Search

Supersonic wind tunnel test data is compared with a nonlinear aeroelastic computational model, considering a freestream flow with Mach nearly 2, coupled cavity, and no-shock impingement. The measurements include Limit Cycle Oscillations (LCO) for the given set of flow and structural parameters. The effect of a static pressure differential, temperature differential, and the type of LCO is also studied: periodic or chaotic. A fully coupled aero-thermal–acoustic–elastic analysis was carried out, where the strong dependency of the dynamic instability upon the in-plane boundary stiffness was discovered. Associated with aerodynamic and acoustic formulations, the nonlinearities from the structure are found to be the strongest factor determining the occurrence and character of LCO. At the same time, the cavity effect was found to be negligible for this experimental configuration. Additionally, a comparison analysis was performed on the aerodynamic model for this flow condition between the Linear Piston Theory and the Full Potential Aerodynamics. The calculation of the critical flutter boundary was also performed in terms of the static pressure differential between the cavity pressure and the panel surface pressure, presenting regions where the panel is in static deformation due to the pressure differential versus LCO response. The comparison with the analytical model showed good agreement with the measured data, achieving a similar panel displacement at three-quarters of the panel length as well as the expected behavior of periodicity or chaos for specific values of temperature and static pressure differential. [ABSTRACT FROM AUTHOR]

Abstract Oblique shockwaves may impinge on supersonic vehicles internally in an engine or externally on the outer mold line. They create a severe loading environment and may induce dynamic instabilities such as panel flutter. This study computationally explores the effect of shock-induced panel flutter response in 3D, inviscid, Mach 2 flow. Flutter behavior of a square panel is compared across several incident shock angles and inflow dynamic pressures. The presence of an impinging shockwave is found to produce panel flutter that is characteristically different than the shock-free condition. The response contains significantly larger local pressure gradients, larger spanwise variations, and higher-order modal activity in the panel. Results are also compared to two-dimensional inviscid flow over an infinite-span panel. The 3D centerline is found to compare closely with 2D simulations. However, away from the centerline, 3D effects have a significant influence on the solution. In general, stronger oblique shockwaves raise flutter amplitude and frequency, while weaker shockwaves stabilize the panel response. These latter findings indicate important considerations for both structural lifing and flow control applications. [ABSTRACT FROM AUTHOR]

Abstract Lean blowoff characteristics of an ethylene-fueled model scramjet combustor with cavity flameholder are investigated under the inflow conditions of Ma = 2.52 and T 0 = 1600 K. It is observed that, lean blowoff limits increase with increasing injection distance and which for the single-orifice cases are found to be higher than those for the multiple-orifice cases. For the multiple-orifice cases studied, once the flame is ignited, it can always be stabilized by the cavity as long as the fuel supply is constant. For the single-orifice cases, however, the flame can be extinguished intermittently even if the fuel is served continuously. That is, the lean flames are more stable for the multiple-orifice cases. Near-blowoff dynamics are then analyzed for the less unstable single-orifice cases. When the lean blowoff limits are approached, the cavity flames become less and less stable and may be partially extinguished. Nevertheless, the residual flame within the cavity may reignite the combustible mixture outside the cavity and the entire flame may restabilize. When the equivalence ratio is further decreased, ultimate blowoff takes place and is found to occur in multiple steps - the shear-layer flame becomes weaker, the flame is partially extinguished near the trailing edge, the flame shrinks into the latter part of the cavity, the flame moves towards the cavity front wall and is subsequently extinguished completely. Highlights • Near-blowoff cavity flames in a supersonic flow are studied experimentally. • Local extinction and reignition first appears in the cavity shear layer. • Cavity flame is blown off in multiple steps and extinguished in recirculation zone. [ABSTRACT FROM AUTHOR]

