Aeroacoustic Feedback Research Articles

Aeroacoustic analysis of flow induced resonance in a gas pipeline

Flow induced noise and vibration in the pipework of gas production platforms can lead to fatigue failures of attached small-bore pipes, and this can constrain the operating conditions of the platform. Typically such problems comprise a source of turbulence in the flow, an aeroacoustic feedback mechanism due to acoustic resonances inside the pipework, and mechanical resonances which amplify the resulting vibration. The problem described here occurred in the gas metering system of the platform, with the source located in the header which comprised a bifurcation into two metering lines and a number of sharp bends. Because of the cost of rectifying the problem an in-depth study was carried out to analyse the flow and induct acoustic conditions. The predictions will support recommendations for an efficient redesign of the entire system to control the problem at source. Various levels of CFD modelling (RANS and URANS) were carried out using the Openfoam code to predict the exact location and mechanism of the vortex shedding. The acoustic response of the complex ductwork was then modelled using the ACTRAN code to understand the feedback mechanism which led to resonance and high levels of response.

Deep cavity noise suppression by exploiting aeroacoustic–structural interaction of multiple elastic panels

This paper reports a numerical study of a novel methodology for passive suppression of deep cavity noise by means of strategically designed and arrangements of multiple elastic panels and examines its underlying aeroacoustic–structural interaction physics. The study is conducted with a freestream, at Mach number 0.09 and Reynolds number of 4 × 104 based on the cavity length, past a two-dimensional cavity by means of direct aeroacoustic simulation coupled with a panel dynamic solver in monolithic fashion. For each cavity-panel configuration, the fluid-loaded panel natural frequencies are harmonized with the characteristic aeroacoustic processes of the original/modified cavity aeroacoustic feedback loop. This promotes panel aeroacoustic-structural resonance for absorption of feedback flow and acoustic fluctuation energy for achieving less eventual cavity noise. The most effective configuration gives a remarkable noise power reduction by 15 dB from a rigid cavity. Inadvertently, it reduces cavity drag by almost 15%. Simultaneous reduction of both cavity noise and drag is unprecedented among similar attempts in the literature. In-depth spatiotemporal analyses of aeroacoustic–structural interaction results elucidate the intricate interplay between cavity flow, panel vibration responses, and cavity acoustic modes, leading to noise reduction in all cavity-panel configurations studied. Essentially, the vertical panel acts to curtail the efficacy of coupling between growing shear layer and cavity acoustic modes whose sustenance is further impeded by an acoustically induced resonant panel at the cavity bottom. The proposed methodology is confirmed to be feasible yet effective, which holds great potential for fluid-moving applications in which a quiet and energy-efficient cavity configuration is desired.

Experimental study on acoustic resonance of subsonic and slightly underexpanded impinging jets

The aeroacoustic feedback loops in high-speed circular jets that impinge on a large flat plate are investigated via acoustic measurements and schlieren visualizations. In the present experiments, the nozzle pressure ratio ranges from 1.39 to 2.20, the corresponding ideally expanded jet Mach number $M_j$ is from 0.70 to 1.12 and the nozzle-to-plate distance ( $H$ ) is from 4.0 $D$ to 6.0 $D$ , where $D$ is the nozzle exit diameter. The results of acoustic measurements show that the strongest tones are generated in a limited frequency band. The empirical dispersion relations obtained from the fluctuating greyscales along the jet centreline of time-resolved schlieren images have good agreement with the dispersion relations from the vortex-sheet model. The coherent flow structures at tonal frequencies are extracted by spectral proper orthogonal decomposition and are analysed in detail. For the $M_j<0.82$ jets, the upstream-propagating guided jet mode is progressively confined to the potential core of jets with increasing tonal frequency, which provides the first direct experimental support for theoretical results. The evolution in the structures of acoustic resonance loops is studied along a single frequency stage of axisymmetric impinging tones. When the acoustic resonance between the upstream- and downstream-propagating guided jet modes is formed at tonal frequencies, the impinging tones are intenser. Slightly underexpanded impinging jets can simultaneously produce impingement tones and screech tones. Shock-cell structures have modulatory effects on the downstream-propagating Kelvin–Helmholtz wavepacket and the upstream- and downstream-propagating guided jet modes. Due to the interaction between the flow structures at the frequencies of impinging and screech tones, tones of axisymmetric modes can be produced outside the frequency ranges in which the axisymmetric upstream-propagating guided jet modes are supported by jets.

