Showing papers in "Journal of Aircraft in 2003"

Probabilistic Two-Phase Wake Vortex Decay and Transport Model

Abstract: A new parametric wake vortex transport and decay model is proposed that predicts probabilistic wake vortex behavior as a function of aircraft and environmental parameters in real time. The probabilistic two-phase wake vortex decay model (P2P) accounts for the effects of wind, turbulence, stable stratie cation, and ground proximity. The model equations are derived from the analytical solution of the spatiotemporal circulation evolution of the decaying potential vortex and are adapted to wake vortex behavior as observed in large-eddy simulations. Vortex decay progresses in two phases, a diffusion phase followed by rapid decay. Vortex descent is a nonlinear function of vortex strength. Probabilistic components account for deviations from deterministic vortex behavior inherently caused by the stochastic nature of turbulence, vortex instabilities, and deformations, as well as uncertainties and e uctuations that arise from environmental and aircraft parameters. The output of P2P consists of cone dence intervals for vortex position and strength. To assign a dee ned degree of probability to the predictions reliably, the model design allows for the continuous adjustment of decay parameters and uncertainty allowances, based on a growing amount of data. The application of a deterministic version of P2P to the Memphis wake vortex database yields favorable agreement with measurements.

Future of Airplane Aeroelasticity

Abstract: Aeroelasticity is still dynamic, challenging, and a key part of cutting-edge airplane technology. Emerging trends, as well as challenges and needs in the field of airplane aeroelasticity, are surveyed and discussed. The paper complements other overview papers on various aspects of the fixed-wing aeroelastic problem, published recently for the centennial year of flight. It includes an extensive bibliography and emphasizes those aspects of aeroelastic technology development not covered thoroughly elsewhere.

Data Summary from Second AIAA Computational Fluid Dynamics Drag Prediction Workshop

Abstract: Results from the Second AIAA Drag PredictionWorkshop are summarized. The workshop focused on absolute and configuration delta drag prediction of the DLR, German Aerospace Research Center F6 geometry, which is representative of transport aircraft designed for transonic flight. Both wing–body and wing–body–nacelle–pylon configurations are considered. Comparisons are made using industry relevant test cases that include single-point conditions, drag polars, and drag-rise curves. Drag, lift, and pitching moment predictions from several different Reynolds averagedNavier–Stokes computational fluid dynamics codes are presented and compared to experimental data. Solutions on multiblock structured, unstructured, and overset structured grids using a variety of turbulence models are considered. Results of a grid-refinement study and a comparison of tripped transition vs fully turbulent boundary-layer computations are reported.

FENSAP-ICE's Three-Dimensional In-Flight Ice Accretion Module: ICE3D

Abstract: Two-dimensional and quasi-three-dimensional in-flight ice accretion simulation codes have been increasingly used by the aerospace industry in the last two decades as an aid to the certification process. Such codes predict two-dimensional sectional ice shapes, which are then manufactured from a light material and attached as disposable profiles on test aircraft to investigate them for stability under icing encounters. Although efficient for calculating ice shapes on simple geometries, current codes encounter major difficulties or simply cannot simulate ice shapes on truly three-dimensional geometries such as nonaxisymetric nacelles, high-lift wings, engine intakes, or systems that combine external and internal flows. Modern computational fluid dynamics approaches may not encounter or engender these difficulties, and FENSAP-ICE is a combination of four modules forming a complete and generic in-flight icing simulation system, built in a way to solve successively impingement, ice accretion, heat loads, and performance degradation. The set of equations describing FENSAP-ICE's airflow solver, FENSAP, its impingement module, DROP3D, and its accretion module, ICE3D, are presented

Aerodynamic and Aeroelastic Characteristics of Wings with Conformal Control Surfaces for Morphing Aircraft

Abstract: Investigations are conducted on lifting surfaces with conventional and conformal trailing-edge control surfaces. The Sohngen inversion formula is used with the thin-airfoil integral equation to determine the aerodynamic pressure for various control surface chord-to-airfoil chord ratios. Comparisons to a conventional control surface show increases in lift and pitching moment of the airfoil with a conformal control surface. Aerodynamic pressure distributions acting on a wing with control surfaces are determined with the vortex lattice technique. Predicted aerodynamic pressures and roll moments are compared to available wind-tunnel data and provide a more general understanding of theaerodynamicbehavior observed there. Roll performance of a rectangular wing is determined for various control surface chord-to-wing chord ratios. It is found that the maximum roll rate is greater for a wing with a conformal control surface, but has a lower reversal dynamic pressure than the wing with a conventional control surface. The aerodynamic and aeroelastic results obtained from this investigation provide some insight for wings designed with conformal control surfaces.

