Showing papers in "Journal of Aircraft in 1998"

Accuracy and Coupling Issues of Aeroelastic Navier-Stokes Solutions on Deforming Meshes

Abstract: An implicit time-accurate approach to aeroelastic simulation was developed with particular attention paid to the issues of time accuracy, structural coupling, grid-deformation strategy, and geometric conservation. A Beam ‐Warming, approximate-factored algorithm, modie ed to include Newton-like subiterations was coupled with a structural model, also in subiteration form. With a sufe cient number of subiterations, this approach becomes a fully implicit, e rst- or second-order-accurate aeroelastic solver. The solver was used to compute time-accurate solutions of an elastically mounted cylinder. The fully implicit coupling allowed the overall scheme to become second-order accurate in time, signie cantly reducing the workload for a given accuracy. A new algebraic grid deformation strategy was developed that preserves grid orthogonality near the surface under large deformations. Finally, the oscillatory behavior of an elastically mounted cylinder was reproduced accurately by the present approach, and results compared favorably to previous experiments and simulations.

Aeroelastic Response of a Rigid Wing Supported by Nonlinear Springs

Abstract: Nonlinear aeroelastic behavior is examined. This research extends the efforts of several recent investigations that address freeplay or piecewise nonlinearities in aeroelastic systems; however, in the studies described herein the authors address continuous nonlinearities such as those found in structural systems that exhibit spring hardening or softening effects. The authors describe a unique e utter test apparatus designed to permit experimental investigations of prescribed nonlinear response. The results of complementary analytical and experimental studies are presented for a nonlinear aeroelastic system limited to two degrees of freedom. Nomenclature

Gurney Flap Experiments on Airfoils, Wings, and Reflection Plane Model

Abstract: The effect of Gurney e aps on two-dimensional airfoils, three-dimensional wings, and a ree ection plane model were investigated. There have been a number of studies on Gurney e aps in recent years, but these studies have been limited to two-dimensional airfoil sections. A comprehensive investigation on the effect of Gurney e aps for a wide range of cone gurations and test conditions was conducted at Wichita State University. A symmetric NACA 0011 and a cambered GA (W)-2 airfoil were used during the single-element airfoil part of this investigation. The GA (W)-2 airfoil was also used during the two-element airfoil study with a 25% chord slotted e ap dee ected at 10, 20, and 30 deg. Straight and tapered ree ection plane wings with natural laminar e ow (NLF) airfoil sections were tested for the three-dimensional wing part of this investigation. A fuselage and engine were attached to the tapered NLF wing for the ree ection plane model investigation. In all cases the Gurney e ap improved the maximum lift coefe cient compared to the baseline clean cone guration. However, there was a drag penalty associated with this lift increase.

Further Refinement of the Subsonic Doublet-Lattice Method

Abstract: The doublet-lattice method (DLM) is in use worldwide for e utter and dynamic response analyses of aircraft at subsonic speeds. The present paper develops a further ree nement to extend its frequency limits for applications to higher frequency e utter, e.g., for aeroservoelastic systems with high-frequency control surfaces, and dynamic response, e.g., for short wavelength gusts. The DLM is an aerodynamic e nite element method for modeling oscillating interfering lifting surfaces in subsonic e ows. It reduces to the vortex-lattice method at zero-reduced frequency. The number of e nite elements (boxes) required for accurate results depends on aspect ratio and reduced frequency, among other parameters. At high reduced frequency, the chordwise dimension of the boxes must be small. However, an approximation in the method, viz., that the variation of the numerator of the incremental oscillatory kernel function is parabolic across the span of the box bound vortex, restricts the box aspect ratio to about 3. Hence, high-frequency requirements bring an associated requirement for a large number of boxes in the aerodynamic ideali- zation. If a higher-order approximation is used for the spanwise variation of the numerator of the incre- mental oscillatory kernel, the limitation on box aspect ratio can be relaxed and the number of spanwise divisions required in high-frequency analyses will be reduced signie cantly, leading to a reduction in the total number of boxes. This paper replaces the original parabolic approximation by a quartic approxi- mation. The quartic curve-e tting coefe cients are determined for the planar and nonplanar kernels, and the new integrals for the planar and nonplanar normalwash factors are evaluated. The ree nement is incorporated into a DLM code previously known as N5KA, and convergence studies on typical cone gu- rations are presented that attempt to specify a higher limit for practical box aspect ratios.

