Showing papers on "Supersonic speed published in 2007"

Direct numerical simulation of supersonic turbulent boundary layer over a compression ramp

Abstract: A direct numerical simulation of shock wave and turbulent boundary layer interaction for a 24 deg compression ramp configuration at Mach 2.9 and Re θ 2300 is performed. A modified weighted, essentially nonoscillatory scheme is used. The direct numerical simulation results are compared with the experiments of Bookey et al. at the same flow conditions. The upstream boundary layer, the mean wall-pressure distribution, the size of the separation bubble, and the velocity profile downstream of the interaction are predicted within the experimental uncertainty. The change of the mean and fluctuating properties throughout the interaction region is studied. The low frequency motion of the shock is inferred from the wall-pressure signal and freestream mass-flux measurement.

Direct numerical simulation of hypersonic turbulent boundary layers. Part 1. Initialization and comparison with experiments

Abstract: A systematic procedure for initializing supersonic and hypersonic turbulent boundary layers at controlled Mach number and Reynolds number conditions is described. The initialization is done by locally transforming a true direct numerical simulation flow field, and results in a nearly realistic initial magnitude of turbulent fluctuations, turbulence structure and energy distribution. The time scales necessary to forget the initial condition are studied. The experimental conditions of previous studies are simulated. The magnitude of velocity and temperature fluctuations, as well as the turbulent shear stresses given by the direct numerical simulations are in agreement with the experimental data.

LES of supersonic combustion in a scramjet engine model

Abstract: In this study, Large Eddy Simulation (LES) has been used to examine supersonic flow and combustion in a model scramjet combustor. The LES model is based on an unstructured finite volume discretization, using total variational diminishing flux reconstruction, of the filtered continuity, momentum, enthalpy, and passive/reactive scalar equations, used to describe the combustion process. The configuration used is similar to the laboratory scrarmjet at the Institute for Chemical Propulsion of the German Aerospace Center (DLR) and consists of a one-sided divergent channel with a wedge-shaped flameholder at the base of which hydrogen is injected. Here, we investigate supersonic flow with hydrogen injection and supersonic flow with hydrogen injection and combustion. For the purpose of validation, the LES results are compared with experimental data for velocity and temperature at different cross-sections. In addition, qualitative comparisons are also made between predicted and measured shadowgraph images. The LES computations are capable of predicting both the non-reacting and reacting flowfields reasonably well-in particular we notice that the LES model identifies and differentiates between peculiarities of the flowfields found in the experiments. (c) 2006 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

The forced response of choked nozzles and supersonic diffusers

Abstract: The response of choked nozzles and supersonic diffusers to one-dimensional flow perturbations is investigated. Following previous arguments in the literature, small flow perturbations in a duct of spatially linear steady velocity distribution are determined by solution of a hyper-geometric differential equation. A set of boundary conditions is then developed that extends the existing work to a nozzle of arbitrary geometry. This analysis accommodates the motion of a plane shock wave and makes no assumption about the nozzle compactness. Numerical simulations of the unsteady, quasi-one-dimensional Euler equations are performed to validate this analysis and also to indicate the conditions under which the perturbations remain approximately linear. The nonlinear response of compact choked nozzles and supersonic diffusers is also investigated. Simple analyses are performed to determine the reflected and transmitted waveforms, as well as conditions for unchoke, 'over-choke' and unstart. This analysis is also supported with results from numerical simulations of the Euler equations. © Cambridge University Press 2007.

Can Protostellar Jets Drive Supersonic Turbulence in Molecular Clouds

Abstract: Jets and outflows from young stellar objects have been proposed as candidates to drive supersonic turbulence in molecular clouds. Here we present the results from multidimensional jet simulations where we investigate in detail the energy and momentum deposition from jets into their surrounding environment and quantify the character of the excited turbulence with velocity probability density functions. Our study includes jet-clump interaction, transient jets, and magnetized jets. We find that collimated supersonic jets do not excite supersonic motions far from the vicinity of the jet. Supersonic fluctuations are damped quickly and do not spread into the parent cloud. Instead, subsonic, noncompressional modes occupy most of the excited volume. This is a generic feature that cannot be fully circumvented with overdense jets or magnetic fields. Nevertheless, jets are able to leave strong imprints in the cloud structure and can disrupt dense clumps. Our results question the ability of collimated jets to sustain supersonic turbulence in molecular clouds.

