Values Of Mach Number Research Articles

Comparison of Magnetic and Plasma Overshoots of Interplanetary Shocks

The structure of interplanetary shock fronts was studied based on the data from the BMSW plasma spectrometer, installed on the SPEKTR-R spacecraft, supplemented by magnetic field measurements on the WIND spacecraft. Special attention was paid to periodic growths (overshoots) in the value of the ion flux or magnetic field relative to their mean values outside the ramp. A comparison was performed with the overshoot in the magnetic field, with the Mach number, and with the β parameter. Based on an analysis of 18 intersections of interplanetary shock fronts, in which the overshoots in the ion flux and magnetic field value were observed, it was shown that the value of the magnetic field overshoot is, on the average, less than a similar value in the solar wind’s ion flux, which is associated with different time resolution of measurements. The ion flux overshoot value is found to grow with the growth of the Mach number, in the same way, as the value of the magnetic field overshoot. It is shown that overshoots are formed not only in the supercritical shocks, but also in those with Mach numbers that are less than the value of the first critical Mach number. It is also found that the estimates of the wavelength of the ion flux and magnetic field oscillations behind the ramp well correlate with the value of a gyroradius of captured ions.

Gas Flow in Microchannels and Nanochannels With Variable Cross Section for All Knudsen and All Mach Number Values

AbstractThe analytical solution for steady viscous pressure-driven compressible isothermal gas flow through micro- and nanochannels with variable cross section for all Knudsen and all Mach number values is presented in this paper. The continuum one-dimensional governing equations are solved using the friction factor that is established in a special way to provide solutions for mass flow rate, pressure, and velocity distribution through the microchannels and nanochannels in the entire rarefaction regime. The friction factor, defined by the general boundary condition and generalized diffusion coefficient proposed by Beskok and Karniadakis (1999, “A Model for Flows in Channels, Pipes, and Ducts at Micro and Nano Scales,” J. Microscale Thermophys. Eng., 3, pp. 43–77), spreads the solution application to all rarefaction regimes from continuum to free molecular flow. The correlation between the product of friction factor and Reynolds number (Poiseuille number) and Knudsen number is established explicitly in the paper. Moreover, the obtained solution includes the inertia effect, which allows the application of the solution to both subsonic and supersonic gas flows, which was not shown earlier. The presented solution confirms the existence of the Knudsen minimum in the diverging, converging, and microchannels and nanochannels with constant cross section. The proposed solution is verified by comparison with experimental, analytical, and numerical results available in literature.

Vibro-acoustic characteristics of eccentrically stiffened functionally graded material sandwich cylindrical shell under external mean fluid

The main objective of this research work is focused on acoustic radiation analysis of functionally graded material (FGM) sandwich cylindrical shell structure reinforced by periodically eccentrically ring stiffeners. In the analysis, two common types of FGM sandwich cylindrical shell are considered, which include one with FGM face shell and homogeneous ceramic or metal core, and the other with FGM core and homogeneous face shell. Given the structure is immersed in a mean flowing fluid of velocity tangential to the accoustically deformed boundary and it is excited by a radial point force. The ring stiffeners are assumed to be identical and uniformly spaced, and the in-plane flexural and out-of-plane torsional motions of the ring stiffeners are employed to accurately describe the force moment coupling between the stiffeners and the shell. Fluid-structure coupling is considered by imposing velocity continuity condition at fluid-structure interfaces. By applying the Poisson summation formula and the Fourier transformation technique for periodic ring stiffeners FGM shell structures, the sound pressure level (SPL) about structural acoustic radiation performance index is expressed in a superposition form of space harmonics for a given wave number. The validity of the proposed theoretical model is verified by comparisons with previously published numerical results. Based on the developed theoretical mode, the influences of the FGM properties, gradient index, sandwich type, the size of ring stiffener, external mean flow and skin-core-skin ratios type on the structural acoustic radiation performance have been investigated. In addition, the results show that under the same skin-core-skin ratio and thickness, the acoustic radiation performances of type C seem better, and the larger the section size of ring stiffened, the greater the influence of structure acoustic radiation characteristics, whereas the magnitudes of the SPL curves increase with the increasing values of Mach number.

