The exhaust plume produced by a planar, multistream supersonic nozzle is studied with and without the effect of an aft deck and sidewalls. Measurements encompass static wall pressure of the internal flow, shadowgraphy images of the exhaust plume, and far-field acoustics for a range of pressure ratios. An experimentally validated Reynolds averaged Navier–Stokes model of the internal flow is then used to reveal how the aft deck and sidewalls allow the expanding gas to persist longer, thereby extending the location of the separation shock. Where the far-field sound is concerned, the aft deck and sidewall attachments cause a noticeable reduction in overall sound pressure levels by as much as 5 dB at shallow angles and sideline observer positions. Many of the tones associated with screech and broadband shock associated noise are also absent with the addition of the aft deck and sidewalls.
The design and construction of various high-speed digital optical flow techniques forstudying the aeroacoustics and turbulence dynamics, as it relates to supersonic/hypersonicflow phenomena, is discussed. The three systems comprise retroreflective shadowgraphy, a z-type schlieren and a focused schlieren system. The performance of the three systems are examined using an axisymmetric Mach 3 flow, which has received considerable attention at the University of Texas at Austin for understanding the sound produced by flows with supersonic convective acoustic Mach numbers. Various techniques to aid in the setup of optical elements in conjunction with the high resolution digital cameras, capable of producing a million frames per second, are described. Various analysis methods are then employed (wavenumer-frequency transforms and wavelet analysis) to quantify the dynamical nature of the flow and sound field produced by this supersonic nozzle.
A first principles understanding of the sound field produced by multirotor drones in hover is presented. Propeller diameters ranging from 8 to 12 in. are examined and with configurations comprising an isolated rotor, quadcopter, and hexacopter configuration. The drone pitch, defined as the ratio of drone diameter to rotor diameter, is the same for all multirotor configurations and is valued at 2.25. A six-degree-of-freedom load cell is used to assess the aerodynamic performance of each configuration, whereas an azimuthal array of 1∕2 in. microphones, placed between two and three hub-center diameters from the drone center, is used to assess the acoustic near field. The analysis is performed using standard statistical metrics such as sound pressure level and overall sound pressure level and is presented to demonstrate the relationship between the number of rotors, the drone rotor size, and its aerodynamic performance (thrust) relative to the near-field acoustics.
The design, construction and preliminary measurements of a new test stand for accurately assessing the shear stress acting at the fluid surface interface of wall bounded flows is discussed. This stand is based on control volume analysis whereby a fully developed turbulent velocity profile produces shear forces which equate to the pressure drop measured between fixed points in a constant area pipe. The calibration stand is designed to facilitate both subsonic and supersonic flow. Subsonic flow conditions are achieved by placing different diameter nozzles at the exhaust of the test section thereby permitting different free stream velocities and mass flow rates for a given ratio of the total pressure to static pressure in the pipe. The advantages of this facility is in its ability to produce a broadrange of Reynolds numbers (based on centerline velocity and pipe diameter) and elevatedpressures that are required to gauge the sensitivity of modern shear stress sensors.
Reduced-order models of the airwake produced by the flow over a simple frigate shipare developed using POD based methods. The focus is to understand the trade spacebetween cost and accuracy, where different forms of the POD technique are concerned.Of particular importance is the upfront expense of employing ‘classical’ or snapshot formsof the POD technique in both scalar and vector forms using either time suppressed data(conventional-POD), or kernels constructed from cross-spectral densities of the fluctuatingvelocity. The latter approach is referred to as harmonic-POD so as not to exclude harmonictransforms in space. The flow over a simple frigate ship is an ideal test bed given that it isunsteady, three-dimensional, inhomogenous in all spatial directions, and stationary in time. The spatial modes from all three techniques are shown to correspond to unique time-scales, thereby demonstrating how the preservation of the temporal behavior associated with a particular spatial scale is not unique to the harmonic-POD approach.
A framework for using continuous wavelet transforms to isolate and extract blade–vortex interaction noise from helicopter acoustic signals is described. The extraction method allows for the investigation of blade–vortex interactions independent of other sound sources. Experimentally acquired acoustic data from full-scale helicopter flyover tests are first transformed into time-frequency space through the wavelet transformation, with blade–vortex interactions identified and filtered by their high-amplitude, high-frequency impulsive content. The filtered wavelet coefficients are then used to create a pressure signal solely related to blade–vortex interactions. Analysis of a synthetic data set is conducted and shows that blade–vortex interactions can be accurately extracted so long as the blade–vortex interaction wavelet energy is comparable to the wavelet energy in the main rotor harmonic.
