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At least 685 records · Page 38
All sourcesNASA NTRSDOE OSTI
NASA NTRS2002

Results from a Sting Whip Correction Verification Test at the Langley 16-Foot Transonic Tunnel

In recent years, great strides have been made toward correcting the largest error in inertial Angle of Attack (AoA) measurements in wind tunnel models. This error source is commonly referred to as 'sting whip' and is caused by aerodynamically induced forces imparting dynamics on sting-mounted models. These aerodynamic forces cause the model to whip through an arc section in the pitch and/or yaw planes, thus generating a centrifugal acceleration and creating a bias error in the AoA measurement. It has been shown that, under certain conditions, this induced AoA error can be greater than one third of a degree. An error of this magnitude far exceeds the target AoA goal of 0.01 deg established at NASA Langley Research Center (LaRC) and elsewhere. New sting whip correction techniques being developed at LaRC are able to measure and reduce this sting whip error by an order of magnitude. With this increase of accuracy, the 0.01 deg AoA target is achievable under all but the most severe conditions.

Crawford, B. L.↗
NASA NTRS2002

Prediction of Aerodynamic Coefficients using Neural Networks for Sparse Data

Basic aerodynamic coefficients are modeled as functions of angles of attack and sideslip with vehicle lateral symmetry and compressibility effects. Most of the aerodynamic parameters can be well-fitted using polynomial functions. In this paper a fast, reliable way of predicting aerodynamic coefficients is produced using a neural network. The training data for the neural network is derived from wind tunnel test and numerical simulations. The coefficients of lift, drag, pitching moment are expressed as a function of alpha (angle of attack) and Mach number. The results produced from preliminary neural network analysis are very good.

Rajkumar, T.↗
NASA NTRS2002

Performance Study of a Ducted Fan System

An experimental investigation was completed in the NASA Ames 7- by 10-Foot Wind Tunnel with the objective of determining the performance characteristics of a ducted fan. The model was an annular duct with a 38-in diameter, 10-in chord, and a 5-bladed fixed-pitch fan. Model variations included duct angle of attack, exit vane flap length, flap deflection angle, and duct chord length. Duct performance data were obtained for axial and forward flight test conditions. Axial flow test data showed figure of merit decreases with increasing advance ratio. Forward flight data showed an increasing propulsive force with decreasing duct angle of attack. Exit vane flap deflection angle and flap chord length were shown to be an effective way of providing side force. Extending the duct chord did not effect the duct performance.

Abrego, Anita I.↗
NASA NTRS2002

Sliding Mode Control Applied to Reconfigurable Flight Control Design

Sliding mode control is applied to the design of a flight control system capable of operating with limited bandwidth actuators and in the presence of significant damage to the airframe and/or control effector actuators. Although inherently robust, sliding mode control algorithms have been hampered by their sensitivity to the effects of parasitic unmodeled dynamics, such as those associated with actuators and structural modes. It is known that asymptotic observers can alleviate this sensitivity while still allowing the system to exhibit significant robustness. This approach is demonstrated. The selection of the sliding manifold as well as the interpretation of the linear design that results after introduction of a boundary layer is accomplished in the frequency domain. The design technique is exercised on a pitch-axis controller for a simple short-period model of the High Angle of Attack F-18 vehicle via computer simulation. Stability and performance is compared to that of a system incorporating a controller designed by classical loop-shaping techniques.

Hess, R. A.↗
NASA NTRS2003

Training Data Requirement for a Neural Network to Predict Aerodynamic Coefficients

Basic aerodynamic coefficients are modeled as functions of angle of attack, speed brake deflection angle, Mach number, and side slip angle. Most of the aerodynamic parameters can be well-fitted using polynomial functions. We previously demonstrated that a neural network is a fast, reliable way of predicting aerodynamic coefficients. We encountered few under fitted and/or over fitted results during prediction. The training data for the neural network are derived from wind tunnel test measurements and numerical simulations. The basic questions that arise are: how many training data points are required to produce an efficient neural network prediction, and which type of transfer functions should be used between the input-hidden layer and hidden-output layer. In this paper, a comparative study of the efficiency of neural network prediction based on different transfer functions and training dataset sizes is presented. The results of the neural network prediction reflect the sensitivity of the architecture, transfer functions, and training dataset size.

