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Showing papers on "Pitching moment published in 2007"


Journal ArticleDOI
TL;DR: In this article, the equations of motion of an insect with flapping wings were derived and then simplified to that of a flying body using the "rigid body" assumption, and the longitudinal dynamic flight stability of four insects (hoverfly, cranefly, dronefly and hawkmoth) in hovering flight was studied.
Abstract: The equations of motion of an insect with flapping wings are derived and then simplified to that of a flying body using the “rigid body” assumption. On the basis of the simplified equations of motion, the longitudinal dynamic flight stability of four insects (hoverfly, cranefly, dronefly and hawkmoth) in hovering flight is studied (the mass of the insects ranging from 11 to 1,648 mg and wingbeat frequency from 26 to 157 Hz). The method of computational fluid dynamics is used to compute the aerodynamic derivatives and the techniques of eigenvalue and eigenvector analysis are used to solve the equations of motion. The validity of the “rigid body” assumption is tested and how differences in size and wing kinematics influence the applicability of the “rigid body” assumption is investigated. The primary findings are: (1) For insects considered in the present study and those with relatively high wingbeat frequency (hoverfly, drone fly and bumblebee), the “rigid body” assumption is reasonable, and for those with relatively low wingbeat frequency (cranefly and howkmoth), the applicability of the “rigid body” assumption is questionable. (2) The same three natural modes of motion as those reported recently for a bumblebee are identified, i.e., one unstable oscillatory mode, one stable fast subsidence mode and one stable slow subsidence mode. (3) Approximate analytical expressions of the eigenvalues, which give physical insight into the genesis of the natural modes of motion, are derived. The expressions identify the speed derivative Mu (pitching moment produced by unit horizontal speed) as the primary source of the unstable oscillatory mode and the stable fast subsidence mode and Zw (vertical force produced by unit vertical speed) as the primary source of the stable slow subsidence mode.

201 citations


Book
05 Feb 2007
TL;DR: In this article, a subsonic business jet is used to calculate lift, drag, pitching moment, and stability derivatives of a single-passenger aircraft in a vertical plane.
Abstract: Flight mechanics is the application of Newton's laws to the study of vehicle trajectories (performance), stability, and aerodynamic control. This text is concerned with the derivation of analytical solutions of airplane flight mechanics problems associated with flight in a vertical plane. Algorithms are presented for calculating lift, drag, pitching moment, and stability derivatives. Flight mechanics is a discipline. As such, it has equations of motion, acceptable approximations, and solution techniques for the approximate equations of motion. Once an analytical solution has been obtained, numbers are calculated in order to compare the answer with the assumptions used to derive it and to acquaint students with the sizes of the numbers. A subsonic business jet is used for these calculations.

185 citations


Proceedings ArticleDOI
01 Jan 2007
TL;DR: In this paper, the predicted absolute and differential drag levels for wing-body and wing-alone configurations of transonic transport aircraft were predicted using the Reynolds-Averaged Navier-Stokes computational fluid Dynamics Methods.
Abstract: The workshop focused on the prediction of both absolute and differential drag levels for wing-body and wing-al;one configurations of that are representative of transonic transport aircraft. The baseline DLR-F6 wing-body geometry, previously utilized in DPW-II, is also augmented with a side-body fairing to help reduce the complexity of the flow physics in the wing-body juncture region. In addition, two new wing-alone geometries have been developed for the DPW-II. Numerical calculations are performed using industry-relevant test cases that include lift-specific and fixed-alpha flight conditions, as well as full drag polars. Drag, lift, and pitching moment predictions from previous Reynolds-Averaged Navier-Stokes computational fluid Dynamics Methods are presented, focused on fully-turbulent flows. Solutions are performed on structured, unstructured, and hybrid grid systems. The structured grid sets include point-matched multi-block meshes and over-set grid systems. The unstructured and hybrid grid sets are comprised of tetrahedral, pyramid, and prismatic elements. Effort was made to provide a high-quality and parametrically consistent family of grids for each grid type about each configuration under study. The wing-body families are comprised of a coarse, medium, and fine grid, while the wing-alone families also include an extra-fine mesh. These mesh sequences are utilized to help determine how the provided flow solutions fair with respect to asymptotic grid convergence, and are used to estimate an absolute drag of each configuration.