Airfoil Tonal Noise Prediction Using Urans

Abstract To examine the feasibility of the laminar boundary layer (LBL), vortex shedding (VS) tonal noise modelling using unsteady Reynolds-averaged Navier–Stokes (URANS) was investigated for the non-symmetric S834 airfoil. A transition SST turbulence model was used to model the laminar-turbulent transition and its vital influence on the laminar bubble and hydrodynamic instabilities generation. The influence of turbulence on the unsteady vortex patterns was investigated. Hence, the hybrid aeroacoustic analysis with Lighthill analogy was conducted to obtain the acoustic pressure field. The approach allowed us to model hydrodynamic instabilities and the resulting VS tonal noise. The frequency of VS matched the experimental data, giving the same 1/3 octave tonal peak only for a limited freestream turbulence regime. The simplification of the present method did not allow us to model the aeroacoustic feedback loop, and resulted in lack of instabilities for higher freestream turbulence.

Dynamic Pressure Measurements in a Rectangular Cavity with Multiple Stores

An experimental investigation is carried out to study the effects of Mach number on the acoustic resonance characteristics of a rectangular cavity with and without multiple stores. Measurements are conducted for a cavity with a length-to-depth ratio of 4.5 at Mach numbers of 2, 3, 4, and 5. Dynamic pressures are measured within the cavity and on the store surfaces. An increase in Mach number results in a decrease in overall sound pressure levels at all measurement locations within the cavity. Pressure spectra indicate that increasing the Mach number decreases both the peak and broadband amplitudes of the cavity pressure fluctuations. At Mach 5, the amplitude of the cavity modes is significantly reduced in the upstream half of the cavity, and the modes are either indistinguishable from broadband levels or absent in the downstream half. The addition of stores reduces cavity acoustics, with two stores offering greater reduction than a single store. The effect of Mach number and the presence of stores on the aeroacoustic feedback loop in a cavity align with literature. Additionally, the pressure fluctuations on the windward store surface scale with Mach number in a similar trend as the cavity acoustics.

Mach Number Dependence of Tone Generation by Impinging Round Jets

The Mach number dependence of tone generation by jets impinging on a plate is investigated using large-eddy simulations, for seven jets at Mach numbers between 0.6 and 1.3 and a nozzle-to-plate distance of 8 nozzle radii. For , the upstream sound radiation is broadband, whereas for , tones emerge in the near-nozzle spectra, highlighting the establishment of feedback loops. The tone frequencies are consistent with those for aeroacoustic feedback loops between the nozzle and the plate, and exhibit a staging behavior with the Mach number, staying inside or close to the ranges of the upstream-traveling free-stream guided jet waves predicted by a vortex-sheet model. The azimuthal and radial structures of the pressure fields at the tone frequencies agree with those of these waves closing the loops. The selection of the dominant tone and the staging behavior are discussed by evaluating the power gains of the shear-layer instability waves between the nozzle and the plate using linear stability. In most cases, the dominant tones emerge at the possible tone frequencies with the highest gain. For , the gain of the instability waves at the frequency of the guided jet waves most likely to close feedback mechanisms is lower than the gain peak values, possibly contributing to the absence of resonance.

Suppression of deep cavity aeroacoustics at low Mach number by localized surface compliance

A unique concept of utilizing localized surface compliance is proposed to suppress deep cavity aeroacoustics at a low Mach number. The core idea is to provide local absorption of the energy of aeroacoustic processes supporting cavity flow self-sustained feedback loop responsible for tonal noise generation. The concept is studied with a flow past cavity of length-to-depth ratio of 0.4 at freestream Mach number 0.09 and Reynolds number based on cavity length 4 × 104 using high-fidelity, two-dimensional direct aeroacoustic simulation. Having confirmed the replication of key aeroacoustic processes in the numerical solution through careful validation, localized surface compliance in the form of an elastic panel is strategically introduced to modify every process for cavity noise suppression. The panel natural frequency is set equal to the feedback loop characteristic frequency to facilitate its flow-induced structural resonance for energy absorption. Suppression of cavity noise pressure and power levels by 3.8 and 4.8 dB, respectively, is successfully achieved, together with an unforeseen cavity drag reduction by almost 19%. Comprehensive wavenumber–frequency analyses of the coupled aeroacoustics and flow-induced panel vibration are conducted to uncover the physical mechanism of noise suppression. The results show that the same type of aeroacoustic feedback loop occurs, but its efficacy is significantly reduced due to the exhaustion of aeroacoustic process energy to the flow-induced vibrating panel. The proposed concept is confirmed to be feasible in terms of giving remarkable cavity noise and drag suppression, yet it retains the basic problem geometry intact, which are considered important in many practical applications.