Discrete Roughness Transition for Hypersonic Flight Vehicles

Abstract: The importance of discrete roughness and the correlations developed to predict the onset of boundary layer transition on hypersonic flight vehicles are discussed. The paper is organized by hypersonic vehicle applications characterized in a general sense by the boundary layer: slender with hypersonic conditions at the edge of the boundary layer, moderately blunt with supersonic, and blunt with subsonic. This paper is intended to be a review of recent discrete roughness transition work completed at NASA Langley Research Center in support of agency flight test programs. First, a review is provided of discrete roughness wind tunnel data and the resulting correlations that were developed. Then, results obtained from flight vehicles, in particular the recently flown Hyper-X and Shuttle missions, are discussed and compared to the ground-based correlations.

Effects of high-speed airflows on a surface dielectric barrier discharge

Abstract: A detailed study of the interaction between high-speed gas flows and surface dielectric barrier discharges (DBD) is presented. In the present paper, it is demonstrated that a DBD can be sustained in transonic airflows, up to isentropic Mach numbers of 1.1. The plasma is characterized electrically, as well as optically with a CCD camera and a photomultiplier tube. Different airflow velocities, plasma excitation frequencies and voltages are investigated. The airflow has a significant influence on the plasma characteristics: the glow component is reduced, the discharge becomes more filamentary and most importantly, the light emission duration from individual microdischarges is reduced by more than a factor of ten at high flow velocities. Large edge effects play a key role in the interaction between the flow and the plasma. These results offer new perspectives for the use of dielectric barrier discharges in transonic and supersonic gas flows and their applications to airflow control and to plasma-assisted combustion.

Supersonic dislocations observed in a plasma crystal.

Abstract: Experimental results on the dislocation dynamics in a two-dimensional plasma crystal are presented. Edge dislocations were created in pairs in lattice locations where the internal shear stress exceeded a threshold and then moved apart in the glide plane at a speed higher than the sound speed of shear waves, ${C}_{T}$. The experimental system, a plasma crystal, allowed observation of this process at an atomistic (kinetic) level. The early stage of this process is identified as a stacking fault. At a later stage, supersonically moving dislocations generated shear-wave Mach cones.

An atomic coilgun: using pulsed magnetic fields to slow a supersonic beam

Abstract: We report the experimental demonstration of a novel method to slow atoms and molecules with permanent magnetic moments using pulsed magnetic fields. In our experiments, we observe the slowing of a supersonic beam of metastable neon from 461.0 ± 7.7 to 403 ± 16 m s−1 in 18 stages, where the slowed peak is clearly separated from the initial distribution. This method has broad applications as it may easily be generalized, using seeding and entrainment into supersonic beams, to all paramagnetic atoms and molecules.

Supersonic flow onto a solid wedge

Abstract: We consider the problem of 2D supersonic flow onto a solid wedge, or equivalently in a concave corner formed by two solid walls. For mild corners, there are two possible steady state solutions, one with a strong and one with a weak shock emanating from the corner. The weak shock is observed in supersonic flights. A long-standing natural conjecture is that the strong shock is unstable in some sense. We resolve this issue by showing that a sharp wedge will eventually produce weak shocks at the tip when accelerated to a supersonic speed. More precisely we prove that for upstream state as initial data in the entire domain, the time-dependent solution is self-similar, with a weak shock at the tip of the wedge. We construct analytic solutions for self-similar potential flow, both isothermal and isentropic with arbitrary $\gamma\geq 1$. In the process of constructing the self-similar solution, we develop a large number of theoretical tools for these elliptic regions. These tools allow us to establish large-data results rather than a small perturbation. We show that the wave pattern persists as long as the weak shock is supersonic-supersonic; when this is no longer true, numerics show a physical change of behaviour. In addition we obtain rather detailed information about the elliptic region, including analyticity as well as bounds for velocity components and shock tangents.