104 citations


Journal ArticleDOI
TL;DR: In this article, a thin-airfoil theory is applied to the lift problem of an airfoil with a Gurney flap, and the lift and pitching moment coefficient increments are given as a square-root function of the relative Gurny flap height, and they are proportionally related.
Abstract: Thin-airfoil theory is applied to the lift problem of an airfoil with a Gurney flap. The lift and pitching moment coefficient increments are given as a square-root function of the relative Gurney flap height, and they are proportionally related. This model interprets the Gurney flap lift enhancement as a special camber effect. The theoretical relations are in good agreement with experimental and numerical data for several different wings. The theoretical method developed in this paper can be applied to similar trailing-edge devices for lift enhancement, and it is useful in the preliminary design of these flow control devices.

81 citations


Journal ArticleDOI
01 Jul 2007
TL;DR: In this paper, the authors investigated the effect of microtabs and microflaps on wind turbine rotors and their time-dependent effect on sectional lift, drag, and pitching moment.
Abstract: The cost of wind-generated electricity can be reduced by mitigating fatigue loads acting on the blades of wind turbine rotors. One way to accomplish this is with active aerodynamic load control devices that supplement the load control obtainable with current full-span pitch control. Techniques to actively mitigate blade loads that are being considered include individual blade pitch control, trailing-edge flaps, and other much smaller trailing-edge devices such as microtabs and microflaps. The focus of this paper is on the latter aerodynamic devices, their time-dependent effect on sectional lift, drag, and pitching moment, and their effectiveness in mitigating high frequency loads on the wind turbine. Although these small devices show promise for this application, significant challenges must be overcome before they can be demonstrated to be a viable, cost-effective technology.

77 citations


Journal ArticleDOI
TL;DR: In this article, a study of aerodynamic loadings on a NACA 0012 airfoil with a static and an oscillating trailing-edge Gurney flap was made.
Abstract: A study of aerodynamic loadings on a NACA 0012 airfoil with a static and an oscillating trailing-edge Gurney flap was made. The focus is on the experimental measurement of the static and dynamic-pressure distributions on the airfoil surface. The experimental results are also correlated with theoretical results obtained using the Navier-Stokes code INS2D, developed by NASA. A Reynolds number of 348,000, a flow velocity of 20 m/s (65.6 ft/s), and a reduced frequency from 0 to 0.4 based upon half-chord b and freestream velocity U are used. The experimental results show that the effect of the static and oscillating strips located near the trailing edge of the airfoil is to enhance the maximum lift and pitching-moment coefficients for both unstalled and stalled angles of attack. An increase of the oscillating frequency also enhances the aerodynamic loading. Reasonably good agreement between the experiment and theory is obtained. The experimental results confirm the idea that an oscillating small strip located near the trailing edge can be a useful tool for active aerodynamic flow control for a wing.

58 citations


Journal ArticleDOI
01 Jan 2007
TL;DR: In this article, the effects of wing sweep on the aerodynamic performance of a blended wing body (BWB) aircraft that is based on an aerodynamically optimized design with a fixed planform and pitching moment constraint were evaluated.
Abstract: This article presents a study of the effects of wing sweep on the aerodynamic performance of a blended wing body (BWB) aircraft that is based on an aerodynamically optimized design with a fixed planform and pitching moment constraint. Sixteen BWB geometries with varying wing sweep angles ranging from −40° (forward sweep) to 55° sweep were evaluated, while keeping the aerofoil profiles and twist distribution unchanged from the original optimized geometry. This gives some insight into the effects of one of the key planform design parameters.Numerical simulations were carried out using Euler solutions of the flow field with adaptive unstructured grids. Grid sensitivity studies were carried out, along with grid adaptation, on all geometries for solution accuracy. Results show that within 10° −55° sweep, there is significant variation in the lift-to-drag ratio (LID), whereas for sweep values from −40° to 10°, the L/D ratio remained relatively constant. To maintain the design lift for varying wing sweep...

