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NASA NTRS1989

Low Reynolds number numerical solutions of chaotic flow

Numerical computations of two-dimensional flow past an airfoil at low Mach number, large angle of attack, and low Reynolds number are reported which show a sequence of flow states leading from single-period vortex shedding to chaos via the period-doubling mechanism. Analysis of the flow in terms of phase diagrams, Poincare sections, and flowfield variables are used to substantiate these results. The critical Reynolds number for the period-doubling bifurcations is shown to be sensitive to mesh refinement and the influence of large amounts of numerical dissipation. In extreme cases, large amounts of added dissipation can delay or completely eliminate the chaotic response. The effect of artificial dissipation at these low Reynolds numbers is to produce a new effective Reynolds number for the computations.

Pulliam, Thomas H.↗

NASA NTRS1987

Analysis of unsteady wake of a circular cylinder using Navier-Stokes equations

An unsteady Navier-Stokes analysis of the separated flow in the wake of a circular cylinder is presented in order to study the wake structure and the various interactions between its scales at low Re. The quasi-periodic and chaotic flows noted by Sreenivasan (1985) are searched for. The analysis is used to characterize the secondary vortex formation in the early stages of the flow development and its interaction with the primary vortex.

Ghia, K. N.↗

NASA NTRS1997

Vortex leading edge flap assembly for supersonic airplanes

A leading edge flap (16) for supersonic transport airplanes is disclosed. In its stowed position, the leading edge flap forms the lower surface of the wing leading edge up to the horizontal center of the leading edge radius. For low speed operation, the vortex leading edge flap moves forward and rotates down. The upward curve of the flap leading edge triggers flow separation on the flap and rotational flow on the upper surface of the flap (vortex). The rounded shape of the upper fixed leading edge provides the conditions for a controlled reattachment of the flow on the upper wing surface and therefore a stable vortex. The vortex generates lift and a nose-up pitching moment. This improves maximum lift at low speed, reduces attitude for a given lift coefficient and improves lift to drag ratio. The mechanism (27) to move the vortex flap consists of two spanwise supports (24) with two diverging straight tracks (64 and 68) each and a screw drive mechanism (62) in the center of the flap panel (29). The flap motion is essentially normal to the airloads and therefore requires only low actuation forces.

Rudolph, Peter K. C.↗

NASA NTRSDate not supplied

Dynamic Analysis of Reynolds Number Effects on Trailing Edge Transonic Vortex Shedding and Its Impact on Turbine Blade Aerodynamic Performance

Time-resolved, high-speed self-aligned focusing schlieren images were acquired in the NASA Glenn Research Center Transonic Turbine Blade Cascade facility to help understand the aerodynamic behavior of high-pressure, thick trailing edge turbine blades. The trailing edge thickness of 9% of axial chord tested represents simulated ceramic matrix composite fabrication constraints, which was verified previously to possess a high-loss flow regime at high inlet turbulence conditions over a narrow range of Reynolds numbers and at a fixed design exit Mach number of 0.74. Our high-speed images, which were acquired at 10 distinct Reynolds numbers, show a significant increase in energy from flow oscillations due to transonic vortex shedding at Reynolds numbers corresponding to the high loss conditions. For those conditions, strong acoustic waves turn into shock waves. Spectral Proper Orthogonal Decomposition of the high-speed images shows acoustic waves from trailing edge vortex shedding at all conditions, with increased spectral energy at the high-loss conditions and slightly increasing frequency (about 6%) as a function of Reynolds number. Analysis of potential feedback timing is performed using velocity fields from a previous LES simulation, considering different feedback mechanisms. Most noteworthy is the acoustic/shock-boundary layer interaction mechanism on the suction surface at the blade geometric throat, which likely plays an important role in realistic curved blade passages.

Trailing Edge↗

NASA NTRSDate not supplied

The Generation of Mixing Across the Stratospheric Polarvortex Edge By Breaking Gravity Waves

Modern high resolution global atmospheric analyses are capable of resolving larger gravity waves, including their generation, propagation, and breaking. Orographically generated gravity waves that break in the middle stratosphere often appear as disturbances in maps of potential vorticity, a key stratospheric dynamical quantity. Using analyses from the NASA Global Modeling and Assimilation Office, we show how gravity waves not only disturb the local flow above them but can generate flow instabilities leading to eddies that propagate around the stratospheric polar vortex, creating significant mixing across the vortex edge and into the polar region. These disturbances and mixing events are especially noticeable during NH winter months with low planetary scale wave activity such as in January 2022. During January 2022, the initially circular flow around the polar vortex was disrupted by breaking gravity waves, leading to unstable flow that created a chain of eddies around the vortex edge. The mixing associated with this process led to increased ozone in the upper stratosphere polar region.