Multirotor drones are becoming increasingly popular in both the civilian and military sectors of our society. These compact gadgets come in a variety of sizes with the smallest ones measuring less than two inches in diameter, while larger ones can be in excess of five feet. Surprisingly, very little is known about their acoustical footprint, which is becoming a topic of broad importance given that these vehicles most often operate in populated areas. Thus, the objective of this paper is to provide a first principles understanding of the acoustical characteristics of hovering drones. To accomplish this, a new test stand was constructed at the Applied Research Laboratories at The University of Texas at Austin for studying various multirotor drone configurations. The drone test stand is capable of powering up to eight DC electric motors with adjustable arms that allow different rotor diameters to be tested. Rotor diameters ranging from 8 in to 12 in are studied and with configurations comprised of an isolated rotor, a quadcopter configuration and a hexacopter configuration. A six degree-of-freedom load cell is used to assess the aerodynamic performance of each drone configuration. Meanwhile, an azimuthal array of 1/2-inch microphones is placed between 2 and 3 hub-center diameters from the drone center thereby allowing the acoustic near-field to be quantified. The analysis is performed using standard statistical metrics such as Sound Pressure Level and Overall Sound Pressure Level and is presented to demonstrate the relationship between the number of rotors, the drone rotor size and it’s aerodynamic performance (thrust) relative to the far-field noise.
The vibroacoustic loads that form during the startup of both rigid and compliant wallhigh area ratio nozzles is investigated. The rigid wall nozzle is fabricated from 6061 aluminum while the compliant wall nozzles are formed from urethane-based elastomers in order to invoke aeroelastic coupling between the nozzle wall and the internal flow. Single point measurements of the nozzle lip displacement are synchronized with a pressure field microphone located behind the nozzle where the base of a vehicle would reside. Particularattention is drawn to the sound field during transition from free-shock separated flow torestricted shock separated flow, as well as the end-effects regime loads. The findings revealthe sensitivity of the vibroacoustic loads to the aeroelasticity of the nozzle wall duringcritical stages in the startup process.
The use of helium-air mixtures to simulate the effects of elevated temperatures in aeroacoustics is plagued by the inability to match exactly the density and sound speed ratios between the jet flow and the ambient field, all the while maintaining the same gas dynamic Mach number and jet exit velocity. Real heated jet flows are typically achieved using either propane combustion in air or kerosene combustion in air, which results in the formation of carbon-dioxide and water vapor byproducts. In an effort to level the playing field between the heat simulated helium-air mixture system and the air breathing combustion system, a theoretical model is developed to isolate the effect of combustion byproducts on these aeroacoustic parameters to see if similar discrepancies arise. The motivation is to narrow the gap between laboratory and full-scale jet noise testing. Gas properties from the new combustion model are validated by laboratory measurements of a real propane combustion system as well as outputs from NASA’s Chemical Equilibrium with Applications code. The findings reveal how the additional combustion byproducts from propane combustion in air and kerosene combustion in air have a negligible effect on the parameters relevant to jet noise. Closer inspection of the helium-air mixture system demonstrates how variations in the Mach wave radiation angle at moderate pressure and temperature ratios of the nozzle is accurate to within a couple of degrees, relative to a pure heated air system. Similar accuracy is reported with the far-field intensity.
The unsteady wall pressure on the aft deck of a multi-stream, planar supersonic nozzle is studied over a range of nozzle operating conditions corresponding to independent changes to the core and bypass stream pressure ratios. The data are processed using time-frequency analysis and reveal various tones corresponding to transonic resonance as wellunsteady interactions of both separation and reflection shocks with the developing boundary layer. The position of the separation shock is shown to experience significant hysteresis effects, which subside at pressure ratios well above the design pressure ratio of the nozzle. Shadowgraphy images of the exhaust plume are also presented, which are then analyzed using the snapshot form of proper orthogonal decomposition. The findings from this low-dimensional analysis demonstrates how the first most energetic mode highlights the shock cell patterns whereas the second most energetic mode elucidates turbulence motions in the plume.
