SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS

A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 44, ISSUE 13 - JULY 5, 2006

02 AERODYNAMICS
Includes aerodynamics of flight vehicles, test bodies, airframe components and combinations, wings, and control surfaces. Also includes aerodynamics of rotors, stators, fans, and other elements of turbomachinery. For related information see also 34 Fluid Mechanics and Thermodynamics.

20060017022 NASA Langley Research Center, Hampton, VA, USA

Apparent Transition Behavior of Widely-Used Turbulence Models

Rumsey, Christopher L.; [2006]; 19 pp.; In English; 36th AIAA Fluid Dynamics Conference and Exhibit, 5-8 Jun. 2006, San Francisco, CA, USA; Original contains color illustrations Contract(s)/Grant(s): 581.02.08 Report No.(s): AIAA Paper 2006-3906; No Copyright; Avail.: CASI: A03, Hardcopy

The Spalart-Allmaras and the Menter SST kappa-omega turbulence models are shown to have the undesirable characteristic that, for fully turbulent computations, a transition region can occur whose extent varies with grid density. Extremely fine two-dimensional grids over the front portion of an airfoil are used to demonstrate the effect. As the grid density is increased, the laminar region near the nose becomes larger. In the Spalart-Allmaras model this behavior is due to convergence to a laminar-behavior fixed point that occurs in practice when freestream turbulence is below some threshold. It is the result of a feature purposefully added to the original model in conjunction with a special trip function. This degenerate fixed point can also cause nonuniqueness regarding where transition initiates on a given grid. Consistent fully turbulent results can easily be achieved by either using a higher freestream turbulence level or by making a simple change to one of the model constants. Two-equation kappa-omega models, including the SST model, exhibit strong sensitivity to numerical resolution near the area where turbulence initiates. Thus, inconsistent apparent transition behavior with grid refinement in this case does not appear to stem from the presence of a degenerate fixed point. Rather, it is a fundamental property of the kappa-omega model itself, and is not easily remedied. Author

Turbulence Models; Convergence; Airfoils

20060017108 Army Missile Command, Redstone Arsenal, AL USA MissileLab User's Guide

Auman, Lamar M; Feb 2006; 84 pp.; In English; Original contains color illustrations Report No.(s): AD-A444198; RDECOM-TR-AMR-SS-06-12; No Copyright; Avail.: CASI: A05, Hardcopy

A tool has been developed to assist the applied aerodynamicist in the generation of the aerodynamic characteristics of missile configurations. MissileLab provides the ability to enter a single set of vehicle geometry and flight conditions, and then execute as many as six engineering-level aerodynamic prediction engines (APE). The user must have licensed copies of the various aero-prediction engines, as MissileLab simply provides a common geometry front-end to the respective codes. Currently, MissileLab supports input to the Air Force Missile DATCOM, Naval Surface Weapon Center (NSWC) AP98, Aero-Prediction AP02, AP05, and Nielsen Engineering and Research (NEAR) MISL3-March 2004. The features and capabilities of MissileLab are presented. DTIC

Aerodynamic Characteristics; Manuals; Missiles

20060017491 Rice Univ., Houston, TX USA

Optimal Aeroacoustic Shape Design Using the Surrogate Management Framework

Marsden, Alison L; Wang, Meng; Dennis Jr, John E; Moin, Parviz; Feb 9, 2004; 25 pp.; In English; Original contains color illustrations Contract(s)/Grant(s): N00014-01-1-0423 Report No.(s): AD-A444758; No Copyright; Avail.: CASI: A03, Hardcopy

Shape optimization is applied to time-dependent trailing-edge flow in order to minimize aerodynamic noise. Optimization is performed using the surrogate management framework (SMF), a non-gradient based pattern search method chosen for its efficiency and rigorous convergence properties. Using SMF, design space exploration is performed not with the expensive actual function but with an inexpensive surrogate function. The use of a polling step in the SMF guarantees that the lgorithm generates a convergent subsequence of mesh points, each iterate of which is a local minimizer of the cost function on a mesh in the parameter space. Results are presented for an unsteady laminar flow past an acoustically compact airfoil. Constraints on lift and drag are handled within SMF by applying the filter pattern search method of Audet and Dennis, within which a penalty function is used to form and optimize a surrogate function. Optimal shapes that minimize noise have been identified for the trailing-edge problem in constrained and unconstrained cases. Results show a significant reduction (as much as 80%) in acoustic power with reasonable computational cost using several shape parameters. Physical mechanisms for noise reduction are discussed. DTIC

Acoustics; Aeroacoustics; Aerodynamics; Aircraft Design; Shapes; Time Dependence

20060017516 Army Research Lab., Aberdeen Proving Ground, MD USA

Human Factors Assessment of the UH-60M Crew Station During the Limited User Test (LUT)

Havir, Thomas J; Durbin, David B; Frederick, Lorraine J; Hicks, Jamison S; Feb 2006; 100 pp.; In English Contract(s)/Grant(s): Proj-62716AH70 Report No.(s): AD-A444913; ARL-TR-3730; No Copyright; Avail.: CASI: A05, Hardcopy

