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NASA NTRS1990

Incompressible spatially-developing free-shear flows

The first goal of this research is to understand the dynamics of three dimensional vortical structures, especially the interaction of the ribs and two dimensional rollers. Because of vortex stretching effects, this requires accurate and well resolved simulations. Another goal is to distinguish temporally- and spatially-growing shear layers in terms of the dynamics and observed vortical structures. This is important since the former is much easier to compute, but the error incurred in using temporal results for certain applications is unknown. The third goal is to investigate the effects of asymmetry on passive scalar mixing and fast-chemistry product formation. In this report a brief summary of a new numerical method is presented, followed by results for mixing layers and wake flows.

Buell, J. C.↗

NASA NTRS1991

An experimental investigation of boundary layer transition in an adverse pressure gradient

The evolution of a small perturbation introduced periodically into a boundary layer with a moderately strong adverse pressure gradient (APG) was investigated experimentally using the hot-wire method. The magnitude of the disturbance is found to decay with the streamwise distance in the APG before undergoing rapid growth and triggering transition in a very repetitive manner. The evidence suggests that the laminar layer separates and that the transition mechanism is an inviscid shear-layer type of instability. Contours of spanwise vorticity in the plane of the disturbance show the formation of roll-ups in the initial stages. Subsequent interaction with the wall leads to the formation of large-scale vortex loops which merge into a single vortex loop further downstream.

Watmuff, Jonathan H.↗

NASA NTRS1992

Supersonic jet mixing enhancement by 'delta-tabs'

The results of a continuing investigation of the effect of vortex generators, in the form of small tabs at the nozzle exit, on the evolution of a jet are reported. Primarily, tabs of triangular shape are considered, and the effect is studied up to an equivalent jet Mach number of 1.8. By changing the orientation of the tab with respect to the nozzle exit plane, streamwise vortex pairs of opposite sign were generated. This resulted in either an outward selection of jet core fluid into the ambient or an inward indentation of the mixing layer into the core of the jet. A triangular shaped tab with its apex leaning downstream, referred to as a delta tab, was found to be the most effective in influencing the jet evolution. Two delta tabs, spaced 180 degrees apart, completely bifurcated the jet. Four delta tabs increased jet mixing substantially, more than by various other methods tried previously; the mass flux at fourteen jet diameters downstream from the nozzle increased by about 50 percent over that for the no tab case. The tabs were found to be effective in jets with laminar or turbulent boundary layers as well as in jets with low or high core turbulence intensities.

Zaman, K. B. M. Q.↗

NASA NTRS1992

Advances in tilt rotor noise prediction

The two most serious tilt rotor external noise problems, hover noise and blade-vortex interaction noise, are studied. The results of flow visualization and inflow velocity measurements document a complex, recirculating highly unsteady and turbulent flow due to the rotor-wing-body interactions characteristic of tilt rotors. The wing under the rotor is found to obstruct the inflow, causing a deficit in the inflow velocities over the inboard region of the rotor. Discrete frequency harmonic thickness and loading noise mechanisms in hover are examined by first modeling tilt rotor hover aerodynamics and then applying various noise prediction methods using the WOPWOP code. The analysis indicates that the partial ground plane created by the wing below the rotor results in a primary sound source for hover.

George, A. R.↗

NASA NTRS2001

Rotor Broadband Noise Prediction with Comparison to Model Data

This paper reports an analysis and prediction development of rotor broadband noise. The two primary components of this noise are Blade-Wake Interaction (BWI) noise, due to the blades' interaction with the turbulent wakes of the preceding blades, and "Self" noise, due to the development and shedding of turbulence within the blades' boundary layers. Emphasized in this report is the new code development for Self noise. The analysis and validation employs data from the HART program, a model BO-105 rotor wind tunnel test conducted in the German-Dutch Wind Tunnel (DNW). The BWI noise predictions are based on measured pressure response coherence functions using cross-spectral methods. The Self noise predictions are based on previously reported semiempirical modeling of Self noise obtained from isolated airfoil sections and the use of CAMRAD.Modl to define rotor performance and local blade segment flow conditions. Both BWI and Self noise from individual blade segments are Doppler shifted and summed at the observer positions. Prediction comparisons with measurements show good agreement for a range of rotor operating conditions from climb to steep descent. The broadband noise predictions, along with those of harmonic and impulsive Blade-Vortex Interaction (BVI) noise predictions, demonstrate a significant advance in predictive capability for main rotor noise.