Dual-Pump Coherent Anti-Stokes Raman Scattering Measurements in a Supersonic Combustor

Abstract: The dual-pump coherent anti-Stokes Raman scattering (CARS) method was used to measure temperature and the mole fractions of N 2 and O 2 in a supersonic combustor. Experiments were conducted in NASA Langley Research Center's Direct-Connect Supersonic Combustion Test Facility. In this facility, H 2 - and oxygen-enriched air burn to increase the enthalpy of the simulated air test gas. This gas is expanded through a Mach 2 nozzle and into a combustor model consisting of a short constant-area section followed by a small rearward-facing step and another constant-area section. At the end of this straight section, H 2 fuel is injected at Mach 2 and at a 30-deg angle with respect to the freestream. One wall of the duct then expands at a 3-deg angle for over 1 m. The ensuing combustion is probed optically through ports in the side of the combustor. Dual-pump CARS measurements were performed at the facility nozzle exit and at four planes downstream of fuel injection. Maps are presented of the mean temperature, as well as N 2 and O 2 mean mole-fraction fields. Correlations between fluctuations of the different measured parameters are also presented.

Overview of PIV in supersonic Flows

Abstract: The extension of particle image velocimetry to supersonic and hypersonic wind-tunnel flows has been achieved in the last decade. This was mainly possible with the advent of short interframing-time CCD cameras with temporal resolution allowing to obtain correlated particle images at flow velocities exceeding 1000m∕s. The most challenging aspects of PIV experiments in supersonic flows are still recognized as the seeding-particle-selection and seeding-distribution techniques. Also, the optical access for illumination and imaging require a specific attention since pressurized facilities offer limited optical access. The presence of shock waves in supersonic flows introduces regions where particle tracers slip with respect to the surrounding flow. Moreover, the particle seeding density becomes strongly nonuniform and particle-image blur can occur as a result of the strong refractive index variations. The present chapter reviews the physical and technical problems of PIV experiments and discusses the potential of such techniques on the basis of recent experiments performed in high-speed wind tunnels: double compression ramp at Mach 7 and shock-wave turbulent boundary interaction at Mach 2.

Existence and Stability of Multidimensional Transonic Flows through an Infinite Nozzle of Arbitrary Cross-Sections

Abstract: We establish the existence and stability of multidimensional steady transonic flows with transonic shocks through an infinite nozzle of arbitrary cross-sections, including a slowly varying de Laval nozzle. The transonic flow is governed by the inviscid potential flow equation with supersonic upstream flow at the entrance, uniform subsonic downstream flow at the exit at infinity, and the slip boundary condition on the nozzle boundary. Our results indicate that, if the supersonic upstream flow at the entrance is sufficiently close to a uniform flow, there exists a solution that consists of a C1,α subsonic flow in the unbounded downstream region, converging to a uniform velocity state at infinity, and a C1,α multidimensional transonic shock separating the subsonic flow from the supersonic upstream flow; the uniform velocity state at the exit at infinity in the downstream direction is uniquely determined by the supersonic upstream flow; and the shock is orthogonal to the nozzle boundary at every point of their intersection. In order to construct such a transonic flow, we reformulate the multidimensional transonic nozzle problem into a free boundary problem for the subsonic phase, in which the equation is elliptic and the free boundary is a transonic shock. The free boundary conditions are determined by the Rankine–Hugoniot conditions along the shock. We further develop a nonlinear iteration approach and employ its advantages to deal with such a free boundary problem in the unbounded domain. We also prove that the transonic flow with a transonic shock is unique and stable with respect to the nozzle boundary and the smooth supersonic upstream flow at the entrance.