Wing Optimization Using Design of Experiment, Response Surface, and Data Fusion Methods

Abstract: An empirical drag prediction model plus design of experiment, response surface, and data-fusionmethods are brought together with computational fluid dynamics (CFD) to provide a wing optimization system. This system allows high-quality designs to be found using a full three-dimensional CFD code without the expense of direct searches. The metamodels built are shown to be more accurate than the initial empirical model or than simple response surfaces based on the CFD data alone. Data fusion is achieved by building a response surface kriging of the differences between the two drag prediction tools, which are working at varying levels of fidelity. The kriging is then used with the empirical tool to predict the drags coming from the CFD code. This process is much quicker to use than direct searches of the CFD.

Aeroelasticity of Nonconventional Airplane Configurations-Past and Future

Abstract: At the end of the e rst century of manned, powered e ight, it is worthwhile to look backward to understand how innovation in airplane design required developments in aeroelasticity and how aeroelasticity has played a role in shaping the e rst 100 years of aircraft design. The insights gained will help to predict how and where aeroelasticity and aeroservoelasticity will ine uence the future development of efe cient, more capable, innovative air vehicles, and dee nethe needs fortechnology and tools to enable this future. By dee nition, all new aircraft begin as unconventional to a certain extent. Designs that never see universal use remain curiosities, but still help our questforbettervehiclesandguidethedevelopmentofanalysis, design, and testing tools. Innovative, nontraditional designsaffected by aeroelasticconsiderations haveincluded obliquewing aircraft, forward-swept wing aircraft,Xwings,e ying wings, andlargejoined wings. Designsthat wereunusually innovativeat thetimeoftheirintroduction but later became widespread include the swept-back wing jet, the T-tail, and the e y-by-wire control cone gured vehicle. Control and exploitation of aeroelasticity depends on the continued development of new materials, new structuralandaerodynamicconcepts,sensors,actuators,andactivecontroltechniques.Suchdevelopmentsmustbe accompanied by properintegratedanalysis/designtools,and,most importantly, by thesamehumaninquisitiveness and creativity that has driven aircraft design for over a century. This paper uses the history of nonconventional airplane cone gurations to review some of the steps taken during the past century to establish aeroelastic effects as integrated design features that must be anticipated, controlled, and exploited. The paper goes on to discuss the potential impact of past lessons on emerging airplane cone gurations currently in various stages of study and development.

Computational Aeroelasticity: Success, Progress, Challenge

Abstract: This article asserts a much broader dee nition of the term, one in whichCAEencompassesalllevelsofaeroelasticanalysis.Aeroelastic tools based on both linear unsteady aerodynamics and nonlinear CFD methods have been developed and successfully applied. We refer to both of these methodologies as components of CAE. Likewisestructuralmodelingassimpleasbeamtheorytostate-of-the-art e nite element modeling (FEM) have been incorporated into aeroelastic tools, and these techniques should also be included under the CAE heading. It is not the intention of this article to provide an exhaustive history of the development and application of CAE over the past 70 years. Rather, the subject will be examined in the context of two primary themes: 1) aeroelastic problems requiring theoretical

Vibration Control of a Landing Gear System Featuring Electrorheological/Magnetorheological Fluids

Abstract: The feasibility and effectiveness of electrorheological (ER) and magnetorheological (MR) fluid-based landing gear systems on attenuating dynamic load and vibration due to the landing impact are demonstrated. First, the theoretical model for ER/MR shock struts, which are the main components of the landing gear system,is developed based on experimental data. The analysis of a telescopic-type landing gear system using the ER/MR shock struts is theoretically constructed, and its governing equation is derived. A sliding mode controller, designed to be robust against parameter variations and external disturbances, is formulated, and controlled performance of the simulated ER/MR landing gear system is theoretically evaluated during touchdown of the aircraft.

