Showing papers on "Supersonic speed published in 2005"

Noise Prediction for Increasingly Complex Jets. Part II: Applications:

Abstract: The numerical system described in Part I (Ref. 1) is applied to a variety of cases which increase difficulty, and progress in the direction of the complete simulation of an airliner engine. The grids have on the order of 1 million points. In many cases, the system meets the 2-3 dB accuracy target both in terms of directivity and of spectrum, up to a Strouhal number of about 1.5. The jet Mach number is varied from 0.3 to slightly supersonic with under-expansion, generating shock cells and greatly increasing side-line noise. For heated jets, the cross-effect between the acoustic Mach number and the temperature is correctly reproduced. Jets placed in a co-flowing stream with velocity up to 60% of the jet's are studied and found to sustain natural transition without unsteady forcing; the noise trends are correct. Finally, "synthetic chevrons" are added by altering the inflow conditions, and found to reduce low-frequency noise while increasing mid-frequency noise. In total, about fifteen meaningfully different...

An experimental study of the oscillatory flow structure of tone-producing supersonic impinging jets

Abstract: An experimental investigation into the structure of a supersonic jet impinging on a large plate is presented. Digital particle image velocimetry (DPIV), shadowgraph photography and acoustic measurements are used to understand the relationship between the unsteady jet structure and the production of tones for nozzle-to-plate spacings between 1 and 5 nozzle exit diameters at a nozzle–pressure ratio equal to 4. Results indicate that the instability of the jet depends on the location of the plate in the shock cell structure of the corresponding free jet and the strength of the standoff shock wave, rather than on the occurrence of recirculation zones in the impingement region. Phase-locked studies show streamwise displacements of the stand-off shock wave, a moving recirculation zone in the subsonic flow in front of the plate, and significant oscillations of both the compression and expansion regions in the peripheral supersonic flow when tones are produced. Sound is shown to be generated by periodic pulsing of the wall jet boundary resulting from periodic motion of the flow in the impingement and near-wall regions of the flow.

Hypersonic Boundary Layer Stability Analysis Using PSE-Chem

Abstract: In this paper we introduce the boundary layer stability equation solver PSE-Chem. This software employs a physical model approach to studying supersonic and hypersonic boundary layers and it incorporates the relevant physics including thermal and chemical nonequilibrium eects. Reentry F, a Mach 20 5 cone with very small nose bluntness, is used as a test case to determine the N factor at transition. Our results are compared with published results from other researchers. Results are also presented which show, using the e N method, the stabilizing eect of increasing nose radius under these conditions. The sensitivity of the methods is illustrated in the ability to calculate how small nose radius changes aect the transition location at stations hundreds of nose radii downstream.

Stability limits of cavity-stabilized flames in supersonic flow

Abstract: Experiments were performed to examine the stability of hydrocarbon-fueled flames in cavity flameholders in supersonic airflows. Methane and ethylene were burned in two different cavity configurations having aft walls ramped at 22.5° and 90°. Air stagnation temperatures were 590 K at Mach 2 and 640 K at Mach 3. Lean blowout limits showed dependence on the air mass flowrates, cavity geometry, fuel injection scheme, Mach number, and fuel type. Large differences were noted between cavity floor and cavity ramp injection schemes. Visual observations, planar laser-induced fluorescence of nitric oxide, and shadowgraph imaging were used to investigate these phenomena. Cavity ramp injection provided better performance near the lean blowout limit, whereas injection from the cavity floor resulted in more stable flames near the rich limit. Ethylene flames have a wider range of stable operations than methane in all conditions. Lean blowout limits were not significantly different between the Mach 2 and Mach 3 cases at the lean limit; however, variation in Mach number had a measurable effect near the rich limit. Fuel flowrates at ignition were much greater than the lean blowout limit, but showed similar dependence on air mass flowrate.

