Swept Wing Research Articles

Aircraft suitability model applied to a laminar flow control flight experiment

When planning a flight experiment, the test team must analyse the suitability of the testbed aircraft. Often, critical design decisions for the test apparatus are made without high-fidelity models of aircraft performance since few testbed aircraft are designed and built for the flight experiment. This study develops a process to adapt rough performance data to determine the suitability of an airplane to operate at a given angle of attack across a wide range of flight conditions within operational limits. The process illustrates the careful balance of factors in the design of a swept wing laminar flow control wing glove. The process also establishes an important flight planning tool that allows the test team to assess the impact of aircraft performance on the flight experiment. Further, the tool can be updated with flight data when available and leveraged as a planning tool to promote safe and efficient flight research.

DESIGN AND ANALYSIS OF AN STATIC AEROELASTIC EXPERIMENT

Static aeroelastic experiments are very common in the United States and Russia. The objective of static aeroelastic experiments is to investigate deformation and loads of elastic structure in flow field. Generally speaking, prerequisite of this experiment is that the stiffness distribution of structure is known. This paper describes a method for designing experimental models, in the case where the stiffness distribution and boundary condition of a real aircraft are both uncertain. The stiffness distribution form of the structure can be calculated via finite element modeling and simulation calculation and F141 steels and rigid foam are used to make elastic model. In this paper, the design and manufacturing process of static aeroelastic models is presented and a set of experiment model was designed to simulate the stiffness of the designed wings, a set of experiments was designed to check the results. The test results show that the experimental method can effectively complete the design work of elastic model. This paper introduces the whole process of the static aeroelastic experiment, and the experimental results are analyzed. This paper developed a static aeroelasticity experiment technique and established an experiment model targeting at the swept wing of a certain kind of large aspect ratio aircraft.

Aerodynamic parametric analysis of an unconventional joined-wing aircraft configuration

In this paper, the experimental results of an unconventional joined-wing aircraft configuration are presented. The test model uses two different wings, forward and rear, both joined in tandem and forming diamond shapes both in plant and front views. The wings are joined in such a way that it is possible to change the rear wing dihedral angle values and the rear wing sweep angle values in 25 different positions that modify the relative distance and the relative height between the wings. To measure the system aerodynamic coefficients it is necessary to perform wind tunnel tests. The data presented corresponds to the lift, drag and induced drag aerodynamic coefficients, as well as the aerodynamic efficiency and the parameter for minimum required power, from the calculated values of the lift and drag time series measured by a 6-axis force and torque sensor. The results show the influence on the aerodynamic coefficients of the rear wing sweep and dihedral angles parameters. As a main result, it can be concluded that, in general terms, the lift and induced drag aerodynamic coefficients values decrease as both the distance and height between the wings increase, on the other hand, the total drag aerodynamic coefficient decreases if the distance between the wings increases, but nevertheless shows a slight tendency to increase if the height of the rear wing increases, whereas the aerodynamic efficiency and the parameter for minimum required power increase if the distance between the wings increases.

Wing rock motion induced by forebody asymmetric vortices in pitch-up

Aiming to the issue that the angle of attack was usually fixed rather than pitched-up in the previous researches about the wing rock induced by forebody asymmetric vortices,series of wind tunnel experiments were conducted in Beihang University D4 wind tunnel over the configuration of a pointed ogive-cylindrical body with 30° swept wings. Effect of pitch rate on wing rock motion as well as the mechanism of that was first studied through wing rock experiments in different pitch rates. The flow mechanism of wing rock in high rate pitch-up was then studied through the surface pressure measurements during the dynamic wing rock.Experimental results show that,as the wing rock motion time reduces with increase of pitch rate,the motion patterns of wing rock are different in three different intervals of pitch rate. The sinusoidal-like motion,which is totally different from the motion at static angle of attack,would occur in the third interval with high rate pitch-up. The variation of forebody flow with angle of attack,instead of the variation of forebody flow with roll angle,is found to be responsible for the sinusoidal-like motion in high rate pitch-up. Hence,the flow mechanism of wing rock in pitch-up is greatly different from that at static angle of attack.

The Definition of the Load Spectrum for SU-22 Fighter-Bomber Full Scale Fatigue Test

Abstract The Su-22 fighter-bomber is a military aircraft used in the Polish Air Force since the mid 1980’s. By the decision of the Polish Ministry of Defense the predicted service life for this type of aircraft will be extended to 3200 flight hours. Due to the fact that some aircraft were nearing the end of the service life guaranteed by the manufacturer, the actual service life, determined based on the flight profile in the Polish Air Force, had to be validated. Consequently, the Full Scale Fatigue Test (FSFT) had to be carried out in order to verify that the required service life was attainable. This article describes the process of preparation of the load spectra used in the Su-22 FSFT. Due to the fact that the Su-22 has a variable sweep wing the whole test was divided into three Stages (landing, flight and flap loads) carried out at different wing sweep angles (30°/45°/30°). The spectra were developed using the historical data gathered from Flight Data Recorders (FDR), strain signals acquired during the Operational Load Monitoring program (OLM) and aerodynamic calculations.

