Supersonic nozzles are not always operated at design conditions. The total pressure, temperature, and velocity distributions at the nozzle inlet plane are often characterized by inhomogeneities, conditions dictated by the operating regime of the turbine or combustion chamber. In particular, a swirling flow motion can be induced by these components. While homogeneous inflow conditions are well documented for a large range of supersonic nozzles, data on the aeroacoustics of supersonic swirling jets is scarce. Large eddy simulations are deployed to simulate the swirling flow of a nonideally expanded three-dimensional, cold, axisymmetric aerospike nozzle at a nozzle pressure ratio (NPR) of 3. Three swirl numbers are considered and compared with the baseline case. Near-field acoustic analyses are completed by far-field acoustic computations based on the Ffowcs Williams-Hawkings (FWH) equation. Swirling flow shortens the potential core of the jet and leads to an annular shock cell length increase. Two-point space-time cross correlations of pressure data acquired in the annular shear layer indicate an enhancement of the azimuthal modes. Similar cross correlations in the circular jet shear layer further downstream show that screech tones are suppressed. Power spectral density of the radial velocity at monitoring points in the vicinity of the nozzle trailing edge allows to identify the oscillation modes of the annular shock cell structure. The far-field spectra exhibit lower mixing noise with the increasing swirl number. The global sound pressure level (SPL) decreases, while the nozzle thrust remains at 99% of the baseline thrust at low swirl numbers.

This paper presents the design of a transonic low-pressure turbine (LPT) for the next-generation fighter air-breathing engines. The study focuses on the design of a cascade profile representative of an LPT Nozzle Guide Vane (NGV), that follows conventional literature guidelines for transonic turbine airfoils. The paper reports on the numerical and experimental methods that will be employed for a detailed understanding of the flow physics for this baseline solution. The experimental setup includes a high-speed linear cascade that can be operated at a wide range of inlet turbulence levels (Tu = 5% - 8%) and outlet Mach numbers (M = 0.8 - 1.2). The test section inlet is equipped for hot-wire anemometry measurements, while a purposely designed multi-hole probe is traversed to measure the aerodynamic flow quantities at the cascade outlet. The central passage airfoils feature arrays of pneumatic pressure taps to evaluate the blade loading and hot films to study the status of the boundary layer. Optical sidewalls enable full-field Schlieren and Background-Oriented Schlieren imagery to study cascade shock patterns and unsteady shock-boundary layer interactions. The test section is designed with provision for time-resolved stereo-PIV measurements to cross-validate the cascade velocity field and quantify the turbulence statistics and transport mechanisms through the transonic LPT passage. A detailed planning for high-fidelity flow simulations (LES and DNS) is presented in the second part of the paper. State-of-the-art computational methodologies will be employed along with advanced post-processing techniques including mode decomposition to enhance the understanding of the flow physics, assess the limitations of traditional numerical methods and complement the experimental findings.

In supersonic aerospace applications, aerospike nozzles have been subject of growing interest. This study sheds light on the noise components of a cold jet exhausting an aerospike nozzle. Implicit large eddy simulations (ILES) are deployed to simulate the jet at a nozzle pressure ratio (NPR)=3. For far-field acoustic computation, the Ffowcs Williams-Hawk-ings (FWH) equation is applied. A mesh sensitivity study is performed and the jet instantaneous and time-averaged flow characteristics are analyzed. The annular shock structure displays short non-attached shock-cells and longer attached shock-cells. Downstream of the aerospike, a circular shock-cell structure is formed with long shock-cells. Two-point cross-correlations of data acquired at monitoring points located along the shear layers allow to identify upstream propagating waves associated to screech. Power spectral density at monitoring points in the annular shock-cell structure allows to identify its radial oscillation modes. Furthermore, a vortex sheet model is adapted to predict the annular shock-cells length and the BBSAN central frequency. High sound pressure levels (SPL) are detected at the determined BBSAN central frequencies. Finally, high SPL are obtained at the radial oscillation frequencies for the annular shock-cell structure.

