Nozzle Lip Research Articles

J1002-2-5 機械式デバイスによる騒音低減(流体関連の騒音と振動2)

A variable nozzle causing the divergence of nozzle cross-section and the slit at the nozzle lip is one of the potential mechanical devices that suppress mixing noise as well as shock-associated noise. In JAXA, the noise tests using a jet engine have been conducted with the variable nozzle. The engine test suggests that the device might reduce mixing noise to some extent, however, is insufficient to further understanding of the acoustic characteristics of the variable nozzle. That is why a noise test on a scaled nozzle of that used in the engine test was executed in the anechoic facility. This paper describes the noise test result using the scaled nozzle.

The near pressure field of co-axial subsonic jets

Results are presented from pressure measurements performed in the irrotational near field of unbounded co-axial jets. Measurements were made for a variety of velocity and temperature ratios, and configurations both with and without serrations on the secondary nozzle lip. The principal objective of the study is to better understand the near pressure field of the jet, what it can tell us regarding the underlying turbulence structure, and in particular how it can be related to the source mechanisms of the flow.A preliminary analysis of the axial, temporal and azimuthal structure of the pressure field shows it to be highly organized, with axial spatial modes (obtained by proper orthogonal decomposition) which resemble Fourier modes. The effects of serrations on the pressure fluctuations comprise a global reduction in level, a change in the axial energy distribution, and a modification of the evolution of the characteristic time scales.A further analysis in frequency–wavenumber space is then performed, and a filtering operation is used to separate the convective and propagative footprints of the pressure field. This operation reveals two distinct signatures in the propagating component of the field: a low-frequency component which radiates at small angles to the flow axis and is characterized by extensive axial coherence, and a less-coherent high-frequency component which primarily radiates in sideline directions. The serrations are found to reduce the energy of the axially coherent propagating component, but its structure remains fundamentally unchanged; the high-frequency component is found to be enhanced. A further effect of the serrations involves a relative increase of the mean-square pressure level of the acoustic component – integrated over the measurement domain – with respect to the hydrodynamic component. The effect of increasing the velocity and temperature of the primary jet involves a relative increase in the acoustic component of the near field, while the hydrodynamic component remains relatively unchanged: this shows that the additional acoustic energy is generated by the mixing region which is produced by the interaction of the inner and the outer shear layers, whereas the hydrodynamic component of the near field is primarily driven by the outer shear layer.

Flow visualization of vortex structure in a pulsed rectangular jet

Pulsating jet is visualized using hydrogen bubble method to clarify the vortex nature in the near field of the jet. This study focused on the development in space and time of vortex structures evolution in low aspect-ratio rectangular jet with pulsation. Pulsation means large-amplitude, low-frequency excitation which is expected to increase the mixing and spreading of the jet and to accelerate its transition from a rectangular form to an axisymmetric form. It was deemed appropriate to investigate whether jet characteristics of a pulsating, submerged jet flow can be altered by including pulsations. The difference of the vortex deformation process is discussed in relation to pulsating conditions. Consequently, the pulsation leads to the formation of vortices at regular intervals, which are larger than those occurring in a steady jet. The results show that the streamwise interaction, between leading vortex and trailing vortex rolled up at nozzle lips, strengthens with increasing pulsating frequency. The spanwise drift of the vortex becomes strongly apparent at large amplitude and high frequency conditions. The drifting start position does not change regardless of pulsating condition. The convection velocity of vortex increases at lower frequency and larger amplitude.

Large-eddy simulation of variable-density turbulent axisymmetric jets

Three cases of variable-density turbulent round jets discharging from a straight circular pipe into a weakly confined low-speed co-flowing air stream are studied with the aid of large-eddy simulation. The density ratios considered are 0.14 [Helium/air], 1.0 [air/air] and 1.52 [CO 2/air], with Reynolds numbers of 7000, 21,000 and 32,000, respectively. Detailed comparisons of the statistics show good agreement with the corresponding experiments. They confirm that a lower-density jet develops more rapidly than a denser jet with the same exit momentum flux. Pseudo-similarity behavior in the three variable-density round jets is well reproduced in the simulation. The coherent structures of the three jets are investigated by visualization of the iso-surface of pressure fluctuations and vorticity. In the developing stage of the Kelvin–Helmholtz instability, large finger-shape regions of vorticity are observed for the helium jet close to the nozzle lip. This feature, however, is not found in the air and the CO 2 jet. The occurrence of strong streamwise vorticities across the shear layer in the helium jet is demonstrated by a characteristic quantity related to the orientation of the vorticity.