The utility helicopter (UH)-60M Product Manager requested the U.S. Army Research Laboratory's Human Research and Engineering Directorate to participate in the Limited User Test for the UH-60M Black Hawk. ARL conducted a human factors evaluation during the LUT, which assessed workload, situation awareness, simulator sickness, pilot-vehicle interface, and eye tracker data. The data were used to identify characteristics of the UH-60M that enhance or degrade pilot performance. Characteristics that degrade pilot performance were included in the Manpower and Personnel Integration (MANPRINT) assessment for the system s milestone decision and should be considered for future design changes at the earliest opportunity. Three UH-60 crews (six pilots) each conducted six mission scenarios for a total of 18 flights. The conditions of each mission were systematically varied and designed to become progressively more difficult as the pilots became more proficient at flying the aircraft. The pilots completed the simulator sickness questionnaire before and after each flight. They completed the Bedford Workload Rating Scale, Situation Awareness Rating Technique, and the Pilot-Vehicle Interface Questionnaire after each mission. In addition to pilot data, a tactical steering committee (TSC) performed an independent assessment of workload, situation awareness, and mission success. The TSC completed a survey after each mission. The data were analyzed with the use of the Wilcoxon Signed Ranks Test to compare pilot ratings between seat position and results between UH-60M and UH-60A/L model aircraft. The mean workload rating for all tasks for the UH-60M was 2.71, indicating that the pilots typically had enough workload capacity for all desirable additional tasks. The mean situation awareness rating provided by the pilots was 28.25. This SA rating indicates that the pilots felt they had high levels of situation awareness during the missions. DTIC

Crew Workstations; Flight Tests; Helicopters; Human Factors Engineering

20060017645 Texas Univ., Arlington, TX USA

Flight Control and Simulation for Aerial Refueling

Dogan, Atilla; Sato, Shinya; Aug 18, 2005; 16 pp.; In English; Original contains color illustrations Report No.(s): AD-A445129; AIAA-2005-6264; No Copyright; Avail.: CASI: A03, Hardcopy

This paper addresses the problem of controlling the receiver aircraft to achieve a successful aerial refueling. For the performance verification of the controller, a new set of nonlinear, 6-DOF, rigid body equations of motion for the receiver aircraft has been derived. The equations are developed using the reference frame as one that is attached to, translate and rotates with the tanker aircraft. Furthermore, the nonlinear equations contain the wind effect terms and their time derivatives to represent the aerodynamic coupling involved between the two aircraft. These wind terms are obtained using an averaging technique that computes the effective induced wind components and wind gradients in the receiver aircraft's body frame. Dynamics of the engine and the actuators are also included in the study.Alinear position-tracking controller has been designed using a combination of integral control and optimal LQR design. The controller does not use the information of the tanker's vortex induced wind effects acting on the receiver aircraft as well as the mass change that occurs during refueling. The performance of the controller is evaluated in the high fidelity simulation environment employing the new sets of equations of motion. The simulation and control design are applied to a tailless fighter aircraft with innovative control effectors and thrust vectoring capability. Various allocation schemes for redundant control variables are analyzed in a realistic approach maneuver to the refueling contact position behind the tanker aircraft. In this paper, the performance evaluation is presented only during the initial phase of aerial refueling maneuver when the receiver aircraft maneuvers to reach the refueling contact position. DTIC

Air to Air Refueling; Control Simulation; Flight Control; Refueling

20060018461 Air Force Inst. of Tech., Wright-Patterson AFB, OH USA

Dynamic Aeroelastic Analysis of Wing/Store Configurations

Parker, Gregory H; Dec 2005; 124 pp.; In English; Original contains color illustrations Report No.(s): AD-A445218; AFIT/DS/ENY/06-06; No Copyright; ONLINE: http://hdl.handle.net/100.2/ADA445218; Avail.: CASI: A06, Hardcopy

Limit-cycle oscillation, or LCO, is an aeroelastic phenomenon characterized by limited amplitude, self-sustaining oscillations produced by fluid-structure interactions. In order to study this phenomenon, code was developed to interface a modal structural model with a commercial computational fluid dynamics program. LCO was simulated for a rectangular wing, referred to as the Goland+ wing. It was determined that the aerodynamic nonlinearity responsible for LCO in the Goland+ wing was the combination of strong trailing-edge and lambda shocks which periodically appear and disappear. This mechanism limited the flow of energy into the structure which quenched the growth of the flutter, resulting in a steady CO. Under-wing and tip stores were added to the Goland+ wing to determine how stores affected limit-cycle oscillation. It was found that aerodynamic store shapes affect LCO in two off-setting ways: under-wing stores interfere with the airflow on the lower surface of the wing which decreases LCO amplitudes, whereas, aerodynamic forces on both under-wing and tip stores directly increase LCO amplitudes. DTIC

Aeroelasticity; Dynamic Structural Analysis; External Stores; Wing Tanks; Wings

20060018858 Glasgow Univ., UK

Fast Prediction of Wing Rock Onset Based on Computational Fluid Dynamics

Badcock, Ken; Sep 14, 2005; 40 pp.; In English; Original contains color illustrations Contract(s)/Grant(s): FA8655-03-1-3044 Report No.(s): AD-A445823; No Copyright; Avail.: Defense Technical Information Center (DTIC)