Расчетное исследование сверхзвукового воздухозаборного устройства в широком диапазоне параметров набегающего потока

This paper examines operation modes of a mixed compression air intake with a rectangular cross-section at Mach number 2.0. The perfect gas model was used for the calculation. Calculations were performed for three values of Mach numbers: 1.8, 2.0 and 3.0. k–ε turbulence model was chosen for describing flows with large adverse pressure gradients. Two-dimensional and three-dimensional configurations of the air intake device were examined. Versions of geometry with and without the boundary layer drain system were considered. The influence of the boundary layer drain system on the flow in the air intake and its characteristics was established. Throttle characteristic curves were formed for all the considered modes with regard to the averaged flow parameters. A comparison of the calculation and experimental data showed a good agreement of the results. The obtained results can serve as a basis for further optimization and improvement of the efficiency of the aircraft design layout, increase in the stability margin of air intakes, as well as development of software systems for regulating supersonic input devices.

Compressible Flow Analysis through Spreadsheet

The compressible flow analysis is traditionally done by referring to tables where various ratios are listed for various values of Mach number M. Here a different approach is presented which does not require any reference to any tables and the problems can be solved much more comprehensibly and which gives more accurate values.

Propulsive modelling for JT9D-3, JT15D-4C and TF-30 turbofan engines using particle swarm optimization

Purpose The purpose of this paper is to create high-accuracy thrust modelling for cruise flight using particle swarm optimization (PSO) algorithm. Design/methodology/approach In this study, using PSO, new thrust models with high accuracy for the cruise flight stages of Pratt & Whitney JT9D-3, JT15D-4C and TF-30 engines were created. For this aim, real Mach number, flight altitude and thrust values taken from the engine manufacturers were used. In the model, thrust is given as a function of altitude and Mach number. The sensitivity of the results given by the PSO thrust model has been examined using several different error types. Finally, the effect of some PSO parameters on the created models is examined. Findings It was observed that the model created predicted real thrust values with high precision. Practical implications The PSO thrust model can be used in the trajectory estimates of today’s aircraft with the use of accurate scaling factors. In addition, using the developed PSO thrust model together with a correct aerodynamic model provides more effective management of air traffic flow in air traffic management applications. Combining the PSO model with fuel flow-rate models will significantly increase engine efficiency and performance; thus, making a major contribution to reducing engine emissions. Originality/value The originality of this study is that it is the first thrust modelling made with PSO in the literature for turbofan engines. The use of real data in the study and the creation of models for several different turbofan engines are important for the reliability of thrust models.

On the total enthalpy behavior inside a shock wave

The total enthalpy behavior inside a shock wave in a dilute monatomic gas has been numerically studied for various values of Mach and Prandtl numbers with the continuum (the Navier–Stokes–Fourier equations) and kinetic (the Shakhov model and the direct simulation Monte Carlo method) approaches. A significant difference between the results by the continuum and kinetic approaches has been observed for the internal shock wave structure. In a wide range of the free-stream Mach numbers, the continuum approach predicts qualitatively similar behavior of total enthalpy distributions that can be of a concave, constant, or convex shape depending on the Prandtl number. The more sophisticated kinetic approach predicts a more complicated form of total enthalpy profiles: e.g., an inflection point for Mach numbers around two and Prandtl numbers close to unity. The evolution of the total enthalpy in the shock is determined by the balance of heat conduction and mechanical work of normal viscous stress—processes that are predicted inaccurately by using the Navier–Stokes–Fourier equations at high Mach numbers.

Control of a Supersonic Inlet in Off-Design Conditions with Plasma Actuators and Bleed

Supersonic inlets are a key component of present and future air-breathing propulsion systems for high-speed flight. The inlet design is challenging because of several phenomena that must be taken under control: shock waves, boundary layer separation and unsteadiness. Furthermore, the intensity of these phenomena is strongly influenced by the working conditions and so active control systems can be particularly useful in off-design conditions. In this work, a mixed compression supersonic inlet with a double wedge ramp is considered. The flow field was numerically investigated at different values of Mach number. The simulations show that large separations appear at the higher Mach numbers on both the upper and lower walls of the duct. In order to improve the performances of the inlet two different control strategies were investigated: plasma actuators and bleed. Different locations of the plasma actuator are considered in order to also apply this technology to configurations with a diverter which prevents boundary layer ingestion. The potential of the proposed solutions is investigated in terms of total pressure recovery, flow distortion and power consumption.

