Showing papers in "Journal of Aircraft in 1996"

Delay of Airfoil Stall by Periodic Excitation

Abstract: It was recently demonstrated that oscillatory blowing can delay separation from a symmetrical airfoil much more effectively than the steady blowing used traditionally for this purpose. Experiments carried out on different airfoils revealed that this flow depends on many parameters such as, the location of the blowing slot, the steady and oscillatory momentum coefficients of the jet, the frequency of imposed oscillations, and the shape and incidence of the particular airfoil. In airfoils equipped with slotted flaps, the flow is also dependent on the geometry of the slot and on the Reynolds number in addition to the flap deflection that is considered as a part of the airfoil shape. The incremental improvements in single element airfoil characteristics are generally insensitive to a change in Reynolds number, provided the latter is sufficiently large. The imposed oscillations do not generate large oscillatory lift nor do they cause a periodic meander of the c.p. C* C D = dp Ct =

Aircraft Anti-Icing and De-Icing Techniques and Modeling

Abstract: Nomenclature Ac = accumulation parameter Cd drag coefficient Ci = lift coefficient Cm = moment coefficient Cp = pressure coefficient c = specific heat at constant pressure, J/(kg-K); airfoil chord, m D = propeller diameter, m; flexural stiffness, N-m D drag force, N d = droplet diameter, m E = total collection efficiency / = freezing fraction g = acceleration caused from gravity, m/s Hi = ice thickness, m Hp = plate thickness, m h = airfoil projected height, m hc = convective heat transfer coefficient, W/(m-K) hfg = heat of vaporization, J/kg hsi = heat of fusion, J/kg / = airfoil drag constant K = thermal conductivity, W/(m-K); inertia parameter K0 = modified inertia parameter k = roughness diameter, m LWC = liquid water content, kg/m M = local Mach number MVD = median volume droplet diameter, m m = mass, kg ra = mass flow rate, kg/s m' = mass flow rate per unit width, kg/(m-s) m" = mass flux, kg/(m-s) n = normal direction P = pressure, Pa p spatial pressure distribution, N/m <2 = heat rate, W q = normal pressure distribution, N/m

Evolution of flight vehicle system identification

Abstract: Synopsis: From the nostalgic remembrance of the first dynamic flight test this Survey Paper traces several milestones in the history of flight vehicle system identification. A comprehensive account of modern system identification techniques is provided. Several challenging examples bring out the fact that these techniques have reached a high level of maturity, making them a sophisticated and powerful tool not only for research purposes, but also to support the needs of the aircraft industry. This survey paper includes 183 references and provides a consolidated list of publications on the subject.

Parallel Genetic Algorithm Implementation in Multidisciplinary Rotor Blade Design

Abstract: : The present paper describes an adaptation of genetic algorithms in the design of large-scale multidisciplinary optimization problems. A hingeless composite rotor blade is used as the test problem, where the formulation of the objective and constraint functions requires the consideration of disciplines of aerodynamics, performance, dynamics, and structures. A rational decomposition approach is proposed for partitioning the large-scale multidisciplinary design problem into smaller, more tractable subproblems. A design method based on a parallel implementation of genetic algorithms is shown to be an effective strategy, providing increased computational efficiency, and a natural approach to account for the coupling between temporarily decoupled subproblems. A central element of the proposed approach is the use of artificial neural networks for identifying a topology for problem decomposition and for generating global function approximations for use in optimization. (AN)

Unsteady aerodynamics of a flapped airfoil in subsonic flow by indicial concepts

Abstract: An approach based on indicial concepts is described to model the unsteady airloads on a thin airfoil in subsonic compressible flow caused by the arbitrary motion of a trailing-edge flap. Exact indicial aerodynamic responses at small values of time as a result of flap deflection and angular deflection rate about the flap hinge are obtained from linear unsteady subsonic theory in conjunction with the aerodynamic reverse flow theorems. Using the known exact initial (piston theory) and asymptotic values of the airloads, along with an assumed analytic form for the indicial functions, these exact results are used to help obtain complete approximations for the respective indicial responses. The airloads from arbitrary flap motion in subsonic flow are subsequently obtained in state - space form. Validation of the method is conducted with experimental data for time-dependent flap motions.

Development of a finite element aeroelastic analysis capability

Abstract: Extensions of the general-purpose finite element (FE) structural analysis program STARS for computational fluid dynamics (CFD) based aeroelastic analysis are described. Previous capabilities include structural as well as aeroelastic and aeroservoelastic analyses using linear aerodynamic theories. The current extension involves FE-based CFD solution techniques for aeroelastic analysis, and this article describes the development and application of this integrated, multidisciplinary FE analysis tool for effective modeling and simulation of aerospace vehicles. Numerical examples of flutter solution of two representative problems, namely a panel and a 45-deg swept-back wing are presented in this article, along with comparisons of computed and experimental results that testify to the efficacy of the presently developed numerical techniques.