Motion Simulation Capabilities of Three-Degree-of-Freedom Flight Simulators

Abstract: We present the results of a study aimed at determining the simulation realism that might be achieved using reduced-degree-of-freedom flight simulator motion bases. More specifically, the quality of motion produced by two different three-degree-of-freedom platforms was compared to that produced by a conventional six-degree-of-freedom Stewart platform. The three-degree-of-freedom motion bases investigated were a spherical mechanism allowing only rotational motions, as well as a motion base capable of heave, pitch, and roll motions. To compare the different motion bases, three characteristic maneuvers were simulated using a nonlinear model of a Boeing 747. The aircraft motions were then simulated on nine different combinations of virtual motion platforms and motion base drive algorithms. The motion cues (specific forces and angular velocities) produced in this manner were then graphically compared. The analysis revealed that, in most cases, a three-degree-of-freedom simulator is capable of producing motion simulation quality comparable to that produced by a six-degree-of-freedom Stewart platform

Power Requirements for Large-Amplitude Flapping Flight

Abstract: In this paper, a method is presented for computing the circulation distribution along the span of a e apping wing that minimizes the power required to generate a prescribed lift and thrust. The power is composed of three parts: useful thrust power, induced power, and proe le power. Here, the thrust and induced power are expressed in terms of the Kelvin impulse and kinetic energy associated with the sheet of trailing and shed vorticity left behind the e apping wing. The proe le power is computed using a quasisteady approximation of the two-dimensional viscous drag polar at each spanwise station of the wing. A variational principle is then formed to determine the necessary conditions for the circulation distribution to be optimal. Included in the variational principle is a constraint that the wing not stall. This variational principle, which is essentially the viscous extension of the well-known Betz criterion for optimal propellers, is discretized using a vortex-lattice model of the wake, and the optimum solution is computed numerically. The present method is used to analyze a conventional propeller as well as a rigid wing in forward-e ight e apping about a hinge point on the longitudinal axis.

A Transition Closure Model for Predicting Transition Onset

Abstract: A unified approach which makes it possible to determine the extent and onset of transition in one calculation is presented. It treats the laminar fluctuations in a manner similar to that used in describing turbulence. As a result, the complete flowfield can be calculated using existing CFD codes and without the use of stability codes. The method is validated by comparing the results for flat plates, airfoils, and infinite swept wings with available experiments. In general, good agreement is indicated.

Drag Reduction and Shape Optimization of Airship Bodies

Abstract: A tool for the numerical shape optimization of axisymmetric bodies submerged in incompressible flow at zero incidence has been developed. Contrary to the usual approach, the geometry of the body is not optimized in a direct way with this method. Instead, a source distribution on the body axis was chosen to model the body contour and the corresponding inviscid flowfield, with the source strengths being used as design variables for the optimization process. Boundary-layer calculation is performed by means of a proved integral method. To determine the transition location, a semiempirical method based on linear stability theory (e n method) was implemented. A commercially available hybrid optimizer as well as an evolution strategy with covariance matrix adaption of the mutation distribution are applied as optimization algorithms. Shape optimizations of airship hulls were performed for different Reynolds number regimes. The objective was to minimize the drag for a given volume of the envelope and a prescribed airspeed range

Cloud particle measurements in thunderstorm anvils and possible weather threat to aviation