Receptivity of a supersonic boundary layer over a flat plate. Part 3. Effects of different types of free-stream disturbances

Abstract: Supersonic boundary-layer receptivity to different types of free-stream disturbance is studied for a Mach 4.5 boundary-layer flow over a flat plate by using the approaches of both direct numerical simulation and linear stability theory. This paper is Part 3 of a three-part study of the receptivity of supersonic boundary layers to free-stream disturbances. The present paper investigates receptivity to four types of different free-stream disturbances, i.e. slow and fast acoustic waves, entropy waves, and vorticity waves. A high-order shock-fitting scheme is used in the numerical simulation in order to accurately account for the effects of interactions between free-stream disturbance waves and the oblique shock wave. Numerical results on the generation of fast acoustic waves by free-stream entropy waves or vorticity waves are compared with those of a linear theory. Good agreement is obtained in both wave angles and amplitudes immediately behind the bow shock. It is found that the second-mode receptivity to free-stream slow acoustic waves is several times stronger than that to free-stream fast acoustic waves. This is because free-stream slow acoustic waves can directly induce and interact with the first and second Mack modes, while free-stream fast acoustic waves cannot. Instead, the free-stream fast acoustic waves can only induce and interact with stable mode I waves, which in turn induce unstable Mack modes. In the cases of receptivity to free-stream entropy waves and vorticity waves, it is found that the oblique shock wave created by the displacement of the boundary layer plays an important role because boundary-layer disturbances are mainly induced by fast acoustic waves generated behind the shock by free-stream forcing waves. As a result, mechanisms of the receptivity to free-stream entropy and vorticity waves are very similar to those of the receptivity to free-stream fast acoustic waves.

Penetration of a Transverse Supersonic Jet into a Subsonic Compressible Crossflow

Abstract: Particle image velocimetry data have been acquired in the far field of the interaction generated by an overexpanded axisymmetric supersonic jet exhausting transversely from a flat plate into a subsonic compressible crossflow. Mean velocity fields were found in the streamwise plane along the flowfield centerline for different values of the crossflow Mach number M∞ and the jet-to-freestream dynamic pressure ratio J. The magnitude of the streamwise velocity deficit and the vertical velocity component both decay with downstream distance and were observed to be greater for larger J while M∞ remained constant. Jet trajectories derived independently using the maxima of each of these two velocity components are not identical, but show increasing jet penetration for larger J. Similarity in the normalized velocity field was found for constant J at two different transonic M∞ ,b utat two lower M∞ the jet appeared to interact with the wall boundary layer and data did not collapse. The magnitude and width of the peak in the vertical velocity component both increase with J, suggesting that the strength and size of the counter-rotating vortex pair increase and, thus, may have a stronger influence on aerodynamic surfaces despite further jet penetration from the wall.

Transonic shock in a nozzle I: 2D case

Abstract: In this paper we establish the existence and uniqueness of a transonic shock for the steady flow through a general two-dimensional nozzle with variable sections. The flow is governed by the inviscid potential equation and is supersonic upstream, has no-flow boundary conditions on the nozzle walls, and an appropriate boundary condition at the exit of the exhaust section. The transonic shock is a free boundary dividing two regions of C 1,1-δ 0 flow in the nozzle. The potential equation is hyperbolic upstream where the flow is supersonic, and elliptic in the downstream subsonic region. In particular, our results show that there exists a solution to the corresponding free boundary problem such that the equation is always subsonic in the downstream region of the nozzle when the pressure in the exit of the exhaustion section is appropriately larger than that in the entry. This problem is motivated by the conjecture of Courant and Friedrichs on the transonic phenomena in a nozzle [10]. Furthermore, the stability of the transonic shock is also proven when the upstream supersonic flow is a small steady perturbation for the uniform supersonic flow and the corresponding pressure at the exit has a small perturbation. The main ingredients of our analysis are a generalized hodograph transformation and multiplier methods for elliptic equation with mixed boundary conditions and comer singularities.

Low-temperature supersonic boundary layer control using repetitively pulsed magnetohydrodynamic forcing

Abstract: The paper presents results of magnetohydrodynamic (MHD) supersonic boundary layer control experiments using repetitively pulsed, short-pulse duration, high-voltage discharges in M=3 flows of nitrogen and air in the presence of a magnetic field of B=1.5T. We also have conducted boundary layer flow visualization experiments using laser sheet scattering. Flow visualization results show that as the Reynolds number increases, the boundary layer flow becomes much more chaotic, with the spatial scale of temperature fluctuations decreasing. Combined with density fluctuation spectra measurements using laser differential interferometry (LDI) diagnostics, this behavior suggests that boundary layer transition occurs at stagnation pressures of P0∼200–250Torr. A crossed discharge (pulser+dc sustainer) in M=3 flows of air and nitrogen produced a stable, diffuse, and uniform plasma, with the time-average dc current up to 1.0A in nitrogen and up to 0.8A in air. The electrical conductivity and the Hall parameter in these f...