Korsmeyer, David↗
NASA NTRS1993

Doppler Global Velocimetry Measurements of the Vortical Flow Above an F/A-18

A Doppler global velocimeter was used to investigate the vortical flow above an F/A-18 model at 25-degrees angle of attack. The measurements indicate that the flow had the same characteristics as the vortical flow above a standard delta wing. The flow pattern indicating transition from stable to burst conditions found above the delta wing was also found at the 440 station above the F/A-18. Measurements downstream at the 524 station found that the flow velocity varied considerably, with standard deviations reaching 30 percent of free stream. However, individual data images indicated that the flow was spatially coherent, and not chaotic as expected.

Lee, Joseph W.↗
NASA NTRS1989

A Patched-Grid Algorithm for Complex Configurations Directed Towards the F/A-18 Aircraft

A patched-grid algorithm for the analysis of complex configurations with an implicit, upwind-biased Navier-Stokes solver is presented. Results from both a spatial-flux and a time-flux conservation approach to patching across zonal boundaries are presented. A generalized coordinate transformation with a biquadratic geometric element is used at the zonal interface in order to treat highly stretched viscous grids and arbitrarily-shaped zonal boundaries. Applications are made to the F-18 forebody-strake configuration at subsonic, high-alpha conditions. Computed surface flow patterns compare well with ground-based and flight-test results; the large effect of Reynolds number on the forebody flow-field is shown.

High alpha research vehicle↗
NASA NTRS1953

An Exploratory Investigation of Some Types of Aeroelastic Instability of Open and Closed Bodies of Revolution Mounted on Slender Struts

Aeroelastic instability phenomena of isolated open and closed rigid bodies of revolution free to move under elastic restraint have been investigated experimentally at low speeds by means of models suspended at zero angles of attack and yaw on slender flexible struts from a wind tunnel ceiling. Three types of instability were observed - flutter similar to classical bending-torsion flutter, divergence, and an uncoupled oscillatory instability which consists in nonviolent continuous or intermittent small-amplitude oscillations involving only angular deformations. The speeds at which this oscillatory instability starts were found to be as low as about one-third of the speed at flutter or divergence and to depend on the shape of the body, particularly that of the afterbody, and on the relative location of the elastic axis. An attempt has been made to calculate the airspeeds and, in the case of the oscillatory phenomena, the frequencies at which these instabilities occur by using slender-body theory for the aerodynamic forces on the bodies and neglecting the aerodynamic forces on the struts. However, the agreement between the speeds and frequencies calculated in this manner and those actually observed has been found to be generally unsatisfactory; with the exception of the frequencies of the uncoupled oscillations which could be predicted with fair accuracy. The nature of the observed phenomena and of the forces on bodies of revolution suggests that a significant improvement in the accuracy of analytical predictions of these aeroelastic instabilities can be had only by taking into account the effects of boundary-layer separation on the aerodynamic forces.

Clevenson, S. A.↗
NASA NTRS2011

Orion Crew Module Aerodynamic Testing

The Apollo-derived Orion Crew Exploration Vehicle (CEV), part of NASA s now-cancelled Constellation Program, has become the reference design for the new Multi-Purpose Crew Vehicle (MPCV). The MPCV will serve as the exploration vehicle for all near-term human space missions. A strategic wind-tunnel test program has been executed at numerous facilities throughout the country to support several phases of aerodynamic database development for the Orion spacecraft. This paper presents a summary of the experimental static aerodynamic data collected to-date for the Orion Crew Module (CM) capsule. The test program described herein involved personnel and resources from NASA Langley Research Center, NASA Ames Research Center, NASA Johnson Space Flight Center, Arnold Engineering and Development Center, Lockheed Martin Space Sciences, and Orbital Sciences. Data has been compiled from eight different wind tunnel tests in the CEV Aerosciences Program. Comparisons are made as appropriate to highlight effects of angle of attack, Mach number, Reynolds number, and model support system effects.

Murphy, Kelly J.↗
NASA NTRS1980

Flow visualization study of the HiMAT RPRV

Water tunnel studies were performed to qualitatively define the flow field of the highly maneuverable aircraft technology remotely piloted research vehicle (HiMAT RPRV). Particular emphasis was placed on defining the vortex flows generated at high angles of attack. The flow visualization tests were conducted in the Northrop water tunnel using a 1/15 scale model of the HiMAT RPRV. Flow visualization photographs were obtained for angles of attack up to 40 deg and sideslip angles up to 5 deg. The HiMAT model was investigated in detail to determine the canard and wing vortex flow field development, vortex paths, and vortex breakdown characteristics as a function of angle of attack and sideslip. The presence of the canard caused the wing vortex to form further outboard and delayed the breakdown of the wing vortex to higher angles of attack. An increase in leading edge camber of the maneuver configuration delayed both the formation and the breakdown of the wing and canard vortices. Additional tests showed that the canard vortex was sensitive to variations in inlet mass flow ratio and canard flap deflection angle.