Abstract: Since 1990, there have been at least 10 known incidents where jet aircraft have experienced loss of thrust in one or more turbofan engines while maneuvering in the anvil region near the central core of a thunderstorm. The exact cause of the uncommanded thrust reduction, commonly called engine rollback, is still under investigation. It appears that the rollback incidents may be associated with ingestion of high mass concentrations of ice particles, snow, and possibly small concentrations of supercooled liquid water in the anvil region. The characteristics of cloud particles in thunderstorm anvils have not been extensively studied. Results from analysis of aircraft observations in the anvils of midlatitude and tropical thunderstorms are discussed. Aircraft and limited radar observations show that most anvils associated with small, garden-variety thunderstorms contain low ( < ∼ 0.4 g m -3 ) mass concentrations of ice particles. In larger, more intense midlatitude storms, anvils may contain ice water contents from 1 to 3 g m -3 . The mean of the maximum particle dimension in the anvil region of the more intense storms showed a strong modal size of about 2 mm. The particles themselves appear to be ice crystals and aggregates of ice crystals, i.e., snowflakes

Estimation of Aircraft Lateral-Directional Parameters Using Neural Networks

Abstract: A recently proposed method (christened ‘ ‘ the Delta method’ ’) of estimating aircraft parameters from e ight data using feed-forward neural networks is applied for the extraction of lateral ‐directional parameters from simulated as well as real-e ight data. The neural network is trained using aircraft motion and control variables as the network inputs and aerodynamic coefe cients as the network outputs; the trained network is used to predict aerodynamic coefe cients for a suitably modie ed input e le. An appropriate interpretation and manipulation of such input ‐output e les yields the estimates of the parameters. Flight data for lateral ‐directional dynamics are analyzed for various combinations and types of control inputs, and suitable control input forms are identie ed for better estimation via the proposed method. Robustness of the method with respect to measurement noise is demonstrated by its applicability to simulated e ight data with pseudomeasurement noise, and to real-e ight data.

Experimental Active Control of a Typical Section Using a Trailing-Edge Flap

Abstract: This paper presents an experimental implementation of an active control system used to suppress flutter in a typical section airfoil. The H2 optimal control system design is based on experimental system identifications of the transfer functions between three measured system variables - pitch, plunge, and flap position - and a single control signal that commands the flap of the airfoil. Closed-loop response of the airfoil demonstrated gust alleviation below the open-loop flutter boundary. In addition, the flutter boundary was extended by 12.4% through the application of active control. Cursory robustness tests demonstrate stable control for variations in flow speed of ± 10%.

Ice Accretion Prediction on Multielement Airfoils

Abstract: The aim of this paper is to present the Italian icing computational environment code, I 2 CE, developed by CIRA, and discuss its evaluation for ice accretion on single and multielement airfoils. The aerodynamic module of the code is based on a potential panel method, whereas the thermodynamic module is based on the classic Messinger model. Different ways to solve the time-dependent ice accretion problem have been taken into account. A comparison between theoretical calculations using multi-time-step, single-time-step, or predictor-corrector procedures and experimental data has been carried out. The effect of the flowfield viscosity on droplet trajectories and the influence of different approaches for the convective heat exchange coefficient calculation have been considered. The influence on the impingement of the actual droplet size distribution has been taken into account and a comparison with a standard median volumetric diameter calculation is presented

Vortex-generator installation studies on steady-state and dynamic distortion

Abstract: The theoretical and experimental work carried out under the NASA/MOD joint aeronautical program has shown that computational fluid dynamics (CFD) vortex-generator installation designs successfully managed inlet-duct now distortion and that significant benefits in flow unsteadiness at the engine face were also present. The main conclusions to date from the collaborative effort between the NASA Lewis Research Center and the Defence Research Agency in Bedford are as follows: 1) Vortex-generator installations can be designed to be effective over a wide range of inlet operating conditions using CFD and formal optimization procedures, 2) reductions in steady-state engine face distortion of up to 80% have been measured in the M2129 inlet S-duct using CFD-designed vortex-generator installations, 3) reductions in flow unsteadiness of up to 80% have been measured in the M2129 inlet S-duct using CFD-designed vortex-generator installations, and 4) the reduced Navier-Stokes code RNS3D is a useful tool to design vortex-generator installations to manage engine-face distortions over a wide range of inlet operating conditions

