Implicit Large Eddy Simulations (ILES) are deployed to characterize the effect of swirling boundary conditions on vorticity transport as well as on the Lighthill's tensor for a cold, supersonic aerospike nozzle jet. Four jets are simulated at a Nozzle Pressure Ratio (NPR) = 3, one jet without swirl and three jets with swirl numbers S=0.10,0.20,0.30. Swirling boundary conditions lead to an increase in vorticity tilting and stretching downstream of the aerospike bluff body, potentially enhancing sound generation. An exact decomposition of the Lighthill's tensor is undertaken and the magnitude of the obtained source terms is compared. The dominant acoustic source terms are amplified under swirling boundary conditions. Terms describing the interactions between dilatation fields and density gradients balance each other in shock regions. At higher swirl numbers, the convection of density gradients along the flow direction leads to an imbalance that contributes to increased sound generation. Finally, cross-correlations between the near-field pressure and individual source terms reveal that enstrophy correlates more strongly with the near-field acoustics at higher swirl numbers.

In this work, we propose a multi-fidelity approach to optimize the geometry of a small fixed-pitch propeller to reduce its acoustic impact. This way, the advantages of low-fidelity methods (low cost and fast results) can be combined with those of higher-order ones (better resolution and more reliable results) to obtain an optimized solution within a reasonable timeframe. We train a fully connected network surrogate that maps propeller geometry (chord and twist) and thrust to optimal rotating speed, torque, and sound pressure level (SPL) at the blade passing frequency (BPF) at 17 different angles, and benchmark it on three datasets: low-order blade element momentum theory (BEMT), higher-order vortex particle method (VPM), and a balanced Mixed set.

This work examines how inviscid interface-flux choices and Local Discontinuous Galerkin (LDG) viscous-flux parameters affect solution sensitivity and computational cost in high-order Discontinuous Galerkin Flux Reconstruction (DG-FR) simulations of compressible turbomachinery flows. A transonic low-pressure turbine cascade at M2is=1.22 and Re2is=1.5×105 is considered using a range of inviscid Riemann solvers, including Rusanov, HLLC-type solvers, RoeM and an Exact solver. The LDG discretization is varied through the penalty parameter τ and the directional parameter β. In addition to standard formulations, where the same inviscid flux is used on all interfaces, hybrid HF-R-X formulations are investigated, with Rusanov applied at wall boundaries and solver X applied on interior interfaces. The comparison is based on integrated aerodynamic forces, surface distributions of isentropic Mach number and skin friction, loss coefficients evaluated from inlet and outlet sampling planes, base pressure and normalized wall-clock time. The results show that, for the operating condition and resolution considered, lift and drag are only weakly sensitive to the inviscid Riemann solver, with variations remaining below approximately half a percent across the converged cases. Surface pressure and skin-friction trends are also nearly unchanged among the approximate solvers. Loss-related quantities are more sensitive to the numerical formulation and provide a clearer distinction between inviscid-flux and LDG-parameter choices. From a performance perspective, approximate inviscid solvers introduce only marginal overhead relative to the Rusanov baseline, whereas the Exact solver increases the wall-clock time more noticeably. The hybrid HF-R-E formulation reduces part of this cost penalty by avoiding Exact flux evaluations at wall boundaries. The dominant runtime sensitivity is associated with the LDG directional parameter: centered viscous paths with β=0 produce a substantial cost increase compared with biased paths with β=±0.5. Overall, the results indicate that inexpensive approximate inviscid fluxes are sufficient for integrated load prediction in this case, while loss prediction and computational efficiency are more strongly affected by the LDG viscous-flux construction and by the use of hybrid flux strategies.

Implicit large eddy simulations are deployed to simulate the flow of an aerospike nozzle jet at a nozzle pressure ratio = 3 and four temperature ratios (TR) = 1, 3, 4, and 7. The large eddy simulation (LES) calculations are completed by aeroacoustic computations based on the Ffowcs Williams-Hawkings equation. In the supersonic jet exhausting the aerospike nozzle, two shock-cell structures are observed: an annular and a circular one. Higher TR leads to longer shock cells in the annular jet region. In the meantime, the shock cell count in the circular part of the jet decreases with increasing jet TR. Supersonic convection velocities of turbulent structures are detected with increasing TR by means of two-point cross-correlations, indicating the presence of Mach waves. Furthermore, high skewness and kurtosis of the pressure signals along the Mach wave propagation direction indicate crackle noise at higher TR. Screech noise is observed in the near-field signature of the supersonic jets. This sound component has a pronounced multi-peak behavior. Spectral proper orthogonal decomposition of the azimuthally decomposed pressure fields and space–time Fourier transforms of the pressure fluctuations are used to identify the most amplified wavepackets propagating downstream and the upstream waves responsible for closing the screech feedback loop. The simulation results exhibit close correspondence with the theoretical neutral waves predicted by the vortex sheet model. In the cold case, the main screech feedback loop involves only the annular shock-cell structure, whereas at higher TR, both shock-cell structures contribute to the feedback mechanism.

The study examines supersonic square jets in a twin nozzle configuration with the aim of identifying and characterising emergent instability modes during overexpanded operation. Unlike screeching rectangular jets that undergo strong fluctuations normal to the wider jet dimension, the equilateral nature of the exit geometry in square nozzles leads to multiple instability states dictated by shock–turbulence interactions and nozzle operating conditions. Furthermore, strong coupling modes between the jets were identified that led to either phase locked or out of phase interactions of the inner shear layers. Results from experimental studies were examined using spatial and temporal decomposition techniques based on spectral methods to identify the resultants from triadic shock–turbulence interactions. The primary instability mode across both operating conditions were driven by optimal interactions while the harmonics were found to be associated with the suboptimal shock–turbulence interactions.

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.

In this study, we focus on the simulation of transonic flows in the context of turbomachinery applications, employing a high-order Discontinuous Galerkin Flux Reconstruction (DG-FR) methodology. Conducting high-order simulations within the context of transonic turbines is challenging due to the presence of extreme unsteadiness in the flow accompanied by fluid compressibility effects, such as transonic vortex shedding and unsteady shock wave-boundary layer interactions. These phenomena significantly influence the aerodynamic performance of the airfoils playing a major role also in the determination of the losses in the flow. The primary objective of this work is to evaluate the impact of the formulated non-reflective boundary conditions and of a recently proposed shock capturing scheme on crucial aerodynamic predictions, including the isentropic Mach number and the skin friction distribution. The non-reflective boundary condition is specifically designed to leverage modern Graphical Processing Unit (GPU) architectures, aiming to minimize the necessity for blocking communication. The outcomes of this new outlet boundary condition will be thoroughly analyzed in terms of both computational cost and its influence on the resolved flowfield. The developed physics-based shock-capturing scheme is particularly attractive due to its reduced computational cost and ease of implementation and usage. The focus will be on assessing its performance in the context of highly unsteady flows, such as the current low-pressure turbine cascade. The effectiveness of the proposed computational setup will be rigorously examined to ensure its suitability for simulating complex transonic flows in turbomachinery applications.

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.