Showing papers on "Wing root published in 1992"

Application of adaptive technology to static aeroelastic control of wing structures

Abstract: The static aeroelastic behavior of adaptive swept-forward wing structures modeled as thin-walled beams and incorporating piezoelectric effects is investigated. Based on the converse piezoelectric effect, the system of piezoelectric layers, embedded or bonded to the wing, yields control of both divergence instability and, in the subcritical speed range, of aeroelastic lift distribution.

Effect of canard deflection on close-coupled canard-wing-body aerodynamics

Abstract: The thin-layer Navier-Stokes equations are solved for the flow about a canard-wing-body configuration at transonic Mach numbers of 0.85 and 0.90, angles of attack from -4 to 10 degrees and canard deflection angles from -10 to +10 degrees. Effects of canard deflection on aerodynamic performance, including canard-wing vortex interaction, are investigated. Comparisons with experimental measurements of surface pressures, lift, drag and pitching moments are made to verify the accuracy of the computations. The results of the study show that the deflected canard downwash not only influences the formation of the wing leading-edge vortex, but can cause the formation of an unfavorable vortex on the wing lower surface as well.

Navier-Stokes simulation of a close-coupled canard-wing-body configuration

Abstract: The thin-layer Navier-Stokes equations are solved for the flow about a coplanar close-coupled canard-wing-body configuration at a transonic Mach number of 0.90 and at angles of attack ranging from 0 to 12 degrees. The influence of the canard on the wing flowfield, including canard-wing vortex interaction and wing vortex breakdown, is investigated. A study of canard downwash and canard leading-edge vortex effects, which are the primary mechanisms of the canard-wing interaction, is emphasized. Comparisons between the computations and experimental measurements of surface pressure coefficients, lift, drag and pitching moment data are favorable. A grid refinement study for configurations with and without canard shows that accurate results are obtained using a refined grid for angles of attack where vortex burst is present. At an angle of attack of approximately 12 deg, favorable canard-wing interaction which delays wing vortex breakdown is indicated by the computations and is in good agreement with experimental findings.

Effect of Tunnel Walls on Vortex Breakdown Location over Delta Wings

Abstract: Vortex breakdown has been the subject of many investigations during the past few decades. Many of the investigations were performed by visualizing the vortical flowfield above delta wings in water or wind tunnels. In spite of the extensive use of this technique, little attention has been paid to the possible influence of the test section walls on the measured location of the vortex breakdown. The present work suggests a possible model by which the walls may affect the vortex breakdown location. The suggested model is associated with the upwash induced on the wing surface due to the presence of the walls. This upwash was found to be relatively small near the wing's apex and larger on the trailing edge, thus creating an effectively cambered wing. The effective camber tends to shift the vortex breakdown location downstream as compared to a flat wing with the same projected geometry. The influence of the walls was tested in a series of experiments in a water tunnel using delta wings with different sizes, relative to the test section dimensions. The anticipated trend was observed in the experimental results. Nomenclature Cr = wing root chord length H = test section height V = undisturbed velocity S =wing span W = test section width Xbd - vortex breakdown location along the wing root chord a = angle of attack ALE = leading-edge sweep angle

Wing mass formula for twin fuselage aircraft

Abstract: A formula is derived to calculate structural wing mass. This formula can be applied to twin fuselage aircraft, conventional single-body aircraft and some other unconventional aircraft (such as the Voyager). The approach is particularly useful in the first stages of preliminary aircraft design and in optimization programs where the wing-mass calculation time is an important characteristic. The concept model assumes a nontapered inboard wing section, a tapered outboard wing section and fuel stored only in the outboard wing. The theory for the wing-mass estimation is described. Unlike the other mass formulae where mass spanwise distribution is considered by an "unloading coefficient," the present method integrates the mass spanwise distribution with the air load spanwise distribution. This allows more precise consideration of the wing geometry and mass unloading. There are no simplifications applied and the formula completely reflects the initial concept model. Good comparison with statistical data for single body aircraft is obtained.

