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Showing papers on "Pitching moment published in 2005"


Journal ArticleDOI
TL;DR: In this article, an industrial cambered controlled-diffusion airfoil is placed at the exit of an open-jet anechoic wind tunnel, with a jet width of about four chord lengths.
Abstract: A previous experimental investigation of the broadband self noise radiated by an industrial cambered controlled-diffusion airfoil embedded in an homogeneous flow at low Mach number has been extended to various aerodynamic loadings. The instrumented airfoil is placed at the exit of an open-jet anechoic wind tunnel, with a jet width of about four chord lengths. Sound is measured in the far field at the same time as the statistical properties of the wall-pressure fluctuations close to the trailing edge. A new set of mean wall-pressure data has been collected on this airfoil at a chord Reynolds number of 2.9 x 105, which provides some insight on the Reynolds-number effect. Two previously investigated flow regimes with different statistical behaviors are investigated by changing the angle of attack from 8 to 15 deg. They respectively correspond to the nearly separated boundary layer with vortex shedding at the trailing edge and to the turbulent boundary layer initiated by a leading-edge separation.

122 citations


Journal ArticleDOI
TL;DR: In this paper, a reduced-order model is used to predict the lift and pitching moment histories accurately throughout the subsonic and transonic regimes of an airfoil to arbitrary shaped gust inputs.
Abstract: I. Abstract Computational Fluid Dynamics (CFD) based simulations, along with parametric and non-parametric Reduced-Order Models (ROM) for gust responses are presented. A CFD code is enhanced to simulate responses of an airfoil to arbitrary shaped gust inputs. Time-domain Auto-Regressive-Moving-Average (ARMA) models are identified based on CFD responses to random gust excitations, using system-identification methods. Responses to discrete gusts of various shapes, amplitudes and gradient lengths are computed via the ROMs and compared to responses simulated directly by the CFD code. The ROMs predict the lift and pitching moment histories accurately throughout the subsonic and transonic regimes. They offer significant savings in computational resources compared to the full CFD simulation, since only one CFD run is required for ROM identification, from which responses to arbitrary shaped gusts can be rapidly estimated. The combination of ROMs and full CFD simulations offers a computationally efficient tool set of various-fidelity time-domain models for gust responses. The ROMs can be used for rapid tuned-gust analyses, and the critical cases can be simulated with a full CFD run, providing pressure distribution for airframe structural design.

87 citations


Journal ArticleDOI
TL;DR: The U.S. Air Force Missile DATCOM (97 version) and the Naval Surface Warfare Center Dahlgren Division Aeroprediction 98 (AP98) are two widely used aerodynamic prediction codes as discussed by the authors.
Abstract: The U.S. Air Force Missile DATCOM (97 version) and the Naval Surface Warfare Center Dahlgren Division Aeroprediction 98 (AP98) are two widely used aerodynamic prediction codes. These codes predict aerodynamic forces, moments, and stability derivatives as a function of angle of attack and Mach number for a wide range of axisymmetric and nonaxisymmetric missile configurations. This study evaluates the accuracy of each code compared to experimental wind-tunnel data for a variety of missile configurations and flight conditions. The missile configurations in this study include axisymmetric body alone, body wing tail, and body tail. The aerodynamic forces under investigation were normal force, pitching moment, axial force, and center-of-pressure location. For the configurations detailed in this paper, these case studies show normal force prediction for both codes to have minimal error. Both AP98 and Missile DATCOM were effective in predicting pitching-moment coefficients, though at times limiting factors were necessary. Finally, both AP98 and DATCOM predict reasonable axial-force coefficients for most cases, though AP98 proved more accurate for the body-wing-tail and body-tail configurations evaluated in this study.

