Showing papers in "Journal of Aircraft in 1992"

Experimental-theoretical investigation of the vibration characteristics of rotating composite box beams

Abstract: This paper presents a theoretical-cum-experimental study of the free vibration characteristics of thin-walled composite box beams with bending-twist and extension-twist coupling under rotating conditions. The governing equations in generalized displacements were derived using a Newtonian approach. The composite structural model in the derivation used a solid-section approach and contained transverse shear-related couplings and appropriate cross-section warping. The free vibration characteristics of composite box beams were determined by the Galerkin method. In order to validate the theory, glass-epoxy, kevlar-epoxy and graphite-epoxy symmetric and antisymmetric box beams were fabricated using an autoclave molding technique, and tested in an in-vacuo rotor test facility for their vibration characteristics. Beam excitation in the rotating condition was effected by means of induced-strai n actuation with the help of piezoceramic bending elements. Strain gages were used to measure the response of the first three modes over a range of rotational speeds up to 1000 rpm. It was determined that the experimental frequencies and mode shapes correlated satisfactorily with the theoretical results. It is shown also that bending-shear coupling influences the flexural vibration frequencies of antisymmetric box beams significantly. Extension-shear coupling, on the other hand, does not influence the flexural-torsion vibration frequencies of symmetric box beams significantly.

Forcing level effects of internal acoustic excitation on the improvement of airfoil performance

Abstract: The effects of internal acoustic excitation on the leading-edge, separated boundary layers and the aerodynamic performance of NACA 633-018 cross section airfoil are examined as a function of forcing level and forcing frequency of the introduced acoustics. Tests are separately conducted in two suction, open-typed wind tunnels at the Reynolds number of 3.0 x 10 s for the measurements and 1.0 x 10 4 for the visualization. Results indicate that the flow separation is delayed at the angles of attack higher than the stalled angle of small level excitation with the forcing frequency fe near the shear layer instability frequency ft. As the forcing level is increased to some extent, the velocity fluctuations around the slot exit are demonstrated to be the primary governing parameter for modifying the separated boundary layers. Data also show that the effective forcing frequency (and the Strouhal number, 50 extends over wider range as compared to the lower level excitation. Meanwhile, the pressure distributions on the airfoil surface exhibit recovery behaviors with different forcing frequencies. The corresponding boundary layers are visualized to be reattached to the surface to form a recirculation region when the airfoil is around at the stalled angles.

Assessment of Microburst Models for Downdraft Estimation

Abstract: The effectiveness of three microburst models in estimating the downdraft from horizontal velocity measurements is assessed. The development of the models and their characteristics are discussed. The simplest model, a linear one, works well for altitudes below 200 m and near the microburst core. A ring-vortex model and an empirical model give the best overall results. The former is the most complex, requiring four variables to define it. The empirical model uses three variables.

Flutter and Stall Response of a Helicopter Blade with Structural Nonlinearity

Abstract: The purpose of the present paper is to study the flutter instability and forced response of a nonrotating helicopter blade model with a NACA-0012 airfoil and a pitch freeplay structural nonlinearity. In this paper, three typical combinations of linear and nonlinear structure with a linear and nonlinear (ONERA) aerodynamic model are considered. Characteristic results are used to display the limit cycle oscillation and chaotic behavior of both the flutter instability and forced response for all three cases. The effects of various initial disturbance amplitudes on the forced response behavior are discussed. Comparisons of the results for the three cases are helpful in understanding physically the nonlinear aeroelasticity phenomena and chaotic oscillations.

Aerodynamic shape design and optimization - Status and trends

Abstract: A limited number of aerodynamic shape design concepts have been surveyed and an attempt has been made to classify them. Characteristics, both positive and negative, of the more prominent methods were outlined. Future research is expected to concentrate on the use of Navier-Stokes equations and applications to threedimensional configurations. Interdisciplinary constrained optimization is expected to play a more prominent role in the immediate future. Adjoint operator/control theory and its variations are the most promising concepts for interdisciplinary aerodynamic shape design which involves a large number of variables. This theory is expected to constitute the major development area in future research.