Simultaneous Optimization of a Multiple-Aircraft Family

Abstract: Multidisciplinary design optimization is considered in the context of designing a family of aircraft Aframework is developed in which multiple aircraft, each with different missions but sharing common parts, can be optimized simultaneouslyThenewframeworkisusedtogaininsighttotheeffectofdesignvariablescalingontheoptimization algorithm Results are presented for a two-member family whose individual missions differ signie cantly Both missions can be satise ed with common designs Moreover, optimizing both airplanes simultaneously rather than following the traditional baseline plus derivative approach vastly improves the common solution A cost modeling framework is outlined that allows the value of commonality to be quantie ed for design and manufacturing costs A notional example is presented to show the cost benee t that may be achieved by designing a common family of aircraft

The design of winglets for high-performance sailplanes

Abstract: Although theoretical tools for the design of winglets for high-performance sailplanes were initially of limited value, simple methods were used to design winglets that gradually became accepted as benefiting overall sailplane performance. To further these gains, an improved methodology for winglet design has been developed. This methodology incorporates a detailed component drag buildup that includes the ability to interpolate input airfoil drag and moment data across operational lift coefficient, Reynolds number, and flapsetting ranges. Induced drag is initially predicted using a relatively fast multi- lifting line method. In the final stages of the design process, a full panel method, including relaxed-wake modeling, is employed. The drag predictions are used to compute speed polars for both level and turning flight. The predicted performance is in good agreement with flight-test results. The straight and turning flight speed polars are then used to obtain cross-country performance over a range of thermal strengths, sizes, and shapes. Example design cases presented here demonstrate that winglets can provide a small, but important, performance advantage over much of the operating range for both span limited and span unlimited high-performance sailplanes.

Status of Unsteady Aerodynamic Prediction for Flutter of High-Performance Aircraft

Abstract: Current production unsteady aerodynamics codes are based on panel methods. The capabilities present in these methods to model complex shapes represents one of the major advancements in this methodology. A second advancement is that the new codes are now formulated based on a unie ed analytical capability throughout the Mach range. Examples of these capabilities will be demonstrated. Finally, based on the discussion of requirements for e utter analyses of high-performance aircraft, a description of the needs for the future is presented.Itishoped thatthese recommendations willprovide guidance to those investigators working in the e eldof unsteady aerodynamics. II. Background A. Unsteady Aerodynamic Methods Strip theory aerodynamics originated in the early 1940s, 1;2 and this method was the primary aerodynamic tool for e utter analyses for many years.In 1966, Yates 3 proposed modifying CL® toaccount for e nite span effects with the result being referred to as modie ed strip theory. This aerodynamic method, coupled with the normal modeapproach,and the V‐g‐! solution technique formedthe basis for production e utter analyses in the late 1960s, and it was to this methodology that the author was e rst exposed to production e utter analyses.

Natural-Laminar-Flow Airfoil Development for a Lightweight Business Jet

Abstract: A 15% % thick, natural-laminar-e ow airfoil, the SHM-1, has been designed to satisfy requirements derived from the performance specie cations for a lightweight business jet. The airfoil was tested in a low-speed wind tunnel to evaluateitslow-speedperformance.Ae ighttestwasalsoconductedtoevaluatetheperformanceoftheairfoilathigh Reynolds numbers and high Mach numbers. In addition, a transonic wind-tunnel test was conducted to determine the drag-divergence characteristics. The design requirements, methodology, and experimental verie cation are described.

Influences of Mounting Angles and Locations on the Effects of Gurney Flaps

Abstract: Experimental investigations were conducted on a NACA0012 airfoil to determine the influences of mounting angles and mounting locations on the lift-enhancing effects of Gurney flaps at a Reynolds number of 2.1 x 10 6 . The results revealed that all flaps of different mounting angles increased the lift coefficient, and an increment of maximum lift coefficient of 12.3, 15.1, and 17.4% was obtained by 45-, 60-, and 90-deg Gurney flap, respectively. There was a drag penalty associated with the lift enhancement. The best performance was obtained by the 45-deg Gurney flap for all flap deflections tested. When shifted forward from the trailing edge of the airfoil, the Gurney flap led to a decrease in lift, and an increment in drag, and thus a reduction in lift-to-drag ratio