Characteristics of surface roughness associated with leading-edge ice accretion

Abstract: Detailed size measurements of surface roughness associated with leading edge ice accretions are presented to provide information on characteristics of roughness and trends of roughness development with various icing parameters. Data was obtained from icing tests conducted in the Icing Research Tunnel (IRT) at NASA Lewis Research Center (LeRC) using a NACA 0012 airfoil. Measurements include diameters, heights, and spacing of roughness elements along with chordwise icing limits. Results confirm the existence of smooth and rough ice zones and that the boundary between the two zones (surface roughness transition region) moves upstream towards stagnation region with time. The height of roughness grows as the air temperature and the liquid water content increase, however, the airspeed has little effect on the roughness height. Results also show that the roughness in the surface roughness transition region grows during a very early stage of accretion but reaches a critical height and then remains fairly constant. Results also indicate that a uniformly distributed roughness model is only valid at a very initial stage of the ice accretion process.

Full-Scale Wind-Tunnel Studies of F/A-18 Tail Buffet

Abstract: Tail buffet studies were conducted on a full-scale, production F/A-18 fighter aircraft in the 80 by 120 ft Wind Tunnel at NASA Ames Research Center. The F/A-18 was tested over an angle-of-attack range of 18-50 deg, and at wind speeds of up to 168 ft/s, corresponding to a Reynolds number of 12.3x10(exp 6) based on mean aerodynamic chord and a Mach number of 0.15. The port, vertical tail fin was instrumented and the aircraft was equipped with a removable leading-edge extension (LEX) fence. Time-averaged, power-spectral analysis results are presented for the tail fin bending moment derived from the integrated pressure field, for the zero side-slip condition, both with and without the LEX fence. The LEX fence significantly reduces the magnitude of the rms pressures and bending moments. Scaling issues are addressed by comparing full-scale results for pressures at the 60%-span and 45%-chord location with small-scale, F/A-18 tail-buffet data. The comparison shows that the tail buffet frequency scales very well with length and velocity. Root-mean-square pressures and power spectra do not scale as well. The LEX fence is shown to reduce tail buffet loads at all model scales.

Aerodynamic characteristics of a two-dimensional airfoil with ground effect

Abstract: The effect of Reynolds number on the aerodynamic characteristics of an airfoil with ground effect in viscous flow is investigated by numerical method. A numerical scheme, based on the standard k-e turbulence model, generalized body-fixed coordinates and the finite volume method, is developed to solve the two-dimensional wing-in-ground problem hi viscous flow. The steady, incompressible Navier-Stokes equations are solved using a grid generation program developed by the authors, and the PHOENICS code. Some numerical results are presented to show the effects of Reynolds number, ground clearance, and angles of attack on the aerodynamic characteristics of a NACA 4412 airfoil.

Helicopter rotor blade computation in unsteady flows using moving overset grids

Abstract: An overset grid thin-layer Navier-Stokes code has been extended to include dynamic motion of helicopter rotor blades through relative grid motion. The unsteady flowfield and airloads on an AH-IG rotor in forward flight were computed to verify the methodology and to demonstrate the method's potential usefulness towards comprehensive helicopter codes. In addition, the method uses the blade's first harmonics measured in the flight test to prescribe the blade motion. The solution was impulsively started and became periodic in less than three rotor revolutions. Detailed unsteady numerical flow visualization techniques were applied to the entire unsteady data set of five rotor revolutions and exhibited flowfield features such as blade vortex interaction and wake roll-up. The unsteady blade loads and surface pressures compare well against those from flight measurements. Details of the method, a discussion of the resulting predicted flowfield, and requirements for future work are presented. Overall, given the proper blade dynamics, this method can compute the unsteady flowfield of a general helicopter rotor in forward flight.