44 citations


Journal ArticleDOI
TL;DR: In this paper, the aerodynamic characteristics of the Gurney flap were comprehensively investigated in terms of the performance requirements for a helicopter rotor by using two-dimensional Navier-Stokes equations.
Abstract: In the present study, the aerodynamic characteristics of the Gurney flap were comprehensively investigated in terms of the performance requirements for a helicopter rotor by using two-dimensional Navier-Stokes equations. To this end, with the rotor operating flow conditions in mind, the static aerodynamic characteristics of the Gurney flap are thoroughly compared with those of the clean airfoil at various Mach numbers, incidences, and Gurney flap heights. Next, to understand the general dynamic stall features of the airfoils with the Gurney flap, a series of parametric studies are performed with respect to oscillating frequency and amplitude, and the correlations of the aerodynamic coefficients are obtained by using the dynamic stall function of Bousman. It is concluded that there exists some optimum Gurney flap heights, which minimize the drag and pitching moment while maximizing the lift coefficient and lift-to-drag ratio at the same time. In the present study, the 2% Gurney flap would be a good compromise to satisfy all the major rotorcraft design criteria.

41 citations


Proceedings ArticleDOI
25 Jun 2007
TL;DR: In this article, coupled fluid-structure computations of the Navier-Stokes fluid flow and a flexible airfoil in low Reynolds number environments are conducted to probe the aerodynamic implications.
Abstract: The interaction between aerodynamics and structural flexibility in a low Reynolds number environment is of considerable interest to biological and micro air vehicles. In this study, coupled fluid-structure computations of the Navier-Stokes fluid flow and a flexible airfoil in low Reynolds number environments are conducted to probe the aerodynamic implications. While a flexible airfoil deforms in response to the aerodynamic loading, it exhibits an equivalent pitching motion, which modifies the effective angle of attack, causing noticeable differences in lift and thrust generation. Within the range of the flexibility considered, the flow fields are similar in all cases. Even at Re=100, in the plunging motion, the force acting on airfoil is dominated by pressure and the viscous force is of little impact on the overall lift and thrust generation. Detailed airfoil shape is secondary compared to the equivalent angle of attack.

38 citations


Journal ArticleDOI
TL;DR: In this paper, the authors proposed a linear fit coefficient for stall-onset incidences for oscillatory motion of airfoils oscillated by a single airfoil oscillator.
Abstract: cN = normal force coefficient, N=qc cp = pressure coefficient, p=q D = drag f = dimensionless separation point in terms of chord length, x=c L = lift of airfoil section M = Mach number m = pitching moment about the axis of the quarter-chord N = normal force q = dynamic pressure, 0:5 V r = reduced pitch rate, _ c=2V (for oscillatory motion of an airfoil, 0 cos!t) r0 = reduced pitch rate that delimits the dynamic stall and quasi-steady stall (normally, 0:01) s = nondimensional time, 2Vt=c T = time-delay constant for angle of attack t = time V = freestream velocity = angle of attack or incidence 0 = lagged angle of attack cr = critical onset angle (dependent on reduced pitch rate) ds = angle of dynamic-stall onset ds0 = constant critical onset angle ss = static stall angle = a step change in a sampled system 0 = amplitude of airfoil oscillation k = reduced frequency, !c=2V = linear fit coefficient for stall-onset incidences ! = frequency of oscillatory motion of airfoils Introduction

29 citations


Patent
13 Dec 2007
TL;DR: In this article, an airborne mobile platform (10) generally includes a plurality of rotating rotor blades operating in an airflow that forms a boundary layer on each of the rotor blades, and at least one of these rotor blades includes a section that encounters the airflow that includes an unsteady subsonic airflow having at least a varying angle of attack.
Abstract: An airborne mobile platform (10) generally includes a plurality of rotating rotor blades (12) operating in an airflow that forms a boundary layer on each of the rotor blades. At least one of the rotor blades includes a section that encounters the airflow that includes an unsteady subsonic airflow having at least a varying angle of attack. At least one of the rotor blades also includes one or more vortex generators (30) on the at least one of the rotor blades that generate a vortex that interacts with the boundary layer to at least delay an.onset of separation of the boundary layer, to increase a value of an unsteady maximum lift coefficient and to reduce a value of an unsteady pitching moment coefficient for the section.