Numerical study of thermal protection system by opposing jet

Abstract: A numerical study on a reduction of aerodynamic heating by opposing jet in supersonic flow has been conducted. Flow field around a hemisphere model is calculated in supersonic free stream of Mach number 3.96 and the coolant gas is injected through the sonic nozzle at the nose of the model. In numerical analysis, axisymmetric Navier-Stokes equations are solved by an implicit finite difference method, and k-ω turbulence model is used. Significant decreases of surface heat flux are observed and opposing jet is proved to be effective on aerodynamic heating reduction around a stagnation region of the blunt body. In our previous study, experiments on the reduction of aerodynamic heating by opposing jet in supersonic flow were conducted in wind tunnel. The result of numerical analysis shows good agreement with experiment. Numerical results show that the recirculation region plays an important role for the reduction of heat flux. For the reduction of the aerodynamic heating, it is effective to cover the body surface with the cool jet flow and to form strong circulation region.

Navier-Stokes Computations for a Spinning Projectile from Subsonic to Supersonic Speeds

Abstract: A computational study has been undertaken to predict the static-aerodynamic, Magnus-moment, and roll-damping coefficients of a standard spinning projectile using a single, modern, unstructured Navier-Stokes flow solver. Numerical results without engraving and semi-empirical results have been obtained for a wide range of Mach numbers to include subsonic, transonic, and supersonic flight regimes. Effects of 0-, 2-, and 5-deg angles of attack have been investigated. Flowfield characteristics of each flight regime are briefly explored. A comparison of coefficients calculated from the computational fluid dynamics results are made to both experimental range data as well as semi-empirical aeroprediction code results with some success. Good predictive capabilities are found for the static aerodynamic coefficients throughout all of the flight regimes. Discrepancies arise between the computational results and the experimental results for the Magnus moment and roll-damping coefficients due in part to the lack of engraving on the computational model.

Time-Accurate Numerical Prediction of Free-Flight Aerodynamics of a Finned Projectile

Abstract: This article describes a new multidisciplinary computational study undertaken to compute the flight trajectories and simultaneously predict the unsteady free-flight aerodynamics of a finned projectile configuration. Actual flight trajectories are computed using an advanced coupled computational fluid dynamics/rigid body dynamics technique in a body-fixed coordinate system. An advanced time-accurate Navier–Stokes computational technique has been used in computational fluid dynamics to compute the unsteady aerodynamics associated with the free flight of the finned projectile at supersonic speeds. Computed positions and orientations of the projectile have been compared with actual data measured from free-flight tests and are found to be generally in good agreement with the data. Predicted aerodynamics forces and moments also compare well with the forces and moments used in the 6 degree of freedom fits of the results of the same tests. Unsteady numerical results obtained from the coupled method show the flowfield, the aerodynamic coefficients, and the flight paths of the projectile.

Experimental Study of Strut Injectors in a Supersonic Combustor Using OH-PLIF

Abstract: The present experimental study, carried out in the framework of ONERA-JAXA cooperation, deals with the ignition and hydrogen flame development in a supersonic air flow at Mach 2.5, with stagnation conditions of 0.6 MPa and 1620 K. Different types of struts for hydrogen injection (equivalence ratio of 0.45) have been investigated. Some (Alternating-Wedge struts) are specially designed to enhance fuel-air mixing by the creation of streamwise vortices. The principal aim of this study is to obtain information on the flow structures in the supersonic combusting mixing layer, and the location of ignition and combustion zones. Instantaneous maps of the OH radical by means of the OH-PLIF technique in two planes downstream of the strut injector and conventional observations of the flame by video camera have been obtained. These visualizations are completed by global measurements of the ignition delays and heat release by means of wall pressure measurements. The efficiency of the different strategies of fuel injections are then compared and discussed.