Abstract The bleeding control method and analysis of advantages, challenges on wavecatcher (inward turning) intakes are investigated in this paper. Firstly, a quantitative analysis on the substantial advantages of wavecatcher intakes for ramjet is presented comparing to two-dimensional planar symmetry compression. According to the results, with the same parameters in the entrance and exit section, the total pressure recovery of Internal Conical Flow of C increases by 28.0%; the compression surface length decreases by 5%; the wetted area decreases by 15.7% and the pressure drag decreases by 12.1%. However, due to high compressive efficiency of wavecatcher intake, the boundary layer experiences a higher adverse pressure gradient, contributing to enhance shock wave/boundary layer interaction (SBLI). Secondly, the effects of bleeding control on wavecatcher intake are elucidated by numerical simulations. The bleeding control improves the flow structures by decreasing the boundary layer thickness to weaken the SBLI and bleeding the spanwise vortex out. Furthermore, due to the weak interaction, the terminal shock wave is stable closely behind the throat section, increasing the resistance against the back pressure, decreasing the Mach number before the terminal shock wave and improving the total pressure recovery of the exit section by 3.73% relatively. Highlights • The qualitative and quantitative advantages of wavecatcher intake has been obtained. • The large numbers of challengeable APG zones have been found in inward-turning flowpath. • A bleeding control method on section-controllable wavecatcher intake has been designed. • The positive results by bleeding control to weaken SBLI in the wavecatcher intake has been obtained. [ABSTRACT FROM AUTHOR]

Shock waves generated by underwater pulsed discharge have been applied to industrial and bio-medical treatments, and the key technology is to induce high intensity shock waves. Based on the underwater pulsed discharge platform, the effect of the electrical breakdown modes on the shock wave intensity in water is discussed in the positive pin-to-plane configuration. Experimental results demonstrate that the amplitude of applied voltage have a significant influence on the breakdown modes as well as the shock wave intensity. Based on the analysis of discharge images captured by a high-speed camera, it can be concluded that the breakdown modes are classified as subsonic bush-like streamers or supersonic filamentary streamers. The mechanism of subsonic streamer in lower applied voltage is more likely an electro thermal process. The micro-bubbles are firstly generated, and then discharge occurs in the bubble cluster. The bubble volume grows until a breakdown of the underwater gap, and the shock wave is the generated. The time to breakdown of subsonic streamers is commonly within hundreds of microseconds at the experimental condition. In higher applied voltage, the energy loss in the pre-breakdown process is reduced significantly as well. In our experiment, a threshold value of 22.5 kV is necessary for the subsonic streamer mode transition into supersonic streamer mode. The energy conversion efficiency, the time to breakdown and the influence of the applied voltage are discussed at different breakdown modes. The shock wave intensity differs 2–4 times in different breakdown modes even in the same applied voltage, thus increasing the applied voltage may be a remarkable way to enhance the shock wave intensity. [ABSTRACT FROM AUTHOR]

The flow around a backward-facing step in the sub-, trans- and supersonic regimes was investigated at the Trisonic Wind Tunnel Munich with particle image velocimetry and dynamic pressure measurements. These two techniques were combined to simultaneously measure and correlate the velocity fluctuations in a streamwise vertical plane with the pressure fluctuations on the reattachment surface. The results show that the dynamic loads on the reattachment surface increase from subsonic up to the transonic regime while the mean reattachment location moves downstream. As soon as the flow becomes locally supersonic aft of the backward-facing step, the mean reattachment location suddenly moves upstream while the normalized dynamic loads drastically decrease. By correlating the velocity and the dynamic pressure data, it was shown that a clear separation between outer flow and the flow close to the surface aft of the step is responsible for the drastic load reduction. Due to the large difference in pressure/density, the disturbances from the locally supersonic flow do not have an effect on the flow close to the surface. This is also reflected in the power spectral densities of the pressure fluctuations on the surface, showing that at supersonic free-stream Mach numbers a low-frequency pumping motion of the locally subsonic flow is the dominant mode, while in sub-/transonic flow Kelvin-Helmholtz instabilities and a cross-pumping motion of the shear layer dominate the dynamic loads. [ABSTRACT FROM AUTHOR]