Optimized Scramjet Integration on a Waverider

Abstract: Scramjet integrated waverider airframes optimized for maximum thrust margin and lift-to-drag ratio (L/D) are presented. Parameters affecting the success of the scramjet/waverider system are reviewed and the computer code developed to optimize the integrated system is discussed. In this study, conically derived waveriders are specifically designed to supply the scramjet engines with the required properties for effective combustion. The results of this study show that the waverider airframe lends itself well to scramjet engine integration, yielding promising vehicle configurations.

Thermal structures - Four decades of progress

Abstract: Since the first supersonic flight in October 1947, the United States has designed, developed and flown flight vehicles within increasingly severe aerothermal environments. Over this period, major advances in engineering capabilities have occurred that will enable the design of thermal structures for high speed flight vehicles in the twenty-first century. Progress in thermal-structures is surveyed for the last four decades to provide a historical perspective for future efforts.

Physics of Vortical Flows

Abstract: Separation in three-dimensional flows leads to the formation of vortical structures resulting from rolling up of the viscous flow "sheet," initially contained in a thin boundary layer, which springs up from the surface into the outer perfect fluid flow. A clear physical understanding of this phenomenon must be based on a rational analysis of the flowfield structure using the critical-point theory. With the help of this theory, it is possible to interpret correctly the surface flow patterns that constitute the imprints of the outer flow and to give a rational and coherent description of the vortical system generated by separation. This kind of analysis is applied to separated flows forming on typical obstacles, the field of which has been thoroughly studied by means of visualizations and probings using multihole pressure probes and laser velocimetry. Thus, the skin friction line patterns of a transonic channel flow and of a multibody launcher are interpreted. Then, the vortical systems of a delta wing and an afterbody at an incidence are considered. The last two configurations are a missile fuselage-type body and an oblate ellipsoid. I. Introduction F LIGHT at high-incidenc e of combat aircraft or hypersonic vehicles during re-entry, as well as that of tactical missiles, raises practical interest on the study of three-dimensional separated flows. Applications also concern internal flows, in particular air intakes and turbomachines in which the often complex geometry of the channel and the existence of shock waves almost inevitably lead to boundary-laye r separation. In three-dimensional flows, separation entails the formation of vortical structures—frequently, but improperly, called vortices to simplify—form ed by rolling up of the viscous flow "sheet," previously confined in a thin layer attached to the wall, which suddenly springs into the outer nondissipative flow. Although it has been known for a long time, this phenomenon is still incompletely understood from a physical point of view and it is delicate to model due to the flowfield complexity, all the components of which are difficult to capture properly. Many predictive methods are based on perfect fluid models, the first of which use the vortex sheet concept. Such a sheet is defined as a surface of tangential discontinuity for the velocity field. The computational method can use different schemes: doublets, vortex filaments, vortex particles, and so forth. Publications in this domain are too numerous to be cited here. A greater accuracy in flow prediction can be obtained in the solution of the complete Euler equations, which allows, in theory, automatic capture of sheet-like disconti

Sources of High Alpha Vortex Asymmetry at Zero Sideslip

Abstract: A review of existing experimental results for slender bodies and delta wings, tested at high angles of attack, reveals that no physical evidence exists that vortex asymmetry on slender pointed bodies or delta wings has ever occurred through the so-called hydrodynamic instability process. It will be shown that in the numerous tests performed, asymmetric flow separation and/or asymmetric flow reattachment, were the flow mechanisms triggering the vortex asymmetry. Slender wing rock is found to result from a basic lack of roll damping, existing for attached leading-edge vortices, and the vortex-asymmetry is generated at nonzero roll angle, i.e., for asymmetric flow conditions. Nomenclature b = wingspan c = reference length, d CQ = delta wing center chord d = maximum diameter of body of revolution € = rolling moment, coefficient C€ = €/(p00U£/2) Re = Reynolds number based on d and freestream conditions S — reference area, ird2/4 or projected wing area U = horizontal velocity Y = side force, coefficient CY = Y/(pJJl/2) a = angle of attack OA = aPex half-angle Oc = cone half-angle A = leading-edge sweep p = air density = body roll angle Subscripts A = apex c = cone oo = freestream conditions

