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At least 559 records · Page 31

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NASA NTRSDate not supplied

Aeroacoustic Analysis using Dynamic Mode Decomposition of Unsteady Pressure-Sensitive Paint Measurements

In this paper, we present a new method for the diagnosis and analysis of aeroacoustic phenomena. Taking advantage of the well-known property that the measurements of the Unsteady Pressure-Sensitive Paint (uPSP) have much higher spatial resolution compared to those of the conventional pressure transducers, the method is based on the visualization and analysis of the outputs of the Dynamic Mode Decomposition (DMD) of the uPSP measurements. The uPSP measurements were collected with four Phantom high-speed cameras in the Ascent Transient Aerodynamics Test (ATAT) of the Space Launch System (SLS) Block 1 cargo vehicle in the 11-by-11-foot transonic test section of the Unitary Plan Wind Tunnel (UPWT) at NASA Ames Research Center in September 2019. The method presented in this paper is demonstrated by investigating an interesting phenomenon observed in the SLS ATAT – the vortex shedding tone generated by the forward attachment of the Solid Rocket Booster (SRB). As examples, the DMD outputs of the uPSP measurements in a subsonic test in the SLS ATAT are presented. It is shown the information retrieved from the DMD outputs of the uPSP measurements can be effectively used in the identification and diagnosis of the aeroacoustic phenomena. The work described in this paper is a part of NASA’s development of a new state-of-the-art uPSP capability in production wind tunnels. Funding was provided by the NASA Aerosciences Evaluation and Test Capabilities Portfolio Office.

acoustics↗

NASA NTRSDate not supplied

Vortex Beam Sidelobe Suppression via Bayliss Synthesis from a Rectangular Phased Array

We demonstrate the application of the Bayliss synthesis method for suppression of sidelobes in an OAM vortex beam, when generated by a commercial-of-the-shelf, rectangular phased array. The method is optimized for the current application and simulated for a 26.5 GHz signal in CST Microwave Studio. Sidelobes are notably reduced by 16.9 dB, with minimal reduction in gain and increase in beamwidth, as compared to a uniformed stimulated phased array. Further, the OAM modes are found to have greater purity that that of the uniformly stimulated array, as well as for other amplitude tapering methods. The Bayliss synthesis is then applied to a phased array in a test range, verifying the simulation results that demonstrate sidelobe suppression. These results indicate that OAM may be applied to existing phased arrays with minimal effort.

Orbital Angular Momentum (OAM)↗

NASA NTRSDate not supplied

Vortex Beam Sidelobe Suppression via Bayliss Synthesis from a Rectangular Phased Array

Orbital Angular Momentum (OAM)↗

DOE OSTI2024

Vortex bursting and associated twist dynamics on helical vortex tubes and vortex rings

The interaction of opposite-signed twist waves on vortex tubes can lead to vortex bursting, a process where the core expands into a double ring-like structure with strong swirling flows. Previous works have studied vortex bursting on rectilinear vortices by axially perturbing the initial core size to generate the twist waves, and observed largely axisymmetric bursting dynamics. In this work, we numerically study bursting on vortical structures with curved centrelines, analysing the interaction between the centreline dynamics, twist wave generation and propagation, and vortex bursting. We focus on axially perturbed helical vortex tubes with small radius-to-pitch ratios up to$0.0625$, as well as vortex rings with a large radius-to-core size ratio$10$, both at a circulation-based Reynolds number$5000$. The results show that though the initial twist wave propagation speeds are relatively unaffected by the curvature and torsion of the centreline, the bursting process is altered significantly compared with rectilinear vortices. The self-induced rotation of the centreline of the helical tube induces a non-axisymmetric distortion of the bursting structure, which rapidly breaks up the vortex core into small-scale helical structures. A similar destabilization of the bursting structure also occurs on vortex rings. The enstrophy increase and accelerated energy decay associated with bursting are predominantly determined by the twist wave strength, rather than the curvature and torsion of the centreline. Combined, our findings imply that bursting could play an important role in transferring and dissipating energy of vortical structures in wakes, and turbulent flows in general.

Mechanics↗

NASA NTRS1975

Flow visualization of vortices locked by spanwise blowing over wings featuring a unique leading and trailing-edge flap system

Flow visualization studies were conducted to qualitatively determine the effects of active generation and augmentation of vortex flow over wings by blowing a discrete jet in a spanwise direction in the channel formed by extension of upper surface leading- and trailing-edge flaps. Spanwise blowing from a reflection plane over a rectangular wing was found to generate and lock a dual corotating vortex system within the channel and, at sufficient blowing rates, cause the separated flow off the upper end of the leading-edge flap to reattach to the trailing-edge flap. Test parameters included wing angle of attack, jet momentum coefficient, leading- and trailing-edge flap deflection angle, and jet location above the wing surface. Effects due to removal of the leading- and trailing-edge flap were also investigated.