63 citations


Journal ArticleDOI
TL;DR: In this paper, the authors compared the measured data from flight and wind-tunnel tests were compared with calculations obtained using the comprehensive analysis CAMRAD II, and the analysis showed that the aerodynamic tip design (chord length and quarter-chord location) has an important influence on the phase correlation.
Abstract: Blade section normal force and pitching moment were investigated for six rotors operating at transition and high speeds: H-34 in flight and wind tunnel, SA 330 (research Puma), SA 349/2, UH-60A full-scale, and BO-105 model (Higher-Harmonic Acoustics Rotor Test I). The measured data from flight and wind-tunnel tests were compared with calculations obtained using the comprehensive analysis CAMRAD II. The calculations were made using two free-wake models: rolled up and multiple trailer with consolidation models. At transition speed, there is fair to good agreement for the blade section normal force between the test data and analysis for the H-34, research Puma, and SA 349/2 with the rolled-up wake. The calculated airloads differ significantly from the measurements for the UH-60A and BO-105. Better correlation is obtained for the UH-60A and BO-105 by using the multiple trailer with consolidation wake model. In the high-speed condition, the analysis shows generally good agreement with the research Puma flight data in both magnitude and phase. However, poor agreement is obtained for the other rotors examined. The analysis shows that the aerodynamic tip design (chord length and quarter-chord location) of the research Puma has an important influence on the phase correlation.

56 citations


Journal ArticleDOI
TL;DR: In this paper, the authors used the commercial computational fluid dynamics (CFD) code CFX-5 of ANSYS to compute the engine installation drag for the German Aerospace Center (DLR) F6 aircraft configuration as part of the Second AIAA Drag Prediction Workshop.
Abstract: The commercial computational fluid dynamics (CFD) code CFX-5 of ANSYS, Inc., has been used to compute the engine installation drag for the German Aerospace Center (DLR) F6 aircraft configuration as part of the Second AIAA Drag Prediction Workshop. The computations were performed with the standard hexahedral meshes provided by ICEM.CFD to the workshop. The full drag polar for the DLR-F6 configuration has been computed. For all cases, good agreement between the experiment and the predictions were obtained for lift, drag, and pitching moment coefficients. All simulations where based on the shear stress transport turbulence model, and additional computations have indicated that turbulence modeling issues are largely responsible for the overprediction of the lift curve slope that was observed by many of the workshop participants.

46 citations


Proceedings ArticleDOI
18 Apr 2005
TL;DR: The results of a wind tunnel test at NASA Langley in the Transonic Dynamics Tunnel of a gust load alleviation system for a SensorCraft concept and the 4 trailing edge and 1 leading edge control effectors were used to simultaneously control first and second bending.
Abstract: SensorCraft is an Air Force Research Lab (AFRL) concept for a high flying vehicle that will be capable of providing greatly increased Intelligence, Surveillance, and Reconnaissance (ISR) capabilities. Part of the technology SensorCraft will require is high aerodynamic and structural efficiency to accomplish its mission goal of long range and sustained presence. To achieve a long loiter time, it will have a light weight structure with a high aspect ratio wing that carries much of the fuel mass. Hence, the structural modes will be in the same frequency range as the rigid body modes and will be strongly coupled with them. Wing stresses will need to be reduced and the gust load contribution will need to be minimized to keep the structural bending loads low along the span of the wing. This paper documents the results of a wind tunnel test at NASA Langley in the Transonic Dynamics Tunnel of a gust load alleviation system for a SensorCraft concept. The wing was fixed to the wall of the tunnel and the 4 trailing edge and 1 leading edge control effectors were used to simultaneously control first and second bending as well as control the pitching moment at the balance attachment.