Structural optimization for joined-wing synthesis

Abstract: The joined wing is an innovative aircraft configuration with a rear wing, or horizontal tail, that is attached near the top of the vertical tail and sweeps forward to join the trailing edge of the main wing. This study evaluates two structural design methods for application in an aircraft synthesis code and quantifies the differences between these methods for joined wings in terms of weight, stress, direct operating cost, and computational time. A minimum weight optimization method and a fully stressed design method are used to design joined-wing structures. Both methods determine the sizes of 204 structural members, satisfying 1020 stress constraints and five buckling constraints. Monotonic splines are shown to be a very effective way of linking spanwise distributions of material to a few design variables. Five beam buckling constraints for the horizontal tail are included in both design methods. Without this constraint on buckling, the fully stressed design is shown to be very similar to the minimum weight structure. Adding a beam buckling constraint for the horizontal tail increased the structural weight by 13% and produced a fully stressed design that is 0.9% heavier than the minimum weight structure. Using the minimum weight optimization method to design the structure and to save 0.9% in weight required 20 times the computational time. Furthermore, the minimum weight structure produced only a 0.02% savings in direct operating cost. This study suggests that a fully stressed design method based on nonlinear analysis is adequate for a joined-wing synthesis study. The same joined wing considered in this study was shown, in an earlier study, to be slightly more expensive to operate than a conventional configuration designed for the same medium range transport mission. Since the same fully stressed design method was used in this earlier study, this work supports the comparisons of joined-wing and conventional aircraft performance presented in the earlier study. Of course, a different set of mission specifications and design assumptions may produce joined wings that perform significantly better.

Body-axis rolling motion critical states of a 65-degree delta wing

Abstract: Dynamic wind-tunnel test results of a 65-deg swept delta wing are reviewed. These tests involved bodyaxis rolling motions at moderate (15- to 35-deg) angles of attack in both the Institute for Aerospace Research 2 X 3 m low-speed wind tunnel and the 7 X 10 ft Subsonic Aerodynamic Research Laboratory facility at Wright-Patter son Air Force Base. They included static, forced oscillation, and free-to-roll experiments with flow visualization. Multiple trim points (attractors) for body-axis rolling motions and other unusual dynamic behavior were observed. These data are examined in light of the nonlinear indicial response theory. The analysis confirms the existence of critical states with respect to roll angle. When these singularities are encountered in a dynamic situation, large and persistent transients are induced. Conventional means of representing the nonlinear forces and moments hi the aircraft equations of motion, notably the locally linear model, are shown to be inadequate for these cases. Finally, the impact of these findings on dynamic testing techniques is discussed.

Launch Vehicle Acoustics Part 2: Statistics of the Time Domain Data

Abstract: Skewness and kurtosis coefficients were investigated as a means of quantifying the shock content and non-Gaussian characteristics of rocket noise. Rocket noise data measured at 3-5 locations during the launches of four different vehicles with thrusts ranging from 440 to 10,700 kN were analyzed. The kurtosis coefficients of the acoustic pressure data showed no discernible pattern of variation about the Gaussian value of 3. The skewness coefficients ranged from 0.02 to 0.55, all of them greater than the value of 0 expected for Gaussian data. Both the skewness and kurtosis coefficients for the derivative of the acoustic pressure showed much greater deviations from the values expected for Gaussian noise. This study has confirmed that the coefficient of skewness is a useful metric for the characterization of rocket noise. The statistics of the pressure gradient have been shown to be more sensitive indicators of shock content, and these metrics are much less sensitive to low-frequency instrumentation limits.

Experiments and Analysis for a Gust Generator in a Wind Tunnel

Abstract: An experimental investigation was made of the gust field generated by a rotating slotted cylinder installed in the Duke University low-speed, closed-circuit wind tunnel. The system has a very simple configuration with low cost and can produce a controllable single or multiple harmonic gust wave in the lateral and longitudinal directions. It requires minimal power and torque input. A simplified theoretical aerodynamic model and a design estimation of the lateral and longitudinal gust flowfield is also proposed in this article. The design estimate is based on a two-dimensional dynamic lift coefficient that is given by the theoretical and experimental results. An interfering wake vortex effect is the major disadvantage of this system. Nomenclature C(k) = Theodorsen's function ICleql = magnitude of equivalent lift coefficient for rotating slotted cylinder/airfoil c = airfoil chord d = cylinder diameter dLa = airfoil lift force per span length dLrsc = rotating slotted cylinder lift force per span length e, e = gap between the o.d. of the rotating slotted cylinder and trailing edge of the airfoil, elc H, H = vertical position from tunnel bottom, HIHW Hw = height of the tunnel test section

Open- and closed-loop results of a strain-actuated active aeroelastic wing

Abstract: Open- and closed-loop tests for a strain-actuated active aeroelastic wing are summarized. Linear quadratic Gaussian (LQG) control laws as well as robust control laws are designed using sensitivity weighted LQG, classical rationalization, and multiple models. Significant vibration suppression and load alleviation are demonstrated, reducing the power spectral density of the first mode's response by an order of magnitude. The flutter dynamic pressure is increased by 12%. The three major performance limitations are the saturation limit of the piezoelectrics, the choice of performance metric or output sensor, and the changes in the dynamic response of the test article. T

