Showing papers in "Journal of Aircraft in 1990"

Thin-walled composite beams under bending, torsional, and extensional loads

Abstract: Symmetric and antisymmetric layup graphite-epoxy composite beams with thin-walled rectangular cross sections are fabricated using an autoclave molding technique and tested under bending, torsional, and extensional loads. The bending slope and elastic twist at a station are measured using an optical system, and the results correlated with predicted values from a simple beam analysis as well as a refined finite element analysis. A symmetric ply layup results in bending-twist coupling whereas an antisymmetric layup causes extension-twist coupling. Simple analytical results with plane-stress assumption agree better with measured data as well as finite element predictions than with plane-strain assumption. For symmetric layup beams, the bending-induced twist and torsion-induced bending slope are predicted satisfactorily using simple analytical solution. Correlations with measured data, however, are generally improved using a finite element solution. For antisymmetric beams, axial force-induced twist is predicted satisfactorily by both methods.

ASTROS - A multidisciplinary automated structural design tool

Abstract: ASTROS (Automated Structural Optimization System) is a finite-element-based multidisciplinary structural optimization procedure developed under Air Force sponsorship to perform automated preliminary structural design. The design task is the determination of the structural sizes that provide an optimal structure while satisfying numerous constraints from many disciplines. In addition to its automated design features, ASTROS provides a general transient and frequency response capability, as well as a special feature to perform a transient analysis of a vehicle subjected to a nuclear blast. The motivation for the development of a single multidisciplinary design tool is that such a tool can provide improved structural designs in less time than is currently needed. The role of such a tool is even more apparent as modern materials come into widespread use. Balancing conflicting requirements for the structure's strength and stiffness while exploiting the benefits of material anisotropy is perhaps an impossible task without assistance from an automated design tool. Finally, the use of a single tool can bring the design task into better focus among design team members, thereby improving their insight into the overall task.

Sensitivity analysis and multidisciplinary optimization for aircraft design - Recent advances and results

Abstract: Optimization by decomposition, complex system sensitivity analysis, and a rapid growth of disciplinary sensitivity analysis are some of the recent developments that hold promise of a quantum jump in the support engineers receive from computers in the quantitative aspects of design Review of the salient points of these techniques is given and illustrated by examples from aircraft design as a process that combines the best of human intellect and computer power to manipulate data

Vortex Dynamics on a Pitching Delta Wing

Abstract: A study of the dynamic behavior of the leading-edge vortices on a delta wing undergoing oscillatory pitching motions is presented, A sharp-edged, flat-plate, delta wing having a sweep angle of 70 deg was used in this investigation. The wing was sinusoidally pitched about its one-half chord position at reduced frequencies ranging from /r = 2ir/c/i/ = 0.05 to 0.30 at root chord Reynolds numbers between 9xl04 and 3.5 x 10s, for angle-ofattack ranges of a = 29 to 39 deg and a = 0 to 45 deg. During these dynamic motions, visualization of the leadingedge vortices was obtained by injecting TiCl4 through ports located near the model apex. The location of vortex breakdown was recorded using high-speed motion-picture photography. The motion-picture records were analyzed to determine the vortex trajectory and breakdown position as a function of angle of attack. When the wing was sinusoidally pitched, hysteresis was observed in the location of the breakdown position. This hysteresis increased with reduced frequency. The velocity of breakdown propagation along the wing and the phase-lag between model motion and breakdown location were also determined. Detailed information was also obtained on the oscillation of breakdown position in both static and dynamic cases.

Dynamic stall of a constant-rate pitching airfoil

Abstract: A computational study is presented for the dynamic stall of an airfoil that is pitched at a constant rate from zero incidence to a high angle of attack. The unsteady flow is simulated employing the mass-averaged NavierStokes equations and an algebraic turbulent eddy viscosity model. The approach is first validated by comparison of computed and experimental results for a pitching airfoil at low freestream Mach numbers. The computed dynamic stall events, as well as the computed effects of pitch rate and axis location, are found in qualitative agreement with experimental observations. The effect of compressibility on dynamic stall is investigated. As the freestream Mach number increases, the appearance of a supersonic region provides—through the shock/boundarylayer interaction—an additional mechanism in the dynamic stall process. The main effects of compressibility are found to be 1) a change from trailing-edge stall to leading-edge stall and 2) a reduction in the stall delay and in the attained maximum lift.

