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At least 1,459 records · Page 81

Tests at Mach Number 1.62 of a Series of Missile Configurations Having Tandem Cruciform Lifting Surfaces

An investigation at a Mach number of 1.62 was made in the Langley 9-inch supersonic tunnel of a series of missile configurations having tandem lifting surfaces of low aspect ratio and of newly equal span. Some of the variables investigated were interdigitation angle, wing and tail plan form, and longitudinal location of wing with respect to tail. All configurations were tested through an angle-of-attack range from -5 deg to 15 deg at roll angles of 0 deg and 45 deg. Lift, drag, and pitching moment data are presented, together with center-of-pressure locations and tail-lift efficiency factors.

Grigsby, Carl E↗

Supersonic Retropropulsion Experimental Results from the NASA Ames 9- x 7-Foot Supersonic Wind Tunnel

Supersonic retropropulsion was experimentally examined in the Ames Research Center 9x7-Foot Supersonic Wind Tunnel at Mach 1.8 and 2.4. The experimental model, previously designed for and tested in the Langley Research Center Unitary Plan Wind Tunnel at Mach 2.4, 3.5 and 4.6, was a 5-in diameter 70-deg sphere-cone forebody with a 9.55-in long cylindrical aftbody. The forebody was designed to accommodate up to four 4:1 area ratio nozzles, one on the model centerline and the other three on the half radius spaced 120-deg apart. Surface pressure and flow visualization were the primary measurements, including high-speed data to investigate the dynamics of the interactions between the bow and nozzle shocks. Three blowing configurations were tested with thrust coefficients up to 10 and angles of attack up to 20-deg. Preliminary results and observations from the test are provided

Berry, Scott A.↗

Simulation of Sweep-Jet Flow Control, Single Jet and Full Vertical Tail

This work is a simulation technology demonstrator, of sweep jet flow control used to suppress boundary layer separation and increase the maximum achievable load coefficients. A sweep jet is a discrete Coanda jet that oscillates in the plane parallel to an aerodynamic surface. It injects mass and momentum in the approximate streamwise direction. It also generates turbulent eddies at the oscillation frequency, which are typically large relative to the scales of boundary layer turbulence, and which augment mixing across the boundary layer to attack flow separation. Simulations of a fluidic oscillator, the sweep jet emerging from a nozzle downstream of the oscillator, and an array of sweep jets which suppresses boundary layer separation are performed. Simulation results are compared to data from a dedicated validation experiment of a single oscillator and its sweep jet, and from a wind tunnel test of a full-scale Boeing 757 vertical tail augmented with an array of sweep jets. A critical step in the work is the development of realistic time-dependent sweep jet inflow boundary conditions, derived from the results of the single-oscillator simulations, which create the sweep jets in the full-tail simulations. Simulations were performed using the computational fluid dynamics (CFD) solver Overow, with high-order spatial discretization and a range of turbulence modeling. Good results were obtained for all flows simulated, when suitable turbulence modeling was used.

flow control↗

Ice Roughness and Thickness Evolution on a Swept NACA 0012 Airfoil

Several recent studies have been performed in the Icing Research Tunnel (IRT) at NASA Glenn Research Center focusing on the evolution, spatial variations, and proper scaling of ice roughness on airfoils without sweep exposed to icing conditions employed in classical roughness studies. For this study, experiments were performed in the IRT to investigate the ice roughness and thickness evolution on a 91.44-cm (36-in.) chord NACA 0012 airfoil, swept at 30-deg with 0deg angle of attack, and exposed to both Appendix C and Appendix O (SLD) icing conditions. The ice accretion event times used in the study were less than the time required to form substantially three-dimensional structures, such as scallops, on the airfoil surface. Following each ice accretion event, the iced airfoils were scanned using a ROMER Absolute Arm laser-scanning system. The resulting point clouds were then analyzed using the self-organizing map approach of McClain and Kreeger to determine the spatial roughness variations along the surfaces of the iced airfoils. The resulting measurements demonstrate linearly increasing roughness and thickness parameters with ice accretion time. Further, when compared to dimensionless or scaled results from unswept airfoil investigations, the results of this investigation indicate that the mechanisms for early stage roughness and thickness formation on swept wings are similar to those for unswept wings.