Given constant data of density ρ0, velocity -u0er, pressure p0 and electric force -E0er for supersonic flow at the entrance, and constant pressure pex for subsonic flow at the exit, we prove that Euler-Poisson system admits a unique transonic shock solution in a two dimensional convergent nozzle, provided that u0 > 0, E0 > 0, and that E0 is sufficiently large depending on (ρ0, u0, p0) and the length of the nozzle. [ABSTRACT FROM AUTHOR]

In view of the excellent performance of the supersonic separator for natural gas dehydration, a new type of natural gas liquefaction process using the Laval nozzle is proposed in this paper. Theoretical and numerical studies of the supersonic flow and liquefaction process of the methane-ethane binary system in this nozzle are carried out. The effects of the inlet temperature, inlet pressure, back pressure and component composition on the liquefaction process are analyzed. The results show that the critical liquefaction temperature and pressure of the methane-ethane binary system decrease under low inlet temperature or high inlet pressure conditions and the range of the liquid phase region increases, which promotes the liquefaction process. With the increase of the back pressure of the nozzle, the position of the shock wave moves forward and the liquefaction environment is more completely destroyed. For a multi-component natural gas, in which the heavy hydrocarbon content is high, natural gas is more easily liquefied using the Laval nozzle. The liquefaction efficiency range of the newly designed liquefaction process with the Laval nozzle are 0.0795–0.1321 (HYSYS results) and 0.0718–0.1505 (MATLAB results) when the inlet pressure of the process is 2–5 MPa. The nozzle more easily achieves liquefaction compared with a throttle under the same conditions. [ABSTRACT FROM AUTHOR]

A separated turbulent flow in an axisymmetrical nozzle is studied numerically. Two configurations nozzle are investigated. The first one is the truncated ideal contour nozzle, DLR-TIC, is fed with nitrogen. The second configuration is called the thrust optimized contour nozzle or TOC type, ONERA-TOC, where the operating gas is a hot air. The classical pattern of a free shock separation is obtained for different values of the nozzle pressure ratio. The results are compared and validated using experimental data. [ABSTRACT FROM AUTHOR]

Accurate prediction of aerodynamic and thermal environment around a gap has a significant effect on the development of spacecraft. The implicit finite volume schemes are derived and programmed from Navier-Stokes equations. Taking the gap between thermal insulation tiles as an example, a numerical simulation is performed by the finite volume method to obtain the flow characteristic in a gap and then to analyze the heat transfer mechanism. The numerical results are consistent with the experimental ones, which prove the precision of the method used in this paper. Furthermore, the numerical results reveal that the heat convection plays a leading role in heat transfer around a gap. [ABSTRACT FROM AUTHOR]

Introduction: Cirrhosis, understood as a dynamic and reversible process, is the final result of chronically progressive liver diseases of various etiologies. Non-invasive methods based on ultrasound have gradually become useful diagnostic tools for studying liver diseases. Real-time supersonic shearwave elastography (SSWE) study of these patients has recently been introduced to Colombia. Objective: The objective of this study was to evaluate real-time SSWE in patients with cirrhosis at a center for the treatment of liver and digestive diseases in Bogotá. Materials and Methods: This is a retrospective pilot study of a cohort of cirrhotic patients with diagnoses confirmed by liver biopsies, imaging, clinical indicators or a combination of these. Results: Sixty-five patients who had been diagnosed with cirrhosis were included in this study. The median age was 58 years, 61.5% were women, and 38.5% were men. We found that median hepatic stiffness was 16.6 kilopascals (kPs), the interquartile range was 13.9-19.5, the minimum value was 12, and the maximum value was 30.5. There were no statistically significant differences between women and men, within groups and with and without biopsies (p = 0.64 and p = 0.26 respectively). Conclusions: Hepatic rigidity measurements for cirrhotic patients analyzed in this pilot study were within the ranges reported in initial international tests, so this non-invasive test should be considered for patients suspected of having cirrhosis as a way to avoid complications of liver biopsies. [ABSTRACT FROM AUTHOR]