37 citations


Journal ArticleDOI
TL;DR: In this article, a genetic algorithm was used to minimize the total drag at fixed lift subject to various geometrical and aerodynamic constraints, such as pitch moment, pressure and free-stream Mach number.
Abstract: Ar obust and efficient approach to the multiobjective constrained design, previously developed by the authors, is extended to optimization of three-dimensional aerodynamic wings. The objective is to minimize the total drag at fixed lift subject to various geometrical and aerodynamical constraints. The approach employs genetic algorithms (GAs) as an optimization tool in combination with a reduced-order-models (ROM) method, based on linked local databases obtained by full Navier‐Stokes computations. The work focuses on the following issues: geometrical representation of three-dimensional shapes, handling of sensitive nonlinear constraints such as pitching moment, and the influence of flight conditions on the results of optimization. The method, implemented in the computer code OPTIMAS (Optimization of Aerodynamic Shapes), was applied to the problem of multipoint transonic threedimensional wing optimization with nonlinear constraints. The results include a variety of optimization cases for two wings: a classical test case of ONERA M6 wing and a generic cranked transport-type wing. For the investigated class of problems, significant aerodynamic gains have been obtained. Nomenclature C D = total drag coefficient CL = total lift coefficient C M = total pitching-moment coefficient C p = pressure coefficient M = freestream Mach number N D = dimension of the search space Nws = number of sectional airfoils Q = objective function R/c = relative radius of the airfoil leading edge Re = freestream Reynolds number t/c = relative thickness of airfoil α = angle of attack θ = trailing-edge angle of airfoil

37 citations


01 Sep 2005
TL;DR: In this paper, a discrete adjoint of the Navier-Stokes equations has been developed in the unstructured finite-volume solver the DLR-TAU-code.
Abstract: A discrete adjoint of the Navier-Stokes equations has been developed in the unstructured finite-volume solver the DLR-TAU-code. The method consists of the explicit construction of the exact Jacobian of the spatial discretization with respect to the unknown variables allowing the adjoint equations to be formulated and solved. A wide range of the spatial discretizations available in TAU have been differentiated, including the Spallart-Almaras-Edwards one-equation, and the Wilcox k-omega two-equation turbulence models. The aim of this paper is to give an overview of the capabilities of the discrete adjoint to perform aerodynamic shape optimization in viscous flow. The strategy developed is extensively validated on 2D cases. The adjoint based design method is first validated by comparing the gradients of the drag, lift and pitching moment it produces with the approximate gradients obtained by finite-differences. Then the accuracy and efficiency of the approach are demonstrated for transonic airfoil design by considering geometric as well aerodynamic constraints, single- as well as multi-point design. Finally, the flap design of a multi-element airfoil in take off configuration confirms the capability of the discrete adjoint to solve wide range of aerodynamic problems.

31 citations


Journal ArticleDOI
TL;DR: In this article, a series of nearly 90 CFD test cases performed as a contribution to the second Drag Prediction Workshop, held in Orlando, Florida in June 2003, is described and analyzed.

28 citations


Proceedings ArticleDOI
06 Jun 2005
TL;DR: In this article, a trapezoidal three-element high-lift wing obtained with an unstructured Reynolds averaged Navier-Stokes code is presented, and the requirements for grid renemen ts needed to model the pertinent o w physics at various points on the lift curve are identied.
Abstract: Solutions on a trapezoidal three-element high-lift wing obtained with an unstructured Reynolds averaged Navier-Stokes code are presented. Requirements for grid renemen ts needed to model the pertinent o w physics at various points on the lift curve are identied. Improvements to the lift prediction due to grid renemen t are demonstrated. Lift, drag, pitching moment are shown to be in good agreement with experimental data through C Lmax : The eect of slat and ap brackets are shown, although more grid resolution is needed to correctly capture these eects.

26 citations


Journal ArticleDOI
TL;DR: In this article, the authors developed an accurate airfoil analysis lower order code, based on the singularities method, for wind turbine applications using the 2D incompressible potential flow model.