Aeroelastic characteristics of a highly flexible aircraft

Abstract: This paper describes an aeroelastic analysis performed on a human-powered aircraft. The structural characteristics of this aircraft (namely, very flexible wings, high aspect ratio, and low wing loadings) are typical of proposed high-altitude long-endurance aircraft. Because of these unique features, the aircraft exhibits uncommon aeroelastic characteristics. The structural dynamics of the airplane are obtained from a finite element model. In an assumed mode approach, a subset of the natural mode shapes is used to calculate the quasisteady and unsteady generalized modal forces using a two-dimensional strip model, which includes unsteady drag and leading-edge suction forces. The aeroelastic model predicts the airplane to be stable (at both sea level and high altitude), except for a mildly unstable phugoid mode. The results indicate that unsteady aerodynamics are significant for this type of aircraft and that the flexibility of the aircraft must be included to correctly model the dynamics for control purposes. A final conclusion is that the aircraft dynamics are highly dependent on the flight conditions (flight speed and altitude). d,D e J k,K Nomenclature = area of span station i • parasite drag coefficient = slope of lift curve = aerodynamic moment about the aerodynamic center at zero angle of attack = half chord (ct/2) at span station i - damping, damping matrix = Oswald's efficiency factor

Static aeroelastic analysis of fighter aircraft using a three-dimensional Navier-Stokes algorithm

Abstract: An aeroelastic analysis method for fighter aircraft operating at extreme flight conditions has been developed and tested. The method involves the use of state-of-the-art zonal grid generation methods, three-dimensional Reynoldsaveraged Navier-Stokes analysis, and linear structures to analyze the flow over complex, flexible aircraft. The main objective of this effort is to develop a method capable of analyzing aircraft operating at flight conditions where vortices, strong shock waves, separated flow, and even highly unsteady flow may be present. The present application focuses on the static aeroelastic analysis of fighter aircraft operating at high angle of attack and high transonic Mach number. The developed method has been compared against static aeroelastic wind-tunnel data on an aeroelastically tailored wing/fuselage configuration, and the results are very encouraging.

Application of global sensitivity equations in multidisciplinary aircraft synthesis

Abstract: The present paper investigates the applicability of the Global Sensitivity Equation (GSE) method in the multidisciplinary synthesis of aeronautical vehicles. The GSE method provides an efficient approach for representing a large coupled system by smaller subsystems and accounts for the subsystem interactions by means of first-order behavior sensitivities. This approach was applied in an aircraft synthesis problem with performance constraints stemming from the disciplines of structures, aerodynamics, and flight mechanics. Approximation methods were considered in an attempt to reduce problem dimensionality and to improve the efficiency of the optimization process. The influence of efficient constraint representations, the choice of design variables, and design variable scaling on the conditioning of the system matrix was also investigated.

Connection between leading-edge sweep, vortex lift, and vortex strength for delta wings

Abstract: An effort is made to clarify the effect of leading-edge sweep on the vortex lift and leading-edge vortex strength of a slender wing; while it is often assumed that increasing sweep enhances vortex lift and strength, the opposite is the case. The suction analogy is used in association with numerical and experimental data to derive simple formulas yielding the actual relationship for delta wings. The difference between vortex lift and nonlinear lift is highlighted.

Hypersonic Aerospace Sizing Analysis for the Preliminary Design of Aerospace Vehicles

Abstract: A review of the hypersonic literature indicated that a general weight and sizing analysis was not available for hypersonic orbital, transport, and fighter vehicles. The objective here is to develop such a method for the preliminary design of aerospace vehicles. This report describes the developed methodology and provides examples to illustrate the model, entitled the Hypersonic Aerospace Sizing Analysis (HASA). It can be used to predict the size and weight of hypersonic single-stage and two-stage-to-orbit vehicles and transports, and is also relevant for supersonic transports. HASA is a sizing analysis that determines vehicle length and volume, consistent with body, fuel, structural, and payload weights. The vehicle component weights are obtained from statistical equations for the body, wing, tail, thermal protection system, landing gear, thrust structure, engine, fuel tank, hydraulic system, avionics, electral system, equipment payload, and propellant. Sample size and weight predictions are given for the Space Shuttle orbiter and other proposed vehicles, including four hypersonic transports, a Mach 6 fighter, a supersonic transport (SST), a single-stage-to-orbit (SSTO) vehicle, a two-stage Space Shuttle with a booster and an orbiter, and two methane-fueled vehicles.