Erickson, G. E.↗

NASA NTRS1984

Identification of vortex-induced clear-air turbulence using airline flight records

The nature and cause of clear-air turbulence is being investigated, in cooperation with the National Transportation Safety Board, using the flight records available from airline encounters with severe turbulence. This paper presents two case studies of severe turbulence which indicate that the airplanes involved encountered vortex arrays which were generated by destabilized wind-shear layers near the tropopause. In order to identify and analyze vortex patterns (i.e., vortex strength, size, and spacings), potential-flow models of vortex arrays were developed that describe reasonably well the wind patterns derived from the airliner flight records. The results of this analysis indicate that in the two cases studied, the vortex cores had diameters in the range of 900 to 1,200 ft with tangential velocities in the range of 70 to 85 ft/sec. This study presents the first identification and analysis of vortex arrays from airline flight data. The results are compared with theoretical predictions and previous observations.

Parks, E. K.↗

NASA NTRS1979

Three-dimensional simulation of the free shear layer using the vortex-in-cell method

We present numerical simulations of the evolution of a mixing layer from an initial state of uniform vorticity with simple two- and three-dimensional small perturbations. A new method for tracing a large number of three-dimensional vortex filaments is used in the simulations. Vortex tracing by Biot-Savart interaction originally implied ideal (non-viscous) flow, but we use a 3-d mesh, Fourier transforms and filtering for vortex tracing, which implies 'modeling' of subgrid scale motion and hence some viscosity. Streamwise perturbations lead to the usual roll-up of vortex patterns with spanwise uniformity maintained. Remarkably, spanwise perturbations generate streamwise distortions of the vortex filaments and the combination of both perturbations leads to patterns with interesting features discernable in the movies and in the records of enstrophy and energy for the three components of the flow.

Couet, B.↗

NASA NTRS1975

Flight test investigation of the vortex wake characteristics behind a Boeing 727 during two-segment and normal ILS approaches (A joint NASA/FAA report)

Flight tests were performed to evaluate the vortex wake characteristics of a Boeing 727 aircraft during conventional and two-segment instrument landing approaches. Smoke generators were used for vortex marking. The vortex was intentionally intercepted by a Lear Jet and a Piper Comanche aircraft. The vortex location during landing approach was measured using a system of phototheodolites. The tests showed that at a given separation distance there are no readily apparent differences in the upsets resulting from deliberate vortex encounters during the two types of approaches. The effect of the aircraft configuration on the extent and severity of the vortices is discussed.

Barber, M. R.↗

NASA NTRS2003

Spectral Characteristics of Wake Vortex Sound During Roll-Up

This report presents an analysis of the sound spectra generated by a trailing aircraft vortex during its rolling-up process. The study demonstrates that a rolling-up vortex could produce low frequency (less than 100 Hz) sound with very high intensity (60 dB above threshold of human hearing) at a distance of 200 ft from the vortex core. The spectrum then drops o rapidly thereafter. A rigorous analytical approach has been adopted in this report to derive the spectrum of vortex sound. First, the sound pressure was solved from an alternative treatment of the Lighthill s acoustic analogy approach [1]. After the application of Green s function for free space, a tensor analysis was applied to permit the removal of the source term singularity of the wave equation in the far field. Consequently, the sound pressure is expressed in terms of the retarded time that indicates the time history and spacial distribution of the sound source. The Fourier transformation is then applied to the sound pressure to compute its spectrum. As a result, the Fourier transformation greatly simplifies the expression of the vortex sound pressure involving the retarded time, so that the numerical computation is applicable with ease for axisymmetric line vortices during the rolling-up process. The vortex model assumes that the vortex circulation is proportional to the time and the core radius is a constant. In addition, the velocity profile is assumed to be self-similar along the aircraft flight path, so that a benchmark vortex velocity profile can be devised to obtain a closed form solution, which is then used to validate the numerical calculations for other more realistic vortex profiles for which no closed form solutions are available. The study suggests that acoustic sensors operating at low frequency band could be profitably deployed for detecting the vortex sound during the rolling-up process.