Angle-of-attack and Mach number effects on the aeroacoustics of an SD7003 airfoil at Reynolds number 60,000

The aeroacoustics of an SD7003 airfoil at Reynolds number 60,000 is investigated using Large Eddy Simulation. Five simulations are performed to study the effects of angle-of-attack and Mach number at fixed Reynolds number. For the three cases with angle-of-attack equal to 0° (M = 0.1, 0.3 and 0.6), a pure tonal noise associated with a 2D organisation of the flow is obtained. This flow topology is due to the establishment of a well known aeroacoustic feedback loop between the separation point on the suction side of the airfoil and the trailing edge. The occurrence of this loop is corroborated by the presence of a standing wave pattern with characteristic mode number in accordance with Panda’s model. The main effect of the Mach number is to promote flow separation and hence increase separation length and mode number. In addition, the first harmonic and the sub-harmonic of the tone, observed in the far field acoustic spectrum, are found to be generated in the wake, presumably due to non-linear vortex interactions. For the two other angles-of-attack 4° and 8° at M = 0.1, the feedback loop does not establish and a Laminar Separation Bubble (LSB) is observed. When increasing the angle-of-attack, the LSB shrinks with earlier reattachment. For those two cases, far-field spectra are characterized by a low frequency associated with the breathing motion of the LSB and the reattachment point fluctuating in space. The frequency of this fluctuation depends on the curvature of the bubble. Far-field spectra are also characterized by a broadband trailing edge noise whose frequency range decreases with the angle-of-attack. Again, this evolution is found to depend on the curvature of the bubble which may promote a centrifugal instability in the separated shear layer.

Effect of Mach number on the aeroacoustic feedback loop generating airfoil tonal noise

Airfoil tonal noise emission at low-to-moderate Reynolds number and flow conditions featuring a laminar separation bubble close to the trailing edge is often related to an aeroacoustic feedback mechanism and, therefore, the Mach number is a primary parameter for the flow field and noise generation. This study experimentally explores the effect of the Mach number on airfoil tonal noise generation in the nominally incompressible flow regime. Using airfoil profiles of different chord lengths, the Mach number is varied for a constant Reynolds number. Acoustic and flow field measurements for a range of combinations of Reynolds and Mach numbers were conducted. At zero incidence, the tonal noise regime in the Reynolds number domain is found to be sensitive to the Mach number. At non-zero angle of attack (2°), the noise generation is found to be dominated by vortex shedding over a separation bubble on the pressure side. The details of the separation bubble and shedding process depend on the Mach number. The frequencies of the dominant tones and the frequency intervals between tones increase with the Mach number. Moreover, the measured frequency interval can be collapsed using a relation based on the aeroacoustic feedback loop model. The relation is rewritten to separate the effects of the Reynolds and Mach numbers. As a result, the dependence on the Mach number is identified and tested. In contrast, the tonal noise level shows a more complex dependence on details of the laminar separation bubble and the vortex shedding process.

Spectral proper orthogonal decomposition of coupled hydrodynamic and acoustic fields: Application to impinging jet configurations (draft)

A large variety of flow configurations involve an aero-acoustic coupling and the complexity of the interaction can make the analysis difficult. In the purpose to decompose these flows in simpler phenomena, the spectral proper orthogonal decomposition (SPOD) algorithm can be used. Compared to POD, this method adds up a discrete Fourier transform before the singular value decomposition. A theoretical framework for the weighting of the flow field variables used in the SPOD has been proposed to identify hydrodynamic modes by Schmidt et al. However, in flow configurations where both the hydrodynamic and acoustic fields interact, without additional data processing, only hydrodynamic modes that are generally dominants would be identified by the SPOD. This paper proposes a general framework based on a few pre-processing steps applied to the input aero-acoustic fields to facilitate the identification of both the hydrodynamic and acoustic fields by the SPOD, the purpose being to analyze how the hydrodynamic waves lead to the acoustic ones and the possible feedback process. This framework is applied to large-eddy simulations of supersonic ideally expanded impinging planar jet of infinite extent characterized by different nozzle-to-plate distances and incidence angles of the flat plate. The analysis shows that the SPOD algorithm is able to capture two important aero-acoustic feedback loops of impinging jet configurations: the one closed by sound waves traveling upstream (related to the jet impingement on the flat plate) and the one characterized by neutral wave modes of the jet shear layer itself. The analysis has also shown that the symmetrical/anti-symmetrical acoustic patterns radiated with respect to the jet main axis and due to the jet impingement are mainly correlated to the axial velocity fluctuations structures in the jet for the normally impinging cases. For the cases with flat plate incidence, the acoustic radiated is mainly correlated to vertical velocity fluctuations structures developing in the upper shear layer of the jet.