A nonintrusive measure of the exhaust plume and immediate sound field produced by a cluster of two thrust optimized parabolic contour nozzles is studied during two steady-state conditions. The first condition is at a nozzle pressure ratio of 25, at which point the flow is in a restricted-shock separated state. The second condition is at a nozzle pressure ratio of 37 and is when the flow and internal shock pattern transition rapidly between free-shock separated flow and the end-effects regime. These end-effects regime pulsations produce significant vibroacoustic loads due to the intermittent breathing of the last trapped annular separation bubble with the ambient. The exhaust plumes and surrounding sound field are first visualized by way of retroreflective shadowgraphy. Radon transforms of the spatially resolved shadowgraphy images are then used to characterize the statistical behavior of the acoustic wave fronts that reside within the hydrodynamic periphery of the nozzle flow. The findings reveal quantitative evidence of the sources of most intense vibroacoustic loads during the end-effects regime of clustered rockets.
The effect of stagger startup on the vibroacoustic loads that form during the end-effects regime of clustered rockets is studied using both full-scale (hot-gas) and laboratory-scale (cold-gas) data with vehicle geometry. Both configurations comprise three nozzles with thrust-optimized parabolic contours that undergo free-shock separated flow and restricted-shock separated flow as well as an end-effects regime before flowing full. Acoustic pressure waveforms recorded at the base of the nozzle cluster are analyzed using various statistical metrics as well as time-frequency analysis. The findings reveal a significant reduction in end-effects regime loads when engine startups are staggered. However, regardless of stagger, both the skewness and kurtosis of the acoustic pressure time derivative elevate to the same levels, thereby demonstrating the intermittence and impulsiveness of the acoustic waveforms during the end effects regime.
A method for calculating the effective Gol’dberg number for diverging waveforms is presented, which leveragesknown features of a high-speed jet and its associated sound field. The approach employs a ray tube situated along the Mach wave angle where the sound field is not only most intense, but advances from undergoing cylindrical decay to spherical decay. Unlike other efforts, a “piecewise-spreading regime” model is employed, which yields, separately, effective Gol’dberg numbers for the cylindrically and spherically spreading regions in the far field. The new approach is applied to a plethora of experimental databases, encompassing both laboratory- and full-scale jet noise studies. The findings demonstrate how cumulative nonlinear distortion is expected to form in the acoustic near field of laboratory scale round jets where pressure amplitudes decay cylindrically; waveform distortion is not expected in the acoustic far field where waveform amplitudes diverge spherically. On the other hand, where full-scale jet studies are concerned, effective Gol’dberg number calculations demonstrate how cumulative waveform distortion is significant in both the cylindrical- and spherical-spreading regimes. The laboratory-scale studies also reveal a pronounced sensitivity to humidity conditions, relative to the full-scale counterpart.
The plume produced by a cluster of two large area-ratio thrust-optimized paraboliccontour nozzles is visualized over a range of nozzle pressure ratios by way of retrore-flective shadowgraphy. Both nozzles exhibit free-shock separated flow, restricted-shockseparated flow and an end-effects regime prior to flowing full. Transient (startup) op-erations of the nozzles are studied with the primary focus being the pulsations thatform during the end-effects regime. This occurs at a pressure ratio of 37 for thesenozzles and is associated with elevated sound levels in the immediate vicinity of thenozzles and vehicle. The shadowgraphy images reveal the formation of turbulent largescale structures, on the order of the nozzle diameter, during the end-effects regime.These large scale structures are driven by the intermittent opening of the last trappedannular separation bubble to the ambient and grow rapidly within the first two nozzlediameters.
The spatial evolution of acoustic waveforms produced by a laboratory-scale Mach 3 jet are investigated using both 1∕4 in. and 1∕8 in. pressure field microphones located along rays emanating from the postpotential core where the peak sound emission is found to occur. The measurements are acquired in a fully anechoic chamber, where ground or other large surface reflections are minimal. Various statistical metrics are examined along the peak emission path, where they are shown to undergo rapid changes within 2m from the source region. An experimentally validated wave-packet model is then used to confirm the location where the pressure amplitude along the peak emission path transitions from cylindrical to spherical decay. Various source amplitudes, provided by the wave-packet model, are then used to estimate shock formation distance and Gol’dberg numbers for diverging waves. The findings suggest that cumulative nonlinear distortion is likely to occur at laboratory scale near the jet flow, where the waveform amplitude decays cylindrically, but less likely to occur farther from the jet flow, where the waveform amplitude decays spherically. Direct inspection of the raw time series reveals how steepened waveforms are generated by rogue like waves that form from the constructive interference of waves from neighboring sources as opposed to classical cumulative nonlinear distortion.