Effect of Nozzle Lip Shape on Screech Tone in a Supersonic Jet

The effect of nozzle lip shape on screech tone is investigated by a computational aeroacoustic approach. First, the scattered patterns of screech tone are examined without a jet mean flow, where plane waves interact with a nozzle lip wall. Then, an axisymmetric jet screech is simulated in the Mach number range from 1.07 to 1.2, where the axisymmetric mode is a dominant screech mode. The conservative form of the axisymmetric Euler equations written in generalized coordinates are used to treat the complex nozzle lip geometry for the scattering problem as well as the Reynolds Averaged Navier-Stokes equations with the modified Spalart-Allmaras turbulence model. The computed shock-cell structure, screech tone frequency, and sound pressure levels in the near field are in good agreement with existing experimental data.

Near-Field Investigation of Chevron Nozzle Mechanisms

A detailed investigation into the effect of chevron nozzles on the near-field acoustics of a separate flow exhaust system was conducted at the University of Cincinnati Anechoic Test Facility. Chevrons with varying numbers of lobes and levels of penetration were selected to provide insight into the effects of these geometric parameters on the acoustic near field. Tests were conducted at two different nozzle operating conditions and the chevrons were shown to produce substantial modifications to the near field over a wide range of frequencies. The chevrons were most effective at lower frequencies where the peak noise region was reduced by 5-7 dB and dramatically reduced in size. At higher frequencies, the chevrons provided strong noise suppression downstream of approximately seven equivalent nozzle diameters with increases closer to the nozzle lip. The nozzle penetration was shown to have the most significant impact on the acoustic near field with more subtle differences being seen with respect to the number of chevron lobes.

Acoustic properties associated with nozzle lip thickness in screeching jets

Imperfectly expanded supersonic jets generate discrete frequency sound known as screech tones. In the low supersonic Mach number, the screech phenomenon is axisymmetric. In this study the effect of nozzle lip thickness on the axisymmetric jet screech is investigated both by screech scattering computation and screeching jet simulation. The conservative form of the axisymmetric Euler equations in generalized coordinates are used to simulate a scattering problem and the Reynolds-Averaged Navier-Stokes equations with the modified Spalart-Allmaras turbulence model is employed for screeching jet simulation. Numerical results of scattered directivity pattern in scattering computation will be reported. Influence of the nozzle lip thickness on the momentum thickness, shock-cell spacing and screech tone wavelength is examined.

Temporally resolved PIV for space–time correlations in both cold and hot jet flows

Temporally resolved particle image velocimetry (TR-PIV) is the newest and most exciting tool recently developed to support our continuing efforts to characterize and improve our understanding of the decay of turbulence in jet flows—a critical element for understanding the acoustic properties of the flow. A new TR-PIV system has been developed at the NASA Glenn Research Center which is capable of acquiring planar PIV image frame pairs at up to 25 kHz. The data reported here were collected at Mach numbers of 0.5 and 0.9 and at temperature ratios of 0.89 and 1.76. The field of view of the TR-PIV system covered six nozzle diameters along the lip line of the 50.8 mm diameter jet. The cold flow data at Mach 0.5 were compared with hotwire anemometry measurements in order to validate the new TR-PIV technique. The axial turbulence profiles measured across the shear layer using TR-PIV were thinner than those measured using hotwire anemometry and remained centred along the nozzle lip line. The collected TR-PIV data illustrate the differences in the single point statistical flow properties of cold and hot jet flows. The planar, time-resolved velocity records were then used to compute two-point space–time correlations of the flow at the Mach 0.9 flow condition. The TR-PIV results show that there are differences in the convective velocity and growth rate of the turbulent structures between cold and hot flows at the same Mach number.

Hybrid Reynolds-Averaged Navier-Stokes/Large Eddy Simulation Approach for Predicting Jet Noise

Hybrid acoustic prediction methods have an important advantage over the current Reynolds-averaged Navier-Stokes based methods in that they only involve modeling of the relatively universal subscale motion and not the configuration-dependent larger-scale turbulence. Unfortunately, they are unable to account for the high-frequency sound generated by the turbulence in the initial mixing layers. This paper introduces an alternative approach that directly calculates the sound from a hybrid Reynolds-averaged Navier-Stokes/large eddy simulation flow model (which can resolve the steep gradients in the initial mixing layers near the nozzle lip) and adopts modeling techniques similar to those used in current Reynolds-averaged Navier-Stokes based noise prediction methods to determine the unknown sources in the equations for the remaining unresolved components of the sound field. The resulting prediction method would then be intermediate between the current noise prediction codes and previously proposed hybrid noise prediction methods.