Towards integrated multidisciplinary synthesis of actively controlled fiber composite wings

Abstract: The synthesis of actively controlled composite wings is formulated as a multidisciplinary optimization problem. A unique integration of analysis techniques spanning the disciplines of structures, aerodynamics, and controls is described. A rich variety of behavior constraints can be treated including stress, displacement, control surface travel and hinge moment, natural frequency, aeroservoelastic stability, gust response, and handling quality constraints, as well as performance measures in terms of drag/lift coefficients, drag polar shape, required load factor or roll rate, and wing mass. The design space includes a simultaneous treatment of structural, aerodynamic, and control system design variables. The paper sets the stage for multidisciplinary wing optimization by describing the capabilities and discussing the accuracy of the analysis and related behavior sensitivity analysis. Applicability of approximation concepts to the multidisciplinary optimization problem is examined by studying typical aeroservoelastic stability, gust response, and performance-related constraints. The computational efficiency of the combined analysis and sensitivity as well as the quality of key behavior constraint approximations indicate that single-level optimization of composite, actively controlled practical wings is within reach.

Unsteady supersonic computations of arbitrary wing-body configurations including external stores

Abstract: A general harmonic gradient method has been developed for computations of unsteady supersonic flow handling given wing-body combinations, including arbitrary external store arrangements. The harmonic gradient model is adopted so that the total panel number is least affected by the given Mach number and reduced frequencies. To assess the accuracy and effectiveness of the present method, comparison with available data is given including the National Aerospace Laboratory's measurements for underwing store and wing-with-tip missile cases. The accurate prediction and ease of application of the present method suggests that it is fully developed for supersonic aeroelastic applications to realistic aircraft configurations.

Flow visualization studies of the Mach number effects on dynamic stall of an oscillating airfoil

Abstract: Compressibility effects on the dynamic stall of a NACA 0012 airfoil undergoing sinusoidal oscillatory motion were studied using a stroboscopic schlieren system. Schlieren pictures and some quantitative data derived from them are presented and show the influence of free-stream Mach number and reduced frequency on the dynamic-stall vortex. This study shows that a dynamic stall vortex always forms near the leading edge and convects on the airfoil upper surface at approximately 0.3 times the free stream velocity for all cases studied. The results also demonstrate that initiation of the dynamic stall vortex is delayed to higher angles of attack with increased reduced frequency, but that dynamic stall occurs at lower angles of incidence with increasing Mach numbers.

Euler flutter analysis of airfoils using unstructured dynamic meshes

Abstract: Modifications to a two-dimensional unsteady Euler code for the aeroelastic analysis of airfoils are described. The modifications involve including the structural equations of motion and their simultaneous time-integration with the governing flow equations. A novel aspect of the capability is that the solutions are obtained using unstructured grids made up of triangles. Comparisons are made with parallel calculations performed using linear theory and a structured grid Euler code to assess the accuracy of the unstructured grid Euler results. Results are presented for a flat plate airfoil and the NACA 0012 airfoil to demonstrate applications of the Euler code for generalized force computations and aeroelastic analysis. In these comparisons, two different finite-volume discretizations of the Euler equations on unstructured meshes were employed. Sensitivity of the Euler results to changes in numerical parameters were also investigated.

Twin-jet screech suppression

Abstract: A facility test was performed on twin-jet configurations to determine the effectiveness of several concepts in suppressing the supersonic screech tones. Supersonic jet Mach numbers up to 1.75 were tested. The screech suppression concepts were tabs, lateral spacing, axial spacing, and secondary air jets. Acoustic and optical data were obtained. It was found that the twin-jet configuration can result in screech tone amplitudes as much as 20 dB higher than a single jet. Screech tone amplitudes up to 162 dB were measured. Small tabs located at the exit plane were shown to be very effective suppressors if they were large enough or if multiple tabs were installed. Lateral spacing can result in significant tone suppression, however, at certain spacings little suppression was achieved. Axial spacing resulted in essentially no suppression. The secondary air jet was shown to be a very effective suppressor of screech tones from a single jet but was not tested on the twin jet configuration.