Brooks, Thomas F.↗

NASA NTRS2015

Detached Eddy Simulation Results for a Space Launch System Configuration at Liftoff Conditions and Comparison with Experiment

Computational simulations for a Space Launch System configuration at liftoff conditions for incidence angles from 0 to 90 degrees were conducted in order to generate integrated force and moment data and longitudinal lineloads. While the integrated force and moment coefficients can be obtained from wind tunnel testing, computational analyses are indispensable in obtaining the extensive amount of surface information required to generate proper lineloads. However, beyond an incidence angle of about 15 degrees, the effects of massive flow separation on the leeward pressure field is not well captured with state of the art Reynolds Averaged Navier-Stokes methods, necessitating the employment of a Detached Eddy Simulation method. Results from these simulations are compared to the liftoff force and moment database and surface pressure data derived from a test in the NASA Langley 14- by 22-Foot Subsonic Wind Tunnel.

Computational fluid dynamics↗

NASA NTRS1984

Inlet flowfield investigation. Part 2: Computation of the flow about a supercruise forebody at supersonic speeds

A numerical procedure which solves the parabolized Navier-Stokes (PNS) equations on a body fitted mesh was used to compute the flow about the forebody of an advanced tactical supercruise fighter configuration in an effort to explore the use of a PNS method for design of supersonic cruise forebody geometries. Forebody flow fields were computed at Mach numbers of 1.5, 2.0, and 2.5, and at angles-of-attack of 0 deg, 4 deg, and 8 deg. at each Mach number. Computed results are presented at several body stations and include contour plots of Mach number, total pressure, upwash angle, sidewash angle and cross-plane velocity. The computational analysis procedure was found reliable for evaluating forebody flow fields of advanced aircraft configurations for flight conditions where the vortex shed from the wing leading edge is not a dominant flow phenomenon. Static pressure distributions and boundary layer profiles on the forebody and wing were surveyed in a wind tunnel test, and the analytical results are compared to the data. The current status of the parabolized flow flow field code is described along with desirable improvements in the code.

Paynter, G. C.↗

NASA NTRS1984

Report of the Panel on Aerodynamics

Progress in aerodynamics over the past 50 years has been evidenced by the development of increasingly sophisticated and efficient flight vehicles throughout the flight spectrum. Advances have generally arisen in an evolutionary manner from experience gained in wind tunnel testing, flight testing, and improvements in analytical and computational capabilities. As a result of this evolutionary development, both military and commercial vehicles operate at a relatively high efficiency level. This observation plus the fact that airplanes have not changed appreciably in outward appearance over recent years has led some skeptics to conclude incorrectly that aerodynamics is a mature technology, with little to be gained from further developments in the field. It is of interest to note that progress in aerodynamics has occurred without a thorough understanding of the fundamental physics of flow, turbulence, vortex dynamics, and separated flow, for example. The present understanding of transition, turbulence, and boundary layer separation is actually very limited. However, these fundamental flow phenomena provide the key to reducing the viscous drag of aircraft. Drag reduction provides the greatest potential for increased flight efficiency from the standpoint of both saving energy and maximizing performance. Recent advances have led to innovative concepts for reducing turbulent friction drag by modifying the turbulent structure within the boundary layer. Further advances in this basic area should lead to methods for reducing skin friction drag significantly. The current challenges for military aircraft open entirely new fields of investigation for the aerodynamicist. The ability through very high speed information processing technology to totally integrate the flight and propulsion controls can permit an aircraft to fly with "complete abandon," avoiding departure, buffet, and other undesirable characteristics. To utilize these new control concepts, complex aerodynamic phenomena will have to be understood, predicted, and controlled. Current requirements for military aircraft include configuration optimization through a widened envelope from subsonic to supersonic and from low to high angles of attack. This task is further complicated by requirements for control of observables. These challenging new designs do not have the luxury of a large experimental data base from which to optimize for various parameter combinations. Consequently, there exists a strong need for better techniques, both experimental and computational, to permit design optimization in a complete sense.