Polar Vortex↗

NASA NTRS1991

Unsteady flowfield of a propfan at takeoff conditions

The unsteady flowfield of a propfan operation at takeoff conditions with angular inflow is examined by solving the three-dimensional Euler equations. The operating conditions considered are: Mach no. = 0.31, advance ratio = 1.6, and inflow angle to the propfan = 8.3 deg. The predicted results clearly show the cyclic variations of the blade power and thrust coefficients due to angular inflow. The flow changes from blade passage to passage are illustrated in terms of static pressure contours. The predicted blade surface pressure waveforms were compared with flight measurements. The predictions at the inboard radial station, r/R = 0.68, show reasonable agreement with flight data. At the outboard radial station, r/R = 0.95, where the interactions of the tip vortex, the tip-region flow and the blade wake appear to result in a complex nonlinear measured response. The prediction shows poor agreement.

Nallasamy, M.↗

NASA NTRS1989

Numerical simulation of pressure oscillations in a ramjet combustor

Large-eddy simulations of compressible cold flow in a ramjet combustor configuration were performed. The objectives were to investigate the mechanisms for vortex-acoustic interaction in such a device and to develop a simulation method that can be extended to cases with combustion heat release to study combustion instability. From the simulation, it was found that the separated shear layer rolls up into concentrated vortices that merge to form large coherent structures. These vortices interact with the choked nozzle downstream to produce an axial acoustic dipole. The spectrum of the pressure fluctuation at the base of the backward-facing step shows that there are two types of oscillations: an acoustic resonant mode and a vortex-acoustic coupled mode. Based on the flow physics observed in the simulations, a simple one-dimensional model for the vortex-acoustic coupled mode was proposed. The eigenvalue problem based on this model was solved to obtain the frequency of the coupled mode.

Jou, Wen-Huei↗

NASA NTRS1996

Vortex generator installation studies on steady state and dynamic inlet distortion

The theoretical and experimental work carried out under the NASA/MOD Joint Aeronautical Program has shown that CFD vortex generator installations designs successfully managed inlet duct flow distortion and that significant benefits in flow unsteadiness at the engine face were also present. The main conclusions to date from the collaborative effort between NASA/Lewis and DRA/Bedford are as follows: (1) vortex generator installations can be designed to be effective over a wide range of inlet operating conditions using Computational Fluid Dynamics and formal optimization procedures, (2) reductions in steady state engine face distortion of up to 80% have been measured in the M2129 inlet S-duct using CFD designed vortex generator installations, (3) reductions in flow unsteadiness of up to 80% have been measured in the W129 inlet S-duct using CFD designed vortex generator installations, and (4) the Reduced Navier-Stokes code RNS3D is a useful tool to design vortex generator installations to manage engine face distortions over a wide range of inlet operating conditions.

Anderson, Bernhard H.↗

NASA NTRS2016

Tracking Blade Tip Vortices for Numerical Flow Simulations of Hovering Rotorcraft

Blade tip vortices generated by a helicopter rotor blade are a major source of rotor noise and airframe vibration. This occurs when a vortex passes closely by, and interacts with, a rotor blade. The accurate prediction of Blade Vortex Interaction (BVI) continues to be a challenge for Computational Fluid Dynamics (CFD). Though considerable research has been devoted to BVI noise reduction and experimental techniques for measuring the blade tip vortices in a wind tunnel, there are only a handful of post-processing tools available for extracting vortex core lines from CFD simulation data. In order to calculate the vortex core radius, most of these tools require the user to manually select a vortex core to perform the calculation. Furthermore, none of them provide the capability to track the growth of a vortex core, which is a measure of how quickly the vortex diffuses over time. This paper introduces an automated approach for tracking the core growth of a blade tip vortex from CFD simulations of rotorcraft in hover. The proposed approach offers an effective method for the quantification and visualization of blade tip vortices in helicopter rotor wakes. Keywords: vortex core, feature extraction, CFD, numerical flow visualization