Journal ArticleDOI
TL;DR: In this paper, the effect of Gurney flap sizes on NACA 4412 and NACA 0011 airfoils was evaluated using two dimensional steady state Navier-Stokes computations.
Abstract: Two dimensional steady state Navier-Stokes computations were performed to determine the effect of Gurney flap on NACA 4412 and NACA 0011 airfoils. Gurney flap sizes selected for the study range from 0.5% to 4% of the airfoil chord. A compressible Navier-Stokes solver with Baldwin-Lomax turbulence model, JUMBO2D, is used to predict the flow field around the airfoils. Computed results have been compared with available experimental and computational data. There is good correlation observed between computed and experimental data. Addition of Gurney flap increases the lift coefficient significantly with very little drag penalty if proper Gurney flap height is selected. Nose down pitching moment also increases with Gurney flap height. Flow field structure near the trailing edge shows very good resemblance with Liebeckx2019;s hypothesis that provides the possible explanation for the increased aerodynamic performance

Proceedings ArticleDOI
11 Jan 2007
TL;DR: A series of overset grids was generated in response to the 3rd AIAA CFD Drag Prediction Workshop (DPW-III) which preceded the 25th Applied Aerodynamics Conference in June 2006.
Abstract: A series of overset grids was generated in response to the 3rd AIAA CFD Drag Prediction Workshop (DPW-III) which preceded the 25th Applied Aerodynamics Conference in June 2006. DPW-III focused on accurate drag prediction for wing/body and wing-alone configurations. The grid series built for each configuration consists of a coarse, medium, fine, and extra-fine mesh. The medium mesh is first constructed using the current state of best practices for overset grid generation. The medium mesh is then coarsened and enhanced by applying a factor of 1.5 to each (I,J,K) dimension. The resulting set of parametrically equivalent grids increase in size by a factor of roughly 3.5 from one level to the next denser level. CFD simulations were performed on the overset grids using two different RANS flow solvers: CFL3D and OVERFLOW. The results were post-processed using Richardson extrapolation to approximate grid converged values of lift, drag, pitching moment, and angle-of-attack at the design condition. This technique appears to work well if the solution does not contain large regions of separated flow (similar to that seen n the DLR-F6 results) and appropriate grid densities are selected. The extra-fine grid data helped to establish asymptotic grid convergence for both the OVERFLOW FX2B wing/body results and the OVERFLOW DPW-W1/W2 wing-alone results. More CFL3D data is needed to establish grid convergence trends. The medium grid was utilized beyond the grid convergence study by running each configuration at several angles-of-attack so drag polars and lift/pitching moment curves could be evaluated. The alpha sweep results are used to compare data across configurations as well as across flow solvers. With the exception of the wing/body drag polar, the two codes compare well qualitatively showing consistent incremental trends and similar wing pressure comparisons.

Journal ArticleDOI
TL;DR: In this article, an energy-based stability analysis is performed to understand unsteady flow separation using high-accuracy compact schemes to solve the incompressible Navier-Stokes equation.
Abstract: Accelerated flow past a NACA 0015 aerofoil is investigated experimentally and computationally for Reynolds number Re =7968 at an angle of attack α =30◦. Experiments are conducted in a specially designed piston-driven water tunnel capable of producing free-stream velocity with different ramp-type accelerations, and the DPIV technique is used to measure the resulting flow field past the aerofoil. Computations are also performed for other published data on flow past an NACA 0015 aerofoil in the range 5200�Re �35 000, at different angles of attack. One of the motivations is to see if the salient features of the flow captured experimentally can be reproduced numerically. These computations to solve the incompressible Navier–Stokes equation are performed using high-accuracy compact schemes. Load and moment coefficient variations with time are obtained by solving the Poisson equation for the total pressure in the flow field. Results have also been analysed using the proper orthogonal decomposition technique to understand better the evolving vorticity field and its dependence on Reynolds number and angle of attack. An energy-based stability analysis is performed to understand unsteady flow separation.

