Showing papers on "Supersonic speed published in 2013"

On the universality of supersonic turbulence

Abstract: Compressible turbulence shapes the structure of the interstellar medium of our Galaxy and likely plays an important role also during structure formation in the early Universe The density PDF and the power spectrum of such compressible, supersonic turbulence are the key ingredients for theories of star formation However, both the PDF and the spectrum are still a matter of debate, because theoretical predictions are limited and simulations of supersonic turbulence require enormous resolutions to capture the inertial-range scaling To advance our limited knowledge of compressible turbulence, we here present and analyse the world's largest simulations of supersonic turbulence We compare hydrodynamic models with numerical resolutions of 256^3-4096^3 mesh points and with two distinct driving mechanisms, solenoidal (divergence-free) driving and compressive (curl-free) driving We find convergence of the density PDF, with compressive driving exhibiting a much wider and more intermittent density distribution than solenoidal driving Analysing the power spectrum of the turbulence, we find a pure velocity scaling close to Burgers turbulence with P(v) k^(-2) for both driving modes in our hydrodynamical simulations with Mach = 17 The spectrum of the density-weighted velocity rho^(1/3)v, however, does not provide the previously suggested universal scaling for supersonic turbulence We find that the power spectrum P(rho^(1/3)v) scales with wavenumber as k^(-174) for solenoidal driving, close to incompressible Kolmogorov turbulence, k^(-5/3), but is significantly steeper with k^(-210) for compressive driving We show that this is consistent with a recent theoretical model for compressible turbulence that predicts P(rho^(1/3)v) k^(-19/9) in the presence of a strong div(v) component as is produced by compressive driving and remains remarkably constant throughout the supersonic turbulent cascade

Ship Wakes: Kelvin or Mach Angle?

Abstract: From the analysis of a set of airborne images of ship wakes, we show that the wake angles decrease as U(-1) at large velocities, in a way similar to the Mach cone for supersonic airplanes. This previously unnoticed Mach-like regime is in contradiction with the celebrated Kelvin prediction of a constant angle of 19.47° independent of the ship's speed. We propose here a model, confirmed by numerical simulations, in which the finite size of the disturbance explains this transition between the Kelvin and Mach regimes at a Froude number Fr=U/√[gL]~/=0.5, where L is the hull ship length.

Progress in research on mixing techniques for transverse injection flow fields in supersonic crossflows

Abstract: The transverse injection flow field has an important impact on the flowpath design of scramjet engines. At present a combination of the transverse injection scheme and any other flame holder has been widely employed in hypersonic propulsion systems to promote the mixing process between the fuel and the supersonic freestream; combustion efficiency has been improved thereby, as well as engine thrust. Research on mixing techniques for the transverse injection flow field is summarized from four aspects, namely the jet-to-crossflow pressure ratio, the geometric configuration of the injection port, the number of injection ports, and the injection angle. In conclusion, urgent investigations of mixing techniques of the transverse injection flow field are proposed, especially data mining in the quantitative analytical results for transverse injection flow field, based on results from multi-objective design optimization theory.

Predictions of a Supersonic Turbulent Flow in a Square Duct

Abstract: In this paper we have confirmed that eddy viscosity turbulence models are inadequate to predict secondary vortical flows developed from corners in internal flows. To remedy this shortcoming, we have added the quadratic constitutive relation (QCR) of Spalart to the one- and two-equation SA and SST turbulence models, respectively. The results of QCR with the SA and SST turbulence models have been validated against experimental data of Davis and Gessner for supersonic flow through a square duct. The approach is shown to be simple to implement and overall agreement is seen to improve with the use of QCR. Introduction: Supersonic flow through a square duct is altered by secondary vortical flow developing from the corners. These secondary flows are generated by Reynolds stress gradients acting in the corner region and appear to have similar structure to those found in subsonic flow through square ducts (1-3). Such flow is representative of various airplane inlets and therefore it is important to understand and predict the impact of this secondary flow on inlet pressure recovery and distortion. An experimental study was performed at the University of Washington to gain a better understanding of how the secondary flow associated with corners affects local flow conditions in a square duct over the development length (3). This configuration was utilized in this study to validate and improve the RANS turbulence models utilized for corner flows at Boeing. Accurate prediction of the flow around and through the aircraft is essential for design improvement, risk mitigation, and wind-tunnel and flight test reduction. At Boeing, the RANS approach is used routinely for design and analysis of our configurations. While RANS has proven to be a powerful approach for flow-field prediction, it still has some shortcomings even for steady state flow predictions. The one- and two-equation SA (4) and SST (5) eddy viscosity turbulence models, respectively, are still the workhorses for routine applications. The Reynolds