Dynamics of confined premixed flames submitted to upstream acoustic modulations

Dynamics of inverted conical flames anchored on a central bluff-body is investigated in confined and unconfined conditions. The flames are modulated by a driver unit placed upstream. Flame dynamics is strongly coupled to the evolution of large scale structures induced by the modulation and shed from the injector nozzle lip. To study the influence of confinement on the flame motion, the response is measured for different diameters of the duct bounding the flame. This induces changes in the confinement parameter defined as the ratio of the injection duct diameter to the combustor diameter. Natural emission imaging, transfer function measurements and PIV data serve to determine the effect of confinement on the flame response for a set of modulation amplitudes. It is found that when the confinement ratio is low, the time delay between the velocity signal at the burner outlet and heat release perturbations remains essentially constant indicating that the flame is weakly affected by the presence of the wall. As the confinement ratio is increased, the time delay and the gain of the flame transfer function are modified. It is found that interaction with the wall strongly influences the dynamics of the flame tip, the flame wrinkling geometry in the vicinity of the boundary, and to a great extent the amplitude of response. A numerical simulation provides additional information on the fresh stream dynamics inside the tube, and on the dynamical interactions between the flow and the flame.

Bleed Lip Geometry Effects on the Flow in a Hypersonic Wind Tunnel

L AMINAR-turbulent transition in boundary layers is important for the prediction and control of skin friction, heat transfer, and other boundary layer properties. Therefore, it is important to have reliable capabilities in predicting boundary layer transition to optimize hypersonic vehicle performance [1]. The mechanisms leading to transition are poorly understood [2]. Transition experiments have been carried out in conventional ground testing facilities for many decades. However, most of the experimental data obtained from these facilities are not reliable because they have much higher disturbance levels compared with actual flight conditions [3]. Quiet-flow wind tunnels are intended to replicate the low noise conditions of actual flight at hypersonic speed. Reaching quiet flow requires the maintenance of a laminar boundary layer on the nozzle wall to avoid acoustic fluctuations generated by boundary layer turbulence. One method to reduce noise is to delay boundary layer transition using a bleed slot before the nozzle throat. The Boeing/Air Force Office of Scientific Research (AFOSR)Mach 6 wind tunnel at Purdue University has been designed as a quiet tunnel with a bleed slot for which the noise level is an order of magnitude lower than that in conventional wind tunnels. It is a Ludwieg tube that is a long pipe having a converging-diverging nozzle followed by a test section as shown in Fig. 1. A close-up view of the tunnel geometry around the bleed slot lip is shown in Fig. 2. However, the tunnel (which has been operational since 2001) is not yet quiet for the desired range of stagnation pressures of up to 150 psi. Two different nozzles have been fabricated and tested. The tunnel is quiet up to a stagnation pressure of 8 psi with the original electroformed nozzle. The original design of the outer surface of the bleed slot has been modified, and eight different bleed slot designs together with the original one have been tested [4]. A second nozzle throat has been fabricated from aluminum [5]. The tests on the tunnel with this aluminum surrogate throat show that the tunnel is quiet up to a stagnation pressure of 93 psi. Early transition of the nozzle wall boundary layer has been identified as the cause of the test section noise for the tunnel at Purdue University. Separation bubbles on the bleed lip and associated fluctuations induced near the bleed lip were identified as the most likely cause of early transition [4]. The experimental study of Klebanoff and Tidstrom [6] showed that the presence of a separation bubble of sufficient size destabilizes the laminar boundary layer downstream of reattachment thereby leading to an earlier transition to turbulence, i.e., the location of transition moves upstream relative to where it would occur without the separation bubble. This hypothesis regarding a separation bubble was supported by the measurements showing an increase in quiet-flow stagnation pressure from 8 to 93 psia when the electroformed nozzle throat was replaced with the aluminum throat [5]. The bleed lip of the electroformed throat has a 0.001 in. kink that is not present in the aluminum throat, and it appears that the kink in the electroformed throat exacerbates a natural tendency to form a separation bubble near the lip. This separation bubble is highly unsteady and can lead to early transition downstream [7]. However, the separation bubble apparently still exists even at 93 psia, according to the computations presented herein. To achieve quiet flow above 93 psia, and to make the quiet flow less sensitive to the exact shape of the bleed lip, it is desirable to eliminate the separation bubble completely. The situation in the hypersonicwind tunnel at PurdueUniversity is an example illustrating the importance of the bleed lip geometry and the effects of separation bubbles that form around the bleed lip on the quality of the flow at the test section. The objective of this study is to demonstrate the effect of separation bubbles on flow structure by numerically investigating the existence of steady and unsteady separation bubbles on the main-flow or the bleed-flow side of the nozzle lip of the Boeing/AFOSR Mach 6 wind tunnel at Purdue University, and to design a new geometry to eliminate or reduce the size of the separation bubbles.