Different noise generation mechanisms of a controlled diffusion aerofoil and their dependence on Mach number

The self-noise of an isolated controlled-diffusion aerofoil is investigated using direct noise computations. The high-order finite difference solver HiPSTAR is used to conduct four large eddy simulations and one direct numerical simulation. The motivation is to investigate the multiple sources of aerofoil self-noise on a realistic compressor blade geometry, the physical mechanisms leading to the generation of sound, and the influence of compressibility effects. The angle of attack is 8°, the chord-based Reynolds number is 105, and results obtained for four values of the free-stream Mach number [0.2, 0.3, 0.4, 0.5] are compared. For those flow parameters, the pressure side is fully laminar, whereas a separation bubble is present on the suction side close to the leading edge that promotes transition to turbulence. The size of the separation bubble is found to increase with the Mach number. Two noise sources are observed, one at the trailing edge and one in the leading edge transition/reattachment region. The first has a broadband, low frequency spectrum, while the second displays a tone whose frequency depends on the local Mach number. Because the leading edge separation bubble is very small, the associated tone frequency is high and requires a significantly finer grid to faithfully resolve the acoustic propagation than what is typically deemed sufficient. Finally, cross correlations between the surface pressure and the far-field pressure reveal that the pressure fluctuations reaching the trailing edge are initially generated in the transition/reattachment region, which indicates that the trailing edge noise is a consequence of the pressure fluctuations generated by the separation bubble.

Blowing Steam into the Boundary Layer on the Outer Wall of the Axial Exhaust Hood

The paper is devoted to experimental verification of flow on the steam turbine axial exhaust hood model. The main goal of the submitted paper is to evaluate the influence of additional air blowing into the boundary layer in the axial exhaust hood on the change of the pressure recovery coefficient. The experimental model was supplemented with a blowing track and fitted with a centric orifice for mass flow measurement. Several variants were measured with various types of inner struts and supports inside the hood for two values of Mach number at the diffuser inlet.

Pressure-Driven Gas Flows in Micro Channels with a Slip Boundary: A Numerical Investigation

In this paper, flow of slightly rarefied compressible nitrogen in microchannels has been investigated numerically for low values of Reynolds and Mach numbers. The 2D governing equations were solved using Finite Element Method with first-order slip boundary conditions (Comsol Multiphysics software). A validation was performed by comparing with similar configuration from the literature. It was found that our model can accurately predict the pressure driven flow in microchannels. Several interesting findings are reported about the Relative pressure, longitudinal velocity, Mach number, effect of gas rarefaction and flow rate.

Near-sonic pure steam flow with real-gas effects and non-equilibrium and homogeneous condensation around thin airfoils

A small-disturbance asymptotic model is derived to describe the complex nature of a pure water vapour flow with non-equilibrium and homogeneous condensation around a thin airfoil operating at transonic speed and small angle of attack. The van der Waals equation of state provides real-gas relationships among the thermodynamic properties of water vapour. Classical nucleation and droplet growth theory is used to model the condensation process. The similarity parameters which determine the flow and condensation physics are identified. The flow may be described by a nonlinear and non-homogeneous partial differential equation coupled with a set of four ordinary differential equations to model the condensation process. The model problem is used to study the effects of independent variation of the upstream flow and thermodynamic conditions, airfoil geometry and angle of attack on the pressure and condensate mass fraction distributions along the airfoil surfaces and the consequent effect on the wave drag and lift coefficients. Increasing the upstream temperature at fixed values of upstream supersaturation ratio and Mach number results in increased condensation and higher wave drag coefficient. Increasing the upstream supersaturation ratio at fixed values of upstream temperature and Mach number also results in increased condensation and the wave drag coefficient increases nonlinearly. In addition, the effects of varying airfoil geometry with a fixed thickness ratio and chord on flow properties and condensation region are studied. The computed results confirm the similarity law of Zierep & Lin (Forsch. Ing. Wes. A, vol. 33 (6), 1967, pp. 169–172), relating the onset condensation Mach number to upstream stagnation relative humidity, when an effective specific heat ratio is used. The small-disturbance model is a useful tool to analyse the physics of high-speed condensing steam flows around airfoils operating at high pressures and temperatures.

