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NASA NTRS1990

Forebody vortex control as a complement to thrust vectoring

The desire to enhance the controllability of fighter aircraft at high angles of attack, particularly yaw control, has fostered an interest in both vectored thrust and active control of forebody vortices. This paper reviews several methods of forebody vortex control that have been investigated with water and wind tunnel tests of both generic and actual fighter configurations. The methods investigated include pneumatic or blowing techniques using surface-mounted jets and slots, surface suction, variable-height deployable strakes, and rotatable tip strakes. Flow visualization, and force and moment measurements have shown that all of the methods are effective in manipulating the forebody vortices over a wide range of angles of attack and sideslip, primarily through control over flow separation on the surface of the forebody. All are most effective when applied near the forebody tip. The advantages and limitations of the various methods are reviewed.

Malcolm, G. N.↗

DOE OSTI2023

Understanding the mechanisms of anisotropic dissolution in metal oxides by applying radiolysis simulations to liquid-phase TEM

Iron-based redox-active minerals are ubiquitous in soils, sediments, and aquatic systems. Their dissolution is of great importance for microbial impacts on carbon cycling and the biogeochemistry of the lithosphere and hydrosphere. Despite its widespread significance and extensive prior study, the atomic-to-nanoscale mechanisms of dissolution remain poorly understood, particularly the interplay between acidic and reductive processes. Here, we use in situ liquid-phase-transmission electron microscopy (LP-TEM) and simulations of radiolysis to probe and control acidic versus reductive dissolution of akaganeite (β–FeOOH) nanorods. Informed by crystal structure and surface chemistry, the balance between acidic dissolution at rod tips and reductive dissolution at rod sides was systematically varied using pH buffers, background chloride anions, and electron beam dose. We find that buffers, such as bis-tris, effectively inhibited dissolution by consuming radiolytic acidic and reducing species such as superoxides and aqueous electrons. In contrast, chloride anions simultaneously suppressed dissolution at rod tips by stabilizing structural elements while promoting dissolution at rod sides through surface complexation. Dissolution behaviors were systematically varied by shifting the balance between acidic and reductive attacks. The findings show LP-TEM combined with simulations of radiolysis effects can provide a unique and versatile platform for quantitatively investigating dissolution mechanisms, with implications for understanding metal cycling in natural environments and the development of tailored nanomaterials.

37 INORGANIC, ORGANIC, PHYSICAL, AND ANALYTICAL CH↗

DOE OSTI2020

Water Energy Resource Innovation on the Cavitation Characteristics

The main purpose of this study is to numerically correlate the amount of generated vapor over a hydrofoil to the lift and drag coefficients acting on it. As such, cavitation characteristics were investigated of a hydrofoil in the cavitating, sub-cavitating, and non-cavitating flows for different angles of attacks (AoA) with the high upstream flow velocity. The hydrofoil was tested in a square water tunnel with water entering the tunnel at various velocities for each AoA ranges from 9.1 m/s to 12.2 m/s. It was found that lift and drag forces acting on the hydrofoil follow the trend of the experimental data quite closely. While the cavitation can be identified by a unique number (averaged vapor volume fraction), the work done created an inverse correlation between this number and the cavitation number at the same angle of attack. The lift force declines with the increase in the vapor content on the hydrofoil surface, meanwhile the drag force peaks at certain vapor volume fraction, and then, a huge reduction occurs with the considerable decrease in the corresponding cavitation number. A fourth-order correlation generated between the lift to drag (L/D) and the cavitation number (σ). It was found the lift-to-drag ratio decreases by the formation of the cavitation over the hydrofoil, thus causing a drop in the efficiency of the turbomachines.

42 ENGINEERING↗

DOE OSTI2025

Developing Machine Learning Interatomic Potential for Fe-Cr-Ni Alloys

Accurate prediction of creep and fatigue behavior of stainless steel at elevated temperatures in hydrogen environment requires fundamental understanding of alloy-hydrogen interaction at cross-scale including bulk lattice and key defects such as vacancies, grain boundaries, surfaces, stacking faults, dislocations, and precipitates. This project aims to predict creep behavior of 347H stainless steel with H using machine learning interatomic potentials based on first-principles density functional theory simulations. The Moment Tensor Potentials platform is adopted for this work since it demonstrates a fine balance between model accuracy and computational efficiency. The potential is well trained based on large amount of high-fidelity density functional theory calculations. The validation is carried out by comparing various important properties including short range order, coefficient of thermal expansion, elastic properties, stacking fault energy, grain boundary energy, and surface energy. This work lays the foundation for reliable atomistic simulation of high temperature hydrogen attack of stainless steel.