Lorincz, D. J.↗
NASA NTRS1954

Flight Tests of a 0.13-Scale Model of the Convair XFY-1 Vertically Rising Airplane with the Lower Vertical Tail Removed, TED No.DE 368

An experimental investigation has been conducted to determine the dynamic stability and control characteristics in hovering and transition flight of a 0.13-scale flying model of the Convair XFY-1 vertically rising airplane with the lower vertical tail removed. The purpose of the tests was to obtain a general indication of the behavior of a vertically rising airplane of the same general type as the XFY-1 but without a lower vertical tail in order to simplify power-off belly landings in an emergency. The model was flown satisfactorily in hovering flight and in the transition from hovering to normal unstalled forward flight (angle of attack approximately 30deg). From an angle of attack of about 30 down to the lowest angle of attack covered in the flight tests (approximately 15deg) the model became progressively more difficult to control. These control difficulties were attributed partly to a lightly damped Dutch roll oscillation and partly to the fact that the control deflections required for hovering and transition flight were too great for smooth flight at high speeds. In the low-angle-of-attack range not covered in the flight tests, force tests have indicated very low static directional stability which would probably result in poor flight characteristics. It appears, therefore, that the attainment of satisfactory directional stability, at angles of attack less than 10deg, rather than in the hovering and transition ranges of flight is the critical factor in the design of the vertical tail for such a configuration.

Lovell, Powell M., Jr.↗
NASA NTRS1984

Airfoil trailing edge flow measurements and comparison with theory, incorporating open wind tunnel corrections

Trailing edge data for boundary layer-near wake thickness parameters are given for airfoils and flat plates. Reynolds number effects are examined as a function of model size, velocity and boundary layer tripping. These data expand that presented previously by the authors particularly for airfoil non-zero angles of attack. Comparisons are made here with boundary layer calculations using potential flow modeling and a well documented two-dimensional finite-difference method for laminar and turbulent boundary layers. Open wind tunnel corrections to angle of attack and camber are developed and are incorporated in the potential flow modeling to assure correct comparisons for non-zero angles of attack. It was found that although the open tunnel flow turbulence affected boundary layer transition for the higher velocities the theory successfully 'brackets' the data. Comparisons demonstrate the degree of accuracy one might expect for the prediction of boundary layer thickness parameters when given only geometry and nominal flow conditions as input to boundary layer codes.

Brooks, T. F.↗
NASA NTRS1986

Subsonic aerodynamic characteristics of a circular body earth-to-orbit transport

To reduce the weight and improve the performance of future earth-to-orbit transports, the use of circular cross sections in the fuselage bodies of these vehicles is being considered at the Langley Research Center. Structurally, circular cross sections are stronger and lighter than other shapes. A study has been made applying the circular body concept to a vertical-takeoff, delta-winged, single-stage-to-orbit transport. A 52 in., 0.022-scale model of the circular body vehicle was tested at a Mach number of 0.3 in the 7 x 10 ft High Speed Wind Tunnel at the Langley Research Center to obtain aerodynamic forces and moments. Oil-flow photographs were taken at several angles of attack to aid in the aerodynamic analysis. Model control surfaces included elevons and ailerons for the evaluation of pitch and roll characteristics and either wing-tip fins, a nose mounted dorsal fin, or a conventional vertical tail for the evaluation of yaw characteristics. Other deflecting surfaces included speedbrakes and body flaps. Basic data on longitudinal flight characteristics are shown, including lift, drag, and pitching moments. Comparisons of the directional stability and control effectiveness of the three directional control devices are also shown.

Lepsch, R. A., Jr.↗
NASA NTRS1986

Vortex cloud model for body vortex shedding and tracking

The discrete vortex cloud approach models a missile airframe's vortex wake by means that are capable of treating a variety of configurations over a range of flow conditions. Attention is given to the sheets of vorticity formed on the lee side of a missile at moderate angles-of-attack. While three-dimensional attached flow models are used to represent the missile body, two-dimensional, incompressible, separated flow models are used to represent the separated vortex wake. The predicted pressure distribution of the body under the influence of the freestream and the separation vortex wake are used to calculate aerodynamic loads on the body. The separation vortex wake is represented by clouds of discrete vortices in cross flow planes normal to the body axis.