Investigation of confluent boundary layers in high-lift flows

Abstract: The confluent boundary layers over a three-element high-lift airfoil are studied using both numerical and experimental approaches. The results suggest that wake prediction is crucial to the convergence and accuracy of the overall solution. At maximum lift, unsteadiness is observed in the experiment, which is not captured by computations. However, solutions at maximum lift indicate that, although the flow is attached over the flap, the separation bubble at the leading edge of the slat upper surface is coupled with inviscid flow reaching the compressibility limit. The thickened slat wake results in a displacement of near-surface flow over the main element and limits the main element from gaining more lift. The trends in the confluent boundary layers development require all aspects of the physics be modeled appropriately, including transition, turbulence, and inviscid-viscous interaction

Airfoil Drag Prediction and Decomposition

Abstract: The accuracy and the consistency of numerical techniques for the prediction of the aerodynamic drag of airfoils in viscous transonic and subsonic flows are explored. Attention is paid to the calculation of the total drag as well as to the decomposition of the drag into its physical components: viscous drag and wave drag. Two different Reynolds-averaged Navier-Stokes solvers are used to generate the flowfield solutions for the NLF(1)-0416 and the RAE 2822 airfoils. The results show that wake integration can produce results comparable with those the often-used surface integral technique, thus demonstrating that wake integration has great potential in simplifying drag calculations for more complex problems such as multielement airfoils or complex three-dimensional configurations

Reduced-Order Aerodynamic Model and Its Application to a Nonlinear Aeroelastic System

Abstract: Starting from a finite state model for a two-dimensional aerodynamic flow over an airfoil, the eigenmodes of the aerodynamic flow are determined. Using a small number of these aerodynamic eigenmodes, ie., a reduced-order model, the aeroelastic model is formed by coupling them to a typical section structural model with a trailing-edge flap. A free-play nonlinearity is modeled. Results are shown from the finite state model, the reduced-order model, and previous theoretical and experimental work. All results are in good agreement.

Comparison of Boundary-Layer Transition Predictions Using Flight Test Data

Abstract: Flight tests using a Fokker F100 aircraft equipped with a natural laminar  ow glove demonstrated that natural laminar  ow is feasible for transport aircraft with up to 130 passengers. Furthermore, the  ight tests generated a wealth of experimental data. These data have been evaluated with all known variations of the emethod. The aim of this investigation was to compare the different methods and to identify the one that yielded the best correlation. No clear winner turned up in our investigation; each method has its own merit. The envelope method, if based on a compressible stability theory with surface curvature effects, gives a valuable N-factor correlation. With this method, few pathological cases occurred. These were characterized by a measured transition behind the maximum of the computed N factors. Methods using the N-factor pairs (NTS, N b ) and (NTS, NCF) computed with incompressible or compressible stability theory without curvature effects are also suitable. With these methods, however, many pathological cases occur. A satisfactory correlation can only be obtained if the pathological cases are excluded.

Optimal Design of Grid-Stiffened Composite Panels

Abstract: A design strategy for optimal design of composite grid-stiffened panels subjected to global and local buckling constraints is developed using a discrete optimizer. An improved smeared stiffener theory is used for the global buckling analysis. Local buckling of skin segments is assessed using a Rayleigh-Ritz method that accounts for material anisotropy and transverse shear flexibility. The local buckling of stiffener segments is also assessed. Design variables are the axial and transverse stiffener spacing, stiffener height and thickness, skin laminate, and stiffening configuration, where the stiffening configuration is herein defined as a design variable that indicates the combination of axial, transverse, and diagonal stiffeners in the stiffened panel. The design optimization process is adapted to identify the lightest-weight stiffening configuration and stiffener spacing for grid-stiffened composite panels given the overall panel dimensions, in-plane design loads, material properties, and boundary conditions of the grid-stiffened panel.