Component-based empirical model for high-lift system noise prediction

Abstract: This paper presents an empirical model for predicting noise from high lift systems, derived from a large database of airframe noise tests, involving various airplane models at various operating conditions The model correlates noise not only to gross airplane parameters such as the dimensions of the high lift system and flight Mach number, but also to flow quantities that are physically responsible for the noise generation Noise data used in the development of the model were acquired by using phased microphone arrays, which enables the decomposition of the total noise into components, relating the noise to the six individual components of the wing/high lift system The methodology and results of this component-based model is presented, including source identification by source strength maps, component integration to derive far field spectra, validation/calibration of the integrated spectra by conventional free field microphone data, extrapolation of small-scale model test data to full-scale conditions with Reynolds number dependent scaling laws and the correlation between noise and flow quantities Validations of the predictions with flight test data are also given to show the accuracy of the developed prediction tool

Optimization of a wing structure for gust response and aileron effectiveness

Abstract: Reliability-based weight optimization of a generic, e ghter-like wing structure is conducted for gust response and aileron effectiveness constraints. The formulation accounts for parametric uncertainties in these aeroelastic response quantities. Reliability indices measure the probability of satisfying each constraint, and a preliminary design procedure is developed in which constraints are enforced on these indices. This framework integrates ASTROS for structural and loads analysis, object-oriented MATLAB ® tools for reliability analysis, and DOT for optimization and most probable point estimation. The reliability analysis algorithm takes advantage of adaptive nonlinear approximations to compensate for nonlinearity of the failure surfaces. The wing structure is modeled with e nite elements, each of which is assumed to have random thickness of known standard deviation. Young’ s modulus of the wing skin material is also assumed to be random. Mean thickness values are taken as design variables. Linear unsteady aerodynamics is used to estimate frequency response functions caused by continuous gust loads. Reliability index constraints are enforced for gust-induced bending moment and shear at the wing’ s root, and also for aileron effectiveness. Redistribution of structural mass by the optimizer produces designs with improved aeroelastic performance reliability and relatively small weight penalties.

Effect of Leading-Edge Curvature on Airfoil Separation Control

Abstract: Separation control on NACA 0012 and NACA 0015 airfoils was compared under incompressible conditions, using leading-edgeperiodicexcitation, in orderto assess the effect of leading-edgecurvature. Both lift and moment coefe cients were considered to compare and analyse control effectiveness. In contrast to the relatively mild NACA 0015 trailing-edge stall, NACA 0012 stall was dominated by a leading-edge bubble-bursting mechanism that gave rise to alternating intervals of partial attachment and separation, but with no regular frequency. Low-amplitude excitation downstream of the bubble enhanced poststall lift and signie cantly attenuated the associated unsteadiness. In general, larger momentum coefe cients were required for NACA 0012 separation control due to the large centrifugal acceleration of the e ow around the leading edge. Because of the different stalling characteristics, relatively high- and low-excitation frequencies were effectivefor the NACA 0012 and NACA 0015 airfoils, respectively. However, the combination of high-excitation amplitudes with relatively low frequencieswas effective on the NACA 0012, and this was believed to be associated with the large harmonic content of the evolving perturbations.

Flight-Test Evaluation of Flutter Prediction Methods

Abstract: Several methods have been formulated to predict the onset of e utter during e ight testing. These methods have been demonstrated using data from simulations; however, a rigorous evaluation that includes data from e ight testing must be performed. The ability of several methods to predict the onset of e utter by analyzing data from e ight tests of the aerostructures test wing is evaluated. The evaluated methods include data-based approaches that use damping extrapolation, an envelope function, the Zimmerman‐ Weissenburger e utter margin, and a discretetime autoregressive moving-average model. Also, a model-based approach that uses the π-method e utterometer is evaluated. The data-based methods are demonstrated to be unable to predict e utter accurately using data from low-speed test points, but converge to the accurate solution as airspeed is increased. Conversely, the e utterometer isdemonstrated tobeimmediatelyconservativeusing data from low-speed testpoints,butthesepredictionsremain conservative and do not converge to the true e utter speed as the envelope is expanded. The operation of a e ight test should note theproperties of each method to perhaps adjust test points based on the predicted e utter margins.