glaze icing↗

Characterization of Slat Noise Radiation from a High-Lift Common Research Model

A test campaign was conducted in the open-jet test section of the NASA Langley 14- by 22-Foot Subsonic Tunnel on a 10%-scale, semispan version of a High-Lift Common Research Model (CRM-HL) incorporating a leading-edge slat, trailing-edge flap and removable high-fidelity main landing gear. Aerodynamic and far field acoustic measurements were obtained on baseline and acoustically treated model configurations using an ensemble of time-averaged and unsteady surface pressure sensors paired with a traversing 97-microphone phased array viewing the pressure side of the airframe. A modified version of the DAMAS deconvolution method incorporating corrections for shear layer acoustic wave decorrelation was employed to determine the locations and strengths of relevant noise sources along the span of the slat and in the vicinity of the flap edges and landing gear. The main goal of the test campaign was to use the surface pressure and array data to evaluate the effectiveness of various slat noise reduction concepts. It was found that slat-gap fillers produced significant broadband noise reduction with minimal aerodynamic performance penalties for all Mach numbers and airfoil angles of attack that were investigated. The test campaign clearly demonstrated the noise reduction benefits that can be obtained by applying appropriate treatments to leading-edge slats on commercial transport-class aircraft.

Aeronautics Research Mission Directorate↗

3D reconstruction and neural rendering for adversarial machine learning

While evasion attacks on computer vision systems have been widely studied, creating attacks that remain effective under significant changes in viewpoint continues to be challenging. Traditional approaches often rely on affine transformations of images, but these approaches degrade at larger perspective shifts and often produce unrealistic or ineffective perturbations. Recent methods use differentiable renderers to improve viewpoint robustness, but they typically depend on manually constructed 3D models. We introduce a semi-automated pipeline that generates physically printable and perspective-invariant adversarial patches using only a small set of 2D images. Our method integrates 3D reconstruction, neural rendering, adversarial patch optimization, and an object detection victim model into a unified workflow. We use 2D Gaussian Splatting for high fidelity mesh reconstruction and FlexPara for surface parameterization that produces texture maps suitable for patch editing. Together, these components form a fully differentiable pipeline in PyTorch3D that links texture modification to model outputs, enabling efficient optimization of patches that remain effective across many viewpoints. The complete process, from image capture to patch printing and physical evaluation, can be completed within a few hours. We demonstrate the effectiveness of the resulting patches through attacks on the YOLOv8 object detection model and discuss remaining challenges and opportunities for improving robustness and scalability.

Singhvi, Vivaan [ORNL] (ORCID:0009000586288221)↗

Results of tests (OA64 and IA35) of an 0.015-scale model (36-OTS) of the space shuttle configuration 140A/B in the NASA/LaRC unitary plan wind tunnel

Supersonic aerodynamic tests were conducted in the NASA/LaRC Unitary Plan Wind Tunnel on a 0.015-scale model of the Space Shuttle Vehicle 4 configuration. Surface pressure data were obtained for both the launch (test IA35) and entry (test OA64) configurations at Mach numbers from 2.5 to 4.5. The surface pressure was measured in the vicinity of the cargo bay door hinge and parting lines and on the side of the fuselage at the crew compartment and below the OMS pods at the aft compartment. Data were obtained for angles of attack and sideslip consistent with the expected excursions about the nominal trajectory values at the test Mach numbers. These tests were a part of a test series supporting the orbiter venting analysis over a Mach number range from 0.6 to 10.4

Thornton, D. E.↗

Lateral stability and control tests of the XP-77 airplane in the NACA full-scale tunnel, 16 June 1944

Tests were made in the NACA full-scale tunnel to determine the lateral stability and control characteristics of the XP-77 airplane. Measurements were made of the forces and moments on the airplane at various angles of attack and angles of yaw. The measurements were made with the propeller removed and with the propeller installed and operating at various thrust coefficients, and with the landing flaps retracted and deflected. The effects of aileron, elevator, and rudder deflection on control surface effectiveness and hinge moments were determined. The tests were planned to obtain the data required to evaluate as completely as possible the Army Air Force requirements on lateral stability and control for pursuit-type airplanes.