Stability of a Towed Body

Abstract: This paper presents a mathematical model for the efe cient computation of the stability of bodies subject to e uid ‐ dynamic forces while constrained by a e exible, extensible cable. The way the cable is represented permits the body to be heavier or lighter than air and to have a steady-state lift. The model is applied to the case of a pendant vehicle towed by a cable attached to an aircraft, a case of considerable practical interest. It is shown that inherent instabilities are present and that they can be eliminated by the correct method of cable attachment. The paper emphasizes the physics of the system and the reasons for the instabilities.

Navier-Stokes Analysis of the Flow About a Flap Edge

Abstract: The current study computationally examines one of the principal three-dimensional features of the e ow over a high-lift system, the e ow associated with a e ap edge. Structured, overset grids were used in conjunction with an incompressible Navier ‐Stokes solver to compute the e ow over a two-element highlift cone guration. The computations were run in a fully turbulent mode using the one-equation Baldwin‐Barth model. Specie c emphasis was given to the details of the e ow in the vicinity of the e ap edge, and so the geometry was simplie ed to isolate this region. The geometry consisted of an unswept wing, which spanned a wind-tunnel test section, equipped with a single-element e ap. Two e ap cone gurations were computed: a full-span and a half-span Fowler e ap. The chord-based Reynolds number was 3.7 3 10 6 for all cases. The results for the full-span e ap agreed with two-dimensional experimental results and verie ed the method. Grid topologies and related issues for the half-span e ap geometry are discussed. Results of the half-span e ap case are compared with three-dimensional experimental results, with emphasis on the e ow features associated with the e ap edge. The results show the effect of the vortex created by the e ap edge, including the impact on e ow separation and spanwise lift distribution.

Incorporating Flight Data into a Robust Aeroelastic Model

Abstract: Stability analysis and control synthesis for high-performance aircraft must account for errors in the aircraft model. This paper introduces a method to update a theoretical model using measured flight data. Variations between the flight data and model are represented as uncertainty operators in a robust stability framework. The structured singular value can directly account for these uncertainty operators to compute stability margins robust to the associated dynamical variations. This procedure is used to formulate an uncertain model of an F/A-18 fighter aircraft and compute stability margins that indicate the worst-case flutter conditions.

Approximate performance of periodic hypersonic cruise trajectories for global reach

Abstract: This paper develops the analytical framework to compare the approximate performance of periodic hypersonic cruise trajectories with previously proposed hypersonic trajectory profiles for global reach. Specifically, range, ΔV, and payload-carrying capacity are evaluated for various trajectory types to illustrate the enhanced performance achieved by flying periodic hypersonic cruise trajectories with existing hypersonic vehicle aerodynamic, propulsion, and structures technology. Analytical results reveal that periodic hypersonic cruise trajectories achieve better fuel-consumption savings and deliver more payload over long distances (≅20,000 km) than other trajectory types proposed for high-speed flight. Over 20% improvement in fuel consumption savings is possible for a Mach 10 vehicle with a modest L/D of 4, and a curve-fitted rocket-based combined cycle engine model.

Frequency Effects on Dynamic Stability Derivatives Obtained from Small-Amplitude Oscillatory Testing

Abstract: As a result of a continuing program of work to establish a basic understanding of the aerodynamic phenomena that ine uence agility, stability, and control of future combat aircraft cone gurations, it has become clear that the conventional stability or aerodynamic derivative model for the representation of aerodynamic loads in the aircraft equations of motion is no longer adequate. This paper discusses the limitations of the stability derivative model for modern combat aircraft maneuvers, in particular the problem of motion frequency effects in static and dynamic derivatives derived from small-amplitude oscillatory wind-tunnel tests, and presents an alternative modeling technique based on the concept of an aerodynamic transfer function. Nomenclature A = in-phase component of rolling moment response AR = amplitude ratio ai, bi, ci = constants in generalized transfer functions B = in-quadrature component of rolling moment response b = wingspan, m Cl = rolling moment coefe cient, L/qSb Clb,att = attached e ow component of steady-state derivative Clb,0