Proceedings ArticleDOI
17 May 2005
TL;DR: In this paper, a small-scale expandable morphing wing is presented, which is separated into inner and outer wings as a typical bird wing and the wing can change its aspect ratio from 4.7 to 8.5 in about 2 seconds and the speed can be controlled.
Abstract: In this paper, we present design, manufacturing, and wind tunnel test for a small-scale expandable morphing wing. The wing is separated into inner and outer wings as a typical bird wing. The part from leading edge of the wing chord is made of carbon composite strip and balsa. The remaining part is covered with curved thin carbon fiber composite mimicking wing feathers. The expandable wing is driven by a small DC motor, reduction gear, and fiber reinforced composite linkages. Rotation of the motor is switched to push-pull linear motion by a screw and the linear motion of the screw is transferred to linkages to create wing expansion and folding motions. The wing can change its aspect ratio from 4.7 to 8.5 in about 2 seconds and the speed can be controlled. Two LIPCAs (Lightweight Piezo-Composite Actuators) are attached under the inner wing section and activated on the expanded wing state to modify camber of the wing. In the wind tunnel test, change of lift, drag, and pitching moment during wing expansion have been investigated for various angles of attack. The LIPCA activation has created significant additional lift.

Journal ArticleDOI
TL;DR: In this paper, the yawing moment acting on reinforced concrete bridges constructed using the balanced cantilever method during the erection stage has been experimentally analyzed by testing different types of bridge cross-sections.

Journal ArticleDOI
TL;DR: In this paper, the relationship between boundary-layer transition and the freestream turbulence intensities for a 33.8% thickness elliptic cylinder at zero angle of attack was investigated, and the results showed that at low Reynolds numbers, aerodynamic characteristics can be dominated by laminar flow phenomena including transition and separation bubbles.
Abstract: Introduction R ECENTLY unmanned vertical takeoff and landing vehicles have attracted aeronautical engineers’ attention. One of those is a canard rotor/wing (CRW) aircraft, and several studies have been reported.1−3 CRW aircraft usually employs a two-bladed rotor to takeoff the ground; after taking off, it locks the rotor to act as a fixed wing, allowing it to cruise at high speed. This requires that the cross section of the rotor blades be elliptic. Flow past an elliptic cylinder or ellipse cross-sectioned airfoil has been studied for a long time because of its significance in fundamental flow physics and its practical importance. However, most studies so far have been concerned with unsteady flow characteristics and/or heat transfer at very low Reynolds numbers. Aerodynamic characteristics such as lift, drag, and pitching moment variations with angle of attack have rarely been reported. Zahm et al.4 reported the drag and surface pressures of four elliptic cylinders of different thickness ratio for various yaw angles. Schubauer5 carried out an intensive study of the boundary layer around an elliptic cylinder. This study, however, focused on the relationship between the boundary-layer transition and the freestream turbulence intensities for a 33.8% thickness elliptic cylinder at zero angle of attack. Hoerner and Borst6 reported lift coefficients of elliptic airfoils having various thickness ratios for the Reynolds numbers of 2 × 106 and 7 × 106. White7 compiled the drag coefficients of several elliptic cylinders with various thicknesses for zero angle of attack in laminar and turbulent flow conditions. Wings for unmanned vehicles are often operated at low Reynolds number. At such low Reynolds numbers, aerodynamic characteristics can be dominated by laminar flow phenomena including transition and separation bubbles. Gleyzes et al.8 reported boundary-layer

01 May 2005
TL;DR: Aerodynamic characteristics for the aircraft model with NACA (National Advisory Committee for Aeronautics) wing No. 653-218 have been studied using subsonic wind tunnel of 1000 mm x 1000 mm rectangular test section and 2500 mm long of Aerodynamics Laboratory Faculty of Engineering (Universiti Putra Malaysia).
Abstract: Aerodynamic characteristics for the aircraft model with NACA (National Advisory Committee for Aeronautics) wing No. 653-218 have been studied using subsonic wind tunnel of 1000 mm x 1000 mm rectangular test section and 2500 mm long of Aerodynamics Laboratory Faculty of Engineering (Universiti Putra Malaysia). Six components wind tunnel balance is used for measuring lift, drag and pitching moment. Tests are conducted on the aircraft model with and without winglet of two configurations at Reynolds numbers 1.7 x 10 5 , 2.1 x 10 5 , and 2.5 x 10 5 . Lift curve slope increases more with the addition of the elliptical winglet and at the same time the drag decreases more for the aircraft model with elliptical shaped winglet giving an edge over the aircraft model without winglet as far as Lift/Drag ratio for the elliptical winglet is considered. Elliptical winglet of configuration 2 (Winglet inclination 60 0 ) has, overall, the best performance, giving about 6% increase in lift curve

