Range Of Subsonic Mach Numbers Research Articles

On the subsonic near-wake of a space launcher configuration with exhaust jet

The Ariane 5 failure flight 157 made clear that the loads in the base region of space launcher configurations were underestimated and its near-wake dynamics required more attention. In the recent years, many studies have been published on buffet/buffeting in the critical high subsonic flow regime. Nevertheless, not much experimental data are available on the interaction of the ambient flow with an exhaust jet over a wide subsonic Mach number range. Further, a preceding study without exhaust jet revealed questions regarding a similar distribution of the velocity and Reynolds stress in the near-wake if scaled with the reattachment length. Consequently, a generic space launcher configuration featuring a cold, supersonic, over-expanded jet is investigated experimentally in the vertical test section Cologne (VMK) by means of particle image velocimetry (PIV) for five subsonic Mach numbers ranging from 0.5 to 0.9 with corresponding Reynolds numbers between Re_{text {D}}=0.8times 10^6 to 1.6times 10^6. The velocity and Reynolds stress distribution are provided for the near-wake flow and additionally for the incoming boundary layer. Just as in the preceding study, self-similar features are found in the flow field as long as the separated shear layer reattaches on the solid nozzle wall. Substantial changes are then measured for an alternating (hybrid) reattachment between the solid nozzle wall and supersonic exhaust jet as found for Mach 0.8, one of them being the increased axial turbulence in the recirculation bubble due to a ‘dancing’ large-scale, clockwise-rotating vortex.Graphic abstract

Aeroacoustic characteristics of subsonic flow from axisymmetric pipe-cavities

A detailed experimental study on the acoustic characteristics of an axisymmetric pipe-cavity jet is conducted for different ratios of cavity depth to length, over a range of subsonic Mach numbers. Power spectral analysis and Scalogram are implemented to unravel the physics of the pipe-cavity resonance. Proper Orthogonal Decomposition analysis is used to understand the flow dynamics of the pipe-cavity jet. The results show that the pipe-cavity resonates close to the first tangential mode for a wide range of Mach numbers. However, an increase in the depth and Mach number leads to the onset of combined resonance modes. Experimentally obtained resonance frequencies are validated against numerical simulations and theoretical predictions. Scalogram results show the existence of nonlinear interactions and mode shifting of resonance modes. A parametric space plot demarcates the domains of operation of various oscillation modes, over the relevant parametric range. The results reveal that the upstream cavity significantly affects the downstream jet flow dynamics, and its far-field noise radiation. Deeper cavities have a stronger influence on the far-field acoustics compared to shallow cavities.

On a two-dimensional mode-matching technique for sound generation and transmission in axial-flow outlet guide vanes

The present work deals with the analytical modeling of two aspects of outlet guide vane aeroacoustics in axial-flow fan and compressor rotor-stator stages. The first addressed mechanism is the downstream transmission of rotor noise through the outlet guide vanes, the second one is the sound generation by the impingement of the rotor wakes on the vanes. The elementary prescribed excitation of the stator is an acoustic wave in the first case and a hydrodynamic gust in the second case. The solution for the response of the stator is derived using the same unified approach in both cases, within the scope of a linearized and compressible inviscid theory. It is provided by a mode-matching technique: modal expressions are written in the various sub-domains upstream and downstream of the stator as well as inside the inter-vane channels, and matched according to the conservation laws of fluid dynamics. This quite simple approach is uniformly valid in the whole range of subsonic Mach numbers and frequencies. It is presented for a two-dimensional rectilinear-cascade of zero-staggered flat-plate vanes and completed by the implementation of a Kutta condition. It is then validated in sound generation and transmission test cases by comparing with a previously reported model based on the Wiener-Hopf technique and with reference numerical simulations. Finally it is used to analyze the tonal rotor-stator interaction noise in a typical low-speed fan architecture. The interest of the mode-matching technique is that it could be easily transposed to a three-dimensional annular cascade in cylindrical coordinates in a future work. This makes it an attractive alternative to the classical strip-theory approach.