Streamlines behind axisymmetric curved shock waves were used to predict the internal surfaces that produced them. Axisymmetric ring wedge models with varying internal radii of curvature and leading-edge angles were used to produce numerical results. Said numerical simulations were validated using experimental shadowgraph results for a series of ring wedge test pieces. The streamlines behind curved shock waves for lower leading-edge angles are examined at Mach 3.4, whereas the highest leading-edge angle cases are explored at Mach 2.8 and 3.4. Numerical and theoretical streamlines are compared for the highest leading-edge angle cases at Mach 3.6. It was found that wall-bounding theoretical streamlines did not match the internal curved surface. This was due to extreme streamline curvature curving the streamlines when the shock angle approached the Mach angle at lower leading-edge angles. Increased Mach number and internal radius of curvature produced more reasonable results. Very good agreement was found between the theoretical and numerical streamlines at lower curvatures before the influence of the trailing edge expansion fan. [ABSTRACT FROM AUTHOR]

A method for rapid computation of supersonic rotary derivatives is developed. It is based upon perturbation relation between rotary derivatives and normal velocity on surface due to angular velocity according to Riemann invariant and Euler equations solution using Cartesian meshes. The resulting formula of supersonic rotary derivatives depends only on steady Euler solution of the basic flow,therefore,all supersonic rotary derivatives of interest can be obtained at the same time after obtaining the CFD steady-state Euler solutions. The international dynamic derivatives standard model ″Basic Finner″ is studied as an example to validate the method. The calculating results are in agreement well with experimental data and reference data appeared in open literature. It is proved that the proposed method has less runtime,higher accuracy and better practicability in engineering,and can be used to rapidly calculate the supersonic rotary derivatives of flying vehicles with complicated configurations due to using Euler solutions of Cartesian meshes [ABSTRACT FROM AUTHOR]

The aim of this work is to highlight the significance of Fluid-Thermal-Structural Interaction (FTSI) as a diagnosis of existing designs, and as a means of preliminary investigation to ensure the feasibility of new designs before conducting experimental and field tests. The novelty of this work lies in the multi-physics simulations, which are, for the first time, performed on rectangular nozzles. An existing experimental supersonic rectangular converging/diverging nozzle geometry is considered for multi-physics 3D simulations. A design that has been improved by eliminating the sharp throat is further investigated to evaluate its structural integrity at design Nozzle Pressure Ratio (NPR 3.67) and off-design (NPR 4.5) conditions. Static structural analysis is performed by unidirectional coupling of pressure loads from steady 3D Computational Fluid Dynamics (CFD) and thermal loads from steady thermal conduction simulations, such that the simulations represent the experimental set up. Structural deformation in the existing design is far less than the boundary layer thickness, because the impact of Shock wave Boundary Layer Interaction (SBLI) is not as severe. FTSI demonstrates that the discharge coefficient of the improved design is 0.99, and its structural integrity remains intact at off-design conditions. This proves the feasibility of the improved design. Although FTSI influence is shown for a nozzle, the approach can be applied to any product design cycle, or as a prelude to building prototypes. [ABSTRACT FROM AUTHOR]

By employing particle image velocimetry, the impact of convex curvature on the supersonic turbulent boundary layer with inflow Mach number of 2.95 is experimentally investigated. The test models are carefully designed to have a zero-pressure-gradient convex boundary layer to isolate the influence of the streamwise pressure gradient while looking at the impact of convex curvature only. The results are compared with the favorable-pressure-gradient case (with pressure gradient parameter β=-0.70). For both of the zero-pressure-gradient and favorable-pressure-gradient convex boundary layers, the log law is found to be well preserved in the profile of streamwise velocity with inner scaling. Both of the mean streamwise velocity and the principal strain rate of the zero-pressure-gradient case collapse after scaling. By accumulating the streamwise velocity along the wall normal direction, the convex curvature is found to have the potential of diminishing the organized motion of the large scale vortices, which contributes to the stabilization of the turbulence. [ABSTRACT FROM AUTHOR]