Effects of ambient turbulence on the decay of a trailing vortex wake

Abstract: Experiments were conducted to investigate the effects of ambient turbulence on the evolution of a trailing vortex wake. The wake was generated by towing an NACA 0012 wing, with a span of 10.2 cm and a chord of 5.1 cm, at an incidence of —10 deg and a speed of 40 cm/s in a tow tank. The chord Reynolds number was 20,400. The ambient turbulence was generated by towing three grids, with square meshes of 1.45, 10.2, and 20.3 cm, upstream of the wing. Turbulence parameters were measured with crossed hot-film probes. The trailing vortex wake, tagged with a fluorescent dye, was visualized from different perspectives and its decay was derived from 16-mm movie records. For weak turbulence with large integral scales compared with the vortex separation, vortex linking is the dominant mode of instability. The dominant wavelength of the linking decreases with increasing turbulence intensity or dissipation rate. As the turbulence intensity increases, vortex bursting appears and eventually replaces linking as the dominant mode of instability. For turbulence with a small integral scale as compared with the vortex separation, vortex instability is predominantly of the bursting type.

Compensating Lags in Head-Coupled Displays Using Head Position Prediction and Image Deflection

Abstract: Images on head-coupled systems are delayed by latencies in measuring head position and generating computer graphics. The objectives of this study were 1) to investigate the effects of time delays on head tracking performance; 2) to evaluate the use of an image deflection technique to reduce deleterious effects of delayed images; and 3) to investigate the application of a head position prediction algorithm to enhance the benefits of image deflection. There were significant decreases in head tracking performance when lags of 40 ms or more were added to a system with an inherent 40 ms lag. Lag compensation by image deflection significantly improved tracking performance with lags up to 380 ms. However, by deflecting the delayed image back to its prelag angular position, part of the picture was displaced beyond the edge of the screen. The amount of deflection required was reduced by a simple means of predicting the position of the head before applying deflection. Improved means of predicting head position would further reduce the required image deflection.

Alleviation of side force on tangent-ogive forebodies using passive porosity

Abstract: An experimental investigation to determine the effectiveness of porosity for alleviating side forces on forebodies was conducted in the NASA Langley Research Center 7 x 10 ft High-Speed Wind Tunnel. The study consisted of a comparison of experimental force, moment, and surface pressure results obtained on a fineness ratio 5.0 tangent-ogive porous forebody with 0.020 in. hole diameter and 22 percent porosity with results obtained on a solid forebody. The forebodies were tested with cylindrical afterbodies. The solid forebody was tested with transition grit to simulate fully turbulent conditions and without transition grit to simulate free transition conditions. The extent of porosity on the forebody was varied to determine the extent of porosity needed to alleviate side forces. Static longitudinal and lateral-directional stability and surface pressure data were obtained at Mach numbers of 0.2, 0.5, and 0.8. The angle of attack range was from 5 to 45 deg and roll angles from -90 to 180 deg. The solid forebody exhibited large asymmetries at moderate to high angles of attack causing large side forces and yawing moments. The porous forebody exhibited no significant side forces or yawing moments at any angle of attack tested.

Nonlinear flutter of two-dimensional simply supported symmetric composite laminated plates

Abstract: The nonlinear flutter behavior of a two-dimensional simply supported symmetric composite laminated plate at high supersonic Mach number has been investigated. Yon Karman's large deflection plate theory and quasisteady aerodynamic theory have been employed. Galerkin's method has been used to reduce the governing equations to a system of nonlinear ordinary differential equations in time, which are then solved by a direct numerical integration method. Nonlinear flutter results are presented with the effects of aerodynamic damping, in-plane force, static pressure differential, and anisotropic properties. Results show that the anisotropic properties such as fiber orientation and elastic modulus ratio have significant effects on the behavior of both limit cycle oscillation and chaotic motion.