Characterizing pulsating flow in high-temperature, high-pressure engine exhaust gas is crucial for the development and optimization of exhaust energy recovery systems. However, the experimental investigation of engine exhaust pulses is challenging due to the difficulties in conducting crank angle-resolved measurements under these unsteady flow conditions. This study contributes to characterizing mass flow pulses from an isolated cylinder exhaust of a heavy-duty diesel engine using a single-pipe measurement system, developed for pulsating flow measurement. A Pitot tube-based approach is adopted to measure exhaust mass flow pulsations, complemented by fast temperature measurements obtained using customized unsheathed thin-wire thermocouples. The on-engine experiment is performed by isolating the in-cylinder trapped mass and the valve opening speed to produce different exhaust pulse waveforms. The adopted approach’s sensitivity in resolving instantaneous mass flows is evaluated analytically and experimentally, considering attenuated temperature measurement effects. Based on exhaust flow measurements, mass flow pulses are analyzed with regard to blow-down and scavenge phases. Under the load sweep, the main waveform change occurs during the blow-down phase, with pulse magnitude increasing with the load. In contrast, as the engine speeds up with a comparable trapped mass, the exhaust mass distribution in the blow-down phase decreases from 75.5% at 700 rpm to 41.9% at 1900 rpm. Additionally, it is observed that cycle-to-cycle variations in mass flow pulses align with combustion stability during the blow-down phase and are predominantly influenced by gas-exchange processes during the scavenge phase.

Obstructive sleep apnea syndrome (OSAS) is referring to partial or complete cessation of airflow during sleep due to upper pharyngeal airway collapse. Snoring is often associated to such sleep-induced apnea as the result of the strong coupling and interaction between the respiratory airflow and the flexible tissue of the upper airway, resulting in self-excited oscillations of the soft tissue. It can occur at the level of the soft palate but also at the lower level of the pharyngeal airway. The knowledge of the tissue behavior subject to a particular airway flow is relevant for understanding the underlying physical mechanism of OSA. However, in-vivo measurements are usually not practical. A 3D fluid-structure interaction model for the human uvula-palatal system relevant to OSAS based on simplified geometries is utilized in the present study. Numerical simulations are performed to assess the influence of gravity and the rigidity of the soft tissue on the oscillatory dynamics. Meanwhile, the vortex dynamics and the structural modal frequency response are investigated for the coupled fluid-structure system as well.

Flow and acoustic fields of a rectangular over-expanded supersonic jet interacting with an adjacent parallel plate are investigated using compressible Large Eddy Simulations (LES). The jet exits from a converging diverging rectangular nozzle of aspect ratio 2 with a design Mach number 1.5. Four distances (0 to 3 equivalent diameters) between the plate and the adjacent lip of the rectangular jet in the minor axis plane are studied. The geometry of the nozzle, the positions of the plate, and the exit conditions are identical to the ones of an experimental study. Snapshots and mean velocity fields are presented. Good agreement with the PIV experimental measurements is obtained. Previously, the corresponding free jet has been found to undergo a strong flapping motion in the minor axis plane due to screech. Here, it is shown that the intensity of the screech increases for certain distances from the plate and decreases for others, as compared to the corresponding free jet. Two points space-time cross correlations of the pressure along the jet’s shear-layers show, in two cases, an amplification of the aeroacoustic feedback mechanism leading to screech noise in the jet shear-layer closer to the plate. This amplification is due to acoustic waves impinging on the plate, and generating propagating waves back towards the jet, thus exciting the shear-layer at the screech frequency, around the tenth shock cell. Moreover, when the jet develops as a wall jet on the plate, the screech frequency and its associated flapping motion is canceled but a symmetrical oscillation of the jet at a lower frequency becomes dominant and radiates in the near acoustic field. This oscillation mode, as the ones associated with the screech tones for the other cases studied, can be explained by the use of a vortex sheet model of the ideally expanded equivalent planar jet.