Shock Waves Asymmetry in a Symmetric Nozzle

The results of the experimental research on the symmetry of supersonic flow in a symmetric convergent-divergent nozzle are presented. The investigations were focused on the fact that for some flow conditions the flow in a precisely symmetric nozzle becomes asymmetric. Starting from a specific value of Mach number, the flow becomes asymmetric in terms of shock wave λ-foot geometry on both sides of a symmetric nozzle. The evolution of the abovementioned asymmetry has been analysed for Mach number value ranging from M = 1.26 to M = 1.59 with the nozzle opening angle of up to 6.5° on each side. The presented results indicate that for the same flow parameters as Mach number and Reynolds number, and for the same geometry of the nozzle, different λ-foot size is formed at each wall. This unexpected behaviour is responsible for the flow asymmetry. Numerical simulations carried out earlier confirm the appearance of shock wave asymmetry. The side in which the asymmetry takes place is accidental, as the full symmetry of simulation mesh and experiment setup was secured. In numerical simulation the asymmetry follows always the same direction. In experiments the direction of asymmetry happens alternatively without any apparent reason. The explanation of the phenomena is provided in this paper.

A study on the performance characteristics of two-stream supersonic diffusers

This paper investigates the mixing of two supersonic streams in a supersonic exhaust diffuser using theoretical, numerical, and experimental approaches. It has focused attention on the startup and stationary operation of the two-stream supersonic diffusers. The diffuser geometry consists of a constant area throat which connects two ducts of convergent and divergent cross-sections. The effects of diffuser contraction ratio (CR) and convergence angle (θ) on the starting process, pressure recovery, and flow features. Experiments have been performed at five different secondary Mach number (Ms) values in the range of 1.8 <Ms < 2.6, at a fixed primary Mach number (Mp) value of 2. Synchronized pressure measurements are used to study the diffuser performance characteristics. Numerical simulations were also carried out to obtain a better insight into the flow field and computed results have been compared with the experimental results. It was found that the starting behavior has a hysteresis. As a result, the tunnel can be operated at lower pressure ratio than the starting pressure ratio. It is observed that the settling chamber pressure required to start a Constant-Area Diffuser (CAD) is decreased with increase in the diffuser length up to an L/D ratio of 9; subsequently, starting pressure remains almost constant with the further increase in the length due to frictional and pressure recovery losses. The present study provides valuable insights into the multi-stream supersonic mixing and diffusion problem.

Experience of the thermodynamic description and optimization of the Ranque effect

A thermodynamic description of the Ranque effect is given. Inclusion of expansion and compression into the energy exchange process provides a rational explanation to the effect observed. These works are performed by small portions of the gas in polytropic processes. Rotational motion of the gas and the presence of irreversibility in the expansion and compression processes are responsible for energy separation. The mathematical model is constructed with the use of conservation laws and polytropic expansion. Optimization calculations are performed for various values of the efficiency and Mach number of the exhausting hot gas. A method of working process optimization hat can be used in practice is proposed. The efficiency of the expansion work is also determined in the course of optimization.

Experimental study of shock wave propagation and its influence on the spontaneous ignition during high-pressure hydrogen release through a tube

An experimental study of shock wave propagation and its influence on the spontaneous ignition during high-pressure hydrogen release through a tube are measured by pressure transducers and light sensors. Results show that the pressure behind a shock wave first increases, and subsequently remains near constant value with an increase of the propagation distance. That is, a certain propagation distance is required to form a stable shock wave in the tube. In the front of the tube, the minimum value of pressure behind the shock wave (Pshock) required for spontaneous ignition decreases with the increase in axial distance to the diaphragm. However, the minimum Pshock remains nearly a constant value in the rear part of the tube. Moreover, the critical values of shock Mach number (MS) for spontaneous ignition decrease with the increase in tube length. And the ignition delay time decreases with the increase of the MS. As the ignition kernel grows in size to a flame, it propagates downstream along the tube with velocity greater than the theoretical flow velocity of the hydrogen-air contact surface. The flame propagation velocity relative to tube wall increases with MS. When the self-sustained flame exits from the tube, a rapid non-premixed turbulent combustion is observed in the chamber. The combustion-wave overpressure increases with the increase of the MS.