A theoretical combustion model is developed to simulate the influence of ideal gas effects on various aeroacoustic parameters over a range of equivalence ratios. The motivation is to narrow the gap between laboratory and full-scale jet noise testing. The combustion model is used to model propane combustion in air and kerosene combustion in air. Gas properties from the combustion model are compared to real lab data acquired at the National Center for Physical Acoustics at the University of Mississippi as well as outputs from NASA’s Chemical Equilibrium Analysis code. Different jet properties are then studied over a range of equivalence ratios and pressure ratios for propane combustion in air, kerosene combustion in air and heated air. The findings reveal negligible differences between the three constituents where the density and sound speed ratios are concerned. Albeit, the area ratio required for perfectly expanded flow is shown to be more sensitive to gas properties, relative to changes in the temperature ratio.
The effect of stagger startup on the vibro-acoustic loads that form during the end-effects regime of clustered rockets is studied using both full-scale (hot-gas) and laboratory scale (cold gas) data with vehicle geometry. Both configurations comprise three nozzles with thrust optimized parabolic contours that undergo free shock separated flow and restricted shock separated flow as well as an end-effects regime prior to flowing full. Acoustic pressure waveforms recorded at the base of the nozzle cluster are analyzed using various statistical metrics as well as time-frequency analysis. The findings reveal a significant reduction in end-effects regime loads when engine ignition is staggered. However, regardless of stagger, both the skewness and kurtosis of the acoustic pressure time derivative elevate to the same levels during the end-effects regime thereby demonstrating the intermittence and impulsiveness of the acoustic waveforms that form during engine startup.
The design, fabrication and calibration of a new thrust stand for conducting thrust andaeroacoustic measurements concurrently in a fully anechoic chamber is discussed. The new thrust stand employs the scale-force measurement technique and is designed to accommodate multi-stream nozzles (core and bypass flow streams). Each stream has a dedicated helium air mixture system thereby permitting a multitude of test conditions. The methodology for designing the thrust stand is described and uses a notch type flexure which demonstrates high repeatability over extended thrust ranges. Calibration is performed with elevated pressure inside the plenum to characterize the effect of increased pressure on the flexure performance. A further qualification of the thrust measurement accuracy is conducted using a small arsenal of nozzles comprising method of characteristics contours. Surveys of the far-field pressure are then conducted during various operating points along the startup curve of a Mach 1.71 rectangular supersonic nozzle.
Shock wave / boundary layer interaction is studied in a large area ratio axisymmetric nozzle comprising a design exit Mach number of 5.58. Shock motion unsteadiness is captured by way of the dynamic wall pressure and is evaluated during overexpanded operations up to a nozzle pressure ratio of 65. Stationary SWBLI is first considered at a nozzle pressure ratio of 28.7 such that the internal flow structure is in a restricted-shock separated state; the mean position of the annular separation shock resides at a fixed position. Conditional averages of the wall pressure fluctuations show how the motion of the incipient separation shock is out of phase with pressure fluctuations measured in the separated region downstream of the shock; pressure decreases when the shock moves downstream and vice versa. This is indicative of a long intermittent region, in terms of the boundary layer thickness, as the observed phenomena can be explained by translating the static wall pressure profile along with the shock motion. Non-stationary SWBLI is then considered by increasing the nozzle pressure ratio over time (transient startup). Under these conditions, the shock pattern varies in strength and structure as it sweeps through the nozzle. A time-frequency analyses of the fluctuating wall pressure during the non-stationary operations, and at the same location that the stationary unsteadiness is analysed, reveals a similar spectral footprint. However, for relatively slower start-ups, the amplitude of the unsteadiness is reduced by a factor of about seven. The findings demonstrate how the rate at which the nozzle pressure ratio increases can have a significant influence on the amplitude of the unsteady shock foot motion.
High area ratio rockets generate strong vibro-acoustic loads primarily during transient operations, like start-up and shut-down of the engine. These loads can adversely affect the launch vehicle and its payload. Thus, an accurate prediction of the loads produced during engine start-up is pertinent to the safety and reliability of the launch vehicle. The present work focuses on developing a robust framework for predicting these loads using laboratory scale rocket nozzles tested in the fully anechoic chamber at The University of Texas at Austin. This encompasses corrections for the observer position relative to the prominent source region, as well as scaling factors to correct for geometric factors. The test campaign encompasses single, two, three and four nozzle clusters, as well as differences in nozzle geometry and operating conditions (nozzle pressure ratio).