Abstract: Because of the rapidly growing use of ultrasound shear wave elastography (SWE) in clinical practices, there is a significant need for development of clinical physics performance assessment methods for this technology. This study aims to report two clinical medical physicists’ tasks: (a) acceptance testing (AT) of SWE function on ten commercial ultrasound systems for clinical liver application and (b) comparison of SWE measurements of targets across vendors for clinical musculoskeletal application. For AT, ten GE LOGIQ E9 XDclear 2.0 scanners with ten C1‐6‐D and ten 9L‐D transducers were studied using two commercial homogenous phantoms. Five measurements were acquired at two depths for each scanner/transducer pair by two operators. Additional tests were performed to access effects of different coupling media, phantom locations and operators. System deviations were less than 5% of group mean or three times standard deviation; therefore, all systems passed AT. A test protocol was provided based on results that no statistically significant difference was observed between using ultrasound gel and salt water for coupling, among different phantom locations, and that interoperator and intraoperator coefficient of variation was less than 3%. For SWE target measurements, two systems were compared — a Supersonic Aixplorer scanner with a SL10‐2 and a SL15‐4 transducer, and an abovementioned GE scanner with 9L‐D transducer. Two stepped cylinders with diameters of 4.05–10.40 mm were measured both longitudinally and transaxially. Target shear wave speed quantification was performed using an in‐house MATLAB program. Using the target shear wave speed deduced from phantom specs as a reference, SL15‐4 performed the best at the measured depth. However, it was challenging to reliably measure a 4.05 mm target for either system. The reported test methods and results could provide important information when dealing with SWE‐related tasks in the clinical environment. [ABSTRACT FROM AUTHOR]

Paper presents the results of numerical studies of hybrid rocket engines operating cycle including unsteady-state transition stage. A mathematical model is developed accounting for the peculiarities of diffusion combustion of fuel in the flow of oxidant, which is composed of oxygen-nitrogen mixture. Three dimensional unsteady-state simulations of chemically reacting gas mixture above thermochemically destructing surface are performed. The results show that the diffusion combustion brings to strongly non-uniform fuel mass regression rate in the flow direction. Diffusive deceleration of chemical reaction brings to the decrease of fuel regression rate in the longitudinal direction. [ABSTRACT FROM AUTHOR]

The endeavor to improve the aerodynamic efficiency of supersonic flying bodies has been constant and significant ever since. Notably it has a significant contribution towards the growth of Aerospace industry. The rigorous experiments carried out on aerodynamic efficiency have led to the optimization of flying body design involving structural, aerodynamic or chemical performance parameters. Aerodynamic efficiency is the basis of energy-efficient flying model, and low drag is the basis of aerodynamic efficiency. The relative impact of drag depends upon the flight regime and specific design requirements. Hence, in the search of improving the aerodynamic efficiency and reducing the undesirable flight properties, specifically in supersonic regime, a channelised study has been carried out, as also being carried out many researchers across the world through different processes. Since shock waves create a considerable amount of drag, which can result in extreme drag on the supersonic body, it forms a significant area of study while aiming at improving the aerodynamic efficiency of a supersonic flying object. Hence, this paper deals with the reduction of wave drag by adopting cascade fin, which is a kind of fin that is still evolving in the aerospace environment. A cascade is a control surface which is composed of an external frame supporting an internal cascade of planar surfaces having small chord length. Various properties of cascade fin impact the aerodynamic efficiency individually and also in consonance with other properties. However specific properties of angle of attack and leading edge shape of cascade fin have been considered in this paper to evaluate their effect in reducing the wave drag of a supersonic flying object. This paper deals with these properties of cascade fin individually in isolation, in order to study and understand their impact in detail. The study on wave drag reduction by varying the angle of attack of cascade fins, in supersonic flow regime, has been carried out by adopting computational fluid dynamics simulation, performed for a Mach number of 2 and angle of attack of 0° and 5°. The study on wave drag reduction by varying the leading edges of cascade fins, in supersonic flow regime, has been carried out by adopting computational fluid dynamics simulation, performed for a Mach number of 2 and angle of attack of 0°. The study of varied leading edges considered have displayed significant variation of wave drag, but in practical application, a particular leading edge needs to be chosen as part of the larger scheme of aerodynamic designing process. [ABSTRACT FROM AUTHOR]