Investigating Shear-Layer Instabilities in Supersonic Impinging Jets Using Dual-Time Particle Image Velocimetry

Flow unsteadiness in supersonic impinging jets due to the aeroacoustic feedback can lead to practical problems like ground erosion in short takeoff and vertical landing aircraft, high noise levels, and high unsteady structural loading. The shear-layer instabilities of the jet play an important role in the flow dynamics. A full study of these instabilities requires time-resolved measurements of the flow velocity field. However, experimentally obtaining time-resolved velocity fields, derived from particle image velocimetry (PIV) techniques, has its challenges in high-speed flow applications due to the simultaneous requirement of high spatial and temporal resolution. A novel approach is presented in this paper, which uses time-unresolved dual-time digital PIV coupled with dynamic mode decomposition (DMD) analysis (DT–DPIV–DMD) of the velocity field pairs and applies it to study the dynamics of the shear-layer instabilities in supersonic impinging jets. DT–DPIV–DMD involves the acquisition of a pair of instantaneous velocity fields separated by a small time delay using two covisual PIV systems. The time delay must be sufficiently small to provide the necessary bandwidth to capture all the relevant dynamics of the flow being investigated. This is followed by DMD analysis of all acquired digital PIV velocity field pairs. A large-eddy simulation of an underexpanded impinging jet at Reynolds number 60,000, with nozzle-to-wall distance was first used to investigate the required measurement scales for the application of DT–DPIV–DMD. Using the optimum measurement scales, the technique was then experimentally demonstrated using DT–DPIV–DMD for two supersonic impinging jets cases and validated with acoustic measurements.

Generation of acoustic tones in round jets at a Mach number of 0.9 impinging on a plate with and without a hole

The generation of acoustic tones in four round jets at a Mach number of 0.9 impinging on a plate at a distance $L=6r_0$ from the nozzle exit, where $r_0$ is the nozzle radius, has been investigated by large-eddy simulation. Three plates are perforated by holes of diameters $h=2r_0$ , $3r_0$ and $4.4r_0$ , centred on the jet axis, whereas the fourth plate has no hole, in order to study the effects of the hole on the jet flow and acoustic fields. In all cases, acoustic tones emerge in the jet near field, upstream but also downstream of the plate for the perforated plates. Their frequencies are similar for all jets and close to those predicted for aeroacoustic feedback loops establishing between the nozzle and the plate, involving flow disturbances convected downstream and waves propagating upstream at the ambient speed of sound. Their levels, however, decrease with the hole diameter, by a few dB for $h\leqslant 3r_0$ but approximately by 40 dB for $h=4.4r_0$ . The features of the feedback loops are identified by computing two-dimensional space–time correlations and frequency–wavenumber spectra of the pressure fluctuations in the jet mixing layers. These loops are found to be closed by free-stream upstream-propagating guided jet waves, produced by the impingement of vortical structures on the plate for $h\leqslant 3r_0$ and by the scattering of the jet aerodynamic pressure fluctuations by the hole edges for $h=4.4r_0$ . Finally, an acoustic analogy is used to show that the contributions of the pressure fluctuations on the plate to the radiated noise are dominant.