Improved Sonic Boom Minimization with Extendable Nose Spike

Abstract: Marketing surveys of traditional business aircraft operators have shown that there is significant interest customer interest in the development of a supersonic business jet. One requirement for such an aircraft is that it be capable of unrestricted supersonic flight over land. This requirement has focused considerable attention on sonic boom minimization for a class of 100,000 pound, Mach 1.6 to 2.0 cruise airplanes. Configuring the design with sufficient low boom characteristics while at the same time including enough cabin volume has proven difficult. Gulfstream has developed a Quiet Spike concept that significantly improves the sonic boom signature while at the same time allowing for additional fuselage volume. The multi stage extendable nose spike projects out of, and retracts into, the forward fuselage of the airplane. Wind tunnel tests of supersonic equivalent area distributions have validated the near field aerodynamics of the spike concept. CFD analysis of a wing/body configured with a multi stage spike have extended the analysis beyond idealized equivalent area distributions and shown a great potential for tailoring the shape of the sonic boom ground signature while simultaneously allowing greater flexibility in the fuselage design.

Computation and flow visualization in high-speed aerodynamics

Abstract: This paper relates to numerical simulations and flow visualizations of internal and external high-speed aerodynamics. Fully compressible Navier–Stokes solvers using high-order shock-capturing schemes and turbulence models are developed for solving gas-dynamics problems. Numerical schlieren pictures, as well as computed interferogram techniques, are used to visualize the major features of physical phenomena occurring in such flows. A variety of test problems encountered in supersonic flows, such as supersonic turbulence including shock/shock and shock/boundary layer interactions, shear-layer instability and transient flows are considered. Some of the numerical visualization results constructed from computed Navier–Stokes flow-fields are directly compared with experimental images. Most of the features observed in the experiment are accurately reproduced by the simulations. The results of this study provide a better understanding of the main characteristics of complex flows that are not easily accessible exp...

Turbulent Characteristics of a Transverse Supersonic Jet in a Subsonic Compressible Crossflow

Abstract: Fluctuating velocity fields have been determined from particle image velocimetry data acquired in the farfield of the interaction generated by an overexpanded axisymmetric supersonic jet exhausting transversely from a flat plate into a subsonic compressible crossflow. Peak magnitudes of the turbulent stresses were found to be larger and located further from the wall for greater values of the jet-to-freestream dynamic pressure ratio J while the crossflow Mach number M∞ remained constant. These stress magnitudes diminish with downstream distance as their peak location moves further from the wall. The vertical positions of the peak normal stresses and shear stress inflection point coincide with the maximum mean streamwise velocity deficit induced by the jet. Instantaneous realizations of the velocity fluctuation fields reveal large-scale structures whose mean diameter is greater for larger J and decreases with downstream distance. The integral length scale calculated from profiles of the correlation coefficient instead shows an increase downstream; the discrepancy between the two length scales results from a low-pass filter effect of the correlation coefficient. Similarity was demonstrated for constant J at two transonic M∞, but not at two lower M∞ where the flowfield does not distinguish between the jet and its wake.

Case for Small Supersonic Civil Aircraft

Abstract: Civil aviation progress in the last 40 years has included a significant expansion of the small civil aircraft market involving regional jets, business jets, and the emerging personal jets. A significant factor in the growth of the small civil aircraft market is the value of time. Recognition of the ever-increasing value of time has lead to increased interest in the feasibility of a small supersonic civil aircraft. The step to supersonic speeds offers the potential of a dramatic decrease in travel time. Feasibility studies of a small quiet supersonic jet (QSJ) have been conducted. Market research, environmental concerns, program and design requirements, and vehicle characteristics are summarized. Areas for concentrated future supersonic aeronautics research and development efforts are highlighted.