Booth, Earl R., Jr.↗

NASA NTRS2001

High Order Difference Method for Low Mach Number Aeroacoustics

A high order finite difference method with improved accuracy and stability properties for computational aeroacoustics (CAA) at low Mach numbers is proposed. The Euler equations are split into a conservative and a symmetric non- conservative portion to allow the derivation of a generalized energy estimate. Since the symmetrization is based on entropy variables, that splitting of the flux derivatives is referred to as entropy splitting. Its discretization by high order central differences was found to need less numerical dissipation than conventional conservative schemes. Owing to the large disparity of acoustic and stagnation quantities in low Mach number aeroacoustics, the split Euler equations are formulated in perturbation form. The unknowns are the small changes of the conservative variables with respect to their large stagnation values. All nonlinearities and the conservation form of the conservative portion of the split flux derivatives can be retained, while cancellation errors are avoided with its discretization opposed to the conventional conservative form. The finite difference method is third-order accurate at the boundary and the conventional central sixth-order accurate stencil in the interior. The difference operator satisfies the summation by parts property analogous to the integration by parts in the continuous energy estimate. Thus, strict stability of the difference method follows automatically. Spurious high frequency oscillations are suppressed by a characteristic-based filter similar to but without limiter. The time derivative is approximated by a 4-stage low-storage second-order explicit Runge-Kutta method. The method has been applied to simulate vortex sound at low Mach numbers. We consider the Kirchhoff vortex, which is an elliptical patch of constant vorticity rotating with constant angular frequency in irrotational flow. The acoustic pressure generated by the Kirchhoff vortex is governed by the 2D Helmholtz equation, which can be solved analytically using separation of variables.

Mueller, B.↗

NASA NTRS1979

Vortex-lift roll-control device

A wing is described for aircraft of cropped, arrow-type planform with thin leading and side edges. The wing has a pivotable tip to alter the crop angle of the wing during flight. Increasing the crop angle causes the wing side edge to become a trailing edge which reduces the strength of the side edge vortex flow. Decreasing the crop angle causes opposite results, in particular the side edge is now a leading edge and can generate a leading edge vortex flow. The wing constitutes a roll control device for aircraft of the stated design particularly effective at higher angles of attack.

Lamar, J. E.↗

NASA NTRS1992

Two-fence concept for efficient trapping of vortices on airfoils

Previous work on the use of a vortex trapped above a wing in order to produce high lift at low angles of attack is extended here. It is first postulated that the optimum way to trap a vortex is to design the airfoil section and wing so that the flow along the vortex core is minimized. It is then shown that a vertical fence both in front of and behind the separation bubble generated by the trapped vortex is an effective way to reduce the mass flow removal and its associated drag to a negligible amount. In order to show that vertical surfaces upstream and downstream of the vortex separation bubble have an opposite effect on the source requirements for vortex trapping, conformal mapping methods are used to obtain the solutions for a variety of simple two-dimensional, inviscid, incompressible flow configurations. Trapped-vortex flowfield solutions for the flow over flat plate and Clark-Y airfoils are then used to demonstrate that the heights of the fences can be tailored to make the required mass withdrawal (and therefore, the drag due to trapping) to be vanishingly small.

Rossow, Vernon J.↗

NASA NTRS1986

Review and evaluation of recent developments in melic inlet dynamic flow distortion prediction and computer program documentation and user's manual estimating maximum instantaneous inlet flow distortion from steady-state total pressure measurements with full, limited, or no dynamic data

A review of the Melick method of inlet flow dynamic distortion prediction by statistical means is provided. These developments include the general Melick approach with full dynamic measurements, a limited dynamic measurement approach, and a turbulence modelling approach which requires no dynamic rms pressure fluctuation measurements. These modifications are evaluated by comparing predicted and measured peak instantaneous distortion levels from provisional inlet data sets. A nonlinear mean-line following vortex model is proposed and evaluated as a potential criterion for improving the peak instantaneous distortion map generated from the conventional linear vortex of the Melick method. The model is simplified to a series of linear vortex segments which lay along the mean line. Maps generated with this new approach are compared with conventionally generated maps, as well as measured peak instantaneous maps. Inlet data sets include subsonic, transonic, and supersonic inlets under various flight conditions.

Schweikhard, W. G.↗

NASA NTRS1987

Blade-vortex interaction noise predictions using measured blade surface pressures

The generation of helicopter noise by blade-vortex interactions during descent under impulsive conditions is investigated analytically. A noise-prediction technique is developed on the basis of the dipole source term of the Ffowcs-Williams/Hawkings equation and applied to data from simultaneous blade-pressure and acoustic measurements obtained by Cowan et al. (1986) on a 10-ft-diameter 4-blade rotor model in a wind tunnel. Preliminary results show that input-blade-airload azimuth resolution of 1 deg or better and computational azimuth step size of 2 deg or less are required to achieve good agreement between predicted and recorded acoustic time histories. The need for more sophisticated methods to model chordwise input data and for a more extensive experimental data base is indicated.