01 Jun 2005
TL;DR: In this article, the authors describe airfoil and blown high lift concepts that also address cruise drag reduction and reductions in mass flow through the use of pulsed pneumatic blowing on a Coanda surface.
Abstract: Circulation Control technologies have been around for 65 years, and have been successfully demonstrated in laboratories and flight vehicles alike, yet there are few production aircraft flying today that implement these advances. Circulation Control techniques may have been overlooked due to perceived unfavorable trade offs of mass flow, pitching moment, cruise drag, noise, etc. Improvements in certain aspects of Circulation Control technology are the focus of this paper. This report will describe airfoil and blown high lift concepts that also address cruise drag reduction and reductions in mass flow through the use of pulsed pneumatic blowing on a Coanda surface. Pulsed concepts demonstrate significant reductions in mass flow requirements cor Circulation Control, as well as cruise drag concepts that equal or exceed conventional airfoil systems.

Book ChapterDOI
10 Jan 2005
TL;DR: In this paper, the authors evaluate the feasibility of controlling the shock/boundary layer (SBL) interaction on airfoils in transonic flow using synthetic jet actuators and thus achieving the desired changes in aerodynamic forces and moments.
Abstract: The objective of this paper is to evaluate, via numerical simulation, the feasibility of controlling the shock/boundary layer (SBL) interaction on airfoils in transonic flow thereby weakening the shock wave(s) and reducing the pockets of supersonic regions using synthetic jet actuators and thus achieving the desired changes in aerodynamic forces and moments. It is shown that by properly choosing the amplitude, frequency, momentum and location of the synthetic jet on the airfoil surface, desired modulation in lift, drag and pitching moment coefficients can be obtained for a given free-stream Mach number, Reynolds number and angle of attack. These calculations indicate the possibility of “hingeless” control of airplanes in the future using active flow control (AFC).

Proceedings ArticleDOI
06 Jun 2005
TL;DR: It is shown that starting from basic information such as the planform, it is able to predict the anticipated performance with sufficient confidence for comparative assessments and further work is proposed in several areas.
Abstract: Currently there is a revival of interest in flying wings for military (and civil) use. The military context has arisen from the future “stealthy” HALE and UCAV aircraft. Questions on aerodynamics, control and structural efficiency arise. Compared with conventional wing / tail arrangements, flying wings have a special set of very different constraints. These are mentioned. Without a trim surface, the constraints on the wing pitching moment dictate the design camber and twist. Control power requirements can be high because of effectively short moment arms. The camber and twist are strongly dependent on trim stability margins. This aspect needs to be understood in detail when comparing different types of planforms. This paper is concerned with the study of wing planforms for UCAV applications. It is inspired by the need to understand a variety of wings (in public domain) that are, at first sight, aimed at similar missions. The main emphasis has been on developing and understanding cruise design camber and twist with Cm constraints of stable, neutral and unstable static margins. . Spanwise lift and drag loadings have also been presented. Camber design has been via attained thrust methods and a modal approach. It is shown that starting from basic information such as the planform, we are able to predict the anticipated performance with sufficient confidence for comparative assessments. Further work is proposed in several areas.

Journal ArticleDOI
TL;DR: In this paper, the stall hysteresis of GA(W)-2 airfoil with 25% slotted flap was examined by using the data obtained for lift, pitching moment, surface pressure distribution and hot film velocity vector.
Abstract: Stall hysteresis discovered in wind tunnel performance of GA(W)-2 airfoil with 25%chord slotted flap is examined further by using the data obtained for lift, pitching moment, surface pressure distribution and hot film velocity vector. Test cases include 30 and 40 deg flap deflections each having an optimum and narrow gap at chord Reynolds number of 2.2 million and Mach number of 0.13. The flap optimized to produce the highest Clmax for each flap angle apparently did not have a proper contour and nose location for the slot flow to function effectively at off-design conditions. It is shown that suction pressures over flap, suppressed by thickening wing wake at stall, are not reversible to their pre-stall values within the decreasing α side of loop. It is suggested that the flap design include the use of a new flap parameter called slot flow angle to describe the slot flow orientation and a pressure recovery factor to select a proper contour for flap upper surface.