density functional theory (DFT)↗

NASA NTRS1971

Sharp flat plate heat transfer in helium at Mach numbers of 22.8 to 86.8 and in corner flow with air at Mach number of 19

Surface heat transfer rates were measured on a sharp flat plate at zero angle of attack in a hypersonic shock tunnel. The density and leading edge Knudsen number were varied to span the continuum to near free molecule regimes. The strong interaction parameter varied from 11 to 16,000 with Knudsen numbers from 0.56 to 17.1 respectively. Local heat transfer rates in the corner flow region produced by the intersection of two perpendicular flat plates with sharp leading edges were determined for various flow densities. The strength of the shock wave from the vertical plate was varied by adjusting the angle of attack from 0 to 5 deg. The unit Reynolds number varied from 1,000 to 17,200 and the Knudsen numbers from 1.6 to 27. The strong interaction parameter varied from 14 to 500.

Nagamatsu, H. T.↗

NASA NTRS1975

Wind tunnel investigation of aerodynamic loads on a large-scale externally blown flap model and comparison with theory

Results from a wind-tunnel investigation of a large-scale externally blown flap model are presented. The model was equipped with four turbofan engines, a triple-slotted flap system, and a T-tail. The wing had a quarter-chord sweep of 25 deg, an aspect ratio of 7.28, and a taper ratio of 0.4. Aerodynamic loads and load distributions were determined from a total of 564 static pressure orifices located on the upper and lower surfaces of the slat, wing, and flaps. Loads are presented for variations of angle of attack, engine thrust setting, and flap deflection angle. In addition, the experimental results are compared with analytical results calculated by using a potential flow analysis.

Perry, B., III↗

NASA NTRS1982

Sharp flat plate heat transfer in helium at Mach numbers of 22.8 to 86.8

Surface heat transfer rates were measured on a sharp flat plate at zero angle of attack with helium over a Mach number range of 22.8 to 86.8 in a hypersonic shock tunnel. The density and leading edge Knudsen numbers were varied to span the continuum to near free molecule regimes. The strong interaction parameter varied from 11 to 16,000 with leading edge Knudsen numbers from 0.56 to 17.1, respectively. For high unit Reynolds number the rate of change of the heat transfer coefficient with the strong interaction parameter agreed well with the strong interaction theory of Li and Nagamatsu (1953, 1955).

Nagamatsu, H. T.↗

NASA NTRS1983

A study of the unsteady flow field of an airfoil with deflected spoiler

The flow field generated by a two-dimensional airfoil with a statically deflected flap type spoiler has been experimentally studied. Subsonic tests ascertained the effect of varying spoiler deflection and angle of attack on the flow field at two Reynolds numbers. Measurements of the mean and fluctuating surface pressures and velocities in the wake, along with schlieren flow visualization of the near wake, were made. The results pertaining to the unsteady characteristics of the spoiler flow field are presented. It is shown that vortex shedding characterized the spoiler wake and determines the unsteady nature of the overall flow field; the character of the vortex shedding changes with spoiler deflection; and the fluctuating surface pressure field (vortex shedding frequency component) is of appreciable magnitude at large spoiler deflections.

Mclachlan, B. G.↗

NASA NTRS1984

Numerical simulation of the tip vortex off a low-aspect-ratio wing at transonic speed

The viscous transonic flow around a low aspect ratio wing was computed by an implicit, three dimensional, thin-layer Navier-Stokes solver. The grid around the geometry of interest is obtained numerically as a solution to a Dirichlet problem for the cube. A low aspect ratio wing with large sweep, twist, taper, and camber is the chosen geometry. The topology chosen to wrap the mesh around the wing with good tip resolution is a C-O type mesh. The flow around the wing was computed for a free stream Mach number of 0.82 at an angle of attack of 5 deg. At this Mach number, an oblique shock forms on the upper surface of the wing, and a tip vortex and three dimensional flow separation off the wind surface are observed. Particle path lines indicate that the three dimensional flow separation on the wing surface is part of the roots of the tip vortex formation. The lifting of the tip vortex before the wing trailing edge is observed by following the trajectory of particles release around the wing tip.

Mansour, N. N.↗

NASA NTRS1984

Numerical simulation of the tip vortex off a low-aspect-ratio wing at transonic speed

The viscous transonic flow around a low-aspect-ratio wing has been computed using an implicit, three-dimensional, 'thin-layer' Navier-Stokes solver. The grid around the geometry of interest is obtained numerically as a solution to a Dirichlet problem for the cube. The geometry chosen for this study is a low-aspect-ratio wing with large sweep, twist, taper, and camber. The topology chosen to wrap the mesh around the wing with good tip resolution is a C-O type mesh. Using this grid, the flow around the wing was computed for a free-stream Mach number of 0.82 at an angle of attack of 5 deg. At this Mach number, an oblique shock forms on the upper surface of the wing, and a tip vortex and three-dimensional flow separation off the wing surface are observed. Particle path lines indicate that the three-dimensional flow separation on the wing surface is part of the roots of the tip-vortex formation. The lifting of the tip vortex before the wing trailing edge is clearly observed by following the trajectory of particles released around the wing tip.