Shock induced fluid structure interaction on a flexible wall in supersonic turbulent flow

Abstract: A detailed knowledge of the flow structure interaction in supersonic flows is important for the design of future space transportation systems. Therefore this work was devoted to the investigation of the shock wave boundary layer interaction on an elastic panel. During the wind tunnel experiments the panel deflection was measured with fast non-intrusive displacement sensors. On the flow side pressure, high-speed Schlieren photography and oil-film technique were used. The flow manipulation due to the panel deflection becomes manifest in a deformation of the impinging shock and the separation zone. The panel deflection consists of a constant and a dynamic component. The experimental results are discussed and compared to numerical results.

Control of stationary cross-flow modes in a Mach 3.5 boundary layer using patterned passive and active roughness

Abstract: Spanwise-periodic roughness designed to excite selected wavelengths of stationary cross- ow modes was investigated in a 3-D boundary layer at Mach 3.5. The test model was a sharp-tipped 14deg right-circular cone. The model and integrated sensor traversing system were placed in the Mach 3.5 Supersonic Low Disturbance Tunnel (SLDT) equipped with a "quiet design" nozzle at the NASA Langley Research Center. The model was oriented at a 4:2deg angle of attack to produce a mean cross-fl ow velocity component in the boundary layer over the cone. Five removable cone tips have been investigated. One has a smooth surface that is used to document the baseline ("natural") conditions. Two had minute (20 - 40 micron) "dimples" that are equally spaced around the circumference, at a streamwise location that is just upstream of the linear stability neutral growth branch for cross- ow modes. The azimuthal mode numbers of the dimpled tips were selected to either enhance the most amplified wave numbers, or to suppress the growth of the most amplified wave numbers. Two of the cone tips had an array of plasma streamwise vortex generators that were designed to simulate the disturbances produced by the passive patterned roughness. The results indicate that the stationary cross-fl ow modes were highly receptive to the patterned roughness of both passive and active types. The patterned passive roughness that was designed to suppress the growth of the most amplified modes had an azimuthal wavelength that was 66% smaller that that of the most amplified stationary cross- ow mode. This had the effect to increase the transition Reynolds number from 25% to 50% depending on the measurement technique. The application of the research is on turbulent transition control on swept wings of supersonic aircraft. The plasma-based roughness has the advantage over the passive roughness of being able to be adaptable to different conditions that would occur during a flight mission.

The in-gas-jet laser ion source: resonance ionization spectroscopy of radioactive atoms in supersonic gas jets

Abstract: New approaches to perform efficient and selective step-wise resonance ionization spectroscopy (RIS) of radioactive atoms in different types of supersonic gas jets are proposed. This novel application results in a major expansion of the in-gas laser ionization and spectroscopy (IGLIS) method developed at KU Leuven. Implementation of resonance ionization in the supersonic gas jet allows to increase the spectral resolution by one order of magnitude in comparison with the currently performed in-gas-cell ionization spectroscopy. Properties of supersonic beams, obtained from the de Laval-, the spike-, and the free jet nozzles that are important for the reduction of the spectral line broadening mechanisms in cold and low density environments are discussed. Requirements for the laser radiation and for the vacuum pumping system are also examined. Finally, first results of high-resolution spectroscopy in the supersonic free jet are presented for the 327.4 nm 3d104s 2S1/2→ 3d104p 2P1/2 transition in the stable 63Cu isotope using an amplified single mode laser radiation.

Characteristics of Oscillations in Supersonic Open Cavity Flows

Abstract: Characteristics of oscillations in supersonic open cavity flows are investigated numerically using hybrid RANS/LES (Reynolds-Averaged Navier-Stokes/Large Eddy Simulation) method. The oscillation regimes and feedback mechanisms for the supersonic cavity flows are identified and analyzed. The calculation captures a mixed shear-layer/wake oscillation mode in the flow of Ma = 1.75, where these two modes occur alternately. The shear-layer mode and wake mode are driven by vortex convection-acoustic feedback and absolute instability, respectively. In particular, the results indicate that the feedback-acoustic-wave in the shear-layer mode is probably generated by the reflection of the downstream-traveling pressure wave, associated with the shed vortex in the shear layer, on the aft wall. The cavity flow of Ma = 2.52 is then simulated to see the influence of Mach number. It is found that the increase of Mach number may decrease the amplitude of the fluctuations in the shear layer, inhibiting the transition to wake mode. Furthermore, the influence of upstream injection is also studied, where the results show that the injection only weakens the oscillations and faintly shifts the resonant frequencies.