Effect of the gas to wall temperature ratio on the bypass transition

The study of the boundary layer transition plays a fundamental role in the field of turbomachinery. The main reason is the strong influence of the transition on the flow field local parameters, such as skin friction and heat transfer, this variation is reflected on the global ones such as efficiency and heat load of the blade row. Turbulent transition models are nowadays commonly used tools in both CFD research and design practice. It is then of particular interest to understand if they are able to predict the effect of temperature on bypass transition and, in case of positive answer, the reasons of their behaviour. This becomes even more interesting as the effect of the flow aero-thermal coupling becomes prominent in the analysis of such phenomena and is typically not accounted for in the validation of turbulence models. In this work we focus our attention on two state of the art transition model that use two radically different approaches to describe transition. To isolate the effects of the temperature ratio on the transition the simulations have been performed keeping the same values of Reynolds and Mach numbers and changing the value of the wall to freestream Temperature Ratio (TR). The results of the two transition models have been compared between them as well as with experimental results. They show that both models are sensitive to TR, though a locally based (rather than correlation based) approach for transition modelling should be favoured.

Transient boundary layer flow of viscous compressible fluid past a heated sphere with the effect of Mach number

Abstract A two-dimensional transient boundary layer flow of an unsteady, viscous, compressible fluid past a heated sphere with the effect of Mach number is considered. The governing boundary layer equations are reduced to nonsimilar form, and these are solved numerically by three methods, namely, the finite difference method for all time regime, perturbation solutions for small time regime, and the asymptotic solutions for large time regime. Comparisons of the results show a good agreement between the above mentioned methods. Effects of various practical values of the surface temperature parameter, α and the Mach number, M 0 on the local shear stress, the rate of heat transfer, streamlines and isoenthalpies are shown graphically. Further, it is observed that increase in the value of Mach number, M 0 , leads to an increase in the local shear stress and the rate of heat transfer; whereas opposite effect is found in case of local shear stress once the point of separation occured. On the other hand size of the vortex found to be decreased with an increase in Mach number.

Flow field behind a complex total pressure distortion screen

The flow field behind a complex total pressure distortion screen is investigated experimentally and numerically. The distortion screen is designed using an established design methodology and fabricated by water-jet cutting technique. The distorted total pressure field behind the screen is quantified by a distortion index parameter, which is evaluated from computations and experiments for several values of inlet Mach number. The root-mean-square error between the target total pressure values and that achieved by the screen design at the aerodynamic interface plane is 4.75%. The evolution of the distorted total pressure field downstream of the screen is presented in detail in terms of radial and circumferential total pressure distributions and their gradients. An alternative interpretation of the distorted total pressure field is made by means of defining a total pressure flux existing behind the screen and expanding it using derivative-moment transformation technique. It is seen that the circumferential vorticity is a major contributing factor to the total pressure flux.

Reliability Analysis of Aero-Engine Compressor Rotor System Considering Cruise Characteristics

This paper develops an improved dynamic reliability model for compressor rotor system, defined as a “unconventional active system” here. First, regarding to the variety of system-component connections, system specific reliability modeling technique is presented, in which the stress and strength are both nonnegative stochastic processes. The new model fully considers the load characteristics of rotor blades, and inherently embodied the effect of the statistical dependence between the failures of the rotor blades. Then, calculated by the new model, the relationship between reliability and time, and that between the hazard rate and time, for a compressor rotor system are discussed, respectively. Comparing to the traditional series system reliability model, the new system reliability model developed in this paper has higher accuracy for compressor rotor system, since the traditional models neglect the difference between compressor rotor system and traditional active system. Moreover, in order to master the cruise characteristics of the aero-plane during reliability analysis of its compressor rotor blade set, a flight cruise is separated into several stages based on the value of Mach number and its change rate, and a subsection solution method is innovatively proposed accordingly. At the end of the paper, a certain cruise of fighter plane is analyzed to verify the accuracy of the new method. The results show that, as to a flight cruise with different stages, the new method is basically reasonable to most of the load histories, especially for the situation with impact loads existing.