Patent
05 Sep 2007
TL;DR: In this paper, the authors present methods and apparatus for reducing sinkage and wetted surface, and thus hydrodynamic drag of a high speed boat through the generation of aerodynamic lift while decreasing overall aerodynamic drag.
Abstract: The present invention provides methods and apparatus for reducing sinkage and wetted surface, and thus hydrodynamic drag of a high-speed boat through the generation of aerodynamic lift while decreasing overall aerodynamic drag. At least one lift-generating front wing proximate a bow section of the boat with at least one corresponding front air channel may be provided. At least one lift-generating rear wing proximate a transom section of the boat with at least one corresponding rear air channel may also be provided. At least one of the wings may be adjustable to generate aerodynamic lift with one of: (1) a neutral; (2) a transom-lifting; and (3) a bow-lifting pitching moment about a center of inertia of the boat. At least one wing proximate the leading edge of the tunnel of a multi-hull boat may be provided to increase the operational envelope.

Journal ArticleDOI
TL;DR: In this article, the effect of camber on the strength of the jet flow through the damage and delayed the onset of strong jet flows to higher incidences was investigated, and it was shown that in terms of damage flow characteristics and changes in lift, drag, and pitching moment coefficients, using a circular hole is a reasonable simulation of battle damage.
Abstract: Wind tunnel tests are reported that investigate three aspects of aerodynamic flows through battle damaged airfoils. The first aspect investigated was the effect of camber. This showed that reducing camber weakened the strength of the jet flow through the damage and delayed the onset of strong jet flows to higher incidences. The second investigation used five hole probe measurements to survey the flow field on a battle damaged flat plate airfoil. The measurements indicated that the use of the jet-to-freestream velocity ratio is a poor criteria for determining whether damage flows have undergone transition to strong jets. Finally, the influence of a star shaped hole to simulate more realistic battle damage was investigated. It was shown that in terms of damage flow characteristics and changes in lift, drag, and pitching moment coefficients, the use of a circular hole is a reasonable simulation of battle damage.

Proceedings ArticleDOI
23 Apr 2007
TL;DR: In this article, a robust aerodynamic airfoil design optimization using DFMOSS successfully showed the trade-off information between maximization and robustness improvement in aerodynamic performance by a single optimization run without careful input parameter tuning.
Abstract: A new optimization approach for robust design, design for multi-objective six sigma (DFMOSS) has been developed and applied to robust aerodynamic airfoil design for Mars exploratory airplane. The present robust aerodynamic airfoil design optimization using DFMOSS successfully showed the trade-off information between maximization and robustness improvement in aerodynamic performance by a single optimization run without careful input parameter tuning. The obtained trade-off information indicated that an airfoil with a smaller maximum camber improves robustness of lift to drag ratio, and that with a larger curvature near the shock wave location improves robustness of pitching moment against the variation of flight Mach number.

Journal ArticleDOI
TL;DR: Optimal 3-D range exceeds that predicted by 2-D models because, although angle of attack and lift are negative initially,3-D motion allows advantageous orientation of lift later in flight, with tilt of the axis of symmetry from vertical becoming much smaller at landing.

Proceedings ArticleDOI
25 Jun 2007
TL;DR: A wind tunnel test was carried out on an aspect ratio 6 wing equipped with Gurney flaps and trailing edge T-strips in this article, and the results showed that the wing achieved a positive increment in lift coefficient, a negative shift in the zero-lift angle of attack, and an increase in the wing maximum lift coefficient.
Abstract: A wind tunnel test was carried out on an aspect ratio 6 wing equipped with Gurney flaps and trailing edge T-strips. The test was conducted at the University of Washington Aeronautical Laboratory’s 8 x 12 foot low-speed wind tunnel at Reynolds numbers of 1.95x10, 1.02x10 and 0.51x10. The NACA 23012 test wing was unswept and untwisted with a 90 inch span and a constant chord of 15 inches. Gurney flap heights of 0.21%, 0.52%, 1.04%, 1.46%, 2.08%, 3.33%, 4.00% & 5.00% chord were tested on the model. T-strip heights of 0.42%, 1.04%, 1.67%, 2.08%, 2.92%, 4.17% & 5.00% chord were also tested. Results showed that Gurney flaps produced a positive increment in lift coefficient, a negative shift in the zero-lift angle of attack, and an increase in the wing maximum lift coefficient. Tstrips produced an increase in the slope of the lift curve and an increase in maximum lift coefficient, but produced no shift in the wing zero-lift angle of attack. Gurney flaps produced a negative (nose-down) shift in the pitching moment curve and a rearward shift in the wing aerodynamic center. T-strips also produced a rearward shift in the wing aerodynamic center, but produced no increment in the pitching moment coefficient near zero lift. Both devices produced a drag increment that was non-linear with device height, larger Gurney flaps and T-strips producing a disproportionately larger drag increment.