Czarnecki, K. R.↗

Burner rig alkali salt corrosion of several high temperature alloys

The hot corrosion of five alloys was studied in cyclic tests in a Mach 0.3 burner rig into whose combustion chamber various aqueous salt solutions were injected. Three nickel-base alloys (IN-792, IN-738, and IN-100), a cobalt-base alloy (MM-509), and an iron-base alloy (304 stainless steel) were studied at temperatures of 700, 800, 900, and 1000 C with various salt concentrations and compositions. The relative resistance of the alloys to hot corrosion attack was found to vary with temperature and with both the concentration and composition of the injected salt solution. Results indicate that the corrosion of these alloys is a function of both the presence of salt condensed as a liquid on the surface and of the composition of the gas phases present.

Deadmore, D.↗

Supersonic aerodynamic characteristics of a tail-control cruciform maneuverable missile with and without wings

The aerodynamic characteristics for a winged and a wingless cruciform missile are examined. The body was an ogive-cylinder with a 3.5 caliber forebody; an overall length-to-diameter ratio of 11.667; and has cruciform tails that were trapexoidal in planform. Tests were made both with and without 72.9 deg cruciform delta wings. The investigation was made for Mach numbers from 1.50 to 4.63, roll attitudes of 0 and 45 deg, angles of attack from -40 to 22 deg, and tail control deflections from 10 to -40 deg. The purpose is to determine the influence of the aerodynamic behavior on the design choice for maneuverable missiles intended primarily for air-to-air or surface-to-surface missions. The results indicate that the winged missile with its more linear aerodynamic characteristics and higher lift-curve slope, should provide the highest maneuverability over a large operational range.

Spearman, M. L.↗

A vortex-lattice method for calculating lifting-surface interference

The interference effect of closely coupled lifting surfaces was investigated. A nonsteady vortex-lattice method was applied to calculate the steady-state and nonsteady lift characteristics of configurations having interference effects. The configurations consist of low-aspect-ratio delta wings, at low and high angles of attack. Leading-edge separation and wake roll-up are simulated by a time-dependent wake-shedding procedure. For steady-state flow conditions, this numerical procedure saves a considerable amount of computer time, compared with iterative methods, and yields the same results. A better understanding of the interference effect can be gained by using the method to study the transient behavior. The nonsteady approach also offers the capability of calculating various nonsteady motions, as is demonstrated in the calculation of the longitudinal damping.

Levin, D.↗

Aerothermodynamic environment of a Titan aerocapture vehicle

The extent of convective and radiative heating for a Titan aerocapture vehicle is investigated. The flow in the shock layer is assumed to be axisymmetric, steady, viscous, and compressible. It is further assumed that the gas is in chemical and local thermodynamic equilibrium and tangent slab approximation is used for the radiative transport. The effect of the slip boundary conditions on the body surface and at the shock wave are included in the analysis of high-altitude entry conditions. The implicit finite difference techniques is used to solve the viscous shock-layer equations for a 45 degree sphere cone at zero angle of attack. Different compositions for the Titan atmosphere are assumed, and results are obtained for the entry conditions specified by the Jet Propulsion Laboratory.

Tiwari, S. N.↗

Comparison of pressure distributions on model and full-scale NACA 64-621 airfoils with ailerons for wind turbine application

The aerodynamic similarity between a small (4-inch chord) wind tunnel model and a full-scale wind turbine blade (24-foot tip section with a 36-inch chord) was evaluated by comparing selected pressure distributions around the geometrically similar cross sections. The airfoils were NACA 64-621 sections, including trailing-edge ailerons with a width equal to 38 percent of the airfoil chord. The model airfoil was tested in the OSU 6- by 12-inch High Reynolds Number Wind Tunnel; the full-scale blade section was tested in the NASA Langley Research Center 30- by 60-foot Subsonic Wind Tunnel. The model airfoil contained 61 pressure taps connected by embedded tubes to pressure transducers. A belt containing 29 pressure taps was fixed to the full-scale section at midspan to obtain surface pressure data. Lift coefficients were obtained by integrating pressures, and corrections were made for the 3-D effects of blade twist and downwash in the blade tip section. The results of the two different experimental methods correlated well for angles of attack from minus 4 to 36 degrees and aileron reflections from 0 to 90 degrees.