Subsonic unsteady aerodynamics caused by gusts using the indicial method

Abstract: Indicial approximations are derived for the lift on an airfoil penetrating a stationary sharp-edge gust in two-dimension al subsonic flow. Using an assumed exponential form, the approximations have been generalized in terms of Mach number alone by means of an optimization algorithm where certain coefficients of the approximations are free parameters. The optimization is subject to prescribed constraints in terms of the initial and asymptotic behavior of the gust response, and by requiring the response closely match the known exact solutions given by subsonic linear theory at earlier values of time. An alternative approximation is obtained by using results from a direct numerical simulation of the gust problem using computational fluid dynamcs (CFD). For an airfoil-vortex interaction problem, comparisons were made with experimental data and CFD results. Finally, the indicial method was integrated into a three-dimensional rotor simulation, and the near- and far-field acoustics were computed using the Ffowcs WilliamsHawkins equation. Good agreement was found with simultaneously measured airloads and acoustics data.

Integrated structural tailoring and control using adaptive materials for advanced aircraft wings

Abstract: This study combines structural tailoring with the adaptive capabilities of piezoelectric materials for the purpose of controlling the vibration and static aeroelastic characteristics of advanced aircraft wings. The structural model consists of a thin/thick-walled closed cross-sectional cantilevered beam whose constituent layers exhibit elastic anisotropic properties and incorporates a number of nonclassical features. Results reveal that a combination of structural tailoring and control using adaptive materials can play a major role in enhancing the vibrational and static aeroelastic response characteristics of aircraft wings. ECAUSE of their outstanding properties, such as high strength/stiffness to weight ratios, fiber-reinforced laminated thick/thin-walled structures are likely to play an increasing role in the design of advanced aircraft wings. In addition, a number of elastic couplings resulting from anisotropy and the ply-angle sequence of composite material structures can be exploited so as to enhance the response characteristics. In this regard, within the last two decades, a technique referred to as structural tailoring has been used with spectacular results.1 It should be noted, however, that structural tailoring is a passive design technique. This implies that the structure cannot respond adaptively to changes in its parameters or external stimuli. To overcome this shortcoming, additional capabilities must be built into the structure. This is particularly true in view of the fact that future generations of flight vehicles are likely to operate under increasingly severe conditions. An approach showing good promise is based upon the incorporation into the structure of materials featuring sensing and actuating capabilities.2"5 Piezoelectric materials are excellent candidates for the role of sensors and actuators. In contrast to passive structures, in which the vibrational and aeroelastic response characteristics are predetermined, in adaptive structures these characteristics can be altered in a known and predictable manner. These adaptive capabilities can be used to prevent structural resonance and/or any other type of instability, as well as to improve the static and dynamic response of the structure. In this article, the task of enhancing the static aeroelastic response and free vibration characteristics of aircraft wingtype structures made of advanced composite materials is accomplished through the synergistic combination of structural tailoring and adaptive materials technology. The structure simulating the aircraft wing consists of a thin/thick-walled closed cross-sectional cantilevered beam whose constituent layers feature elastic anisotropic properties. The control capability is achieved by electrically actuating piezoelectric ele

Reduction of vertical tail buffet response using active control

Abstract: The active vertical tail (AVT) successfully reduced the buffet response of structures by utilizing piezoelectric actuators, strain gauge sensors, and simple control techniques. The AVT is a 5%-scale aeroelastically tailored structure that exhibits vibration response similar to a full-scale aircraft structure, and was designed such that it's piezoelectric actuators could provide control authority in the first two bending modes. The AVT was wind-tunnel tested on a generic twin-tailed double-delta fighter model at angles of attack and dynamic pressures representative of actual aircraft flight envelopes. At high angles of attack, the model's leading-edge vortices impinge upon the AVT. Simple control algorithms were used with piezoelectric actuators and collocated strain gauge sensors to either minimize the acceleration at the AVT's tip or the strain at the root of the tail. Control gains were verified to be a nonlinear function of angle of attack, dynamic pressure, and location of the actuator/sensor pair. Spectral analysis showed that the peak response of the controlled AVT was up to 65% lower than the uncontrolled response. This represents approximately an order of magnitude improvement in the fatigue life of a similar aircraft structure. The rms response below 200 Hz was reduced by over 20%.