Proceedings ArticleDOI
08 Jan 2007
TL;DR: In this paper, a closed-loop pitch control on a moving 1-DOF wing model is investigated in wind tunnel experiments, where the model's attitude is controlled over a broad range of angles of attack when the baseline flow is fully attached using bi-directional pitching moment.
Abstract: Closed-loop pitch control on a moving 1-DOF wing model is investigated in wind tunnel experiments. The model's attitude is controlled over a broad range of angles of attack when the baseline flow is fully attached using bi-directional pitching moment that is effected by flow-controlled trapped vorticity concentrations on the pressure and suction surfaces near the trailing edge. In the present work, the model is trimmed using a position feedback loop and a servomotor actuator. Once the model is trimmed, the position feedback loop is opened and the servomotor acts like an inner loop control to alter the dynamic characteristics and to introduce disturbances. Position control of the model is achieved by the flow control actuation using an arbitrary reference model based adaptive outer loop controller. The control architecture employs a neural network based adaptive element that permits adaptation to both parametric uncertainty and unmodeled dynamics.

Journal ArticleDOI
TL;DR: In this paper, a quantitative model of the nonlinear motion leading to catastrophic yaw of finned missiles is presented, which is represented by pitch-yaw and roll equations of motion that include cubic aerodynamic coefficients and roll orientation-dependent induced moments.
Abstract: A quantitative model of the non-linear motion leading to catastrophic yaw of finned missiles is presented. The coupled roll-yaw dynamics of the missile, acted on by the trim angle of attack due to slight configurational asymmetries, is represented by pitch-yaw and roll equations of motion that include cubic aerodynamic coefficients and roll orientation-dependent induced moments. The steady-state equilibrium points of the system are found and their stability is determined by linearization. The solutions are evaluated by comparison with numerical integration of the equations of motion, proving the capability of the model to predict catastrophic yaw. Nomenclature CD = drag force coefficient CLα = angle of attack coefficient of the lift force CNα = angle of attack coefficient of the normal force CMα = angle of attack coefficient of the pitch moment Cli = induced roll moment coefficient Clp = roll damping moment coefficient Clδ = roll effectiveness moment coefficient CMpα = angle of attack coefficient of the Magnus moment

Journal ArticleDOI
TL;DR: In this paper, the authors investigated the work required by an airfoil to overcome the aerodynamic forces and produce a change in lift and showed that the relationship between this work and the total aerodynamic energy balance is shown to have significant consequences for transient changes in the shape of a wing.
Abstract: This paper investigates the resistance to a change in wing shape due to the aerodynamic forces. In particular, the work required by an airfoil to overcome the aerodynamic forces and produce a change in lift is examined. The relationship between this work and the total aerodynamic energy balance is shown to have significant consequences for transient changes in airfoil shape. Specification of the placement of the actuators and the actuator energetics is shown to be required for the determination of the airfoil shape change, requiring minimum energy input. A general simplified actuator model is adopted in this study, which assigns different values of actuator efficiency for negative and positive power output. Unsteady thin airfoil theory is used to analytically determine the pressure distribution and aerodynamic coefficients as a function of time for a ramp input of control deflection. This allows the required power and work to overcome the aerodynamic forces to be determined for a prescribed change in the airfoil camberline. The energy required for a pitching flat plate, conventional flap, conformal flap, and two variable camber configurations is investigated. For the pitching flat plate, the minimum energy pitching axis is shown to be dependent on the pitch rate and the initial angle of attack. The conformal flap is shown to require less actuator energy than the conventional flap to overcome the aerodynamic forces for a prescribed change in lift. The energy requirements of a variable camber configuration are shown to be sensitive to the layout of the variable camber device. The present analysis shows that the unsteady aerodynamic influence is important only for τ* values less than five. For τ* values larger than this, the present analysis reduces to the steady airfoil results of past studies.