Bradley, Richard G.↗

NASA NTRS1991

Forebody tangential blowing for control at high angles of attack

A feasibility study to determine if the use of tangential leading edge blowing over the forebody could produce effective and practical control of the F-18 HARV aircraft at high angles of attack was conducted. A simplified model of the F-18 configuration using a vortex-lattice model was developed to obtain a better understanding of basic aerodynamic coupling effects and the influence of forebody circulation on lifting surface behavior. The effect of tangential blowing was estimated using existing wind tunnel data on normal forebody blowing and analytical studies of tangential blowing over conical forebodies. Incorporation of forebody blowing into the flight control system was investigated by adding this additional yaw control and sideforce generating actuator into the existing F-18 HARV simulation model. A control law was synthesized using LQG design methods that would schedule blowing rates as a function of vehicle sideslip, angle of attack, and roll and yaw rates.

Kroo, I.↗

NASA NTRS2010

Simulation Modeling for Off-Nominal Conditions - Where Are We Today?

The modeling of aircraft flight characteris4cs in off-nominal or otherwise adverse conditions has become increasingly important for simulation in the loss-of-control arena. Adverse conditions include environmentally-induced upsets such as wind shear or wake vortex encounters; off-nominal flight conditions, such as stall or departure; on-board systems failures; and structural failures or aircraft damage. Spirited discussions in the research community are taking place as to the fidelity and data requirements for adequate representation of vehicle dynamics under such conditions for a host of research areas, including recovery training, flight controls development, trajectory guidance/planning, and envelope limiting. The increasing need for multiple sources of data (empirical, computational, experimental) for modeling across a larger flight envelope leads to challenges in developing methods of appropriately applying or combining such data, particularly in a dynamic flight environment with a physically and/or aerodynamically asymmetric vehicle. Traditional simplifications and symmetry assumptions in current modeling methodology may no longer be valid. Furthermore, once modeled, challenges abound in the validation of flight dynamics characteristics in adverse flight regimes

Shah, Gautam H.↗

NASA NTRS2014

Stability and Control CFD Investigations of a Generic 53 Degree Swept UCAV Configuration

NATO STO Task Group AVT-201 on "Extended Assessment of Reliable Stability & Control Prediction Methods for NATO Air Vehicles" is studying various computational approaches to predict stability and control parameters for aircraft undergoing non-linear flight conditions. This paper contributes an assessment through correlations with wind tunnel data for the state of aerodynamic predictive capability of time-accurate RANS methodology on the group's focus configuration, a 53deg swept and twisted lambda wing UCAV, undergoing a variety of roll, pitch, and yaw motions. The vehicle aerodynamics is dominated by the complex non-linear physics of round leading-edge vortex flow separation. Correlations with experimental data are made for static longitudinal/lateral sweeps, and at varying frequencies of prescribed roll/pitch/yaw sinusoidal motion for the vehicle operating with and without control surfaces. The data and the derived understanding should prove useful to the AVT-201 team and other researchers who are developing techniques for augmenting flight simulation models from low-speed CFD predictions of aircraft traversing non-linear regions of a flight envelope.

Frink, Neal T.↗

NASA NTRS1986

Wave-interactions in three-dimensional boundary layers

A Fourier-Chebyshev spectral method is used to study the nonlinear stability and disturbance wave-interactions in a two-dimensional boundary layer on a curved wall and a three-dimensional boundary layer on a rotating disk. Goertler and crossflow instability, respectively, are the dominant mechanisms in these boundary layers. For the range of parameters considered, the present computational model does not support the Goertler/TS interaction of the type proposed by Nayfeh (1981). According to his model, oblique TS waves with spanwise wavelengths twice that of the finite-amplitude Goertler vortex are excited in the boundary layer. The present nonlinear calculations show that both the Goertler and crosssflow vortices develop strong instabilities in their harmonics with half the wavelength of the primary vortices.

Malik, M. R.↗

NASA NTRS1981

GRID3O: Computer program for fast generation of multilevel, three-dimensional boundary-conforming O-type computational grids

A fast algorithm was developed for accurately generating boundary-conforming, three-dimensional, consecutively refined computational grids applicable to arbitrary wing-body and axial turbomachinery geometries. The method is based on using an analytic function to generate two-dimensional grids on a number of coaxial axisymmetric surfaces positioned between the centerbody and the outer radial boundary. These grids are of the O-type and are characterized by quasi-orthogonality, geometric periodicity, and an adequate resolution throughout the flow field. Because the built-in nonorthogonal coordinate stretching and shearing cause the grid lines leaving the blade or wing trailing edge to end at downstream infinity, the numerical treatment of the three-dimensional trailing vortex sheets is simplified.