Effect of chevrons on the slat noise of straight and swept wings

An experimental study of the airframe noise for small-scale wing models with high-lift devices (slat and flap) is performed. It is shown that installation of chevrons on the lower edge of a slat leads to noise reduction on both straight and swept wings. Simultaneous acoustic and aerodynamic measurements show that chevrons lead to suppression of the slat tonal noise components without significantly affecting the wing aerodynamics.

An efficient method for predicting zero-lift or boundary-layer drag including aeroelastic effects for the design environment

AbstractOne of the aerospace design engineer’s goals aims to reduce drag for increased aircraft performance, in terms of range, endurance, or speed in the various flight regimes. To accomplish this, the designer must have rapid and accurate techniques for computing drag. At subsonic Mach numbers drag is primarily a sum of lift-induced drag and zero-lift drag. While lift-induced drag is easily and efficiently determined by a far field method, using the Trefftz plane analysis, the same cannot be said of zero-lift drag. Zero-lift drag (CD,0) usually requires consideration of the Navier-Stokes equations, the solution of which is as yet unknown except by using approximate numerical techniques with computational fluid dynamics (CFD). The approximate calculation of zero-lift drag from CFD is normally computed with so-called near-field techniques, which can be inaccurate and too time consuming for consideration in the design environment. This paper presents a technique to calculate zero-lift and boundary-layer drag in the subsonic regime that includes aeroelastic effects and is suitable for the design environment. The technique loosely couples a two-dimensional aerofoil boundary-layer model with a 3D aeroelastic solver to compute zero-lift drag. We show results for a rectangular wing (baseline), a swept wing, and a tapered wing. Then compare with a rectangular wing with variable thickness and camber, thinning out from the root to tip (spanwise direction), thus demonstrating the practicality of the technique and its utility for rapid conceptual design.

A systematic method of smooth switching LPV controllers design for a morphing aircraft

This paper is concerned with a systematic method of smooth switching linear parameter-varying (LPV) controllers design for a morphing aircraft with a variable wing sweep angle. The morphing aircraft is modeled as an LPV system, whose scheduling parameter is the variation rate of the wing sweep angle. By dividing the scheduling parameter set into subsets with overlaps, output feedback controllers which consider smooth switching are designed and the controllers in overlapped subsets are interpolated from two adjacent subsets. A switching law without constraint on the average dwell time is obtained which makes the conclusion less conservative. Furthermore, a systematic algorithm is developed to improve the efficiency of the controllers design process. The parameter set is divided into the fewest subsets on the premise that the closed-loop system has a desired performance. Simulation results demonstrate the effectiveness of this approach.

Extension to the Myers Model for Calculation of Three-Dimensional Glaze Icing

To simulate three-dimensional glaze ice accretion on a typical aircraft surface, the model of water film flow and ice accretion on a curved surface based on two coupled partial differential equations proposed by Myers et al. is extended in nonorthogonal curvilinear coordinates. Then, the governing equation for water film flow coupling with a Stefan model for glaze icing is built up. An efficient implicit–explicit numerical method is proposed for solving the governing equations. The glaze ice computations on two-dimensional airfoils and three-dimensional typical geometries in in-flight icing conditions are completely accomplished. Validation results of two-dimensional calculation cases with experimental ice shapes are presented. Three-dimensional simulations of glaze ice accretion on a sphere and a swept wing are successfully carried out and compared with ice profiles from icing tunnel tests and simulations based on the classical Messinger model.

Parametric Exploration of Wing–Body Junction Flow Using Computational Fluid Dynamics

The wing–body junction flow is investigated parametrically using computational fluid dynamics in an attempt to understand the effects of junction flow on aircraft drag, with a focus on application to large business jet or commercial transport aircraft. A Reynolds-averaged Navier–Stokes computational fluid dynamics methodology is validated against detailed experimental data for a junction flow. Computational fluid dynamics results for a wing with a leading-edge strake (an aerodynamic surface designed to reduce flow separation, thereby reducing aircraft drag) are presented, and the effects of scaling this strake are explored. The effectiveness of the strake on a swept wing is compared to the same for a straight wing. Finally, the results from this parametric study are successfully applied to sizing a leading-edge strake for a commercial transport aircraft. It is demonstrated that a systematic approach, starting with a simple, validated model and building up to a realistic aircraft application, can build confidence in computational fluid dynamics results.