We report our recent development of the high-order flux reconstruction adaptive mesh refinement (AMR) method for magnetohydrodynamics (MHD). The resulted framework features a shock-capturing duo of AMR and artificial resistivity (AR), which can robustly capture shocks and rotational and contact discontinuities with a fraction of the cell counts that are usually required. In our previous paper, we have presented a shock-capturing framework on hydrodynamic problems with artificial diffusivity and AMR. Our AMR approach features a tree-free, direct-addressing approach in retrieving data across multiple levels of refinement. In this article, we report an extension to MHD systems that retains the flexibility of using unstructured grids. The challenges due to complex shock structures and divergence-free constraint of magnetic field are more difficult to deal with than those of hydrodynamic systems. The accuracy of our solver hinges on 2 properties to achieve high-order accuracy on MHD systems: removing the divergence error thoroughly and resolving discontinuities accurately. A hyperbolic divergence cleaning method with multiple subiterations is used for the first task. This method drives away the divergence error and preserves conservative forms of the governing equations. The subiteration can be accelerated by absorbing a pseudo time step into the wave speed coefficient, therefore enjoys a relaxed CFL condition. The AMR method rallies multiple levels of refined cells around various shock discontinuities, and it coordinates with the AR method to obtain sharp shock profiles. The physically consistent AR method localizes discontinuities and damps the spurious oscillation arising in the curl of the magnetic field. The effectiveness of the AMR and AR combination is demonstrated to be much more powerful than simply adding AR on finer and finer mesh, since the AMR steeply reduces the required amount of AR and confines the added artificial diffusivity and resistivity to a narrower and narrower region. We are able to verify the designed high-order accuracy in space by using smooth flow test problems on unstructured grids. The efficiency and robustness of this framework are fully demonstrated through a number of two-dimensional nonsmooth ideal MHD tests. We also successfully demonstrate that the AMR method can help significantly save computational cost for the Orszag-Tang vortex problem. [ABSTRACT FROM AUTHOR]

The interaction between shock train and cavity flow oscillations has been investigated experimentally in an open jet facility. Mach 1.71 flow has been passed over a set of rectangular cavities with L/D ratios varying between 5 and 10. Unsteady pressure measurement and schlieren flow visualization is employed to gain insight into the flow physics. Flow visualization reveals the presence of shock train coupled shear layer oscillations at different levels. Confinement variation establishes the importance of cavity depth in affecting the shock train structure, as the increase in depth resulted in decrease in boundary layer thickness. Shock train strength is found to decrease with decrease in L/D and the weak shock system promotes development of generic cavity oscillations and flow features through longitudinal mechanism. The shock train is found to be oscillatory in nature and the mean position of the bifurcated shock shifts downstream with decrease in L/D. Large scale structures and strengthened oscillations increase the mean and RMS pressure level of the cavities respectively. These large scale structures are incoherent for cavities with hardly any oscillations and coherent for cavities with sustained oscillations. Mode switching and temporal variations in pressure fluctuations are absent for shock train coupled cavity oscillations. [ABSTRACT FROM AUTHOR]

Mach 2 jet from a convergent-divergent elliptical nozzle, of aspect ratio 2, has been controlled with limiting flat and arc tabs. The mixing promoting capability of the flat and arc tabs were studied in the presence of different levels of pressure gradient, at the nozzle exit, corresponding to nozzle pressure ratios of 4 to 8, in steps of one. The geometrical blockage of both the tabs is 5% of nozzle exit area. For the flat tab along the minor axis, the waves in the core become weaker and the core length becomes shorter than the uncontrolled jet, at all the NPRs studied. But the flat tab along the major axis promotes mixing only for some NPRs and retards the mixing for the rest of the NPRs studied. At NPR 5, the flat tab along the minor axis causes the largest core length reduction of 86%. For circular arc tab, along the minor axis, the maximum core length reduction is 55% at NPR 6. Arc tab along the major axis protects the core length for the entire range of the nozzle pressure ratios tested and maximum extension in core length is found at NPR 4, which is 40%. [ABSTRACT FROM PUBLISHER]