Collapsible tubes can be employed to study the sound generation mechanism in the human respiratory system. The goals of this work are (a) to determine the airflow characteristics connected to three different collapse states of a physiological tube and (b) to find a relation between the sound power radiated by the tube and its collapse state. The methodology is based on the implementation of computational fluid dynamics simulation on experimentally validated geometries. The flow is characterized by a radical change of behavior before and after the contact of the lumen. The maximum of the sound power radiated corresponds to the post-buckling configuration. The idea of an acoustic tube law is proposed. The presented results are relevant to the study of self-excited oscillations and wheezing sounds in the lungs.

In the present study, the flow and acoustic characteristics of an aerospike nozzle supersonic jet at a Nozzle Pressure Ratio (NPR) = 3, three Temperature Ratios (TR) = 1, 3, 7 and two Swirl Numbers S = 0.10, 0.20 are presented. Implicit Large Eddy Simulations (ILES) are deployed to simulate the aerospike nozzle flow. The far-field aeroacoustic signature is computed based on the Ffowcs Williams-Hawkings (FWH) equation. In the vicinity of the aerospike bluff body, a shock-cell structure is formed for all the configurations. The shock strength and length as well as the pressure fluctuations are primarily affected by the TR in that jet region. The supersonic flow reattaches further downstream towards the aerospike bluff body as the TR increases at a fixed Swirl Number. This influences in particular the flapping motion of the annular shock-cell structure. The latter is characterized by power spectral density of the radial velocity at well-chosen monitoring points in that region. Subsequently, two-point cross-correlations in the annular jet shear layer are computed to detect azimuthal jet modes. The azimuthal jet excitation increases in amplitude with increasing Swirl Number S, leading to high Sound Pressure Levels at the Strouhal numbers observed. Downstream of the aerospike bluff body, a short circular shock-cell structure is observed at Swirl Number S = 0.10 for higher TR while the jet remains annular in the cold case. At S = 0.20, fewer shock cells are formed downstream of the aerospike bluff body. The shortening of the shock-cell structure leads to screech elimination at both Swirl Numbers. Further crosscorrelations for the axial velocity in the jet shear layers show supersonic convection velocities at Temperature Ratios (TR) = 3 and 7 for both Swirl Numbers which confirms the presence of Mach waves observed in the near-field snapshots. The Mach wave radiation features a slight helical propagation pattern in contrast with the baseline case without swirling motion. Furthermore, a skewness larger than 0.4 and a positive kurtosis of the pressure signals along the Mach wave radiation lines indicate crackle noise at TR7. The far-field spectra computed with the Ffowcs Williams-Hawkings equation display mixing noise only for S = 0.10. In the cold cases, high SPL are detected in agreement with the Broadband Shock-Associated Noise (BBSAN) central frequencies which were computed using the annular and circular shock-cell length. Additionally, high SPL are obtained due to Mach wave propagation, at the Strouhal numbers of the azimuthal modes and of radial motion of the annular shock-cell structure.