Thrust Generation Caused by Flapping Airfoils in a Biplane Configuration

Abstract: Unsteady, viscous e ows over e apping airfoils in a biplane cone guration are computed on moving overset grids. The overset grid solutions are obtained in parallel in a distributed memory environment. Unsteady e owe elds are described by particle traces. Time-averaged thrust values are obtained from the integration of the unsteady drag coefe cient. It isshown thatairfoilsin abiplanecone guration andoscillatingina combinedpitchand plungemotion with a proper phase shift between them produce 20 ‐40% more thrust than a single e apping airfoil. Turbulence in the e ow further augments the thrust generation. For a maximum thrust at a given e apping frequency, an optimization of the e apping motion parameters is needed.

Proposed System-Level Multidisciplinary Analysis Technique Based on Exergy Methods

Abstract: It is suggested that it may be time to consider whether we have reached a plateau in terms of the evolutionary nature of e ight vehicle design and optimization. For a time, progress was measured in terms of maximum speed, which is a straightforward metric when the next design is evolved from the preceding model. There are times, however, when we need to depart from the evolutionary process and create a breakthrough design. The question to be asked is whether there is any way to dee ne system-level analysis and optimization techniques to facilitate the vehicle design process with a more global measure of effectiveness. This paper presents such a methodology for the design of the complete integrated system of systems. Work that has been done in energy-based methods is briee y reviewed, sinceenergy is an implicit consideration in many aerospace disciplines. In addition, methods such as exergy and thermoeconomics have been applied in the design of ground power stations and they are currently being studied for application to aircraft subsystems. The objective of this paperis to expand exergy methods to the design of a complete e ight vehicle by dee ning mission requirements as an exergy/work problem cascading down to each component in the same framework. This paper also serves to introduce a special section of this journal devoted to the application of exergy methods to all levels of e ight vehicle design. Overall, the proposed technique provides a method to facilitate system-level optimization at all levels of the design process.

Effects of Parametric Uncertainty on Airfoil Limit Cycle Oscillation

Abstract: Introduction T HE need for revolutionizingmethods of assessing aeroelastic stabilityhas become increasinglypressing in recentyears.This is driven primarily by two factors: 1) the desire to reduce the total cost of certiŽ cation by reducing testing requirements, and 2) the emergence of unique design concepts to provide impressive performance in military applications.A common feature of these designs is that they substantially increase the potential for nonlinear behavior beyond levels that can be adequately addressed by current engineering tools and processes. These needs and concerns were the focus of a recent workshop organizedby the Air Force OfŽ ce of ScientiŽ c Research and the Air Force Research Laboratory. The workshop addressed traditional areas of concern, such as the basic physics and computational requirements of nonlinear aeroelasticity,but it also included sessions on model veriŽ cation and validation (VVnotably,itwas agreed that UQ could provide a common language for promoting communication between analysts and test personnel. This Note is intended to demonstrate the application of standard probabilityconceptsandMonteCarlo simulation(MCS) to the study of airfoil limit cycle oscillation (LCO), which results from a subcriticalHopf bifurcation induced by including a nonlinear spring in the pitch degree of freedom (DOF). Unsteady aerodynamic forces are represented by the R. T. Jones approximation of the circulatory lift. This simple aeroelastic model permits an assessment of aeroelastic performance sensitivity and variability within the context of a well-understood system; furthermore, because each MCS realization requires time integration, the use of a simple model is computationally expedient. Employing such an elementary model is justiŽ ed in this application because our primary goals are to illustrate the qualitative consequences of uncertainty in a nonlinear aeroelastic system and also to provide a simple example of how probabilistic aeroelastic analysesmight be performed in the future. The authors and a colleague have employed essentially the same procedure detailed herein to highlight the in uence of various uncertainties in panel LCO.2i4