Based on Euler-Euler two-fluid model, the gas and liquid flow control equations were established. Combined with the theory of condensation nucleation and droplet growth, the condensation properties of CO2 in a Laval nozzle were simulated. The results showed that both the mathematical models and numerical calculation methods used could reflect the process of gas condensation in the Laval nozzle accurately. The condensation wave of CO2 was not obvious due to its small latent heat of condensation. As the gas flow into the nozzle, the Mach number increased while pressure and temperature decreased, the undercooling increased up to 30 K, and then declined quickly to about 5 K after the occurrence of condensation. The process of nucleation showed a sharp both in time and space. Wilson point is about 2. 21 mm from the throat, the nucleation rate increased sharply from 0 to 2. 04 X 1021 m-3 • s-1 , while the number of droplets reached the order of 1015 instantly. After the formation of condensation core, gas molecules under certain degree of undercooling accumulated and liquefied on the droplets surface, causing droplet radius as well as humidity to increase rapidly. After the process of nucleation, the existing coagulation cores kept growing with the mathematical models and numerical calculation methods used could reflect the process of gas condensation in the Laval nozzle accurately. The condensation wave of CO2 was not obvious due to its small latent heat of condensation. As the gas flow into the nozzle* the Mach number increased while pressure and temperature decreased, the undercooling increased up to 30 K, and then declined quickly to about 5 K after the occurrence of condensation. The process of nucleation showed a sharp both in time and space. Wilson point is about 2. 21 mm from the throat, the nucleation rate increased sharply from 0 to 2. 04 X 1021 m-3 • s-1 , while the number of droplets reached the order of 1015 instantly. After the formation of condensation core, gas molecules under certain degree of undercooling accumulated and liquefied on the droplets surface, causing droplet radius as well as humidity to increase rapidly. After the process of nucleation, the existing coagulation cores kept growing with the droplet radius increasing to 1. 46 X 10-7 m and the humidity up to 0. 093 5 at the exit of the Laval nozzle. [ABSTRACT FROM AUTHOR]

The efficiency of tabs of two geometries in promoting the mixing of a Mach 2 elliptic jet has been studied. Limiting tab of triangular and circular geometry (crosswire) of 5% blockage placed along major and minor axis at the nozzle exit, are tested for nozzle pressure ratio from 4 to 8, in steps of one. Both tabs are efficient mixing promoters, at all the tested NPRs, when placed along the minor axis. But along major axis the crosswire retards the mixing, at all the NPRs. The triangular tab along the major axis is also found to retard the mixing at NPRs 4 and 5, but for nozzle pressure ratios above 5 it causes mixing enhancement even when placed along the major axis. The triangular tab is found to be a better mixing promoter than the crosswire. The maximum core length reduction of 88% is caused by triangular tab along the minor axis is at NPR4. The corresponding core length reduction for the crosswire is only 72 %. Shadowgraph pictures of controlled jets show that both tabs weaken the waves in jet core. The geometry and orientation of the tab and the expansion level influence the mixing caused by the tab. [ABSTRACT FROM AUTHOR]

A study has been made of the possibility for the physical parameters of a source/sink, i.e., for the enthalpy, temperature, total pressure, maximum velocity, and minimum dimension, at a constant radial Mach number to be changed by energy or force action on the gas in a bounded zone. It has been shown that the parameters can be controlled at a subsonic, supersonic, and transonic (sonic in the limit) radial Mach number. In the updated source/sink, all versions of a vortex-source combination can be implemented: into a vacuum, out of a vacuum, into a submerged space, and out of a submerged space, partially or fully. [ABSTRACT FROM AUTHOR]