Ziegenbein, Perry R.↗

NASA NTRS2013

Numerical Simulations of Instabilities in Single-Hole Office Elements

An orifice element is commonly used in liquid rocket engine test facilities either as a flow metering device, a damper for acoustic resonance or to provide a large reduction in pressure over a very small distance in the piping system. While the orifice as a device is largely effective in stepping down pressure, it is also susceptible to a wake-vortex type instability that generates pressure fluctuations that propagate downstream and interact with other elements of the test facility resulting in structural vibrations. Furthermore in piping systems an unstable feedback loop can exist between the vortex shedding and acoustic perturbations from upstream components resulting in an amplification of the modes convecting downstream. Such was the case in several tests conducted at NASA as well as in the Ariane 5 strap-on P230 engine in a static firing test where pressure oscillations of 0.5% resulted in 5% thrust oscillations. Exacerbating the situation in cryogenic test facilities, is the possibility of the formation of vapor clouds when the pressure in the wake falls below the vapor pressure leading to a cavitation instability that has a lower frequency than the primary wake-vortex instability. The cavitation instability has the potential for high amplitude fluctuations that can cause catastrophic damage in the facility. In this paper high-fidelity multi-phase numerical simulations of an orifice element are used to characterize the different instabilities, understand the dominant instability mechanisms and identify the tonal content of the instabilities.

Ahuja, Vineet↗

NASA NTRS1997

Reduction of Helicopter Blade-Vortex Interaction Noise by Active Rotor Control Technology

Helicopter blade-vortex interaction noise is one of the most severe noise sources and is very important both in community annoyance and military detection. Research over the decades has substantially improved basic physical understanding of the mechanisms generating rotor blade-vortex interaction noise and also of controlling techniques, particularly using active rotor control technology. This paper reviews active rotor control techniques currently available for rotor blade vortex interaction noise reduction, including higher harmonic pitch control, individual blade control, and on-blade control technologies. Basic physical mechanisms of each active control technique are reviewed in terms of noise reduction mechanism and controlling aerodynamic or structural parameters of a blade. Active rotor control techniques using smart structures/materials are discussed, including distributed smart actuators to induce local torsional or flapping deformations, Published by Elsevier Science Ltd.

Yu, Yung H.↗

NASA NTRS1998

Wake-Vortex Hazards During Cruise

Even though the hazard posed by lift-generated wakes of subsonic transport aircraft has been studied extensively for approach and departure at airports, only a small amount of effort has gone into the potential hazard at cruise altitude. This paper reports on a studio of the wake-vortex hazard during cruise because encounters may become more prevalent when free-flight becomes available and each aircraft, is free to choose its own route between destinations. In order to address the problem, the various fluid-dynamic stages that vortex wakes usually go through as they age will be described along with estimates of the potential hazard that each stage poses. It appears that a rolling-moment hazard can be just as severe at cruise as for approach at airports, but it only persists for several minutes. However, the hazard posed by the downwash in the wake due to the lift on the generator aircraft persists for tens of minutes in a long narrow region behind the generating aircraft. The hazard consists of severe vertical loads when an encountering aircraft crosses the wake. A technique for avoiding vortex wakes at cruise altitude will be described. To date the hazard posed by lift-generated vortex wakes and their persistence at cruise altitudes has been identified and subdivided into several tasks. Analyses of the loads to be encounter and are underway and should be completed shortly. A review of published literature on the subject has been nearly completed (see text) and photographs of vortex wakes at cruise altitudes have been taken and the various stages of decay have been identified. It remains to study and sort the photographs for those that best illustrate the various stages of decay after they are shed by subsonic transport aircraft at cruise altitudes. The present status of the analysis and the paper are described.

Rossow, Vernon J.↗

DOE OSTI2025

Particle removal from a flat surface using a translating bounded vortex flow

A bounded vortex flow is a hydrodynamic approach for removal of particles from a surface without scattering the particles onto nearby surfaces. The bounded vortex flow field is generated by a nozzle that combines azimuthally tilted jets arranged in a circular pattern and a central suction port. When the nozzle face is directed toward an ‘impingement surface’, the flow develops a wall-normal intake vortex below the suction outlet, which causes high shear stress on the impingement surface. When particles are present on the impingement surface, the high shear stress causes particles to roll along the surface and to be lifted off the surface and transported up the core of the wall-normal vortex into the suction outlet. In typical applications, the nozzle would be translated along the impingement surface to clean particles from the surface. The current paper reports on an experimental study of the effect of nozzle translation on the effectiveness of the bounded vortex flow field for particle mitigation. The effectiveness of particle mitigation was examined as a function of flow rate through the nozzle, particle size, and nozzle translation velocity relative to the impingement surface. As a result, numerical computations are used to relate the flow rate to the maximum shear stress on the impingement surface, which is then used to theoretically predict onset of particle motion.

42 ENGINEERING↗