Aeroacoustic Instability Analysis of Helmholtz Resonators Through High-Order Unsteady Reynolds-Averaged Navier–Stokes Simulations

This paper presents a numerical analysis of the aeroacoustic instabilities in a slit-type Helmholtz resonator subjected to a low Mach number grazing flow. This configuration can be susceptible to nonlinear feedback phenomena due to the mutual coupling between the aerodynamic and acoustic fields in the proximity of the slit. The aeroacoustic flowfield is simulated through a direct approach by solving the unsteady compressible Reynolds-averaged Navier–Stokes equations using a high-order discontinuous Galerkin discretization. The response of the Helmholtz resonator to an incident acoustic field is analyzed using a two-port characterization technique. The onset of the aeroacoustic instabilities is investigated by applying an acoustic energy criterion based on the scattering matrix formulation. The Nyquist stability diagram is studied to determine the system stability and identify the whistling occurrence. The two- and three-dimensional numerical results are compared to experimental data obtained from a dedicated measurement campaign. The comparison between the numerical and experimental data demonstrates a very good agreement, highlighting the potential of the applied numerical approach for the investigation of the aeroacoustic instabilities. This allows a deep description of the aeroacoustic feedback mechanisms leading to the whistling onset, proposing a valid engineering tool for the early design stage.

Aero-acoustic feedback mechanisms in supersonic cavity flow with subcavity

An open rectangular cavity with a length to depth ratio of 2 (L/D = 2) and inlet Mach number 1.71 was considered as a baseline configuration for an unsteady computational analysis. Fast Fourier transform using power spectral density, correlation, and overall sound pressure level (OASPL) was carried out on the unsteady pressure data, to analyze the flow dynamics. Many flow visualization techniques were employed such as density gradient-based numerical schlieren, vorticity contour, and streamlines for revealing major features associated with the flow. Vortex shedding visualization was carried out in terms of the Lambda 2 (λ2) criterion, where the vortex core can be observed moving downstream in the shear layer. In the acoustic pressure contours and OASPL contours, the characteristics of the acoustic waves can be distinctly observed traversing within the cavity. The analysis was extended for different flow conditions and configurations. The front wall subcavity with a ratio of subcavity length (l) to cavity length (L) l/L = 0.20 has already been established as a passive control device experimentally whereas the aft wall subcavity with l/L = 0.20 was found to behave like a passive resonator with distinct harmonic fluid-resonant modes. A more detailed investigation of these configurations with different dimensions and combinations resulted in a comparative and better understanding of the flow features, mode frequencies, and fluid-resonant oscillations in a supersonic cavity flow. Further, a method for designing front wall subcavity for passive control device has been applied to a variety of test cases and has been proven effective in suppressing the oscillations for cavity configurations with longitudinal oscillation mechanisms.

Predicting high-speed feedback mechanisms in rectangular cavities using lattice-Boltzmann very-large eddy simulations

Aeroacoustic feedback mechanisms in rectangular cavities are strongly dependent on length-to-depth ratio and free stream velocity. The capability to predict the effects of these parameters using a scale-resolving transonic lattice-Boltzmann flow simulations is assessed, for the first time, by considering the M219 benchmark configuration for two different length-to-depth ratios (10:1 and 5:1) and two free-stream Mach numbers (0.85 and 1.35). Numerical results are compared and validated using unsteady wall pressure measured in wind tunnel tests. Tonal and broadband components of the pressure fluctuations are properly captured. For the shallow cavity configuration, an increase in Mach number is found to enhance the development of Rossiter modes inside the cavity. A more detailed analysis of these Rossiter modes is carried out for the shallow cavity at supersonic speed, for which physical insight is gained through the usage of spectral proper-orthogonal decomposition and wavelet/Fourier analyses.

Whistling of deep cavities subject to turbulent grazing flow: intermittently unstable aeroacoustic feedback

Abstract

Passive airfoil tonal noise reduction by localized flow-induced vibration of an elastic panel

In this paper a novel passive control method is explored for airfoil tonal noise reduction using localized flow-induced vibration of a short elastic panel flush mounted on suction surface of a NACA 0012 airfoil at low Reynolds number. The key idea is to provide local absorption of energy of natural instabilities evolving in the laminar boundary layer by self-sustained flow-induced vibration of a properly designed panel. The panel ultimately acts to weaken the aeroacoustic feedback loop responsible for airfoil tonal noise radiation. Perturbation evolution method with acoustic broadband excitation is utilized to assess the tonal noise reduction potential with various combinations of elastic panel design parameters and optimal resonating and non-resonating elastic panel configurations are determined. Subsequently, direct aeroacoustic simulation of airfoil flow with optimal panel configurations are carried out to uncover the mechanism of tonal noise reduction using localized flow-induced vibration in a quantitative manner. Extensive analysis of numerical results reveals that a resonating panel located just ahead of the sharp growth of boundary layer instability within the airfoil separation bubble provides the strongest reduction of flow instabilities and an overall tonal noise reduction up to 3 dB. Such significant noise reduction is achieved without any sacrifice in the original aerodynamic characteristics of the airfoil. The outcomes of the study evidently suggest that the proposed passive control method with a localized flow-induced panel vibration is effective in suppressing the fundamental mechanism responsible for tonal noise generation of airfoil flow, making it a promising approach for modifying the acoustics of existing aerodynamic or wing profile operating at low Reynolds numbers.