Mansour, N. N.↗

NASA NTRS1989

Unsteady blade pressure measurements on a model counterrotation propeller

In an exploratory effort an advanced counterrotation propeller instrumented with blade-mounted pressure transducers was tested in the NASA Lewis 9- by 15-Foot Anechoic Wind Tunnel at a simulated takeoff and landing speed of Mach 0.20. The propeller's aft diameter was reduced to investigate possible noise reductions resulting from reduced blade row interaction with the tip vortex. The propeller was tested at three blade row spacings at fixed blade setting angles, at the maximum blade row spacing at higher blade setting angles and at propeller axis angles attack to the flow up to + or - 16 deg. A limited number of unsteady blade surface pressure measurements were made on both rotors of the model counterrotation propeller. Emphasis was placed on determining the effects of rotor-rotor interactions on the blade surface pressures. A unique method of processing the pressure signals was developed that enables even weak interaction waveforms and spectra to be separated from the total signal. The interaction on the aft rotor was many times stronger than that on the forward rotor. The fundamental rotor interaction tone exhibited complicated behavior but generally increased with rotational speed and blade setting angle and decreased with rotor spacing. With the propeller axis at an angle to the flow, the phase response of the aft rotor appeared to be significantly affected by the presence of the forward rotor.

Heidelberg, Laurence J.↗

NASA NTRS1989

Unsteady blade pressure measurements on a model counterrotation propeller

Heidelberg, Laurence J.↗

NASA NTRS1990

Multiple vortex and shock interactions at subsonic, transonic, and supersonic speeds

A wind tunnel experiment was conducted at subsonic, transonic, and supersonic speeds of the vortex-vortex and vortex-shock interactions about a tailless, general research fighter model having chine-like forebody strakes faired into a 55 deg cropped delta wing. The present paper isolates the results obtained at angle of attack 20 deg and free-stream Mach = 0.6 to 1.6, which include off-surface and on-surface flow visualizations, two-component laser velocimeter measurements, and wing upper surface static pressure distributions. Increasing the Mach number decreased the direct interaction (intertwining) of the forebody strake and wing vortex cores. An early bursting of the wing vortex occurred at free-stream Mach = 0.8, where the flow field was in transition from the intertwining vortices characteristic of the lower subsonic speeds to the decoupled vortices at the transonic and supersonic speeds. The vortex interaction and breakdown were sensitive to the character of the secondary boundary layer separation on the wing, which may be shock-induced at free-stream Mach = 0.8 to 0.95.

Erickson, Gary E.↗

NASA NTRS1991

Wind tunnel investigation of vortex flows on F/A-18 configuration at subsonic through transonic speed

A wind tunnel experiment was conducted in the David Taylor Research Center 7- by 10-Foot Transonic Tunnel of the wing leading-edge extension (LEX) and forebody vortex flows at subsonic and transonic speeds about a 0.06-scale model of the F/A-18. The primary goal was to improve the understanding and control of the vortical flows, including the phenomena of vortex breakdown and vortex interactions with the vertical tails. Laser vapor screen flow visualizations, LEX, and forebody surface static pressures, and six-component forces and moments were obtained at angles of attack of 10 to 50 degrees, free-stream Mach numbers of 0.20 to 0.90, and Reynolds numbers based on the wing mean aerodynamic chord of 0.96 x 10(exp 6) to 1.75 x 10(exp 6). The wind tunnel results were correlated with in-flight flow visualizations and handling qualities trends obtained by NASA using an F-18 High-Alpha Research Vehicle (HARV) and by the Navy and McDonnell Douglas on F-18 aircraft with LEX fences added to improve the vertical tail buffet environment. Key issues that were addressed include the sensitivity of the vortical flows to the Reynolds number and Mach number; the reduced vertical tail excitation, and the corresponding flow mechanism, in the presence of the LEX fence; the repeatability of data obtained during high angle-of-attack wind tunnel testing of F-18 models; the effects of particle seeding for flow visualization on the quantitative model measurements; and the interpretation of off-body flow visualizations obtained using different illumination and particle seeding techniques.