Proceedings ArticleDOI
01 Jan 2005
TL;DR: In this paper, a new data analysis technique for the identification of static and dynamic aerodynamic stability coefficients from wind tunnel test video data is presented, which was applied to video data obtained during a parachute wind-tunnel test program conducted in support of the Mars Exploration Rover Mission.
Abstract: A new data analysis technique for the identification of static and dynamic aerodynamic stability coefficients from wind tunnel test video data is presented. This new technique was applied to video data obtained during a parachute wind tunnel test program conducted in support of the Mars Exploration Rover Mission. Total angle-of-attack data obtained from video images were used to determine the static pitching moment curve of the parachute. During the original wind tunnel test program the static pitching moment curve had been determined by forcing the parachute to a specific total angle-of -attack and measuring the forces generated. It is shown with the new technique that this parachute, when free to rotate, trims at an angle-of-attack two degrees lower than was measured during the forced-angle tests. An attempt was also made to extract pitch damping information from the video data. Results suggest that the parachute is dynamically unstable at the static trim point and tends to become dynamically stable away from the trim point. These trends are in agreement with limit-cycle-like behavior observed in the video. However, the chaotic motion of the parachute produced results with large uncertainty bands.

Proceedings ArticleDOI
01 Jan 2005
TL;DR: A pressure sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the effect of wing fillets on the global vortex-induced surface static pressure field about a sharp leading-edge 76o/40o double delta wing, or strake-wing, model at subsonic and transonic speeds as mentioned in this paper.
Abstract: A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the effect of wing fillets on the global vortex-induced surface static pressure field about a sharp leading-edge 76o/40o double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 20 degrees using an in-situ method featuring the simultaneous acquisition of electronically-scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M = 0.50 to 0.85 but increased to several percent at M =0.95 and 1.20. The PSP pressure distributions and pseudo-colored planform view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having a parabolic or diamond planform situated at the strake-wing intersection were designed to manipulate the vortical flows by, respectively, removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.

Patent
30 Nov 2005
TL;DR: In this paper, a fixed-wing aircraft is used to land on its upper side, where the aircraft's back is facing the ground and its underside is facing away from the ground.
Abstract: A method for landing a fixed wing aircraft is provided in which an inversion maneuver is performed so that the aircraft's back is facing the ground, and the aircraft's underside is facing away from the ground. After initiation or completion of this maneuver, deep stall is induced, and the aircraft descends almost vertically to land on its upper side, thus minimizing impact loads or damage on its underside. In a particular aerodynamic arrangement configured for carrying out the method, a flap (24), which may be stowed during normal flight, is deployed in a manner such as to aerodynamically induce a negative pitching moment on the aircraft and deep stall.

Proceedings ArticleDOI
06 Jun 2005
TL;DR: In this paper, a full span NACA 641-412 airfoil with right triangle, inverse triangle and star-shaped damage was experimentally investigated using balance measurements and flow visualisation.
Abstract: The flow on a full span NACA 641-412 airfoil with right triangle, inverse triangle and star shaped damage was experimentally investigated using balance measurements and flow visualisation Generally, when compared with an undamaged model, increasing incidence for a damaged model resulted in increased loss of lift coefficient, increased drag coefficient and more negative pitching moment coefficient The results are compared to each other and also to the results of circle shaped damage from previous work The experiments showed that for all damage cases the flows could be categorised as weak, transitional or strong jets, with the main features of these flows identical to circle shaped damage Adding multiple sharp edges (star) had only limited influence on the observed flow characteristics For all damage cases the jet exited from the rear of the damage and its size was determined by the width of the rear part of the hole