Numerical investigation of performance degradation of wings and rotors due to icing

Abstract: The aerodynamic load characteristics and the performance degradation of moderate aspect ratio wings and rotors with simulated glaze leading-edge ice have been studied using a three-dimensional, compressible Navier-Stokes solver. The effect of a splitter plate at the wing root on both clean and iced wing configurations has been studied and the results are compared with the experiment. A significant difference has been observed with and without splitter plates in the magnitude of flow separation and aerodynamic loading at the inboard stations for the iced wing at 8-deg angle of attack. Inviscid calculations were performed and compared with viscous calculations to investigate whether the performance of iced swept wings can be inexpensively predicted using Euler methods. It is shown that inviscid calculations predict higher aerodynamic loading than viscous calculations, and cannot model separation effects. A typical nonlifting helicopter rotor in forward flight condition is also studied, and the penalty due to the leading-edge ice formation on the required torque is numerically demonstrated.

Wing mass formula for subsonic aircraft

Abstract: lytical Methods Inc., Washington, DC, 1982. Lan, C. E., and S. C. Mehrotra, \"An Improved Woodward's Panel Method for Calculating Leading-Edge and Side-Edge Suction Forces at Subsonic and Supersonic Speeds,\" NASA CR 3205, 1979. Hardy, B. C., and S. P. Fiddes, \"Prediction of Vortex Lift on Non-Planar Wings by the Leading-Edge Suction Analogy,\" Aeronautical Journal, Vol. 92, No. 914, 1988, 154-164.

Wing root aerofoil for forward swept wings

Abstract: A swept forward wing for aircraft comprising an inner wing portion and an outer wing portion in which the upper surface curvature of the inner wing portion is designed to create three dimensional flow thereover to manipulate the sweep of the isobars and prevent desweeping thereof, the inner wing portion including a wing root section (4) having a far aft maximum thickness position (14) coupled with high camber in the region of said maximum thickness position, said wing root section (4) further including a negatively cambered leading edge portion (10) and a nose-down twist configured to suppress high leading edge velocities, the combination of thickness and camber forms aft of the leading edge region (10) causing the flow to accelerate until a maximum velocity is reached relatively far back on the wing.

Aeroelastic effects of spoiler surfaces on a low-aspect-ratio rectangular wing

Abstract: An experimental research study to determine the effectiveness of spoiler surfaces in suppressing flutter onset for a low-aspect-ratio, rectangular wing has been conducted in the Langley Transonic Dynamics Tunnel (TDT). The wing model used in this flutter test consisted of a rigid wing mounted to the wind-tunnel wall by a flexible, rectangular beam. The flexible beam was connected to the wing root and cantilever mounted to the wind-tunnel wall. The wing had a 1.5 aspect ratio based on wing semispan and a NACA 64AGIO airfoil shape. The spoiler surfaces consisted of thin, rectangular aluminum plates that were vertically mounted to the wing surface. The spoiler surface geometry and location on the wing surface were varied to determine the effects of these parameters on the classical flutter of the wing model. Subsonically, the experiment showed that spoiler surfaces increased the flutter dynamic pressure with each successive increase in spoiler height or width. This subsonic increase in flutter dynamic pressure was approximately 15% for the maximum height spoiler configuration and for the maximum width spoiler configuration. At transonic Mach numbers, the flutter dynamic pressure conditions were increased even more substantially than at subsonic Mach numbers for some of the smaller spoiler surfaces. But for larger spoiler sizes (in terms of either height or width) the spoilers forced a torsional instability in the transonic regime that was highly Mach number dependent. This detrimental torsional instability was found at dynamic pressures well below the expected flutter conditions. Variations in the span wise location of the spoiler surfaces on the wing showed little effect on flutter. Flutter analysis was conducted for the basic configuration (clean wing with all spoiler surface mass properties included). The analysis correlated well with the clean wing experimental flutter results.

Wing design for hanggliders having minimum induced drag

Abstract: Introduction T HERE are lifting line design methods for rigid wings having minimum induced drags with practical constraints." Recently, Wohlfahrt and Nickel have analyzed hanggliders with large aspect ratios and small sweep-back angles by using the lifting line theory. It is of practical interest to look for the optimum wing design taking account of a bending moment at the wing root, because the leading edge spars are the most important structural members of hanggliders. This note presents the optimum solutions.