On the Problems of Chaplygin for Mixed Sub-and Supersonic Flows

Abstract: There are investigated the problems of the flow of a supersonic jet out of a vessel with plane side walls and the problem of the supersonic flow about a wedge when there is a zone of local subsonic velocities ahead of the wedge.

Highly focused supersonic microjets: numerical simulations

Abstract: By focusing a laser pulse inside a capillary partially filled with liquid, a vapour bubble is created that emits a pressure wave. This pressure wave travels through the liquid and creates a fast, focused axisymmetric microjet when it is reflected at the meniscus. We numerically investigate the formation of this microjet using axisymmetric boundary integral simulations, where we model the pressure wave as a pressure pulse applied on the bubble. We find a good agreement between the simulations and experimental results in terms of the time evolution of the jet and on all parameters that can be compared directly. We present a simple analytical model that accurately predicts the velocity of the jet after the pressure pulse and its maximum velocity.

Energy cascade and scaling in supersonic isothermal turbulence

Abstract: Supersonic turbulence plays an important role in a number of extreme astrophysical and terrestrial environments, yet its understanding remains rudimentary. We use data from a three-dimensional simulation of supersonic isothermal turbulence to reconstruct an exact fourth-order relation derived analytically from the Navier–Stokes equations (Galtier & Banerjee, Phys. Rev. Lett., vol. 107, 2011, p. 134501). Our analysis supports a Kolmogorov-like inertial energy cascade in supersonic turbulence previously discussed on a phenomenological level. We show that two compressible analogues of the four-fifths law exist describing fifth- and fourth-order correlations, but only the fourth-order relation remains ‘universal’ in a wide range of Mach numbers from incompressible to highly compressible regimes. A new approximate relation valid in the strongly supersonic regime is derived and verified. We also briefly discuss the origin of bottleneck bumps in simulations of compressible turbulence.

General Characteristics of the Flow Through Nozzles at Near Critical Speeds

Abstract: The characteristics of the position and form of the transition surface through the critical velocity are computed for flow through flat and round nozzles from subsonic to supersonic velocity. Corresponding considerations were carried out for the flow about profiles in the vicinity of sonic velocity.

Turbulent boundary layer noise : direct radiation at Mach number 0.5

Abstract: Boundary layers constitute a fundamental source of aerodynamic noise. A turbulent boundary layer over a plane wall can provide an indirect contribution to the noise by exciting the structure and a direct noise contribution. The latter part can play a significant role even if its intensity is very low, explaining why it is difficult to measure. In the present study, the aerodynamic noise generated by a spatially developing turbulent boundary layer is computed directly by solving the compressible Navier–Stokes equations. This numerical experiment aims at giving some insight into the noise radiation characteristics. The acoustic wavefronts have a large wavelength and are oriented in the direction opposite to the flow. Their amplitude is only 0.7 % of the aerodynamic pressure for a flat-plate flow at Mach 0.5. The particular directivity is mainly explained by convection effects by the mean flow, giving an indication about the compactness of the sources. These vortical events correspond to low frequencies and thus have a large lifetime. They cannot be directly associated with the main structures populating the boundary layer such as hairpin or horseshoe vortices. The analysis of the wall pressure can provide a picture of the flow in the wavenumber–frequency space. The main features of wall pressure beneath a turbulent boundary layer as described in the literature are well reproduced. The acoustic domain, corresponding to supersonic wavenumbers, is detectable but can hardly be separated from the convective ridge at this relatively high speed. This is also due to the low frequencies of sound emission as noted previously.

New Method for Computing Performance of Choked Reacting Flows and Ram-to-Scram Transition

Abstract: An improved method has been developed to compute the thrust of a dual-mode scramjet, which is an engine with a combustor that operates both subsonically and supersonically This strategy applies to any internal flow that can be modeled one-dimensionally To handle the mathematical singularity at the location of thermal choking, the simple Shapiro method is expanded to create a new method that includes finite-rate chemistry and high-temperature gas properties A forward shooting method is employed to find appropriate initial conditions for integration of the governing equations, which results in a unique transonic (choked) condition capable of reaching a supersonic state at the end of the domain Solutions of the governing equations are computed using the propulsion code MASIV, which has been integrated into a hypersonic vehicle flight dynamics code Computations for both ram-mode and scram-mode operations are compared to experimental results Predictions are made for flight conditions of a hypersonic vehi