Proceedings ArticleDOI
06 Jun 2005
TL;DR: In this paper, the aerodynamic characteristics of a blended wing body aircraft, constructed using selective laser sintering (SLS), were assessed in the AFIT low-speed wind tunnel.
Abstract: The aerodynamic characteristics of a blended wing body aircraft, constructed using selective laser sintering (SLS), were assessed in the AFIT low-speed wind tunnel The scaled- down model (Rec ~ 10 5 and M = 010 to 020) of a strike tanker consisted of a shaped fuselage and sweptback wings The model evaluation and analysis process included force and moment measurements acquired from a wind tunnel balance, pressure sensitive paint (PSP) measurements, and a comparison to computational fluid dynamics (CFD) solutions One of the most interesting results was the striking difference in the force and moment measurements before and after the paint was applied to the surface The average surface roughness, Ra, was measured and was found to have increased from approximately 03μm to 07μm when the paint was applied Although this value is well below the roughness threshold suggested by 2-D boundary layer theory, the effect was clear and repeatable Force and moment coefficient data suggest that the onset of wing stall was sudden across the entire wing for sufficiently smooth cases but occurs gradually for rougher surfaces Interestingly, the CFD results compared well with the experimental data corresponding to the measurements of the rougher, painted model Both the PSP data and the CFD results indicated that the primary mechanism of lift transitioned from leading-edge suction along the wing to vortex lift near the wing-body junction as angle of attack increased Taken as a whole, the data suggests that, for the range of conditions tested, submicron roughness has a pronounced effect on the transition to the vortex lift mechanism for increasing angle of attack

Proceedings ArticleDOI
10 Jan 2005
TL;DR: In this paper, a series of low-speed wind tunnel tests were carried out on an oscillating airfoil fitted with two rows of air-jet vortex generators (AJVGs).
Abstract: A series of low-speed wind tunnel tests were carried out on an oscillating airfoil fitted with two rows of air-jet vortex generators (AJVGs). The airfoil used had an RAE 9645 section and the two spanwise arrays of AJVGs were located at x/c=0.12 and 0.62. The devices and their distribution were chosen to assess their ability to modify/control dynamic stall; the goal being to enhance the aerodynamic performance of helicopter rotors on the retreating blade side of the disc. The model was pitched about the quarter chord with a reduced frequency (k) of 0.1 in a sinusoidal motion defined by a=15 o +10 o sinω t. The measured data indicate that, for continuous blowing from the front row of AJVGs with a momentum blowing coefficient (Cµ) greater than 0.008, modifications to the stalling process are encouraging. In particular, the pitching moment behavior exhibits delayed stall and there is a marked reduction in the normal force hysteresis.

Proceedings ArticleDOI
15 Aug 2005
TL;DR: In this article, the authors analyzed data from the Flight Data Recorder of a turboprop transport airplane involved in an icing mishap by modeling the aerodynamic characteristics along the flight path.
Abstract: Data from the Flight Data Recorder of a turboprop transport airplane involved in an icing mishap are analyzed by modeling the aerodynamic characteristics along the flight path. The analysis results in the prediction of stability and control derivatives along the flight trajectory. Based on these predicted derivatives, the roll excursion that precedes the accident is interpreted as being caused by divergent wing rock mechanisms, having nonlinear unstable dihedral effect and nonlinear unstable roll damping. To monitor the icing accretion conditions, the aerodynamic normal force and pitching moment coefficients in suspected icing conditions are compared with those in normal flight conditions. For this purpose, the typical aerodynamic models for the normal force and pitching moment coefficients in flight conditions without significant icing are established. The aerodynamic coefficients in flight with suspected icing are obtained through compatibility analysis of recorded flight variables, such as the angle of attack and normal acceleration. By comparing these aerodynamic coefficients with the typical model-predicted aerodynamic coefficients based on the flight variables in suspected icing, it is possible to predict approximately when significant icing may start on the wing and/or the horizontal tail. Therefore, the scheme is capable of timely detection of significant ice built-up.