Supersonic nozzles are not always operated at design conditions. This leads to the formation of shock-cell structures and potentially higher Sound Pressure Levels as compared with the design conditions. Additionally, the total pressure, temperature and velocity distributions at the nozzle inlet plane are characterized by strong inhomogeneities, conditions dictated by the operating regime of the turbine or combustion chamber. In particular, a swirling flow motion can be induced by these components. Using flow control to induce a swirling motion to the jet can also act as a noise suppression technology. While homogeneous inflow conditions are well documented for a large range of supersonic nozzles, data on the aeroacoustic signature of supersonic swirling jets is scarce. Implicit Large Eddy Simulations are deployed to simulate the swirling flow of a non-ideally expanded three-dimensional, axisymmetric aerospike nozzle at a Nozzle Pressure Ratio (NPR) = 3 and Temperature Ratio (TR) =1. Three swirl numbers are considered. Near-field acoustic analyses are completed by far-field acoustic computations based on the Ffowcs Williams-Hawkings (FWH) equation. The swirling flow cases are compared with the baseline case without swirling. The lowest swirl number considered (8 = 0.10) leads to a shortening of the potential core of the jet and a reduced shock cell count. At higher swirl numbers, the circular shock-cell structure is eliminated. Moreover, the shock cell length in the annular part of the jet is increased compared to the baseline case. Because of the additional swirling component, the flow reattaches further downstream on the aerospike bluff body. Two-point space-time cross-correlations of pressure data acquired in the annular shear layer indicate an enhancement of the azimuthal modes, leading to helical acoustic wave propagation in upstream direction. The corresponding Strouhal number increases with increasing swirl number. Furthermore, two-point space-time cross-correlations of pressure data acquired in the circular jet shear layer downstream of the aerospike bluff body show that swirling suppresses screech tones. Power spectral density (PSD) of the radial velocity at monitoring points at the location of the separation bubble and shocks in the vicinity of the nozzle trailing edge allows to identify the oscillation modes of the annular shock-cell structure. The lengthening of the separation bubble leads to a reduced Strouhal number of the radial oscillation mode. Moreover, the passage of vortical structures through the shock cells leads to Broadband-Shock Associated Noise (BB SAN) whose central frequency depends on the shock cell length and on the convection velocity of vortical structures in the shear layer. The latter decreases with increasing swirl number. The far-field spectra display mixing noise at low Strouhal numbers for the baseline case and the lowest swirl number, as well as BBSAN. High Sound Pressure Levels (SPL) are detected in agreement with the BBSAN central frequencies. Moreover, high SPL are obtained at the radial oscillation frequencies of the annular shock-cell structure and at the Strouhal number of the azimuthal modes. The global SPL decreases while the nozzle thrust remains at 99 % of the thrust of the baseline case for the low swirl numbers.

In this paper, the flow and acoustic characteristics of an aerospike nozzle jet at a Nozzle Pressure Ratio (NPR) = 3 and three different Temperature Ratios (TR) = 1, 3, 7 are presented. Implicit Large Eddy Simulations (ILES) are deployed to simulate the flow of the aerospike nozzle. The LES calculations are completed by aeroacoustic computations based on the Ffowcs Williams-Hawkings (FWH) equation. In the supersonic jet exhausting an aerospike nozzle, two shock-cell structures are observed: an annular and a circular one. In the direct vicinity of the annular nozzle, the annular jet is non-attached and reattaches further downstream at increasing distance with increasing jet Temperature Ratio (TR). The observed shock cells in the non-attached annular jet structure are longer for TR3 and TR7 compared to TR1. The shock strength in the annular jet is increasing with increasing jet TR. In the meantime, the total number of shock cells in the circular part of the jet decreases with increasing jet TR. Pressure spectra in the near-field show the presence of a strong tonal noise at upstream angles corresponding to screech tones. Two-point cross-correlations of pressure data acquired in monitoring points located along axial lines in the circular shear layer are computed to quantify the upstream propagating waves associated to this tonal component. Power spectral density of the radial velocity at several monitoring points located at the shock cells as well as at the separation bubble, highlights the main oscillation modes of the annular shock-cell structure. Supersonic convection velocities of turbulent structures are detected with increasing jet temperature by means of two-point cross-correlation. This confirms the presence of Mach waves observed in the instantaneous snapshots. The Mach waves radiation angles are in agreement with existing models. In the far-field spectra, the highest Sound Pressure Levels (SPL) are associated to those Mach waves at angles around 140° (TR3) and 120° (TR7). High skewness and kurtosis in the pressure signals indicate crackle noise at higher jet temperatures. Additionally, the shock-cell length is used to predict the central frequency of Broadband Shock-Associated Noise (BBSAN) as a function of observation angles. The far-field spectra display mixing noise as well as BBSAN, related to the interaction between the convected vortices in the shear layers and the shock-cell structure. With increasing jet temperature, higher SPL are detected in agreement with the BBSAN central frequencies which were computed using the annular and circular shock-cell length.