Various sources of wing rock

Abstract: Limit cycle oscillations in roll of advanced aircraft can result from three different fluid mechanical flow processes The so-called slender wing rock is caused by asymmetric vortex shedding from highly swept wing leading edges A completely different flow mechanism causes wing rock for aircraft with moderately swept leading edges In this case, the causative mechanism is dynamic airfoil stall

Flight and wind-tunnel investigations on boundary-layer transition

Abstract: Flight and wind-tunnel experiments have been carried out to measure the pressure distribution and the transition location on a special wing glove. By means of the linear stability theory of laminar boundary layers limiting TV values of Tollmien-Schlichting waves at the transition location have been evaluated. The values of TV-13.5 are nearly independent of Reynolds number and are the same in flight and wind-tunnel tests.

Effect of detailed surface geometry on riblet drag reduction performance

Abstract: A comparison is made of the effect of small changes in v-groove geometry, for several riblet films applicable for drag reduction to commercial transport aircraft, whose nominal v-groove dimension is of the order of 0002 inch The films were tested in a water towing-tank facility The results obtained indicate that small riblet peak geometry variations can result in a deterioration of riblet drag-reduction efficacy of as much as 40 percent, while interriblet valley curvature was found not to be critical to riblet performance

Zonal Navier-Stokes methodology for flow simulation about a complete aircraft

Abstract: Transonic Navier-Stokes flow simulations are presented for the F-16A fighter aircraft using a zonal grid approach. This approach subdivides the physical space about the aircraft into an ensemble of simple geometric shapes, thus mitigating many of the difficulties of generating a single grid about a complex shape, e.g., providing adequate grid refinement near all body surfaces to capture the boundary layer. Information is propagated between zones via grid overlapping and a spatial interpolation procedure. Computational Cp compare well with experimental values on the wing, horizontal and vertical tails, fuselage centerline, and the inlet/diverter region. The average y+ one grid point off the wing is 3. The experimental lift is underpredicted by 2.6%, and the experimental drag is overpredicted by 1.6%. The flexibility of the zonal approach is demonstrated by adding additional zones inside the inlet up to the compressor face to model flow spillage, and downwind of the exhaust nozzle to model power-on conditions. Computations are also presented for the F-16A in sideslip. These results demonstrate that the present zonal approach provides a flexible and viable means of simulating flowfields about complex geometries.

Minimizing life cycle cost for subsonic commercial aircraft

Abstract: The inclusion of life cycle cost (LCC) as early as possible in the conceptual design process is necessary because of the strong impact early design effort has on the total cost of an aircraft program. Considering LCC for military aircraft is the current industry standard; for commercial aircraft it is also necessary to weigh the merit of decreases in operating costs against increases in acquisition cost and vice versa. A methodology has been developed that makes it possible to identify an aircraft concept that will meet the mission requirements and have the lowest LCC. The methodology consists of an LCC module composed of elements to calculate RDTE however, the models chosen permit alternatives to be treated consistently so that relative comparisons are valid. Provision is made in the methodology for sensitivities to advanced technologies to also be investigated. The conceptual design system is applied to short-, medium-, and medium-to-long range subsonic commercial airplanes. The aircraft are optimized for minimum gross weight, fuel burned, acquisition cost, and DOC to show that different concepts result when LCC is considered. Sensitivities of the aircraft to economic and technology variables are illustrated.

Measurements on an oscillating 70-deg delta wing in subsonic flow

Abstract: A series of low-speed wind tunnel tests on a 70-deg sharp leading-edged delta wing at both static and dynamic conditions were performed to investigate the aerodynamic forces and moments. Forces and moments were obtained from a six-component internal strain-gauge balance. Large amplitude dynamic motion was produced by sinusoidally oscillating the model over a range of reduced frequencies. Static results compared well with previous experimental findings. Significant Reynolds number effects were present in the experimental measurements. Reynolds number effects are reduced, but still present when a sharper leading-edge delta wing was tested. Large hysteresis loops and a delay in dynamic stall were seen in the dynamic data. Dynamic forces and moments were a strong function of reduced frequency. Nonzero sideslip created complex rolling moment and lift behavior due to asymmetric vortex bursting.