Journal ArticleDOI
TL;DR: In this article, the effects of upward ramp rate, actuation start time, and duration of a moveable trailing-edge flap on the critical aerodynamic values of an oscillating wing were investigated.
Abstract: The effects of upward ramp rate, actuation start time, and duration of a moveable trailing-edge flap on the critical aerodynamic values of an oscillating wing were investigated. The largest improvement in the peak negative pitching moment was obtained with a fast ramp rate and a start time near the mean angle. The largest value of net work coefficient, however, was generated with a delayed start time and a slower ramp rate. The peak lift coefficient decreased with increasing ramp rate and decreasing start time. The inclusion of a short steady-state portion in the flap motion was beneficial. An upward flap deflection initiated slightly after the mean angle during pitch-up, moved to its maximum deflection at a moderate speed, remained steady for a rather short period, and then returned to its initial position, which spanned half the oscillation cycle, was found to provide a best compromise between the various aerodynamic requirements.

Proceedings ArticleDOI
08 Jan 2007
TL;DR: In this paper, the use of belly-flaps to enhance the lift and pitching moment during takeoff and landing of a blended-wing-body (BWB) was proposed.
Abstract: Since the beginning of flight with the Wright brothers, the shape of an airplane hasn’t really changed. An airplane has always been a fuselage with wings, a tail, engines and a landing gear. A couple of people, like the Horton brothers or Northrop, tried in the mid 30’s to completely change the airplane configuration with an innovative design of a tailless allwing airplane. However that concept was not consummated until the late 80’s when the B-2, the only flying wing to enter production to date, illustrated its benefits. Since then the airplane industry has attempted to adopt this design for future passenger airplanes because of its increased efficiency over a conventional tailed airplane. But there are some major challenges that must be resolved before one will see a Blended-Wing-Body (BWB) takeoff at a major International airport. One of the issues is its difficulty to rotate due to the missing tail. This paper presents a possible solution, namely the use of belly-flaps to enhance the lift and pitching moment during takeoff and landing. Wind tunnel tests on a generic BWBmodel have shown an increase up to 35% of the lift-off CL and an increase of 10% in pitching moment using belly-flaps mounted on the underside of the model.

Journal ArticleDOI
TL;DR: In this article, a rotatable tail mechanism was applied to a small unmanned aerial vehicle and measured force and moment coefficient measurements for the actual vehicle at a typical flight speed indicated that a rotable tail provided a sufficient yaw moment for turning.
Abstract: An experimental study of a rotatable tail mechanism applied to a small unmanned aerial vehicle was performed using a six-component wind-tunnel balance in the U.S. Air Force Institute of Technology low-speed wind tunnel. Attributes of the control and stability characteristics of the original vehicle, which were documented in an earlier study, are compared with those of a unique control methodology, a tail consisting of a single surface, with controllable elevation and rotation. An advantage of this change is a reduction in the storage length of the vehicle. Because there are similarities in the rotatable tail mechanism and the tail of many birds, the rotatable tail reflects a biomimetric feature. Measured force and moment coefficient measurements for the actual vehicle at a typical flight speed indicated that a rotatable tail provides a sufficient yaw moment for turning. For example, yaw moment coefficients C n , ranging from -0.02 to +0.02, which is typical for a rudder, were achievable as long as the absolute value of the tail elevation angle was large. The dependence of the yaw moment coefficient on the elevator angle and angle of attack, in addition to the tail rotation angle, indicates that there would be significant challenges in applying a robust flight control scheme with the current actuator configuration. An additional feature of the tail design is that by deflecting the tail upward, it could also function effectively as an air brake. A more than twofold increase in drag coefficient for constant angle of attack was measured when the tail elevation angle was increased to nearly 70 deg.