Dulikravich, D. S.↗

NASA NTRS1951

Integrals and integral equations in linearized wing theory

The formulas of subsonic and supersonic wing theory for source, doublet, and vortex distributions are reviewed and a systematic presentation is provided which relates these distributions to the pressure and to the vertical induced velocity in the plane of the wing. It is shown that care must be used in treating the singularities involved in the analysis and that the order of integration is not always reversible. Concepts suggested by the irreversibility of order of integration are shown to be useful in the inversion of singular integral equations when operational techniques are used. A number of examples are given to illustrate the methods presented, attention being directed to supersonic flight speed.

Lomax, Harvard↗

NASA NTRS1981

Direct numerical simulations of turbulent shear flows

Numerical simulations of wakes of axisymmetric bodies and of turbulent mixing layers are reported. The flows were assumed to be statistically homogeneous in the mean flow direction, in concert with experimental data and the self-similarity theorem. The nonlinear Navier-Stokes equations were solved by a pseudo-spectral numerical method using a 32 x 32 x 33 point grid and an algorithm for fast Fourier transforms and inverse transforms. Leapfrog time differencing was employed on nonlinear terms and time differencing on viscous terms. Towed wakes and wakes behind a self-propelled body were simulated, showing that the towed wakes exhibited a proper temporal behavior after an initial period of adjustment, including the development of a kurtosis near the wake edge, which is experimentally verifiable. The mixing-layer simulation displayed the laboratory demonstrated presence of large scale features such as vortex cores, while the lateral coherence was weak.

Metcalfe, R. W.↗

NASA NTRS1988

Navier-Stokes simulations of tip vortices for fixed and rotating helicopter blades

Flowfield and tip vortex results are presented for a hovering rotor blade at subcritical and supercritical flow conditions for both nonlifting and lifting configurations. These results are calculated numerically by solving, in a time-accurate fashion, the unsteady thin layer Navier-Stokes equations written in rotor coordinates. The lifting calculations use an induced downwash correction, estimated from a free wake analysis, to the geometric angle of attack of the blades to account for the wake effect. Comparison of numerical results with the experimental data shows very good agreement for all cases considered. Alternate methods of calculating hovering rotor flowfield as steady state flowfield on isolated fixed-blade that have the same circulation distribution as that of rotor in hover are also explored.

Srinivasan, G. R.↗

NASA NTRS1991

A Navier-Stokes study of shock-boundary layer interaction and flow separation inside a transonic compressor

A numerical study to evaluate a three-dimensional Navier-Stokes method as a tool to predict the detailed flowfield inside a low-aspect-ratio compressor at off-design rotor speed has been conducted. The flow field inside a state-of-the-art transonic compressor is used for the purpose of the evaluation. The experimental study shows that the rotor has higher peak efficiency at 90 percent rotor speed than at 100 percent rotor speed. The details of the flow structure inside the low-aspect-ratio compressor (three-dimensional shock structure, shock-boundary layer interaction, tip leakage vortex, etc.) and the overall aerodynamic performance at various operating conditions are numerically analyzed, and the results are compared with the available experimental data. The numerical results also indicate that the rotor has higher peak efficiency at 90 percent rotor speed than at 100 percent rotor speed. This is due to the reduced shock strength and the reduced interactions among passage shock, tip-clearance vortex and blade boundary layer.

Hah, Chunill↗

NASA NTRS1983

Small Gas Turbine Combustor Primary Zone Study

A development process is described which consists of design, fabrication, and preliminary test evaluations of three approaches to internal aerodynamic primary zone flow patterns: (1) conventional double vortex swirl stabilization; (2) reverse flow swirl stabilization; and (3) large single vortex flow system. Each concept incorporates special design features aimed at extending the performance capability of the small engine combustor. Since inherent geometry of these combustors result in small combustion zone height and high surface area to volume ratio, design features focus on internal aerodynamics, fuel placement, and advanced cooling. The combustors are evaluated on a full scale annular combustor rig. A correlation of the primary zone performance with the overall performance is accomplished using three intrusion type gas sampling probes located at the exit of the primary zone section. Empirical and numerical methods are used for designing and predicting the performance of the three combustor concepts and their subsequent modifications. The calibration of analytical procedures with actual test results permits an updating of the analytical design techniques applicable to small reverse flow annular combustors.

Sullivan, R. E.↗