Computational Modeling of Total-Temperature Probes

A sensor for high-temperature applications is the thermocouple; however, this probe is characterized by errors due to convection, conduction, and radiation heat transfer to/from the sensing element. Experimental testing of total-temperature probes has been essential to identifying key design and performance parameters. With computational simulations that allow for detailed results, these performance parameters can be studied in depth. This work assesses the performance of total-temperature thermocouple probes with respect to measurement error sources, employing multiphysics, computational simulations. Using data from legacy and recent probe designs over a range of subsonic Mach numbers and temperatures, a computational tool is developed and validated to include the effects of conduction, convection, and radiation. Radiation errors are shown to be 50–150°R for the legacy probe at temperatures of 2000–2500°R. Heat transfer coefficients are tabulated as Nusselt numbers, and a new correlation for a shielded total-temperature probe is developed. The heat transfer at the probe tip is found to be three to five times higher than along the sides. The effect of a mounting strut is found to have a large effect on the internal flow patterns and probe performance. This work highlights the usefulness of computer simulations in the design and analysis of small total-temperature probes.

The role of vortices in enhancing the action of a pulsed axisymmetric flow upon an obstacle

It is experimentally established that the pressure at the central point of an obstacle oriented perpendicular to the axis of a gas flow emerging from a channel increases when a fluid nozzle formed by the moving primary (head) and secondary vortices reaches the obstacle surface. This phenomenon has been observed in air-flow at a distance of ∼1–4 channel diameters behind the shock wave in the range of subsonic Mach numbers. The obtained experimental data are generalized in the form of dimensionless relationships.

An algorithm for low Mach number unsteady flows

An algorithm is proposed for the simulation of unsteady viscous stratified compressible flows. The advantage of the method is its capability to deal with a broad range of subsonic Mach numbers, including nearly incompressible flows, with a single modelling, based on the fully compressible Navier–Stokes equations. The method is second-order accurate both in space and time. To remove the stiffness of the numerical problem due to the large disparity between the flow and the acoustic wave speeds at low Mach number, an approximate Newton method, based on artificial compressibility, is used. After a detailed description of the method, the accuracy of the algorithm is checked by computing compressible and incompressible benchmark test cases. Results of flows over a backward facing step with or without stratification effects for a Reynolds number of 800 are compared with the steady solution of incompressible methods. The computation of a compressible natural convection allows to evaluate the accuracy of the method for flow including significant compressibility effects, whereas numerical results of a Poiseuille–Bénard channel flow for a Reynolds number of 10 demonstrate the efficiency of the present algorithm to compute unsteady flows.

Cascade flutter analysis with transient response aerodynamics

Two methods for calculating linear frequency domain unsteady aerodynamic coefficients from a time-marching full-potential cascade solver are developed and verified. In the first method, the influence coefficient method, solutions to elemental problems are superposed to obtain the solutions for a cascade in which all blades are vibrating with a constant interblade phase angle. The elemental problem consists of a single blade in the cascade oscillating while the other blades remain stationary. In the second method, the pulse response method, the response to the transient motion of a blade is used to calculate influence coefficients. This is done by calculating the Fourier transforms of the blade motion and the response. Both methods are validated by comparison with the harmonic oscillation method, in which all the airfoils are oscillated, and are found to give accurate results. The aerodynamic coefficients obtained from these methods are used for frequency domain flutter calculations involving a typical section blade structural model. Flutter calculations are performed for two examples over a range of subsonic Mach numbers using both flat plates and actual airfoils.

A perturbation solution for compressible viscous channel flows

The two-dimensional compressible Navier-Stokes equations are solved by a perturbation expansion in the parameter (Mach number)2/Reynolds number. A fortieth-order solution is generated by a computer algorithm. These series are then summed as convergent series of diagonal Pade approximants. Effectively-exact solutions have been found for Reynolds numbers between zero and 1000 and a range of subsonic Mach numbers in the case of fully-developed isothermal flow between parallel side walls. Choking of the flow is shown to occur for a moderate value of channel Reynolds number. The two-dimensional velocity and pressure fields are obtained. The engineering assumption that friction factor is sensibly independent of Mach number may lead to significant underprediction of head loss in the laminar flow regime.