Lifting-Line Analysis of Roll Control and Variable Twist

Abstract: A more practical form of an analytical solution that can be used to predict the roll response for a wing of arbitrary planform with arbitrary spanwise variation of control surface deflection and wing twist is presented. This infinite series solution is based on Prandtl 's classical lifting-line theory and the Fourier coefficients are presented in a form that only depends on wing geometry. The solution can be used to predict rolling and yawing moments as well as the lift and induced drag, which result from control surface deflection, rolling rate, and wing twist. The analytical solution can be applied to wings with conventional ailerons or to wings utilizing wing-warping control. The method is also applied to full-span twisting control surfaces, named "twisterons," which can be simultaneously used to provide roll control, high-lift, and minimum induced drag. A closed-form solution for optimum twist in a wing with linear taper is also presented. Nomenclature An = coefficients in the infinite series solution to the lifting-line equation an = planform contribution to the coefficients in the infinite series solution to the lifting-line equation b =w ingspan bn = twist contribution to the coefficients in the infinite series solution to the lifting-line equation Di C = induced drag coefficient L C = lift coefficient α , L C = wing lift slope α , ~ L C = airfoil section lift slope f L C δ , = change in wing lift coefficient with respect to flap deflection t L C δ , = change in wing lift coefficient with respect to twisteron deflection " C = rolling moment coefficient p C , " = change in rolling moment coefficient with respect to dimensionless rolling rate δ , " C = change in rolling moment coefficient with respect to control surface deflection m

Rotary wing aeroelasticity - A historical perspective

Abstract: This paper provides a historical perspective of the fundamental developments that have played a central role in rotary-wing dynamics and aeroelasticity and have had a major impact on the design of rotary-wing aircraft. The paper describes a historical progression starting with the classical flap-pitch problem that emulated fixed-wing behavior and describes the evolution of the dynamic and aeroelastic problems into those that are unique to rotoreraft, such as the flap-lag problem, the lag-pitch problem, and the coupled flap-lag-torsional problem. Subsequently, the coupled rotor/fuselage aeromechanical problems such as ground and air resonance are considered. A description of the evolution of the methodology used in the formulation and solution of these types of problems is also provided, emphasizing the structural and aerodynamic models required for their effective formulation and solution. The mathematical techniques used for solving the rotary-wing aeroelastic problems in hover and forward flight are also described. The primary emphasis of the paper is on aeroelastic stability, and aeroelastic response is only treated briefly. The paper focuses on contributions that have historical value because they represent landmark treatments. Because of the large amount of material available, an all-inclusive treatment of the research done in this field is impractical, and the paper has unavoidable omissions.

Exergy Approach to Decision-Based Design of Integrated Aircraft Thermal Systems

Abstract: The concept of using an exergy-based approach as a thermal design methodology tool for integrated aircraft thermal systems is introduced. An exergy-based approach was applied to the design of an environmental control system (ECS) of an advanced aircraft. Concurrently, a traditional energy-based approach was applied to the same system. Simplified analytical models of the ECS were developed for each method and compared to assess the ability of each method to suggest optimal design paths. The study identified some roadblocks to assessing the value of using an exergy-based approach: Energy and exergy methods seek answers to different questions, making direct comparisons awkward, and high-entropy generating devices can dominate the design objective of the exergy approach. Nonetheless, exergy methods do approach design differently, providing a ready estimate for efficiency on a component and system basis. Multiobjective optimization tradeoff studies between design weight and entropy generated were used to determine optimal design points. The results from the two analyses provide similar but different decision solutions. The methodology and its implementation are illustrated

Off-Surface Aerodynamic Measurements of a Wing in Ground Effect

Abstract: The off-surface aerodynamic characteristics of a wing in ground effect are investigated using a number of methods including laser Doppler anemometry and particle image velocimetry. The study focuses on two aspects of the flow: turbulent wake and edge vortex. These features are closely associated with the behavior of the aerodynamic force in ground effect. The size of the wake increases in proximity to the ground. A downward shift of the path of the wake is also observed. Discrete vortex shedding is seen to occur behind the wing. As the wing height is reduced, separation occurred on the suction surface of the wing, and the spanwise vortex shedding is found to couple with a flapping motion of the wake in the transverse direction. An edge vortex is also observed off the edge of the end plate of the wing, which contributes to force enhancement and helps to define the force behavior in the force enhancement region. The rate of change in the downforce vs height curve is linked to the strength of the edge vortex. The vortex breakdown signals a slowdown in the force enhancement. When the maximum downforce height is reached, the edge vortex breaks down completely.