Velocity and Temperature Measurement in Supersonic Free Jets Using Spectrally Resolved Rayleigh Scattering

Abstract: The flow fields of unheated, supersonic free jets from convergent and convergent-divergent nozzles operating at M = 0.99, 1.4, and 1.6 were measured using spectrally resolved Rayleigh scattering technique. The axial component of velocity and temperature data as well as density data obtained from a previous experiment are presented in a systematic way with the goal of producing a database useful for validating computational fluid dynamics codes. The Rayleigh scattering process from air molecules provides a fundamental means of measuring flow properties in a non-intrusive, particle free manner. In the spectrally resolved application, laser light scattered by the air molecules is collected and analyzed using a Fabry-Perot interferometer (FPI). The difference between the incident laser frequency and the peak of the Rayleigh spectrum provides a measure of gas velocity. The temperature is measured from the spectral broadening caused by the random thermal motion and density is measured from the total light intensity. The present point measurement technique uses a CW laser, a scanning FPI and photon counting electronics. The 1 mm long probe volume is moved from point to point to survey the flow fields. Additional arrangements were made to remove particles from the main as well as the entrained flow and to isolate FPI from the high sound and vibration levels produced by the supersonic jets. In general, velocity is measured within +/- 10 m/s accuracy and temperature within +/- 10 K accuracy.

Dynamic Characteristics of Combustion Mode Transitions in a Strut-Based Scramjet Combustor Model

Abstract: T HE supersonic combustion ramjet (scramjet) engine is expected to be the most efficient propulsion system in the hypersonic flight regime [1]. Given the broad range of aerothermodynamic conditions experienced during hypersonic flight, the scramjet would operate under different combustionmodes [2], andmode transition is a critical phenomenon in designing such engines because the thrust and specific impulse of the fuel in each mode varies considerably. In much of the previous work, researchers experimentally achievedmode transition and investigated the static characteristics of different combustion modes. In the open literature, Billig [3] first demonstrated mode transition in ground tests. Heiser and Pratt [4] used a one-dimensional (1-D) analysis approach to comprehend the complex aerothermodynamics of a dual-mode combustion system. The flowfield can be illustrated for threemodes: scramjet with shockfree isolator and oblique shock train, and ramjet with normal shock train. Takahashi et al. [5] and Kouchi et al. [6] observed four different combustion modes with respect to the fuel flow rate, namely, blowout,weak combustion, strong combustion, and thermal choking. As the mode transition occurred, thrust and heat-flux distribution [7] varies considerably. Sullins [8] experimentally achieved the mode transition from a scramjet with a precombustion shock system having a high pressure ratio to a scramjet with no precombustion shock system by increasing the total temperature of airflow to simulate a real acceleration process. Micka and Driscoll [9] reported two distinct reaction zones in a combustor with wall injection and a cavity flameholder corresponding to jet wake stabilization and cavity stabilization. The reaction zonewas found to only appear in the cavity stabilized mode in the scramjet mode, even for conditions where the ramjet modewas jet-wake stabilized. Also, the spreading of scramjet mode combustion is significantly less than that of the ramjet mode. Masumoto et al. [10] investigated the effect of combustor length and total temperature on combustion modes and suggested the minimum combustor length to attain supersonic or dual-mode combustion. However, there have been few studies on the dynamic characteristics of combustion mode transition, and the open literature did not fully investigate the combustor performance changes with the fuel flow rate small changes (∼1 g∕s) near the critical conditions. One interesting phenomenon, rather different from the results available in the open literature, is that the wall pressure and thrust show obvious catastrophe near the critical point of combustion mode transition. The combustion mode transition depends on the path taken (i.e. the fuel flow rate is increasedordecreased).With the sameexternal parameters, the scramjet engine may be a different combustion mode, known as hysteresis effect according to the nonlinear dynamics theory. During hypersonic flight, itmay bringgreat difficulties to the precise control of the vehicle, have a great impact on the flight safety, and even cause a flight accident [11]. Therefore, the successful development of a scramjet engine depends on further understanding and control of the nonlinear mode transition process. In this research, particular attention was focused on the dynamic characteristics of combustion mode through ground tests, especially the nonlinear catastrophic and hysteresis phenomena. As known, the transition between ramjet and scramjet mode is determined from the magnitude of ΔT0∕T0;air (either by decreasing or increasing the amount of heat release). In this paper, we linearly changed the fuel mass flow rate along two adverse paths; that is, increased and decreased the fuel equivalence ratiowhile the rate of changewas held approximately constant. In particular, to obtain performance of the model combustor around the critical conditions in detail, the heat release was changed little by little every time (corresponding to an increase in fuel equivalence ratio of 0.0125). Compared to strut injection in the center of the combustor, the transverse wall injection disturbs the boundary layer significantly. The wall injection plume forms a barrel shock, and induces a bow shock that leads to separation and the formation of a recirculation region in front of the wall injection location. These unnecessary disturbances make it difficult to determine the exact mode transition mechanism [12]. Therefore, a central strut injector has been employed, which also improved fuel mixing in the supersonic core stream and combustion performance in supersonic combustors. Because the liquid hydrocarbon fuel has greater fuel densities and endothermic cooling capabilities than hydrogen, particularly for hypersonic vehicles limited to Mach 8 flight, kerosene was used as the fuel in this research.