Study of Secondary Flows in Blade Cascades of Turbomachines

Experiments have been performed with two cascades of turbomachines. The first cascade is composed of highly loaded turbine blades, and has been used in the low subsonic Mach number range. The second cascade consists of highly loaded compressor blades, of the DCA type. The Mach number was then in the high subsonic range. The experimental results are presented in the form of mean values in the pitch direction. Detailed local values are also described. The growth of a passage vortex and a corner effect are presented in the compressor case. Their interactions with the whole flow are analyzed. In the turbine case, the passage vortex is found to be a dominant effect. Results obtained with a theoretical method of calculation of the flow in the blade passage are used to complete the analysis.

The vortex-street structure of turbulent jets. Part 2

This paper is an extension of the work reported earlier as part 1 in a companion paper. It endeavours first to discover how the vortex-street structure in the round jet develops during its passage downstream, and secondly to resolve some of the questions concerning the flow structure in the outer part of the mixing region. It appears that the main vortex street that was reported in part 1 has a trajectory that converges on the jet axis at about two potential core lengths from the nozzle. This brings the vortex cores towards the axis and contributes to the erosion of the potential core. A branch vortex street apparently forms on the outer part of the jet mixing region downstream of x/D = 1.5, and as the main vortex street approaches the axis, the branch vortex street moves outwards. The vortices in the branch are jostled occasionally by those in the main vortex street, and various effects are caused, including the frequency modulation of the signals in the outer part of the mixing region. Measurements at higher Mach numbers confirm the appropriateness of Strouhal number scaling of the jet data for the full range of subsonic Mach numbers and suggest that the structure described in this paper is valid up to the sonic speed.

Subsonic Axisymmetric Near-Wake Studies

The turbulent near-wake of a cylindrical blunt-based body aligned with a uniform freestream was experimentally investigated. Tests were conducted over the entire subsonic Mach number range in a wind tunnel utilizing an upstream model support system. Results indicate that the separation process influences the flow approaching the blunt base for at least 3 model diameters. The base pressure coefficient is constant for Mach numbers less than 0.8, but drops rapidly at near-sonic speeds. The size of the near-wake increases with Mach number. A similarity expression for the near-wake centerline velocity distribution is developed.

The role of shear-layer instability waves in jet exhaust noise

Large-scale structures in the form of instability waves are an inherent part of a shearlayer mixing process. Such structures are shown to be present in an acoustically and aerodynamically well behaved jet even at high Mach numbers. They do not directly radiate significant acoustic power in a subsonic jet, but do govern the production of the turbulent fluctuations which radiate broad-band jet noise. Over the whole subsonic Mach number range, a significant increase in jet noise can be produced by exciting the shear layer with a fluctuating pressure at the nozzle of only 0·08 % of the jet dynamic head but with the correct Strouhal number. Such excitation by internal acoustic, aerodynamic or thermal fluctuations could explain the variability of jet noise measurements between different rigs and could also be responsible for some components of ‘excess’ noise.

An interferometric study of the shock- reflection process in ionizing argon.

without degrading the performance of the single elements. In this manner, the parameters of temperature, pressure, and flow direction can be measured simultaneously and in close proximity in the flowfield. The probe head can pass through a 0.6-cm hole, and the probe may be used in a fixed position or in a linear- and rotational-actuator system to survey an area. Probe flow characteristics have been determined over a range of subsonic Mach numbers (0.2 < M < 1) as well as for Mach number 1.4. Reynolds number ranged from 3 X 106 to 3 X 10 7/m. General design considerations for this type of probe as well as detailed performance curves are presented in Ref. 1. Although the probe was developed primarily for use in rotating-machinery research, it is applicable wherever restrictions on size and requirements of multiple measurements exist.