Correlation Between Vortex Strength and Axial Velocity in a Trailing Vortex

Abstract: The vortices that trail from the wingtips of a large aircraft provide a significant hazard to an aircraft that follows in its wake. The objective is to contribute to the understanding of these vortices by identifying the conditions where an axial velocity in excess of the freestream value will be generated in the core of a trailing vortex. The axial velocity near the core of a trailing vortex was measured using a triple-sensor hot-wire probe and compared with measured values of vortex circulation strength. The vortex was generated in a wind tunnel using a NACA 0015 wing model with a semispan aspect ratio of 0.80. A linear relationship between the axial velocity and a nondimensional circulation parameter is indicated. For small values of the circulation parameter, the axial velocity shows a velocity deficit

Computation of limit-cycle oscillations of a delta wing

Abstract: This paper presents computational simulations of limit cycle oscillations of a cropped delta wing. A newly developed aeroelastic solver which couples a well validated Navier-Stokes code with a nonlinear nite element method for the von Karman plate equations is employed. Previous computations using a linear structural modal solver to model the delta wing produced limit cycle oscillations with significantly larger amplitudes than the experimentally measured limit cycle response. The present computations with the nonlinear structural model produce limit cycle amplitudes commensurate with the experimental measurements. The computations presented in the paper demonstrate that the geometric nonlinearities in the structural model provide the proper nonlinear mechanism for the development of the limit cycle response observed in the experiments.

Six-Degree-of-Freedom Model of a Controlled Circular Parachute

Abstract: : The paper continues a series of publications devoted to modern advances in aerodynamic decelerator system technology started recently (Journal of Aircraft, Vol. 38, No. 5, 2001) and addresses the development of a six-degree- of-freedom model of a guided circular parachute. The paper reviews existing circular parachute models and discusses several modeling issues unresolved within the frame of existing approaches or completely ignored so far. These issues include using data obtained in the aerodynamic experiments and computational-fluid-dynamics modeling for both undistorted (uncontrolled) and distorted (controlled) canopy shapes, introducing and computing control derivatives, and providing comparison with the real flight data. The paper provides step-by-step development of the mathematical model of circular parachute that includes the basic equations of motion, analysis and computation of the aerodynamic forces and moments, and investigation with modeling of special modes observed in flight. It then introduces a new application of a two-step aerodynamic parameters identification algorithm that is based on comparison with two types of the air-drop data (uncontrolled set and controlled one). The paper ends with summary of the obtained results and proposes a vital direction for the further elaboration of the developed model.

Design of a Martian Autonomous Rotary-Wing Vehicle

Abstract: The design of a Martian autonomous rotary wing vehicle (MARV) is described. MARV is a 50-kg gross takeoff mass, coaxial helicopter designed for Mars exploration. Powered by a fuel cell system, it carries a payload of 10.8 kg over a range of 25 km with an endurance of 39 min including hover capability for 1 min. MARV is designed in response to the Request For Proposal from NASA/Sikorsky for the Year 2000 American Helicopter Society student design competition. The design covers aerodynamic and structural design of rotor blades, vehicle power plant, fuselage and landing gear, control system, transmission, and vehicle lander communications. A detailed mechanism for autonomous deployment of the vehicle from the lander is also described. This preliminary design study indicates that controlled vertical flight on Mars is feasible with existing technology

Investigation of Factors Affecting Iced-Airfoil Aerodynamics

Abstract: A summary of the effects the ice-accretion geometry, size, and location; the airfoil geometry; and the e ight Reynolds number on iced-airfoil aerodynamics, based on the e ndings of the recent University of Illinois investigations,ispresented.Fourairfoilsweretestedwithsimulatedglaze-icehornandspanwiseridgeice.Increasing theiceshapeheightgenerally resultedin moresevereperformancedegradation.Theexceptionwaswhentheiceshapewas locatedattheleadingedgeoftheairfoil,whereincreasedice-shapeheightdidnotsignie cantlydegradeperformance. Varying the leading-edge radius of glaze-icehorn did not have a large effect on airfoil performance. The variations in the geometry of the simulated ridge ice had some effect on airfoil aerodynamics, with (of the shapes tested ) the half-round shape having a signie cantly higher maximum lift. Iced-airfoil aerodynamics were relatively insensitive to Reynolds number variations. Large differences in iced-airfoil aerodynamics were observed between different airfoil geometries. The e ndings showed that an airfoil’ s sensitivity to ridge-ice accretions (which usually forms between 10 and 20% chord )was largely dependent on its load distribution. The airfoil that was very front-loaded, with large leading-edge suction, had the most severe performance degradation due to this type of ice accretion.