Erickson, Gary E.↗

NASA NTRS1992

Preliminary subsonic aerodynamic model for simulation studies of the HL-20 lifting body

A nonlinear, six-degree-of-freedom aerodynamic model for an early version of the HL-20 lifting body is described and compared with wind tunnel data upon which it is based. Polynomial functions describing most of the aerodynamic parameters are given and tables of these functions are presented. Techniques used to arrive at these functions are described. Basic aerodynamic coefficients were modeled as functions of angles of attack and sideslip. Vehicle lateral symmetry was assumed. Compressibility (Mach) effects were ignored. Control-surface effectiveness was assumed to vary linearly with angle of deflection and was assumed to be invariant with the angle of sideslip. Dynamic derivatives were obtained from predictive aerodynamic codes. Landing-gear and ground effects were scaled from Space Shuttle data. The model described is provided to support pilot-in-the-loop simulation studies of the HL-20. By providing the data in tabular format, the model is suitable for the data interpolation architecture of many existing engineering simulation facilities. Because of the preliminary nature of the data, however, this model is not recommended for study of the absolute performance of the HL-20.

Jackson, E. Bruce↗

NASA NTRS1992

Experimental results on a wall interference correction method with interface measurements

A wall interference assessment and correction method for subsonic two-dimensional wind tunnel testing is presented. This method calculates a pressure coefficient and angle of attack correction based on velocity measurements on interfaces inside of the wind tunnel. A mathematical representation of the test article is not required. An experimental verification of the suggested technique is given. A NACA 0012 airfoil is tested at a Mach number of 0.70 and at two different angles of attack. Calculated blockage corrections show reasonable agreement with results based on Hackett's method. Corrected surface pressures compare favorably to free-air flow field data if the tunnel flow field is subsonic. The present wall interference correction method can be applied to transonic tunnel flow fields with some restrictions. Errors are estimated and it is shown that the expected error in calculating the pressure coefficient correction on the model surface is in the order of the error of pressure coefficient measurement on interfaces. Necessary testing equipment in existing tunnels can easily be modified if the present method is applied.

Lo, C. F.↗

NASA NTRS1993

Space environmental effects on LDEF composites: A leading edge coated graphite epoxy panel

The electronics module cover for the leading edge (Row D 9) experiment M0003-8 was fabricated from T300 graphite/934 epoxy unidirectional prepreg tape in a (O(sub 2), +/- 45, O(sub 2), +/- 45, 90, 0)(sub s) layup. This 11.75 in x 16.75 in panel was covered with thermal control coatings in three of the four quadrants with the fourth quadrant uncoated. The composite panel experienced different thermal cycling extremes in each quadrant due to the different optical properties of the coatings and bare composite. The panel also experienced ultraviolet (UV) and atomic oxygen (AO) attack as well as micrometeoroid and space debris impacts. An AO reactivity of 0.99 x 10(exp -24) cm(sup 3)/atom was calculated for the bare composite based on thickness loss. The white urethane thermal control coatings (A276 and BMS 1060) prevented AO attack of the composite substrate. However, the black urethane thermal control coating (Z306) was severely eroded by AO, allowing some AO attack of the composite substrate. An interesting banding pattern on the AO eroded bare composite surface was investigated and found to match the dimensions of the graphite fiber tow widths as prepregged. Also, erosion depths were greater in the darker bands. Five micrometeoroid/space debris impacts were cross sectioned to investigate possible structural damage as well as impact/AO interactions. Local crushing and delaminations were found to some extent in all of the impacts. No signs of coating undercutting were observed despite the extensive AO erosion patterns seen in the exposed composite material at the impact sites. An extensive microcrack study was performed on the panel along with modeling of the thermal environment to estimate temperature extremes and thermal shock. The white coated composite substrate displayed almost no microcracking while the black coated and bare composite showed extensive microcracking. Significant AO erosion was seen in many of the cracks in the bare composite.

George, Pete E.↗

NASA NTRS1958

Estimation of Directional Stability Derivatives at Small Angles and Subsonic and Supersonic Speeds

Methods are presented for estimating the directional stability derivative increments contributed by the stabilizing surfaces of subsonic and supersonic aircraft. These methods are strictly applicable at zero angle of attack and small angles of sideslip. The procedure of totaling the incremental coefficients to obtain an estimation of the total empennage side-force and yawing-moment coefficient derivatives is also shown, together with numerical examples. A correlation is presented between estimated and experimental incremental coefficients which indicates that the methods of this report generally estimate the increment of side force gained by the addition of a panel to within +/-10 percent of the experimental value while the yawing-moment increment is generally estimated to within +/-20 percent. This is true for both subsonic and supersonic Mach numbers. An example application of the methods to one of the problems in directional stability, that of minimizing the effect of Mach number on the side-force coefficient derivative of the empennage, is discussed.

Goodwin, Frederick K.↗