Numerical↗

NASA NTRS1986

Flow-separation patterns on symmetric forebodies

Flow-visualization studies of ogival, parabolic, and conical forebodies were made in a comprehensive investigation of the various types of flow patterns. Schlieren, vapor-screen, oil-flow, and sublimation flow-visualization tests were conducted over an angle-of-attack range from 0 deg. to 88 deg., over a Reynolds-number range from 0.3X10(6) to 2.0X10(6) (based on base diameter), and over a Mach number range from 0.1 to 2. The principal effects of angle of attack, Reynolds number, and Mach number on the occurrence of vortices, the position of vortex shedding, the principal surface-flow-separation patterns, the magnitude of surface-flow angles, and the extent of laminar and turbulent flow for symmetric, asymmetric, and wake-like flow-separation regimes are presented. It was found that the two-dimensional cylinder analogy was helpful in a qualitative sense in analyzing both the surface-flow patterns and the external flow field. The oil-flow studies showed three types of primary separation patterns at the higher Reynolds numbers owing to the influence of boundary-layer transition. The effect of angle of attack and Reynolds number is to change the axial location of the onset and extent of the primary transitional and turbulent separation regions. Crossflow inflectional-instability vortices were observed on the windward surface at angles of attack from 5 deg. to 55 deg. Their effect is to promote early transition. At low angles of attack, near 10 deg., an unexpected laminar-separation bubble occurs over the forward half of the forebody. At high angles of attack, at which vortex asymmetry occurs, the results support the proposition that the principal cause of vortex asymmetry is the hydrodynamic instability of the inviscid flow field. On the other hand, boundary-layer asymmetries also occur, especially at transitional Reynolds numbers. The position of asymmetric vortex shedding moves forward with increasing angle of attack and with increasing Reynolds number, and moves rearward with increasing Mach number.

Keener, Earl R.↗

NASA NTRS1991

A model for the evaporation of clusters of drops embedded in jet vortices. I - Steady injection of identical clusters

A model is developed for describing the interaction of vortex-drop clusters in a flowing gaseous jet, convecting downstream from an injection location. Results are presented for a stationary case representing the situation when identical clusters are continuously injected and the injection rate is constant. The results indicate that, in a rich mixture high-drop-number density regime, the mass evaporated from the drops controls the velocity of the cluster-in-vortex as it propagates downstream.

Bellan, J.↗

NASA NTRS1984

Large-scale turbulence in the Jovian atmosphere

Voyager 1 and 2 image pairs taken one rotation period (10 hr) apart are analyzed in order to compute average cloud velocities in the Jovian atmosphere, as well as to study the global energetics of the processes, in particular the sign and the magnitude of the energy transfer from the fluctuating to the mean flow. Particular attention is paid to the Great Red Spot phenomenon; it is suggested to have originated as a small vortex extracting energy from the surrounding flow. Since the eddy, presently in its near-equilibrium state, has reached a considerable size, while the mean shear lies close to that of a neutrally stable profile, it can be concluded that the presence of one large eddy precludes the growth of any others in the depleted mean shear.

Mitchell, J. L.↗

NASA NTRS1992

Effect of planform and body on supersonic aerodynamics of multibody configurations

An experimental and theoretical investigation of the effect of the wing planform and bodies on the supersonic aerodynamics of a low-fineness-ratio, multibody configuration has been conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.80, 2.00, and 2.16. Force and moment data, flow-visualization data, and surface-pressure data were obtained on eight low-fineness-ratio, twin-body configurations. These configurations varied in inboard wing planform shape, outboard wing planform shape, outboard wing planform size, and presence of the bodies. The force and moment data showed that increasing the ratio of outboard wing area to total wing area or increasing the leading-edge sweep of the inboard wing influenced the aerodynamic characteristics. The flow-visualization data showed a complex flow-field system of shocks, shock-induced separation, and body vortex systems occurring between the side bodies. This flow field was substantially affected by the inboard wing planform shape but minimally affected by the outboard wing planform shape. The flow-visualization and surface-pressure data showed that flow over the outboard wing developed as expected with changes in angle of attack and Mach number and was affected by the leading-edge sweep of the inboard wing and the presence of the bodies. Evaluation of the linear-theory prediction methods revealed their general inability to consistently predict the characteristics of these multibody configurations.