Control of asymmetric vortical flows over delta wings at high anglesof attack

Abstract: Results from a wind tunnel experiment will be used to confirm that Tangential Leading Edge Blowing is capable of controlling the vortical flow over a delta wing to very high angles of attack. The production of significant rolling moments and the ability to unburst a vortex has been demonstrated to 55" angle of attack. Strain gauge balance measurements and upper surface pressure distributions will be presented which illustrate the various modes of operation and which, in particular, identify the uncoupled/coupled nature of the vortex control at pre- and post-stall angles of attack. At angles of attack beyond 40°, rolling moments are produced that exceed those produced by 20" of aileron deflection at 0" angle of attack.

Effect of moving surfaces on the airfoil boundary-layer control

Abstract: The concept of moving surface boundary-layer control, as applied to a Joukowsky airfoil is investigated through a planned experimental program complemented by a flow visualisation study. The moving surface was provided by rotating cylinders located at the leading and trailing edges of the airfoil. The leading-edge rotating cylinders extends the lift curve without substantially affecting its slope, thus effectively increasing the maximum lift and delaying stall. A flow visualisation study substantiates effectiveness of the concept.

Low-speed unsteady aerodynamics of a pitching straked wing at high incidence-Part II: Harmonic analysis

Abstract: Part I of this paper described the wind-tunnel model and test program. This part presents data from the test that are amenable to harmonic analysis that limits the amplitudes to about ±4 or ±6 deg but does not limit mean angles or frequencies. Those results are applicable to flutter and dynamic response problems as well as stability and control characteristics at high incidences. Force, pressure, and flow-visualization data are used to describe the perceived flow phenomena and how they interact with the model to produce the aerodynamic forces*

Effects of Pressure Gradient on Reattaching Flow Downstream of a Rearward-Facing Step

Abstract: Reattachment problems of separated flow on a solid surface such as an airfoil, diffuser, cavity wall, etc. are of interest because a rapid rise of pressure and heat transfer takes place at the reattachment zone. Among the solid surface models for the study of the separated flow reattachment, the rearward-facing step is one of the simplest since the separation point is readily known. A number of investigations for the rearward-facing step is one of the simplest since the separation point is readily known. A number of investigations for the rearward-facing step flow showed that the reattachment length is strongly affected by the pressure gradient. This paper presents the experimental results of reattachment length and wall static pressure distribution affected by the constant streamwise pressure gradients.

Minimum Induced Drag for Wings with Spanwise Camber

Abstract: Linear theory is used to develop optimum circulation distributions and their associated minimum induced drag for wakes from lifting surfaces with various spanwise camber. The work is largely computational and results for cases previously investigated analytically are generally in good agreement. However, some previously published results are found to be in error, and a new solution for the induced drag of a wing with dihedral is given. New results are computed for polynomial and superelliptic camber lines. An empirical correlation is demonstrated between the induced drag factor and the inverse arc length for a variety of optimum cases.

Active flutter suppression for a wing model

Abstract: The paper presents an experimental verification of a method to design active control flutter suppression systems that allows eigenstructure assignment directly within a/?-A flutter approximation. By using a wing model, it is shown how different and simple direct feedback control laws can be effective in producing a substantial improvement of the flutter speed and of the overall damping below the critical speed. Difficulties encountered in correlating designs to test by using a "Maximum Likelihood" identification method are also addressed.

Sensitivity derivatives of flutter characteristics and stability margins for aeroservoelastic design

Abstract: Analytical derivatives of flutter dynamic pressure, flutter frequency, gain margins, and phase margins with respect to various aeroservoelastic design variables are developed. The formulation is based on a first-order time-domain aeroservoelastic mathematical model. The structure is represented in the model by vibration modes, the unsteady aerodymanics by minimum-state rational approximation functions, and the control system is coupled with the aeroelastic system through motion sensors and control surfaces. The sensitivity derivatives are expressed as exact functions of the stability boundary eigenvectors and of the real-valued system matrix derivatives with respect to arbitrary design variables. These expressions may be applied to any aeroservoelastic design variable with respect to which the system matrix derivative is available. System matrix derivatives with respect to structural stiffness and mass parameters, control gains, actuator parameters, and sensor locations are presented. The new sensitivity derivatives facilitate efficient gradient-based aeroservoelastic algorithms, which directly address flutter and stability margin requirements. A numerical example utilizing the mathematical model of the Active Flexible Wing wind-tunnel model is given. A practical control design problem with roll maneuver constraints is used to demonstrate the accuracy and usage of the derivatives.