Computational Study of an Axisymmetric Dual Throat Fluidic Thrust Vectoring Nozzle for a Supersonic Aircraft Application

Abstract: An axisymmetric version of the Dual Throat Nozzle concept with a variable expansion ratio has been studied to determine the impacts on thrust vectoring and nozzle performance. The nozzle design, applicable to a supersonic aircraft, was guided using the unsteady Reynolds-averaged Navier-Stokes computational fluid dynamics code, PAB3D. The axisymmetric Dual Throat Nozzle concept was tested statically in the Jet Exit Test Facility at the NASA Langley Research Center. The nozzle geometric design variables included circumferential span of injection, cavity length, cavity convergence angle, and nozzle expansion ratio for conditions corresponding to take-off and landing, mid climb and cruise. Internal nozzle performance and thrust vectoring performance was determined for nozzle pressure ratios up to 10 with secondary injection rates up to 10 percent of the primary flow rate. The 60 degree span of injection generally performed better than the 90 degree span of injection using an equivalent injection area and number of holes, in agreement with computational results. For injection rates less than 7 percent, thrust vector angle for the 60 degree span of injection was 1.5 to 2 degrees higher than the 90 degree span of injection. Decreasing cavity length improved thrust ratio and discharge coefficient, but decreased thrust vector angle and thrust vectoring efficiency. Increasing cavity convergence angle from 20 to 30 degrees increased thrust vector angle by 1 degree over the range of injection rates tested, but adversely affected system thrust ratio and discharge coefficient. The dual throat nozzle concept generated the best thrust vectoring performance with an expansion ratio of 1.0 (a cavity in between two equal minimum areas). The variable expansion ratio geometry did not provide the expected improvements in discharge coefficient and system thrust ratio throughout the flight envelope of typical a supersonic aircraft. At mid-climb and cruise conditions, the variable geometry design compromised thrust vector angle achieved, but some thrust vector control would be available, potentially for aircraft trim. The fixed area, expansion ratio of 1.0, Dual Throat Nozzle provided the best overall compromise for thrust vectoring and nozzle internal performance over the range of NPR tested compared to the variable geometry Dual Throat Nozzle.

Inviscid and viscous aerodynamics of dense gases

Abstract: A numerical investigation of transonic and low-supersonic flows of dense gases of the Bethe–Zel'dovich–Thompson (BZT) type is presented. BZT gases exhibit, in a range of thermodynamic conditions close to the liquid/vapour coexistence curve, negative values of the fundamental derivative of gasdynamics. This can lead, in the transonic and supersonic regime, to non-classical gasdynamic behaviours, such as rarefaction shock waves, mixed shock/fan waves and shock splitting. In the present work, inviscid and viscous flows of a BZT fluid past an airfoil are investigated using accurate thermo-physical models for gases close to saturation conditions and a third-order centred numerical method. The influence of the upstream kinematic and thermodynamic conditions on the flow patterns and the airfoil aerodynamic performance is analysed, and possible advantages deriving from the use of a non-conventional working fluid are pointed out.

The transient start of supersonic jets

Abstract: This study investigates the initial transient hydrodynamic evolution of highly under-expanded slit and round jets. A closed-form analytic similarity solution is derived for the temporal evolution of temperature, pressure and density at the jet head for vanishing diffusive fluxes, generalizing a previous model of Chekmarev using Chernyi's boundary-layer method for hypersonic flows. Two-dimensional numerical simulations were also performed to investigate the flow field during the initial stages over distances of ~ 1000 orifice radii. The parameters used in the simulations correspond to the release of pressurized hydrogen gas into ambient air, with pressure ratios varying between approximately 100 and 1000. The simulations confirm the similarity laws derived theoretically and indicate that the head of the jet is laminar at early stages, while complex acoustic instabilities are established at the sides of the jet, involving shock interactions within the vortex rings, in good agreement with previous experimental findings. Very good agreement is found between the present model, the numerical simulations and previous experimental results obtained for both slit and round jets during the transient establishment of the jet. Criteria for Rayleigh–Taylor instability of the decelerating density gradients at the jet head are also derived, as well as the formulation of a model addressing the ignition of unsteady expanding diffusive layers formed during the sudden release of reactive gases.