Differential characteristics of the flow field in a plane overexpanded jet in the vicinity of the nozzle lip

A method of theoretical investigation of the flow field in a two-dimensional (plane-parallel or axisymmetric) overexpanded jet of an ideal perfect gas in the vicinity of the nozzle lip is described. The changes in curvature of the shock wave emanating from the lip, as well as the shock-wave intensity and flow parameters behind the shock are analyzed as functions of the Mach number, pressure ratio in the plane jet, and ratio of specific heats of the gas.

An experimental study of the nozzle lip thickness effect on supersonic jet screech tones

It is well known that screech tones of supersonic jet are generated by a feedback loop driven by the instability waves. Near the nozzle lip where the supersonic jet mixing layer is receptive to external excitation, acoustic disturbances impinging on this area excite the instability waves. This fact implies that the nozzle lip thickness can influence the screech tones of supersonic jet. The objective of the present study is to experimentally investigate the effect of nozzle-lip thickness on screech tones of supersonic jets issuing from a convergent-divergent nozzle. A baffle plate was installed at the nozzle exit to change the nozzle-lip thickness. Detailed acoustic measurement and flow visualization were made to specify the screech tones. The results obtained obviously show that nozzle-lip thickness significantly affects the screech tones of supersonic jet, strongly depending on whether the jet at the nozzle exit is over-expanded or under-expanded.

Numerical simulation of the generation mechanism of axisymmetric supersonic jet screech tones

In this paper an axisymmetric computational aeroacoustic procedure is developed to investigate the generation mechanism of axisymmetric supersonic jet screech tones. The axisymmetric Navier-Stokes equations and the two equations standard k-ϵ turbulence model modified by Turpin and Troyes [“Validation of a two-equation turbulence model for axisymmetric reacting and non-reaction flows,” AIAA Paper No. 2000-3463 (2000)] are solved in the generalized curvilinear coordinate system. A generalized wall function is applied in the nozzle exit wall region. The dispersion-relation-preserving scheme is applied for space discretization. The 2N storage low-dissipation and low-dispersion Runge-Kutta scheme is employed for time integration. Much attention is paid to far-field boundary conditions and turbulence model. The underexpanded axisymmetric supersonic jet screech tones are simulated over the Mach number from 1.05 to 1.2. Numerical results are presented and compared with the experimental data by other researchers. The simulated wavelengths of A0, A1, A2, and B modes and part of simulated amplitudes agree very well with the measurement data by Ponton and Seiner [“The effects of nozzle exit lip thickness on plume resonance,” J. Sound Vib. 154, 531 (1992)]. In particular, the phenomena of modes jumping have been captured correctly although the numerical procedure has to be improved to predict the amplitudes of supersonic jet screech tones more accurately. Furthermore, the phenomena of shock motions are analyzed. The predicted splitting and combination of shock cells are similar with the experimental observations of Panda [“Shock oscillation in underexpanded screeching jets,” J. Fluid. Mech. 363, 173 (1998)]. Finally, the receptivity process is numerically studied and analyzed. It is shown that the receptivity zone is associated with the initial thin shear layer, and the incoming and reflected sound waves.

Influence of Feed Back Excitation Signal on the Vortex Structure in a Two-Dimensional Jet (Effect of Excitation Signal with Phase-Shift)

The behavior of the vortical structure in a two-dimensional jet was investigated by the feedback excitation method use the fluctuating velocity with the phase-shift, The experiments were carried out at exit velocity of 10m/sec and two-dimensional nozzle of aspect ratio 14 and Reynolds number is 1.42×104 based on the nozzle height. The jet is excited by the disturbance fed into the jet at the nozzle lips, the excitation intensity u'/Ueo was fixed on 0.04. The effect of excitation was investigated by measuring mean velocities, the fluctuating velocities, and the power spectrum densities. The result suggests that the behavior of the vortical structure is manipulated by use the signal with the phase-shift.