Proceedings ArticleDOI
TL;DR: In this paper, a 7-degree-of-freedom (DOF) dynamic model was developed to provide insight into the flight behavior of Type 200 and other related Lightcraft, and to serve as a research tool for developing future engine-vehicle configurations for laser launching of nano-satellites (1-10+ kg).
Abstract: A seven degree-of-freedom (DOF) dynamic model was developed to provide insight into the flight behavior of Type 200 and other related Lightcraft, and to serve as a research tool for developing future engine-vehicle configurations for laser launching of nano-satellites (1–10+ kg). Accurate engine, beam, and aerodynamics models are included to improve the predictive capability of the 7-DOF code. The aerodynamic forces of lift, drag, and aerodynamic pitching moment were derived from Fluent computational fluid dynamics predictions, and calibrated against limited existing wind tunnel data. To facilitate 7-DOF model validation, simulation results are compared with video analysis of actual flights under comparable conditions. Despite current limitatations of the 7-DOF model, the results compared well with experimental flight trajectory data.

Journal ArticleDOI
TL;DR: In this article, a variable droop leading edge (VDLE) airfoil was used to control compressible dynamic stall through management of its unsteady vorticity.
Abstract: This study reports control of compressible dynamic stall through management of its unsteady vorticity using a variable droop leading edge (VDLE) airfoil. Through dynamic adaptation of the airfoil edge incidence, the formation of a dynamic stall vortex was virtually eliminated for Mach numbers of up to 0.4. Consequently, the leading edge vorticity flux was redistributed enabling retention of the dynamic lift. Of even greater importance was the fact that the drag and pitching moment coefficients were reduced by nearly 50%. The camber variations introduced when the leading edge was drooped are explained to be the source of this benefit. Analysis of the peak vorticity flux levels allowed the determination of minimum necessary airfoil adaptation schedule.

Journal Article
TL;DR: In this paper, a solid obstruction was placed near the leading edge on the suction side of the wing in conjunction with a synthetic jet actuator that was placed farther inboard and was issued normal to the surface.
Abstract: The control effectiveness of active flow control, via arrays of synthetic jet actuators, on the aerodynamic performance of the Stingray UAV at low angles of attack was investigated experimentally in a wind tunnel. Global flow measurements were conducted, where the moments and forces on the vehicle were measured using a six component sting balance. The virtual shape modification technique, which was previously used on 2-D airfoils, was implemented in the present work on the 3-D Stingray configuration. A Solid obstruction was placed near the leading edge on the suction side of the wing in conjunction with a synthetic jet actuator that was placed farther inboard and was issued normal to the surface. Using this technique, the pitching moment was altered where the effect exceeded that of a 5 o deflection of the elevons. This suggests that synthetic-jet based flow control can be used for longitudinal trim control of the UAV during cruise condition, in lieu of the conventional control surfaces.

Proceedings ArticleDOI
25 Jun 2007
TL;DR: In this article, the authors used a novel control surface, a belly-flap, on the under side of the wing to enhance its lift and pitching moment coefficient during landing, go-around and takeoff.
Abstract: During the first century of flight few major changes have been made to the configuration of subsonic airplanes. A distinct fuselage with wings, a tail, engines and a landing gear persist as the dominant arrangement. During WWII some companies developed tailless all- wing airplanes. The concept failed to advance until 1953 when the British Avro Vulcan bomber appeared. This airplane matched, but did not exceed, the aero performance of it's conventional contemporary, the Boeing B-47. In the late 80's the B-2, became the only flying-wing to have entered major production in recent memory. It proved the benefits of all-wing designs, at least for a stealth platform. The advent of the Blended-Wing-Body addresses the historical shortcomings of all-wing designs, specifically poor volume utility, and excess wetted area as a result. The BWB is now poised to become the new paradigm for large subsonic airplanes. Major aerospace companies are studying the concept for next generation passenger airplanes. But there are still challenges. One is the Blended-Wing- Body's short control lever-arm in pitch. This affects rotation and go-around performance. This paper presents a possible solution by using a novel type of control surface, a belly-flap, on the under side of the wing to enhance its lift- and pitching moment coefficient during landing, go-around and takeoff. Increases of up to 30% in lift-off CL and 8% in positive pitching moment have been achieved during wind tunnel tests on a generic BWB-model with a belly-flap.