Supersonic-jet experiments using a high-energy laser.

Abstract: In this Letter, laboratory astrophysical jet experiments performed with the LULI2000 laser facility are presented. High speed plasma jets (150 km.s(-1)) are generated using foam-filled cone targets. Accurate experimental characterization of the plasma jet is performed by measuring its time evolution and exploring various target parameters. Key jet parameters such as propagation and radial velocities, temperature, and density are obtained. For the first time, the required dimensionless quantities are experimentally determined on a single-shot basis. Although the jets evolve in vacuum, most of the scaling parameters are relevant to astrophysical conditions.

Equatorial spread F modeling: Multiple bifurcated structures, secondary instabilities, large density ‘bite-outs,’ and supersonic flows

Abstract: references therein]. These numerical studies have shed light on a number of processes affecting the evolution of ESF: the role of a conducting E-layer, an inhomogeneous neutral wind, seeding conditions, molecular ions, and the interaction of multiple bubbles. Despite the progress made by these studies there are still a number of observations that have not been reported in simulation studies. [4] In this Letter we present new results of the onset and evolution of ESF bubbles using a 2D simulation code (NRLESF2) recently developed at the Naval Research Laboratory. The simulation study shows for the first time multiple bifurcations, secondary instabilities, density ‘biteouts’ of over three orders of magnitude, and supersonic flows within low density channels (V ’ few km/s). These results are consistent with radar and satellite measurements, and all-sky optical images.

Design and operation of a supersonic flow cavity for a non-self-sustained electric discharge pumped oxygen–iodine laser

Abstract: The paper presents results of a high-pressure, non-self-sustained crossed discharge–M = 3 supersonic laser cavity operation. A stable and diffuse pulser–sustainer discharge in O2–He flows is generated at pressures of up to P0 = 120 Torr and discharge powers of up to 2.1 kW. The reduced electric field in the dc sustainer discharge ranges from 0.6 × 10−16 to 1.2 × 10−16 V cm2. Singlet delta oxygen (SDO) yield in the discharge, up to 5.0–5.7% at the flow temperatures of 400-420 K, was inferred from the integrated intensity of the (0, 0) band of the O2(a 1Δ → X 3Σ) infrared emission spectra calibrated using a blackbody source. The yield increases with the discharge power and remains nearly independent of the O2 fraction in the mixture (in the 10–20% range). Static pressure and temperature measurements in the supersonic cavity show that a steady-state M = 3 flow in the cavity can be sustained for up to 20 s, at the flow temperature of T = 120 ± 15 K. The results suggest that the measured SDO yield exceeds the threshold yield at the cavity temperature by up to a factor of 2.5. PLIF iodine vapour visualization in the supersonic cavity, which showed the presence of large-scale structures, suggests the need to improve iodine vapour mixing with the main oxygen–helium flow.

Conceptual Study of a Rocket-Ramjet Combined-Cycle Engine for an Aerospace Plane

Abstract: Operating conditions of a rocket-ramjet combined-cycle engine for a single-stage-to-orbit aerospace plane were studied. The engine was composed of an ejector-jet mode, a ramjet mode, a scramjet mode, and a rocket mode. Characteristics of the engine operating conditions were studied analytically. The thrust augmentation effect of the ejector-jet mode was found to be small at low subsonic speed and to increase with an increase of the flight Mach number. Study of the effective impulse function clarified that higher specific impulse was preferable in supersonic flight, whereas greater thrust coefficient was preferable in hypersonic flight. The mentioned characteristics were examined by simulation of engine operating in an aerospace plane flight. Transportation of a mass into orbit was compared among several engines with different combinations of thrust and specific impulse. The mass which could be carried into orbit was larger with a ramjet mode of higher specific impulse and with a scramjet mode of greater thrust.