B303 感圧・感温塗料を用いた斜め平板衝突噴流の解析

Flow fields of the M2.2 jets impinging on an inclined flat plate at high plate-angles are experimentally investigated using surface pressure measurement with Pressure Sensitive Paint and Schlieren flow visualization. The present experiments show that x/Ls', (the distance from nozzle lip to the point where the jet first impinges on a plate normalized by the distance from nozzle lip to the point where jet expands the most), is still the dominant parameter defining the flow types at high plate angles between 60 and 90 degrees. The present study also shows that the flow type prediction method proposed for the plate angles between 30 and 60 degrees by Nakai et al. requires other considerations when being applied to the flows at higher plate angles. In the present experiments, a new flow type (axisymmetric jet with bubble) is also observed.

A241 超音速噴流騒音の特性に及ぼすノズルリップ厚さの影響

Under certain conditions, supersonic jet issuing from a supersonic nozzle invariably generates supersonic jet noise, which consists of three principal components ; the turbulent mixing noise, the broadband shock-associated noise, and the screech tones. In present study, it was experimentally investigated to the effect of nozzle lip thickness on the characteristics of supersonic jet noise. The convergent-divergent nozzle of design Mach number 2.0 was used in experiment. With two different cases in the nozzle-lip thickness, pressure ratio was varied in the range between 2.0 and 12.0. Acoustic measurements were conducted by microphones in an anechoic room, and the major structures of the supersonic jets were visualized by a Schlieren optical system for the purpose of the effect of nozzle lip thickness. The measured results show that the characteristics of supersonic jet noise, such as overall sound pressure level and screech frequency, depend upon the thickness of nozzle lip.

Large-Eddy Simulation of a Turbulent Compressible Round Jet

A large-eddy simulation (LES) of a high-Reynolds-number jet was performed. A numerical scheme that is uniformly fourth-order accurate in time and space was used to solve the Favre-e ltered Navier ‐Stokes equations. The subgrid-scale stresses were computed using a compressible form of Smagorinsky’ s model. A cold Mach 1.4 nozzlewithaReynoldsnumberof1 :2 ££ 106 wassimulated.Incontrasttosimilarstudies,theentirenozzlegeometry including the nozzle lip was modeled. The LES simulation accurately captures the physics of the turbulent e ow. The agreement with experimental data is relatively good and improves on results in the current literature. This improvement can be attributed to the numerical scheme and the modeling of the nozzle. In addition, a twopoint correlation technique was used to quantify the turbulent structures in the jet mixing layer and showed that computational techniques can be used to characterize such structures for application to Lighthill’ s acoustic analogy. Two-point space correlations were used to obtain a measure of the integral length scale, which proved to be approximately 1 Dj. Two-point space-time correlations were used to estimate the convection velocity for the turbulent structures. This velocity estimates ranged from 0.57 to 0.71 Uj and is in agreement with theory.

The sound production characteristics of high-frequency impinging jet tones

An experimental investigation into the production of high-frequency tones by a supersonic impinging jet using phase-locked shadowgraph photography and high-resolution digital particle image velocimetry (DPIV) is presented. High-frequency tones are often produced for nozzle-to-plate spacings less than two nozzle exit diameters and for nozzle pressure ratios (NPR) less than approximately 3.4, where NPR is equal to the stagnation pressure divided by the pressure at the nozzle lip. The high-frequency tones have lower amplitudes and less impulsive wavefronts than the more dominant low-frequency impinging tones and are presumably produced by a significantly different sound production mechanism. Periodic motion of the jet appears to be confined to the peripheral regions of the jet with no significant motion of the flow in the central regions of the jet behind the standoff shock wave. Unlike the jet motion associated with low-frequency impinging jet tones, recirculating regions are not formed and swept downstream behind the standoff shock wave and stationary stagnation regions are also not observed.

平板に衝突する音速噴流の振動機構に関する数値解析

The unsteady flowfields of the sonic jet impinging on a flat plate are numerically investigated using the two-dimensional axisymmetric Euler equations. The characteristics of the jet oscillations qualitatively agree with the experimental results, and the mechanism of jet oscillation and the physical model of sound wave generation are discussed. The jet oscillations have the same period as the sound waves, because passing of sound waves around the nozzle exit causes oscillations of the jet itself. The mechanism of jet oscillations consists of sound waves passing and the vortex convection. In the mechanism, the time delay between the sound wave passing and the vortex occurrence at the nozzle lip is newly pointed out and corresponds to half wave length of the sound wave. A physical model including the time delay is newly suggested and well agrees with both computational and experimental results.