Large-Eddy Simulation of Broadband Unsteadiness in a Shock/Boundary-Layer Interaction

Abstract: The simulation of low-frequency unsteadiness in shock wave/turbulent boundary-layer interactions constitutes a challenging case insofar as very long time integrations are required to describe these broadband motions at frequencies two orders of magnitude lower than those of the turbulent motions. A relatively low-cost numerical strategy is established in the present study. The use of quasi-spectral centered finite differences in conjunction with high-order selective filtering provides an efficient method for compressible large-eddy simulations based on explicit filtering regularization. This strategy is extended to flows containing discontinuities by switching between the high-order filter used in regular zones and a low-order filter acting selectively near the shock locations. The accuracy of the current strategy is assessed for a developing turbulent supersonic boundary layer. The case of an oblique shock wave impinging on a flat plate is then successfully validated against previous experimental and num...

Alignment and scaling of large-scale fluctuations in the solar wind.

Abstract: Introduction. The solar wind is a hot, tenuous plasma that flows away from the Sun at supersonic speeds. Turbulence transports energy from the driving ‘outer’ scale to smaller scales via non-linear magnetohydrodynamic (MHD) interactions of magnetic B and velocity V fields, until kinetic effects and dissipation become important close to the ion gyroscale. In fast solar wind (|V | > 600 km/s), a ‘1/f’ scaling of magnetic-field power spectra is observed at low spacecraft frequencies, f [1–3]. Slowly evolving structures are advected at supersonic speeds past spacecraft, so the observed spacecraft frequency of a fluctuation is proportional to its characteristic wave number (scale) k [4]. The energy spectrum in the 1/f range is, therefore, expected to scale as E(k) ∝ k −1 . A steeper spectrum close to k −5/3 associated with turbulence is observed at higher spacecraft frequencies in the ‘inertial range’ and there is a spectral break between the two regimes [2, 3, 5, 6]; at 1 AU this typically occurs at f ∼ 10 −3 Hz. Studies have shown [2, 5, 7– 10] that the power spectral density of fluctuations in the low-frequency band decreases with distance from the Sun as R −3 , consistent with these scales containing non

Experimental characterization of the stagnation layer between two obliquely merging supersonic plasma jets.

Abstract: We present spatially resolved measurements characterizing the stagnation layer between two obliquely merging supersonic plasma jets. Intrajet collisionality is very high, but the interjet ion-ion mean free path is of the order of the stagnation layer thickness of a few centimeters. Fast-framing camera images show a double-peaked emission profile transverse to the stagnation layer, with the central emission dip consistent with a density dip in the interferometer data. We demonstrate that our observations are consistent with collisional oblique shocks.

Stabilization of a hypersonic boundary layer using a wavy surface

Abstract: Stability of a supersonic near-wall flow over a shallow grooved plate in the freestream of Mach 6 is investigated by means of numerical simulations and wind-tunnel experiments. Numerical solutions of two-dimensional Navier–Stokes equations are used to model propagation of artificial disturbances of several fixed frequencies generated by an actuator placed on the wall. It is shown that the high-frequency forcing excites unstable waves in the flat-plate boundary layer. These waves are relevant to the second-mode instability. The wavy wall damps the disturbances in a high-frequency band while it enhances them at lower frequencies. Stability experiments are conducted in the Institute of Theoretical and Applied Mechanics Tranzit-M shock tunnel under natural freestream conditions. The measured disturbance spectra are similar to those predicted numerically. They contain a peak associated with the second-mode instability. This peak is damped by the wavy wall, while a marginal increase of the disturbance amplitude...