Journal ArticleDOI
TL;DR: In this paper, an experimental and computational study has been performed for investigation of jet interaction in supersonic flow with a three-dimensional surface ramp located behind a sonic, transverse jet.
Abstract: An experimental and computational study has been performed for investigation of jet interaction in supersonic flow with a three-dimensional surface ramp located behind a sonic, transverse jet. The goal was to reduce the low-pressure region behind the jet, which produces an unwanted nose-down moment on a vehicle. The experimental techniques include conventional pressure taps, schlieren photos, and pressure-sensitive paint (PSP). The numerical solver used in this study is AeroSoft’s structured flow solver GASP Version 4.0. A Mach 4 crossflow with a jet stagnation pressure ratio of 532 was considered, and the three-dimensional downstream ramp was designed by a parametric study using GASP. The computational-fluid-dynamics and PSP results are verified with pressure tap results, and all results are compared between the jet-only case and jet-plus-ramp case. The results showed that the ramp located downstream of the jet decreases the nose-down pitching moment without a net force loss.

Proceedings ArticleDOI
10 Jan 2005
TL;DR: In this paper, the authors developed an active fluidic control system that can effectively manipulate the vortex breakdown location over a highly swept delta wing, which can be incorporated into a feedback loop to input a desired pitching moment based on real-time measured surface pressure.
Abstract: The objective of this work was to develop an active fluidic control system that can effectively manipulate the vortex breakdown location over a highly swept delta wing. By moving the vortex breakdown fore or aft, a pitching moment can be induced on the delta wing without the use of any conventional control surfaces. The active control system can be incorporated into a feedback loop to input a desired pitching moment based on the real-time measured surface pressure. The type of active fluidic control system shown to be the most effective at delaying vortex breakdown was the along-core injection technique. The present study describes an open loop control system, that is, the flow field was observed and measured and then injection conditions were changed manually. The process was repeated until optimal conditions were observed. Initial closed loop feedback control tests were performed to test the response of the system. A 60° delta wing model with a maximum span of 15.5 inches and a root chord of 13.5 inches was mounted in a subsonic wind tunnel. The wing was equipped with six control jets with variable azimuthal and pitch angles on the top surface approximately beneath the vortex core. A thorough optimization process was completed measuring static pressure to determine vortex breakdown location. Other variables in addition to azimuthal and pitch angles were injection momentum, frequency, and duty cycle. Measuring dynamic pressure, that is pressure fluctuation due to vortex shedding and pulsed injection, was also necessary to the development of a control algorithm. Dynamic pressure was measured at four chord-wise locations with uniform and randomly modulated duty cycles in order to ascertain duty cycle sensitivity on the system’s overall effectiveness.

Proceedings ArticleDOI
15 Aug 2005
TL;DR: A/C = aircraft AoA = angle of attack BWB = blended wing body CAP = control anticipation parameter CSAS = control and stability augmentation system cg = centre of gravity CL = lift coefficient Cm = pitching moment about aerodynamic reference point DOF = degrees of freedom DR = Dutch roll mode FCS = flight control system GNC = guidance, navigation and control Gc = compensator structure h = altitude HQ = handling qualities J = performance index Kn = static margin LQR = linear quadratic regulator mac = mean aerodynamic chord MDO = mult
Abstract: A/C = aircraft AoA = angle of attack BWB = blended wing body CAP = control anticipation parameter CSAS = control and stability augmentation system cg = centre of gravity CL = lift coefficient Cm = pitching moment about aerodynamic reference point DOF = degrees of freedom DR = Dutch roll mode FCS = flight control system GNC = guidance, navigation and control Gc = compensator structure h = altitude HQ = handling qualities J = performance index Kn = static margin LQR = linear quadratic regulator mac = mean aerodynamic chord MDO = multidisciplinary design optimization nz = vertical load factor PI = proportional-integral compensator SAS = stability augmentation system SP = short period mode xNP = neutral point position with respect to aerodynamic reference point