCVP) close to the wall surface. Analysis on streamlines passing the separation region shows that the wing of the herringbone separation bubble serves as a micro-ramp vortex generator and streamlines acquire angular momentum downstream to form a secondary surface TCVP in the reattachment valley. Herringbone separation wings disappear in the far field due to the cross-interaction of lateral supersonic flow and the expansion flow in the reattachment valley, which also leads to the vanishing of the secondary TCVP. A three-dimensional schematic of surface trailing wakes is presented and explains the formation mechanisms of the surface TCVPs.

Electrically Driven Supersonic Combustion

Abstract: This manuscript reviews published works related to plasma assistance in supersonic combustion; focusing on mixing enhancement, ignition and flameholding. A special attention is paid for studies, which the author participated in person. The Introduction discusses general trends in plasma-assisted combustion and, specifically, work involving supersonic conditions. In Section 2, the emphasis is placed on different approaches to plasma application for fuel ignition and flame stabilization. Several schemes of plasma-based actuators for supersonic combustion have been tested for flameholding purposes at flow conditions where self-ignition of the fuel/air mixture is not realizable due to low air temperatures. Comparing schemes indicates an obvious benefit of plasma generation in-situ, in the mixing layer of air and fuel. In Section 3, the problem of mixing enhancement using a plasma-based technique is considered. The mechanisms of interaction are discussed from the viewpoint of triggering gasdynamic instabilities promoting the kinematic stretching of the fuel-air interface. Section 4 is related to the description of transitional processes and combustion instabilities observed in plasma-assisted high-speed combustion. The dynamics of ignition and flame extinction are explored. It is shown that the characteristic time for reignition can be as short as 10 ms. Two types of flame instability were described which are related to the evolution of a separation zone and thermoacoustic oscillations, with characteristic times 10 ms and 1 ms correspondingly.

Toward an understanding of supersonic modes in boundary-layer transition for hypersonic flow over blunt cones

Abstract: Realistic flight vehicles for reentry into the Earth’s atmosphere are commonly similar to blunted cones. The main reason for blunting a cone is to mitigate high heat loads at the nose. Another reason for blunting the cone is to delay boundary-layer transition. It is commonly understood that the second mode is damped in flow over a cone as the nose radius is increased. This is thought to lead to the delay in transition. Here, a blunted cone at a realistic reentry trajectory point with significant real-gas effects is studied. It is shown, using linear stability theory and direct numerical simulation, that there exist multiple unstable modal instabilities in the boundary layer. One of these modal instabilities is called the supersonic mode, as its phase velocity is supersonic relative to the flow velocity at the edge of the boundary layer. Its growth rate is found to increase with increasing nose radius until a certain nose radius is reached. After this radius, any further increase in nose radius decreases its growth rate. There is adequate agreement between theory and direct numerical simulation for the growth rate, phase velocity and eigenfunction of the supersonic mode. At the reentry conditions tested, the supersonic mode is more likely the cause of boundary-layer transition than the second mode for blunted cones with a small wall-temperature ratio. Initial parametric studies confirm that a decrease in wall temperature amplifies the supersonic mode. Also, the supersonic mode’s growth rate is shown to be a maximum when its phase velocity is aligned with the flow velocity.

A note on supersonic flow control with nanosecond plasma actuator

Abstract: A concept study on supersonic flow control using nanosecond pulsed plasma actuator is conducted by means of numerical simulation. The nanosecond plasma discharge is characterized by the generation of a micro-shock wave in ambient air and a residual heat in the discharge volume arising from the rapid heating of near-surface gas by the quick discharge. The residual heat has been found to be essential for the flow separation control over aerodynamic bodies like airfoil and backward-facing step. In this study, novel experiment is designed to utilize the other flow feature from discharge, i.e., instant shock wave, to control supersonic flow through shock-shock interaction. Both bow shock in front of a blunt body and attached shock anchored at the tip of supersonic projectile are manipulated via the discharged-induced shock wave in an appropriate manner. It is observed that drag on the blunt body is reduced appreciably. Meanwhile, a lateral force on sharp-edged projectile is produced, which can steer the body a...

Numerical analysis on interactions of vortex, shock wave, and exothermal reaction in a supersonic planar shear layer laden with droplets

Abstract: The analysis on the interactions of a large-scale shearing vortex, an incident oblique shock wave, and a chemical reaction in a planar shear layer is performed by numerical simulations. The reacting flows are obtained by directly solving the multi-species Navier-Stokes equations in the Eulerian frame, and the motions of individual point-mass fuel droplets are tracked in the Lagrangian frame considering the two-way coupling. The influences of shock strength and spray equivalence ratio on the shock-vortex interaction and the induced combustion are further studied. Under the present conditions, the incident shock is distorted by the vortex evolution to form the complicated waves including an incident shock wave, a multi-refracted wave, a reflected wave, and a transmitted wave. The local pressure and temperature are elevated by the shock impingement on the shearing vortex, which carries flammable mixtures. The chemical reaction is mostly accelerated by the refracted shock across the vortex. Two different exot...