Experimental measurements of dynamic control surface effectiveness

Abstract: This article reports on a low Reynolds number experimental study of the unsteady lift characteristics of an NACA 0009 airfoil pitched about its midchord, equipped with a 27% trailing-edge flap that could be independently deflected. The airfoil normal force, obtained from integrated surface pressure measurements, was captured during rapid, trailing-edge flap-only deflections; rapid, arbitrary, pitch-only excursions; and combined rapid pitch and flap-deflection motions. The measured aerodynamic response of the flapped airfoil to the various flap, airfoil, and airfoil flap combination motions was compared to theoretical and panel-code-computed aerodynamic-response predictions. These comparisons, and the results of flow visualization experiments, led to the observation that both the flap's effectiveness and the airfoil's lift-curve slope were higher during their motion than their steady-state values, and essentially matched their inviscid theoretical values.

High-Lift Design Optimization Using Navier-Stokes Equations

Abstract: This article presents a design optimization method for maximizing lift without increasing the drag of multielement airfoils at takeoff and landing configurations. It uses an incompressible Navier-Stokes flow solver (INS2D), a chimera overlaid grid system (PEGSUS), and a constrained numerical optimizer (DOT). Aerodynamic sensitivity derivatives are obtained using finite differencing. The method is first validated with single-element airfoil designs and then applied to three-element airfoil designs. Reliable design results are obtained at reasonable costs. Results demonstrate that numerical optimization can be an attractive design tool for the development of multielement high-lift systems.

Conceptual design of helicopters via genetic algorithm

Abstract: The genetic algorithm (GA) is a computational model of natural selection and reproduction displayed by biological populations. The capabilities of GAs as search and optimization methods make them well suited to perform conceptual design tasks. A GA has been developed and combined with an industry standard sizing code specifically for helicopter conceptual design. This GA-based program was used to generate conceptual designs for three helicopter missions. Results of these efforts are discussed, providing insight into the ability of the GA to perform helicopter conceptual design.

Response of a Research Aircraft to Icing and Evaluation of Severity Indices

Abstract: The relationship between four atmospheric parameters and three measures of flight degradation are investigated using data from an instrumented research aircraft. Data from flights that took place in wintertime stratus clouds over northeastern Colorado are emphasized; additional data points from encounters with large supercooled droplets over northern California and northern Arizona are included. The maximum decrease in coefficient of lift due to icing was 35%, with 68% of cases within 10% of the uniced aircraft value. Coefficient of drag increased by up to 230% as a result of icing and climb capability was reduced by up to 6.9 m/s. Greater performance loss was related to higher liquid water content, median volume diameter, and potential accumulation of ice. A combination of liquid water content > 0.2 g/m3, median volume diameter > 30 /un, and temperature > — 10°C was responsible for the largest performance decreases. Severity indices that were dependent on liquid water content, median droplet volume diameter, and temperature were tested. An index that takes into account the effects of large droplet icing provided the best relation between higher severity level and increased performance degradation.

Vortex model to define safe aircraft separation distances

Abstract: An approximate model of wake vortices behavior is presented. We have introduced modifications in the Greene's model to take into account the effects of ground (divergence, rebound) and crosswind (advection, shear). Direct numerical simulations of laminar flows with and without lateral wind are performed to validate these extensions. The first results are encouraging and efforts are carried on to derive a more reliable model. The capability of this model to mimic the reality is shown by comparison with an experimental test case.

Effect of shear-layer control on leading-edge vortices

Abstract: The effect of shear-layer control on leading-edge vortices over delta wings with sharp leading edges was investigated. By using suction near the separation point, the direction of the shear layer, and the location of the main vortex as well as its structure could be modified. The maximum swirl angle in the core and overall circulation decreased with suction. This caused the vortex breakdown location to move downstream. When the suction slot was located slightly inboard, the swirl angle was found to increase, causing the breakdown location to move upstream. It is shown that extremely small amounts of leadingedge suction can modify the location of the vortex core and can thus be used to induce rolling moment.

Computational fluid-structure interaction model for parachute inflation

Abstract: In parachute research, the canopy inflation process is the least understood and the most complex to model. Unfortunately it is during the opening process that the canopy often experiences the largest deformations and loadings. The complexity of modeling the opening process stems from the coupling between the structural dynamics of the canopy, lines, and payload with the aerodynamics of the surrounding fluid medium. The addition of a computational capability to model the coupled opening behavior would greatly assist in the understanding of the canopy inflation process. This article describes research that involves coupling a computational fluid dynamics code to a mass spring damper parachute structural code. The axisymmetric codes are coupled with an explicit marching method. The current model is described and results for a round parachute are presented. A comparison of the numerical results to experimental data will be presented. The successful solution of these problems gives us confidence that the computational aeroelastic problem for parachute openings can be solved. This solution allows moving the parachute design process from one of cut and try to one based on experimentally verified computational tools and reduces the reliance on costly and time-consuming testing during development.