Mcmillin, S. Naomi↗

NASA NTRS1990

Pinning, flow and plastic deformation of flux vortices in high T(sub c) superconductors

In high temperature superconductor (HTSC) materials the vortices are highly mobile and flexible. This is reflected in different models of melt of a vortex lattice. Another aspect of the problem is stressed: an easy nucleation and high mobility of dislocations in the vortex lattice. Some models of plastic deformation of vortex lattice are considered as a result of its interaction with a real crystal structure. Depinning is interpreted as yield of plastic flow is vortex medium. Effect of macroscopic defects in crystal structures (pores, inclusions, grain and domain boundaries) is being considered in detail. Available experimental facts on magnetization and a critical current in HTSC and conventional superconductors are discussed from the points of view of depinning to vortices vs. plastic flow of vortices vs. plastic flow of vortices medium.

Roytburd, A.↗

NASA NTRS1986

New techniques for experimental generation of two-dimensional blade-vortex interaction at low Reynolds numbers

An experimental investigation of two dimensional blade vortex interaction was held at NASA Langley Research Center. The first phase was a flow visualization study to document the approach process of a two dimensional vortex as it encountered a loaded blade model. To accomplish the flow visualization study, a method for generating two dimensional vortex filaments was required. The numerical study used to define a new vortex generation process and the use of this process in the flow visualization study were documented. Additionally, photographic techniques and data analysis methods used in the flow visualization study are examined.

Booth, E., Jr.↗

NASA NTRS1972

Dynamic flow angle measurements at the diffuser exit of a Mach 2.5 mixed compression inlet

An exploratory investigation was made to determine the magnitude of the dynamic flow angle at a single point in the diffuser exit of a Mach 2.50 inlet. Results indicate that the maximum flow angle variation was + or - 3 degrees except when vortex generators were installed in the subsonic diffuser. The amplitude of these flow angle fluctuations was directly related to the amplitude of the total pressure fluctuations at the probe survey point. Flow angle fluctuations as high as + or - 11 degrees were recorded when vortex generators were installed near the probe survey point.

Prati, W. M.↗

NASA NTRSDate not supplied

On the Implementation and Further Validation of a Time Domain Boundary Element Method Broadband Impedance Boundary Condition

A time domain boundary integral equation with Burton-Miller reformulation is presented for acoustic scattering by surfaces with liners in a uniform mean flow. The Ingard-Myers impedance boundary condition is implemented using a broadband multipole impedance model and converted into time domain differential equations to augment the boundary integral equation. The coupled integral-differential equations are solved numerically by a March-On-in-Time (MOT) scheme. While the Ingard-Myers condition is known to support Kelvin-Helmholtz instability due to its use of a vortex sheet interface between the flow and the liner surface, it is found that by neglecting a second derivative term in the current time domain impedance boundary condition formulation, the instability an be effectively suppressed in computation. The proposed formulation and implementation are validated using a NASA Langley Research Center Grazing Flow Impedance Tube (GFIT) experimental dataset with satisfactory results. Moreover, a minimization procedure for finding the poles and coefficients of the broadband multiple impedance model is formulated in this paper by which, unlike the commonly used vector-fitting method, passivity of the model is ensured. Numerical tests show the proposed minimization approach is effective for modeling liners that are commonly used in aeroacoustic applications.

Time Domain Impedance Boundary Condition↗

NASA NTRS1991

Flight test to determine feasibility of a proposed airborne wake vortex detection concept

This investigation was conducted to determine the radial extent at which aircraft mounted flow vanes or roll rate gyros can sense the circulatory flow field that exists around the lift induced vortex system generated by an aircraft in flight. The probe aircraft was equipped with wingtip sensors for measuring angle of attack and angle of sideslip, and with a fuselage mounted gyroscope for measuring roll rate. Analysis of flight test data indicated that the vortex was detectable at a lateral distance of about 105 feet (best results) using unsophisticated equipment. Measurements were made from the centerline of the probe aircraft to the center of the nearest vortex with the probe aircraft flying between one half and one and one half miles behind the vortex generating aircraft.

Branstetter, James R.↗