This report results from a contract tasking University of Glasgow as follows: The objective is to develop efficient CFD based techniques for the prediction of wing rock instabilities for a delta wing-body. Two approaches in addition to the basic time marching approach will be investigated. First direct calculation of the Hopf point for the onset of the instability (for a fixed sweep angle the incidence is regarded as the bifurcation parameter) will be developed based on techniques previously developed for aeroelasticity. Secondly reduced order models will be developed driven by CFD results to again predict the instability onset efficiently. Finally an assessment of the different methods will be carried out to provide guidance for future application in certification by analysis and design. The derived solutions will also be examined to answer some of the current questions as to the physical mechanisms for wing rock. DTIC

Aerodynamic Characteristics; Computational Fluid Dynamics; Flight Control; Wing Rock

20060018911 Indian Inst. of Tech., Madras, India

Studies on Impingement Effects of Low Density Jets on Surfaces - Determination of Shear Stress and Normal Pressure

Sathian, Sarith P; Kurian, Job; Jul 13, 2005; 7 pp.; In English; Original contains color illustrations Report No.(s): AD-A445973; No Copyright; Avail.: Defense Technical Information Center (DTIC)

This paper presents the results of the Laser Reflection Method (LRM) for the determination of shear stress due to impingement of low-density free jets on flat plate. For thin oil film moving under the action of aerodynamic boundary layer the shear stress at the air-oil interface is equal to the shear stress between the surface and air. A direct and dynamic measurement of the oil film slope is measured using a position sensing detector (PSD). The thinning rate of oil film is directly measured which is the major advantage of the LRM over LISF method. From the oil film slope history, direct calculation of the shear stress is done using a three-point formula. For the full range of experiment conditions Knudsen numbers varied till the continuum limit of the transition regime. The shear stress values for low-density flows in the transition regime are thus obtained using LRM and the measured values of shear show fair agreement with those obtained by other methods. Results of the normal pressure measurements on a flat plate in low-density jets by using thermistors as pressure sensors are also presented in the paper. The normal pressure profiles obtained show the characteristic features of Newtonian impact theory for hypersonic flows. DTIC

Free Jets; Impingement; Jet Flow; Shear Stress; Stress Analysis

20060019110 Ohio State Univ., Columbus, OH USA

Reference Command Tracking for a Linearized Model of an Air-Breathing Hypersonic Vehicle

Groves, Kevin P; Sigthorsson, David O; Serrani, Andrea; Yurkovich, Stephen; Bolender, Michael A; Doman, David B; Aug 2005; 15 pp.; In English; Original contains color illustrations Report No.(s): AD-A445387; No Copyright; ONLINE: http://hdl.handle.net/100.2/ADA445387; Avail.: CASI: A03, Hardcopy

The focus of this paper is on control design and simulation for an air-breathing hypersonic vehicle. The challenges for control design in this class of vehicles lie in the inherent coupling between the propulsion system, and the airframe dynamics, and the presence of strong flexibility effects.Working from a highly nonlinear, dynamically-coupled simulation model, control designs are presented for velocity, angle-of-attack, and altitude command input tracking for a linearized version of a generic air-breathing hypersonic vehicle model linearized about a specific trim condition. Control inputs for this study include elevator deflection, total temperature change across the combustor, and the diffuser area ratio. Two control design methods are presented, both using linear quadratic techniques with integral augmentation, and are implemented in tracking control studies. The first approach focuses on set point tracking control, whereas in the second, a regulator design approach is taken. The effectiveness of each control design is demonstrated in simulation on the full nonlinear model of the generic vehicle. DTIC

Hypersonic Vehicles; Angle of Attack; Airframes; Control Simulation

20060019117 NASA Ames Research Center, Moffett Field, CA, USA

Adjoint Method and Predictive Control for 1-D Flow in NASA Ames 11-Foot Transonic Wind Tunnel

Nguyen, Nhan; Ardema, Mark; [2006]; 16 pp.; In English; AIAA Aerospace Sciences Meeting and Exhibit, 9-12 Jan. 2006, Reno, NV, USA; Original contains black and white illustrations; Copyright; Avail.: CASI: A03, Hardcopy

This paper describes a modeling method and a new optimal control approach to investigate a Mach number control problem for the NASA Ames 11-Foot Transonic Wind Tunnel. The flow in the wind tunnel is modeled by the 1-D unsteady Euler equations whose boundary conditions prescribe a controlling action by a compressor. The boundary control inputs to the compressor are in turn controlled by a drive motor system and an inlet guide vane system whose dynamics are modeled by ordinary differential equations. The resulting Euler equations are thus coupled to the ordinary differential equations via the boundary conditions. Optimality conditions are established by an adjoint method and are used to develop a model predictive linear-quadratic optimal control for regulating the Mach number due to a test model disturbance during a continuous pitch Author

Euler Equations of Motion; Mach Number; Transonic Wind Tunnels; Unsteady Flow; Boundary Conditions; Optimal Control

Source: NASA