Sub-Kelvin Collision Temperatures in Merged Neutral Beams by Correlation in Phase-Space

Abstract: Recent merged neutral beam experiments have introduced the possibility of measuring reactive collisions in the cold regime down to 10 mK. The lowest temperature attained in these experiments cannot be explained using the standard formalism developed for crossed molecular beam scattering. These low temperatures become accessible because pulsed supersonic beams develop a correlation in velocity-position space during free propagation such that the local velocity standard deviation decreases. This effect is responsible for a reduction in the attainable collision energy by more than 2 orders of magnitude along with an order of magnitude improvement in the resolution. We show that supersonic nozzles with short pulsed opening durations compared to the time-of-flight, such as the Even-Lavie valve, have a clear advantage in achieving low collision energies with improved resolution. We discuss possible improvements in the energy resolution by varying the detection time duration.

Modeling of Fuel Sloshing and its Physical Effects on Flutter

Abstract: The sloshing effects of an internal fluid on the flutter envelope of an aeroelastic system have received little attention in the open literature. This issue is nevertheless relevant for many aircraft, especially high-performance fighter jets carrying stores. This paper addresses some aspects of this problem as well as related modeling and analysis issues. These include the importance or insignificance of accounting for the hydroelastic effect when modeling an internal fluid and its container as well as accounting for that container when modeling the aerodynamics of the overall aeroelastic system. The paper also reports on the findings of four independent sets of flutter analyses performed for a wing–store test configuration and various fuel fill levels in the subsonic, transonic, and early supersonic regimes. Two of these sets of numerica l experiments relied on a computational-fluid-dynamics-based computational technology, and two of them on the doublet-lattice method or a supersonic lifting-surface theo...

Receptivity of a supersonic boundary layer to solid particulates

Abstract: Laminar–turbulent transition in the boundary layer at supersonic speeds can be initiated by small solid particles present in the free stream. Particulates interacting with the boundary-layer flow generate unstable wavepackets related to Tollmien–Schlichting (TS) waves. The latter grow downstream and ultimately break down to turbulent spots. This scenario of TS-dominated transition is modelled using the Mack amplitude method. A theoretical model describing the receptivity mechanism is developed to predict the initial spectrum of TS waves. With these initial conditions the downstream growth of TS instability is calculated using the linear stability theory. The transition onset is associated with the point where the disturbance amplitude reaches a threshold value. As an example, calculations are carried out for a 14° half-angle sharp wedge flying in the standard atmosphere at altitude 20 km, Mach number 4 and zero angle of attack. It is shown that spherical particles of radius from to and density can cause transition onset corresponding to the amplification factor , which is in the empirical range of flight data. This indicates that atmospheric particulates may be a major source of TS-dominated transition on aerodynamically smooth surfaces at supersonic speeds. The receptivity model provides a foundation for further treatments of different cases associated with transition in dusty environments. It can also be used for predictions of particle-induced transition at subsonic and hypersonic speeds.

Mode Change Characteristics of a Three-Dimensional Scramjet at Mach 8

Abstract: Experiments were undertaken to investigate the variation in thrust and combustion efficiency of a three-dimensional scramjet operating in different modes. The scramjet flowpath included a rectangular-to-elliptical shape transition inlet and a divergent elliptical combustion chamber, and the experiments replicated conditions for flight at Mach 8 and 32 km altitude. Two combustion chamber configurations were tested with the same overall length and area ratio, but with different lengths of constant area section before divergence. Injection of the hydrogen fuel was through portholes at the combustion chamber entrance. Both engine configurations underwent a significant change in operating character with increasing equivalence ratio that was consistent with a change from “supersonic” to “separated” combustion in a dual-mode scramjet. The configuration with longer constant area section changed mode at an equivalence ratio just less than one, with an abrupt increase in thrust. This was accompanied by a significan...

Comparison of Turbulence Models in the Calculation of Supersonic Separated Flows

Abstract: Air-jet engines are designed with the next generation of calculations on a wide range of speed and altitude, which leads to the use of new technical solutions. Gap generator nozzle ejectors, diffusers with a sudden expansion flow of self-regulation provide a wide range of design parameters. The currents in these devices is accompanied by flow separation from the walls and on certain modes of shock-wave structure vibrations. Calculation of supersonic separated flows is at present some methodological difficulties. Practical application of turbulence models implemented in modern computer packages, the subject of this article.