Large-Eddy Simulation of Transitional Boundary Layer with Impinging Shock Wave

Abstract: The transition of a boundary layer on a flat plate with an impinging shock wave is studied numerically by compressible large-eddy simulation using a hybrid compact/Roe scheme. The numerical code is verified by comparison with experimental observations of shock wave/turbulent boundary-layer interaction and then applied for the analysis of shock wave/transitional boundary-layer interaction. The simulation provides accurate results with respect to the location of reattachment and the boundary-layer properties downstream of reattachment. Large-scale coherent structures such as longitudinal vortex pairs, low-speed streaks, and hairpin vortices are identified in the transitional region, and it is revealed that these coherent structures play important roles in the transition. Thus, it is important to resolve these structures adequately to obtain an accurate prediction of the reattachment point. The subsequent breakdown of these coherent structures have smaller length scales, and resolving the breakdown requires much finer grid resolution. However, the influence of underresolution of breakdown on the downstream flowfield can be largely ignored under the conditions examined. Large-eddy simulation is, therefore, useful for the qualitative analysis of flowfields involving a supersonic transitional boundary layer.

Introduction to Molecular Beams Gas Dynamics

Abstract: Gas Properties The Non-Equilibrium Equations and the Relaxation of the Internal Degrees of Freedom The Fundamental Equations of Gas Dynamics Isoentropic Flow. Characteristic Lines The Method of Characteristics The Shock Waves The Flow in Nozzles and Jets The Supersonic Free Jet Application of the Boltzmann Equation to a Jet of Monoatomic Gas Characterization of a Particle Source and Extraction of the Molecular Beam The Condensation in a Supersonic Free Jet

Large eddy simulation of compressible turbulence using high-resolution methods

Abstract: We present a numerical investigation of high-resolution schemes for solving the compressible Euler and Navier-Stokes equations in the context of implicit large eddy simulation (ILES), also known as monotone integrated LES (MILES). We have employed three high-resolution schemes: a flux vector splitting (FVS), a characteristics-based (Godunov-type) and a hybrid total variation diminishing (TVD) scheme; and carried out computations of: (i) decaying turbulence in a triply periodic cube and (ii) compressible flow around open cavities for low and high Reynolds numbers, at transonic and supersonic speeds

An Experimental Study of Kerosene Combustion in a Supersonic Model Combustor Using Effervescent Atomization

Abstract: Investigation of kerosene combustion in a Mach 2.5 flow was carried out using a model supersonic combustor with cross-section area of 51 mm × 70 mm and different integrated fuel injector/flameholder cavity modules. Experiments with pure liquid atomization and with effervescent atomization were characterized and compared. Direct photography, Schlieren imaging, and planar laser induced fluorescence (PLIF) imaging of OH radical were utilized to examine the cavity characteristics and spray structure. Schlieren images illustrate the effectiveness of gas barbotage in facilitating atomization and the importance of secondary atomization when kerosene sprays interacting with a supersonic crossflow. OH PLIF images further substantiate our previous finding that there exists a local high-temperature radical pool within the cavity flameholder, and this radical pool plays a crucial role in promoting kerosene combustion in a supersonic combustor. Under the same operation conditions, comparison of the measured static pressure distributions along the combustor also shows that effervescent atomization generally leads to better combustion performance than the use of pure liquid atomization. Furthermore, the present results demonstrate that the cavity characteristics can be different in non-reacting and reacting supersonic flows. As such, the conventional definition of cavity characteristics based on non-reacting flows needs to be revised.

Preliminary Design of a 2D Supersonic Inlet to Maximize Total Pressure Recovery

Abstract: *† This paper provides a method of preliminary design for a two-dimensional, mixed compression, two-ramp supersonic inlet to maximize total pressure recovery and match the mass flow demand of the engine. For an on-design condition, the total pressure recovery is maximized according to the optimization criterion, and the dimensions of the inlet in terms of ratios to the engine face diameter are calculated. The optimization criterion is defined such that in a system of (n-1) oblique shocks and one normal shock in two dimensions, the maximum shock pressure recovery is obtained when the shocks are of equal strength. This paper also provides a method to estimate the total pressure recovery for an off-design condition for the specified inlet configuration. For an off-design condition, conservative estimation of the total pressure recovery is given so that performance of the engine at the off-design condition can be estimated. To match the mass flow demand of the engine, the second ramp angle is adjusted and the open/close schedule of a bypass door is determined. The effects of boundary layer are not considered for the supersonic part of the inlet, however friction and expansion losses are considered for the subsonic diffuser. Nomenclature α = Angle of attack j β = The installation angle of the j th ramp γ = The ratio of specific heats j δ = The flow deflection angle of the j th shock (j th ramp half angle) d θ = The half expansion angle of the subsonic diffuser j θ = The shock wave angle of the j th shock * A = The cross section area of flow tube at throat where the flow is sonic j A = The cross section area of flow at j th station point 54 AR = The ratio of inlet cross section areas at station points 5 and 4 5 d = The engine diameter at station point 5 (engine face) 6 d = The engine diameter at station point 6 (fan face) H = Flight altitude c h , 0 h = The captured freestream flow tube height i h = The height of inlet at the entry, measured perpendicular to the flight direction j h = The height of j