Effect of Fuselage Geometry on Delta-Wing Vortex Breakdown

Abstract: A fuselage is usually associated with the use of a delta wing on an actual aircraft. It is also in many cases necessary to use a centerbody of some shape in tunnel tests of pure delta wings. The present paper describes the e ow physics causing the experimentally observed large effect of a fuselage on delta-wing vortex breakdown.

Pitch rate and Reynolds number effects on unsteady boundary-layer transition and separation

Abstract: A NACA 0015 airfoil was pitched at a constant rate through static stall to elevated angles of attack. Shear stress measurements of high spatial and temporal resolution were performed near the airfoil leading edge, in the vicinity of subsequent dynamic stall vortex initiation. Using these data, unsteady boundary-layer reversal and transition were characterized for a range of nondimensional pitch rates and Reynolds numbers. Analyses revealed the independent influences of nondimensional pitch rate and Reynolds number upon unsteady boundary-layer reversal and transition. Temporal and spatial relationships between unsteady boundary-layer reversal and transition imply that unsteady boundary-layer reversal is a precursor and principal determinant in unsteady boundary-layer transition. Comprehension of these and other fundamental unsteady flow physics are crucial for the control of dynamically separated flows generated by maneuvering aircraft, rotorcraft, and wind energy machines

Important Scaling Parameters for Testing Model-Scale Helicopter Rotors

Abstract: An investigation into the effects of aerodynamic and aeroelastic scaling parameters on model scale helicopter rotors has been conducted in the NASA Langley Transonic Dynamics Tunnel. The effect of varying Reynolds number, blade Lock number, and structural elasticity on rotor performance has been studied and the performance results are discussed herin for two different rotor blade sets at two rotor advance rations. One set of rotor blades were rigid and the other set of blades were dynamically scaled to be representative of a main rotor design for a utility class helicopter. The investigation was conducted in forward flight at rotor advance ratios of 0.15 and 0.35. Additionally, the rotors were tested over a range of nominal test medium densities from 0.00382 slugs/ft\super{3} to 0.009 slugs/ft\super{3}. This reange of densities permits the acquisition of data for several Reynolds and Lock nuymber combinations.

Composite patches as reinforcements and crack arrestors in aircraft structures

Abstract: The application of adhesively bonded composite patches as reinforcements and crack arrestors for a multisite damaged aircraft structure is investigated. Experiments are performed to test the ability of a bonded composite reinforcement to prevent cracks from coalescing. With a e nite element model developed for composite patch repairs, the effect of thermal residual stresses on the stress-intensity factor and the resulting fatigue crack growth rate is demonstrated. An effective thermal stress is estimated by comparing experimental results with model predictions. Reinforcement for a multiple-site damage situation is analyzed by modeling an ine nite row of closely spaced cracked rivet holes. The composite reinforcement is shown to dramatically reduce the stress-intensity factor, increase fatigue life, and protect against catastrophic failure.

Simulation of Helicopter Rotor-System Structural Damage, Blade Mistracking, Friction, and Freeplay

Abstract: Selected helicopter rotor faults are simulated by changes in stiffness, inertial, and aerodynamic properties of the damaged or mistracked blade(s). A rotor aeroelastic analysis based on finite elements in space and time and capable of modeling dissimilar blades is used to simulate the damaged rotor in forward flight. The various rotor faults modeled include chordwise imbalance, aerodynamic mistracking, localized cracks, distributed changes in blade stiffness properties modeling a stiffness defect and a manufacturing defect, friction and freeplay in the pitch-control system and the lag damper, and friction in the flap and lag hinge. Changes in blade tip response and rotor hub loads are identified for the selected faults and tables of rotor-system diagnostics are compiled.