A 3D mesh deformation technique for irregular in‐flight ice accretion

SummaryAircraft holding around busy airports may be requested to sustain as much as 45 min of icing before landing or being diverted to another airport. In this paper, a three‐dimensional mesh deformation scheme, based on a structural frame analogy, is proposed for the numerical simulation of ice accretion during extended exposure to adverse weather conditions. The goal is to provide an approach that is robust and efficient enough to delay or altogether avoid re‐meshing while preserving (enforcing) nearly orthogonal elements at the highly distorted ice surface. Robustness is achieved by suitably modifying the axial and torsional stiffness components of the frame elements in order to handle large and irregular grid displacements typical of in‐flight icing. Computational efficiency is obtained by applying the mesh displacement to an automatically selected small subset of the entire computational domain. The methodology is validated first in the case of deformations typical of fluid‐structure interaction problems, including wing bending, a helicopter rotor in forward flight, and the twisting of a high‐lift wing configuration. The approach is then assessed for aero‐icing on two swept wings and compared against experimental measurements where available. Copyright © 2015 John Wiley & Sons, Ltd.

The effect of sweep angle on the limit cycle oscillations of aircraft wings

This study focuses on the limit cycle oscillations (LCOs) of cantilever swept-back wings containing a cubic nonlinearity in an incompressible flow. The governing aeroelastic equations of two degrees-of-freedom swept wings are derived through applying the strip theory and unsteady aerodynamics. In order to apply strip theory, mode shapes of the cantilever beam are used. The harmonic balance method is used to calculate the frequencies of LCOs. Linear flutter analysis is conducted for several values of sweep angles to obtain the flutter boundaries.

Experiments on the Artificial Disturbance Evolution in 2D and 3D Spanwise Modulated Boundary Layers at Mach 2 and 2.5

Experimental data on the nonlinear wave train development in 3D supersonic boundary layer over swept wing and experimental investigation of wave train development in spanwise modulated boundary layer on flat plate and swept wing at Mach number 2 and 2.5 are presented. Artificial disturbances in the boundary layer were excited by periodical glow discharge mainly at 10 and 20kHz. It was found, that some disturbance excitation downstream cannot be explained by linear stability theory. It was obtained that receptivity of supersonic boundary layer to the surface excited controlled pulsations and the wave train development essentially depends on the Mach number.

Non-Parallel-Flow Effects on Stationary Crossflow Vortices at Their Genesis

The linear stability of a laminar boundary-layer flow over a swept wing is studied for a stationary perturbation with spanwise wavenumber ▪, where β is non-dimensionalised by the chordwise distance L to the leading edge, and Re is the Reynolds number based on L This scaling corresponds to the lower-branch regime of the neutral curve, and hence represents the genesis stage of stationary crossflow vortices. We show that in this stage or regime, the non-parallelism of the flow plays a leading order role in the growth of the instability. Non-parallel-flow effects consist of two parts: the streamwise variation of the base-flow profiles, and the distortion of the eigenfunction. These two contributions to the growth rate are found to be of the same order of magnitude, and both are closely related to the 4th-order vertical derivatives of the base-flow profile on the surface of the wing. The implication of the present finding on the description of stationary vortices using the Orr-Sommerfeld equation is discussed.

Weight Prediction of the Lifting System for an Unconventional Aircraft Configuration

In this paper a tool, previously developed and validated for the prediction of wing box structural weight in very early design stages, is improved to better match effects on panel sizing of stability constraints. It is applied to an unconventional configuration aircraft based on the Best Wing System concept, introduced by Ludwig Prandtl in 1924, to achieve minimum induced drag: such lifting system is composed of two swept wings (fore and aft) connected by vertical wings at their tips and two fins connecting the rear wing to the fuselage. The system is over-constrained to the fuselage and, thus, the structural design, as well as the static aeroelasticity and flutter characteristics, totally differs from a conventional aircraft. An optimization method finds out a suitable prediction of the structural weight by defining the wing span stiffness behavior compatible with a mix of global and local design constraints, but without any claim about a structural design. The link among stiffness properties and structural weight is made by means of simplified models of the box cross-sections, suitable also to supply responses for the approximated evaluation of local constraints.