Mendenhall, Michael R.↗
DOE OSTI2023

Advanced Reactor Cyber Analysis and Development Environment (ARCADE) for System-Level Design Analysis

Cybersecurity is a persistent concern to the safety and security of Nuclear Power Plants (NPPs), but has lacked data-driven, evidence-based research. Rigorous cybersecurity analysis is critical for the licensing of advanced reactors using a performance-based approach. One tool that enables cybersecurity analysis is modeling and simulation. The nuclear industry makes extensive use of modeling and simulation throughout the decision process but lacks a method to incorporate cybersecurity analysis with existing models. To meet this need, the Advanced Reactor Cyber Analysis and Development Environment (ARCADE) was developed. ARCADE is a suite of publicly available tools that can be used to develop emulations of industrial control system devices and networks and integrate those emulations with physics simulators. This integration of cyber emulations and physics models enables rigorous cyber-physical analysis of cyber-attacks on NPP systems. This report provides an overview of key considerations for using ARCADE with existing physics models and demonstrates ARCADE’s capabilities for cybersecurity analysis. Using a model of the Small Modular Advanced High Temperature Reactor (SmAHTR), ARCADE was able to determine the sensitivity of the primary heat exchangers (PHX) to coordinated cyber-attacks. The analysis determined that while the PHX’s failures cause disruption to the reactor, they did not cause any safety limits to be exceeded because of the plant design, including passive safety features. Further development of ARCADE will enable rigorous, repeatable, and automated cyber-physical analysis of advanced reactor control systems. These efforts will also help reduce regulatory uncertainty by presenting similar types of cybersecurity analyses in a common format, driving standard approaches and reporting.

22 GENERAL STUDIES OF NUCLEAR REACTORS↗
NASA NTRS1972

Heat transfer and pressure distributions at M equals 8 on 0.029 scale models of the Viking entry vehicle

An investigation in the Langley Research Center Mach-8 Variable Density Hypersonic Tunnel was made of the pressure distributions and heat transfer rate distributions on two 0.029 scale Viking Entry Vehicle models. Comparable ranges of test Reynolds number were exercised for the two tests between run conditions around 4 million and conditions of about 1.6 million. At angles of attack less than 20 degrees the pressure ratio distribution referenced to stagnation pressure appeared invariant with Reynolds number. Increasing angle of attack results in a flatter distribution of both the windward and leeward pressure distributions; in addition, the stagnation point shifted into the windward plane. A subsequent rise in the heating rate profile on the leeward side with further increase in angle of attack is attributed to boundary layer natural transition to turbulent flow. Schlieren photographs were taken for flow field visualization and to correct model angle of attack.

Faye-Petersen, R.↗
NASA NTRS1999

X-33 Hypersonic Boundary Layer Transition

Boundary layer and aeroheating characteristics of several X-33 configurations have been experimentally examined in the Langley 20-Inch Mach 6 Air Tunnel. Global surface heat transfer distributions, surface streamline patterns, and shock shapes were measured on 0.013-scale models at Mach 6 in air. Parametric variations include angles-of-attack of 20-deg, 30-deg, and 40-deg; Reynolds numbers based on model length of 0.9 to 6.6 million; and body-flap deflections of 0, 10 and 20-deg. The effects of discrete and distributed roughness elements on boundary layer transition, which included trip height, size, location, and distribution, both on and off the windward centerline, were investigated. The discrete roughness results on centerline were used to provide a transition correlation for the X-33 flight vehicle that was applicable across the range of reentry angles of attack. The attachment line discrete roughness results were shown to be consistent with the centerline results, as no increased sensitivity to roughness along the attachment line was identified. The effect of bowed panels was qualitatively shown to be less effective than the discrete trips; however, the distributed nature of the bowed panels affected a larger percent of the aft-body windward surface than a single discrete trip.

Berry, Scott A.↗
NASA NTRS1990

Effect of tail size reductions on longitudinal aerodynamic characteristics of a three surface F-15 model with nonaxisymmetric nozzles

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of horizontal and vertical tail size reductions on the longitudinal aerodynamic characteristics of a modified F-15 model with canards and 2-D convergent-divergent nozzles. Quantifying the drag decrease at low angles of attack produced by tail size reductions was the primary focus. The model was tested at Mach numbers of 0.40, 0.90, and 1.20 over an angle of attack of -2 degree to 10 degree. The nozzle exhaust flow was simulated using high pressure air at nozzle pressure ratios varying from 1.0 (jet off) to 7.5. Data were obtained on the baseline configuration with and without tails as well as with reduced horizontal and/or vertical tail sizes that were 75, 50, and 25 percent of the baseline tail areas.

Frassinelli, Mark C.↗