Effect of canard oscillations on the vortical flowfield of a X-31A-like fighter aircraft model

Abstract: : A flow visualization investigation was carried out in the Naval Postgraduate School water tunnel using dye injection technique to study the effects of oscillating a close-coupled canard on a 2.3% scale modal of a X-3 1A- Us fighter This investigation focussed primarily on the effects of canard oscillations on the breakdown characteristics of the wing root vortex for both static and dynamic conditions of the model at zero sideslip angle. The main results of this first of a kind water tunnel visualization that for the static conditions of the model the low frequency canard oscillations tend to destabilize/augment wing vortex core, ie., promote/delay bursting of wing vortex. The dynamic tests indicate that the large amplitude low frequency canard oscillations of the canard interact favorably with the wing vortical flowfield to delay vortex bursting during both pitch-up and pitch-down motions. High angle of attack aerodynamic, effect of pitch rate and canard oscillation s vortex breakdown, flow visualization by dye injection, X-31A-like fighter aircraft model.

Helicopter rotor blade

Abstract: PURPOSE:To suppress noise and vibration by optionally changing the lower camber angle of a helicopter rotor blade during flight to avoid or reduce the interference of the top end vortex of the rotor blade with the following rotor blade. CONSTITUTION:A rotatable shaft disposed in the chord direction on a rotor blade body, a top end wing part 4 whose wing root part is mounted on the rotatable shaft, and a driving device 5 for rotating the rotatable shaft are provided.

Multi-engine airplane

Abstract: PURPOSE:To improve the safety against an engine fire and reduce the noise to a cabin by mounting one engine each at wing root sections of the right and left horizontal tail wings, and mounting one engine at the lower section of a fuselage so that it can be discarded at the time of a fire. CONSTITUTION:One engine 3 is mounted each at wing root sections of the right and left horizontal tail wings 5, two in all, and one engine 3a is mounted via a pylon 4 at the lower section of a fuselage 2 so that it can be discarded at the time of a fire. The engine 3a is used as an option, the engines 3 are normally used, noises hardly reach a passenger compartment, and it is quiet. Even if a fire breaks out on the engines 3, they are sufficiently separated from the passenger compartment, passengers have a sufficient margin for the evacuation time, flames and smoke are discharge backward and hardly enter the passenger compartment. The engine 3a is located on the airframe axis and can be immediately switched for use when a trouble occurs on the engines 3, and it can be discarded when a fire breaks out, thus the safety is improved.

System for Generating Sequences of Phased Gust or Taxi Loadings

Abstract: A method is presented for Monte Carlo generation of phased load conditions for multiple loading durability testing. These load conditions are consistent with the gust and taxi environmental-dynamic model used in aircraft design. Time histories of individual loads from sequences of load conditions are consistent with the load sequences used in standard single load point durability and damage tolerance testing and analysis. Truncation can be applied to a sequence of load conditions so that only the most severe are retained. The truncation process is consistent with the mission analysis approach to design and to durability and damage tolerance. The load conditions can be modified using an induced autocorrelation approach to give variation in the number of zero crossings of the load time histories. OAD sequences generated for durability and damage tolerance analysis are usually applied at a single point. However, the stresses in a component of structure often arise from combinations of multiple loads. Selection of a single application point and direction is often a difficult task. This is particularly true for aircraft structures flying in turbulence. In that case, loads arise from complex combinations of ex- ternal and internal forces. At McDonnell Douglas the full-scale durability test of the C-17 aircraft requires the simultaneous application of test loads which are properly phased by some rational definition of phasing. The following criteria was developed for satisfying this requirement: 1) Time histories for each individual load (e.g., wing root normal bending) must be consistent with load sequences used in durability and damage tolerance testing and analysis. 2) The phasing of the simultaneously applied loads must be consistent with the dynamic gust and taxi models used to design the aircraft. A method is presented here that meets this criteria. The method generates load conditions, that is, multiple loads for simultaneous application at several points on an aircraft struc- ture. The multiple loads are correlated in conformance with the mathematical model of the gust environment and the structure. In this model the magnitude, frequency response, and correlation of loads are based on a flexible linear aircraft response to a forcing function having the characteristics of a stationary Gaussian random process (see Fig. 1). The specifics of the mathematical model of an aircraft flying in turbulence have been summarized by Hoblit in Ref. 1. This method as implemented at Douglas Aircraft Corpo- ration is called the Phased Loading Sequence Generator Sys- tem (PLSGS). This system will be used to generate gust and taxi loadings for the C-17 full-scale durability test. For the C-17 application the taxi forcing function (the runway) was treated as a stationary Gaussian random process, and the