Gregorek, G. M.↗

Flat plate at incidence as a waverider in rarefied hypersonic flow

The physical validity of continuum models and their ability to predict the critical aerothermodynamic properties of a waverider at high altitudes are examined using a flat plate at angle of attack as a generic hypersonic lifting vehicle. For a shock layer far from local translational equilibrium, a theoretical study based on Grad's thirteen-moment equations shows that the Navier-Stokes based solutions can correctly predict the drag, lift, and surface heat transfer rate, with the prediction error comparable to that of the standard shock-layer theory. The conclusion is supported by a comparison with direct simulation Monte Carlo calculations.

Cheng, H. K.↗

Flight Investigation at High Speeds of the Drag of Three Airfoils and a Circular Cylinder Representing Full-Scale Propeller Shanks

Tests have been made at high speeds to determine the drag of models, simulating propeller shanks, in the form of a circular cylinder and three airfoils, the NACA 16-025, the NACA 16-040, and the NACA 16-040 with the rear 25 percent chord cut off. All the models had a maximum thickness of 4 1/2 inches to conform with average propeller-shank dimensions and a span of 20 1/4 inches. For the tests the models were supported perpendicular to the lower surface of the wing of an XP-51 airplane. A wake-survey rake mounted below the wing directly behind the models was used to determine profile drag of Mach numbers of 0.3 to 0.8 over a small range of angle of attack. The drag of the cylinder was also determined from pressure-distribution and force measurements.

Barlow, William H↗

Effect of Compressibility on the Pressure and Forces Acting on a Modified NACA 65,3-019 Airfoil Having A 0.20-Chord Flap

An investigation has been conducted in the Langley rectangular high-speed tunnel to determine the effect of compressibility on the pressure distribution for a modified NACA 65,3-019 airfoil having a 0.20-chord flap. The investigation was made for an angle-of-attack range extending from -2 to 12 deg at .20 flap deflections from 0 to -12 deg. Test data were obtained for Mach numbers from 0.28 to approximately 0.74. The results show that the effectiveness of the trailing-edge-type control surface rapidly decreased and approached zero as the Mach number increased above the critical value.

AIRFOILS-NACA 65,3-019 (MOD )-MOMENTS & ORDINATES ↗

Hot Hydrogen Exposure Degradation of the Strength of Mullite

This study deals with the corrosion of near stoichiometric mullite (3Al2O3-2SiO2) by pure dry hydrogen gas. Exposure of the mullite samples was at temperatures of 1050 and 1250 C for times up to 500 hours. Preferential attack of the alumino-silicate glass present in the grain boundaries of the mullite occurred after 125 hours at 1250 C. Hydrogen scrubbing of the SiO2 from the glassy grain boundaries and the mullite grains yielded a porous alumina-rich surface. The room temperature strength increased after short exposure times at 1250 C (up to 125 hours), then decreased by 53 percent after exposure for 500 hours. At 1050 C, all exposure times (25 to 500 hours) decreased the strength. After 500 hours in hydrogen at 1050 C, the room temperature strength of mullite decreased 22 percent. We also observed a rapid 25 percent strength loss after short exposure times at 1050 C. This is attributed to the calcium/hydrogen assisted crystallization of the glassy, grain-boundary phase.

Herbell, Thomas P.↗

Aerothermodynamic Measurement and Prediction for Modified Orbiter at Mach 6 and 10

Detailed heat-transfer rate distributions measured laterally over the windward surface of an orbiter-like configuration using thin-film resistance heat-transfer gauges and globally using the newly developed relative intensity, two-color thermographic phosphor technique are presented for Mach 6 and 10 in air. The angle of attack was varied from 0 to 40 deg, and the freestream Reynolds number based on the model length was varied from 4 x 10(exp 5) to 6 x 10(exp 6) at Mach 6, corresponding to laminar, transitional, and turbulent boundary layers; the Reynolds number at Mach 10 was 4 x 10(exp 5), corresponding to laminar flow. The primary objective of the present study was to provide detailed benchmark heat-transfer data for the calibration of computational fluid-dynamics codes. Predictions from a Navier-Stokes solver referred to as the Langley aerothermodynamic upwind relaxation algorithm and an approximate boundary-layer solving method known as the axisymmetric analog three-dimensional boundary layer code are compared with measurement. In general, predicted laminar heat-transfer rates are in good agreement with measurements.

Micol, John R.↗