Low-Temperature Supersonic Boundary Layer Control Using Repetitively Pulsed MHD Forcing

Abstract: The paper presents results of magnetohydrodynamic (MHD) supersonic boundary layer control experiments using repetitively pulsed, short pulse duration, high voltage discharges in M=3 flows of nitrogen and air in the presence of a magnetic field of B=1.5 T. We also have conducted boundary layer flow visualization experiments using laser sheet scattering. Flow visualization shows that side wall boundary layers in the supersonic test section are considerably thicker near the center plane of the flow. The results also show that as the Reynolds number increases from Rex=2.7·10 5 to 8.1·10 5 , the boundary layer flow becomes much more chaotic, with the spatial scale of temperature fluctuations decreasing. Combined with density fluctuation spectra measurements using Laser Differential Interferometry (LDI) diagnostics, this behavior suggests that boundary layer transition occurs at stagnation pressures of P0~200-250 torr. Operation of a crossed discharge (pulser + DC sustainer) in M=3 flows of air and nitrogen demonstrated that such a discharge produces a stable, diffuse, and uniform plasma. The time-average DC current achieved in such discharges is up to 1.0 A in nitrogen (conductivity of σ=0.073 mho/m) and up to 0.8 A in air (σ=0.072 mho/m). The electrical conductivity and the Hall parameter in nitrogen and air flows are inferred from the current voltage characteristics of the sustainer discharge. LDI measurements detected MHD effect on the ionized boundary layer density fluctuations at these conditions. Retarding Lorentz force applied to M=3 nitrogen, air, and N2-He flows produces an increase of the density fluctuation intensity by up to 2 dB (about 25%), compared to the accelerating force of the same magnitude. The effect is demonstrated for two possible combinations of the magnetic field and current directions producing the same Lorentz force direction (both for accelerating and retarding force). Comparison with the LDI spectra measured with no MHD force applied showed that the effect on the density fluctuations is produced only by the retarding Lorentz force, while the Joule heat effect appears insignificant.

Structure of Supersonic Twin Jets

Abstract: The flow characteristics of coupled oscillatory motion of M j = 1.5 supersonic twin jets are described using the particle image velocimetry measurements. As a result of self-excitation caused by screech, sinuous oscillation of individual jets is observed. The coupling between the jets, characterized by the presence of coherent large-scale eddies, results in a symmetric mode with respect to the midplane separating the two jets. A control technique using microjets placed at the nozzle exit is used to suppress the unsteady flow

Single-Pulse Performance of the SparkJet Flow Control Actuator

Abstract: A need exists for a robust flow control actuator that is capable of penetrating supersonic boundary layers. Under Air Force Office of Scientific Research sponsorship, The Johns Hopkins University Applied Physics Laboratory is investigating a promising device for high-speed flow control called the SparkJet actuator. This actuator, which produces a synthetic jet with high exhaust velocities, holds the promise of manipulating high-speed flows without active mechanical components. Computational and experimental techniques are being used to investigate the operating characteristics of a cavity SparkJet device for control of supersonic flows over external surfaces. This paper focuses on the detailed characteristics of the SparkJet’s discharge and cooling stages after a single energy deposition pulse. Previous experimental and numerical parametric studies of the current design characterized the performance attributes of the device as a function of orifice size, chamber volume, and energy deposited. Current experimental efforts include the application of high-resolution particle image velocimetry to quantify quiescent air operation of the mm-scale device and to provide validation data for computational models. NOMENCLATURE