Modeling Scramjet Flows with Variable Turbulent Prandtl and Schmidt Numbers

Abstract: A complete turbulence model, where the turbulent Prandtl and Schmidt numbers are calculated as part of the solution and where averages involving chemical source terms are modeled, is presented. The ability of avoiding the use of assumed or evolution Probability Distribution Functions (PDF’s) results in a highly ecient algorithm for reacting flows. The predictions of the model are compared with two sets of experiments involving supersonic mixing and one involving supersonic combustion. The results demonstrate the need for consideration of turbulence/chemistry interactions in supersonic combustion. In general, good agreement with experiment is indicated.

Test Gas Vitiation Effects in a Dual-Mode Scramjet Combustor

Abstract: An experimental study was conducted to characterize the influence of combustion air preheater major vitiate species (H 2 O and CO 2 ) on scramjet combustion. These species were added to an initially clean airflow that was supplied by an electrically heated facility. With dry air, the scramjet combustor operated in the supersonic mode at an equivalence ratio in the range of 0.25-0.32 and transitioned to dual mode over an equivalence ratio range of 0.35-0.37. At an equivalence ratio of 0.27, the combustor operated in the supersonic mode for three cases: 1) dry air, 2) air vitiated with 5% H 2 O by mole, and 3) air vitiated with 5% H 2 O and 2.5% CO 2 by mole. In the second case, the combustor pressure distribution decreased 10% relative to dry air and, in the third case, another 2% decrease was measured. At an equivalence ratio of 0.35, the combustor operated in the dual mode with dry air, but in the supersonic mode with 7 % H 2 O. This is the first demonstration of mode transition solely caused by test gas vitiation. It is therefore important to account for such effects when extrapolating from vitiated ground testing to flight.

Numerical Study of Mixing Enhancement in a Supersonic Round Jet

Abstract: Experiments and computations are used to investigate the mixing enhancement of an underexpanded sonic jet subjected to radial secondary injections in a realistic aircraft-afterbody configuration. In accordance with more academic studies, the present study confirms that secondary jets lead to a strong distortion of the jet interface under the action of the longitudinal vortices. Reynolds-averaged Navier-Stokes computations are validated on experimental data and then used to perform a parametric study. We investigate the influence of parameters such as the number of active injectors, their aspect ratio, and the total pressure and temperature of the injected air. These computations also allow us to investigate the drastic changes in the compressible structure of the flow due to the action of the secondary injection.

Climate impact of supersonic air traffic: an approach to optimize a potential future supersonic fleet – results from the EU-project SCENIC

Abstract: The demand for intercontinental transportation is increasing and people are requesting short travel times, which supersonic air transportation would enable. However, besides noise and sonic boom issues, which we are not referring to in this investigation, emissions from supersonic aircraft are known to alter the atmospheric composition, in particular the ozone layer, and hence affect climate significantly more than subsonic aircraft. Here, we suggest a metric to quantitatively assess different options for supersonic transport with regard to the potential destruction of the ozone layer and climate impacts. Options for fleet size, engine technology (nitrogen oxide emission level), cruising speed, range, and cruising altitude, are analyzed, based on SCENIC emission scenarios for 2050, which underlay the requirements to be as realistic as possible in terms of e.g., economic markets and profitable market penetration. This methodology is based on a number of atmosphere-chemistry and climate models to reduce model dependencies. The model results differ significantly in terms of the response to a replacement of subsonic aircraft by supersonic aircraft, e.g., concerning the ozone impact. However, model differences are smaller when comparing the different options for a supersonic fleet. Those uncertainties were taken into account to make sure that our findings are robust. The base case scenario, where supersonic aircraft get in service in 2015, a first fleet fully operational in 2025 and a second in 2050, leads in our simulations to a near surface temperature increase in 2050 of around 7 mK and with constant emissions afterwards to around 21 mK in 2100. The related total radiative forcing amounts to 22 mW m 2 in 2050, with an uncertainty between 9 and 29 mW m 2 . A reduced supersonic cruise altitude or speed (from Mach 2 to Mach 1.6) reduces both, climate impact and ozone destruction, by around 40%. An increase in the range of the supersonic aircraft leads to more emissions at lower latitudes since more routes to SE Asia are taken into account, which increases ozone depletion, but reduces climate impact compared to the base case.