Sensitivity Analysis of Transonic Flow over J-78 Wings

3D transonic flow over swept and unswept wings with an J-78 airfoil at spanwise sections is studied numerically at negative and vanishing angles of attack. Solutions of the unsteady Reynolds-averaged Navier-Stokes equations are obtained with a finite-volume solver on unstructured meshes. The numerical simulation shows that adverse Mach numbers, at which the lift coefficient is highly sensitive to small perturbations, are larger than those obtained earlier for 2D flow. Due to the larger Mach numbers, there is an onset of self-exciting oscillations of shock waves on the wings. The swept wing exhibits a higher sensitivity to variations of the Mach number than the unswept one.

Effect of Freestream Turbulence on Roughness-induced Crossflow Instability

The effect of freestream turbulence on generation of crossflow disturbances over swept wings is investigated through direct nu- merical simulations. The set up follows the experiments performed by Downs et al. (2012). In these experiments the authors use ASU(67)-0315 wing geometry which promotes growth of crossflow disturbances. Distributed roughness elements are locally placed near the leading edge with a given spanwise wavenumber to excite the corresponding stationary crossflow vortices. In present study, we partially reproduce the isotropic homogenous freestream turbulence through direct numerical simulations using freestream spectrum data from the experiments. The generated freestream fields are then applied as the inflow boundary condition for direct numerical simulation of the wing. The distributed roughness elements are modelled through wing surface deformation and placed near the leading edge to trigger the stationary crossflow disturbances. The effects of the generated freestream turbulence on the initial amplitudes and growth of the boundary layer perturbations are then studied.

Greener Aviation Take-off (Delayed)

In the past fifty years, long-range commercial airliners have changed only incrementally from the paradigmatic design – a tube fuselage with swept wings and mostly-aluminium construction. Reducing the environmental impact of airliners may require radical innovations and a new paradigm, but the transition to a new paradigm is fraught with risks. This paper analyses how key risks have shaped and limited efforts to transition toward three types of radical innovations that would signifi cantly improve airliner fuel effi ciency. We use these three cases to reassess the dominant framework for analysing sociotechnical transitions – the multi-level perspective (MLP) – in light of methods and theoretical perspectives drawn from Science and Technology Studies (STS). We argue that if the MLP is to provide a robust framework for analysing sociotechnical transitions, it must be refi ned in three ways. First, it must ‘open the black box’ to account for the ways that technologically-specific risks shape the transition process. Second, rather than predefi ning particular innovations as radical or conservative, ‘mature’ or ‘immature,’ it should attend to how actors conceive of such terms; an innovation which appears ‘mature’ to one group may appear ‘immature’ to another. Third, the MLP would be strengthened by additional case studies such as ours, which examine incomplete or failed transitions.

Effects of angle of attack on wing rock motion induced by the flows over slender body with low swept wing

The patterns of wing rock motion at 52.5° angle of attack have already been investigated in detail (Rong, 2009; Wang, 2010). These patterns are completely different from those at other angles of attack. This phenomenon indicates that angle of attack affects wing rock motion. The present study aims to examine the different patterns of wing rock motion at different angles of attack. The flow mechanisms of the motion patterns are also revealed, especially the uncommanded lateral motions, including wing rock and lateral deflection, induced by regular asymmetric separated flow from wings at low angles of attack and forebody asymmetric vortices at angles of attack of 27.5°⩽ α ⩽ 70°. The test conditions, including the testing Reynolds number, wind tunnel, experimental techniques, and test model, are all the same as those used in a previous study at α = 52.5°. Finally, the experimental technique of rotating nose of the model to suppress the wing rock or lateral deflection, which is induced by forebody asymmetric vortex flow, is applied. The uncommanded lateral motions are successfully suppressed by this technique.

Targeted Energy Transfer Between a Swept Wing and Winglet-Housed Nonlinear Energy Sink

A prototype nonlinear energy sink, whose design is based upon parameters determined to effectively suppress transonic aeroelastic instabilities of a wind-tunnel wing model (denoted as generic transport wing) via passive targeted energy transfer, is introduced. The lightweight nonlinear energy sink is mounted within a low-profile winglet, which is attached to the tip of the generic transport wing. In addition, safety features and measurement hardware have been built in to the prototype to facilitate a future transonic wind-tunnel experiment. The effects of the nonlinear energy sink on the structural dynamics of the model wing are demonstrated in ground vibration tests, both experimental and computational. Results show that the nonlinear energy sink causes a significant increase in the dissipation rate of energy in the second bending mode of the generic transport wing, even for small wing-tip oscillations. This is a strong indication that the prototype nonlinear energy sink will be effective in wind-tunnel tests because the frequency of the second bending mode is within the range of experimental and computational flutter frequencies of the generic transport wing. Furthermore, accompanying computational analysis is used to show that moderate friction damping in the nonlinear energy sink is unlikely to have a significant effect on the qualitative interaction between the nonlinear energy sink and the generic transport wing.