Application of a POD-Galerkin based method to time resolved and time unresolved data for the determination of the Convective Velocity of Large-Scale Coherent Structures in High Speed Flows

Motivated by the aero-acoustic feedback loop phenomenon in high speed free jets and impinging jets, a thorough examination of a POD (Proper Orthogonal Decomposition)-Galerkin method to determine the average convection velocity of coherent structures in the shear layer is presented in this paper. The technique is shown to be applicable to both time resolved as well as time unresolved data, if the data set meets certain requirements. Using a detailed sensitivity analysis on a synthetic data set, a quantitative estimate on the required time resolution for the technique has been found, which can be useful for both experimental, as well as numerical studies investigating the aero-acoustic feedback loop in high speed flows. Moreover, some innovative ways to apply the technique are also demonstrated using a simulated data set, showing the effectiveness of the technique to any general problem in supersonic jets, heat transfer, combustion or other areas in fluid mechanics, where an advection process can be identified.

Flow dynamics and wall-pressure signatures in a high-Reynolds-number overexpanded nozzle with free shock separation

A delayed detached eddy simulation of an overexpanded nozzle flow with shock-induced separation is carried out at a Reynolds number of , based on nozzle throat diameter and stagnation chamber properties. In this flow, self-sustained shock oscillations induce local unsteady loads on the nozzle wall as well as global off-axis forces. Despite several studies in the last few decades, a clear physical understanding of the factors driving this unsteadiness is still lacking. The geometry under investigation is a subscale truncated ideal contour nozzle, which was experimentally tested at the University of Texas at Austin at a nozzle pressure ratio of 30. Under these conditions, the nozzle operates in a highly overexpanded state and comprises a conical separation shock that merges to form a Mach disk at the nozzle centre. The delayed detached eddy simulation model agrees well with the experimental results in terms of mean and fluctuating wall-pressure statistics. Wall-pressure spectra reveal a large bump at low frequencies associated with an axisymmetric (piston-like) motion of the shock system, followed by a broad and high-amplitude peak at higher frequencies associated with the Mach waves produced by turbulent eddies convecting through the detached shear layer. Moreover, a distinct peak at an intermediate frequency ( ) persists in the wall-pressure spectra downstream of the separation shock. A Fourier-based analysis performed in both time and space (azimuthal wavenumber) reveals that this intermediate-frequency peak is associated with the (non-symmetric) pressure mode and is thus related to the generation of aerodynamic side loads. It is then shown how the unsteady Mach disk motion is characterized by an intense vortex shedding activity that, together with the vortical structures of the annular shear layer, contributes to the sustainment of an aeroacoustic feedback loop occurring within the nozzle.

Uncertainty Quantification for Direct Aeroacoustic Simulations of Cavity Flows

We investigate the influence of uncertain input parameters on the aeroacoustic feedback of cavity flows. The so-called Rossiter feedback requires a direct numerical computation of the acoustic noise, which solves hydrodynamics and acoustics simultaneously, in order to capture the interaction of acoustic waves and the hydrodynamics of the flow. Due to the large bandwidth of spatial and temporal scales, a high-order numerical scheme with low dissipation and dispersion error is necessary to preserve important small scale information. Therefore, the open-source CFD solver FLEXI, which is based on a high-order discontinuous Galerkin spectral element method, is used to perform the aforementioned direct simulations of an open cavity configuration with a laminar upstream boundary layer. To analyze the precision of the deterministic cavity simulation with respect to random input parameters, we establish a framework for uncertainty quantification (UQ). In particular, a nonintrusive spectral projection method with Legendre and Hermite polynomial basis functions is employed in order to treat uniform and normal probability distributions of the random input. The results indicate a strong, nonlinear dependency of the acoustic feedback mechanism on the investigated uncertain input parameters. An analysis of the stochastic results offers new insights into the noise generation process of open cavity flows and reveals the strength of the implemented UQ framework.