Rapid prediction of high-alpha unsteady aerodynamics of slender-wingaircraft

Abstract: Most aerospace vehicles, although designed for hypersonic cruise at low angles of attack, often have to perform rapid maneuvers at high angles of attack from low supersonic down to low subsonic speeds. There is, therefore, a need for rapid prediction of the nonlinear high-alpha vehicle dynamics of slender-wing aircraft in that speed range. The present paper describes such a prediction method, which can account for the nonlinear dynamic effects of leading-edge vortices at angles of attack, yaw, and roll. A comparison with existing experimental results shows that the accuracy of the prediction is satisfactory for preliminary design as long as breakdown of the leading-edge vortices does not occur.

Experimental Study of a Low Reynolds Number Tandem Airfoil Configuration

Abstract: Experiments were concluded with two identical Wortmann FX63-137 airfoils in closely coupled tandem configurations at a Reynolds number of 8.5 x 10 4. For the data presented here, the values of the stagger and gap were 1.5 and 0, respectively. The decalage angles were 0 and ±10 deg. Direct measurement of lift, drag, and 1/4-chord pitching moment, as well as static pressure distributions, were acquired for each airfoil. Flow visualization using kerosene smoke was performed to complement the experimental data. The total drag reduction and lift increase resulted in a significant increase in the lift-to-drag ratio for a number of configurations. Nomenclature Aw = wing area Cd = section drag coefficient, D/(Awqx) Ci = section lift coefficient, L/(AwqJ) Cm = section 1/4-chord moment coefficient, Cp = pressure coefficient, l-q/qx c = chord length D = drag force G = gap, \y\lc L = lift force M = 1/4-chord pitching moment q = dynamic pressure, \pU2 Rc = Reynolds number based on chord, Uxc/v St = stagger, x/c (positive when upstream airfoil is above the downstream airfoil) U = flow velocity x = distance in streamwise direction y - distance normal to streamwise directions a = angle of attack 8 = decalage, aua - ada

Navier-Stokes study of supersonic cavity flowfield with passive control

Abstract: A computational investigation of the supersonic turbulent flow past a two-dimensiona l rectangular cavity with passive venting is described The effect of passive venting was included through the use of a porous surface over a vent chamber in the floor of the cavity The passive venting was numerically simulated by the use of a linear form of the Darcy pressure-velocity law The time-accurate solutions of the two-dimensional, Reynoldsaveraged, Navier-Stokes equations were generated using the explicit MacCormack scheme The capability Of the numerical scheme is first demonstrated by the computations of an open and closed cavity without passive venting The results of these computations also provide a reference case for the passive venting computations The effect of passive venting on the closed cavity is then demonstrated and analyzed These results show that the passive venting dramatically changes the closed cavity flow to nearly an open cavity flow The free shear layer formed between the high-speed outer flow and slower inner flow is seen to bridge the cavity completely j resulting in an open cavity flow The passive venting velocities are determined to be less than 5% of the freestream velocities, and largely confined to the upstream and downstream portions of the cavity floor The computational results show good agreement with available experimental data

Thermal Management for a Mach 5 Cruise Aircraft Using Endothermic Fuel

Abstract: The present thermal management system for a carrier-based Mach 5 cruise-capable aircraft whose propulsion system does not entail cryogenic fuels is predicated on the use of the catalytic endothermic reaction of a petroleum-derived hydrocarbon fuel as the heat sink for engine cooling. The insulation of engine flowpath surfaces reduces cooling requirements. The primary elements of this closed-cycle cooling system are a fuel preheater, a catalytic fuel reactor, and engine wall-cooling panels; a silicone-based liquid polymer is used as the coolant. Structural, weight, and thermal analysis results are presented for each of the primary components.

Design and development of a compressible dynamic stall facility

Abstract: A dynamic stall facility offering a unique new capability for studies of compressibility effects on dynamic stall is described. This facility features complete visual access by mounting the test airfoil between optical-quality glass windows which are rotated in unison to produce the oscillating airfoil motion associated with helicopter rotor dynamic stall. By using the density gradients associated with the rapidly changing dynamic stall flow field, this facility permits simultaneous detailed investigation of the flow on the surface as well as in the flow field surrounding airfoils experiencing dynamic stall.

Experimental Study of Vortex Flows over Delta Wings in Wing-Rock Motion

Abstract: The phenomenon of wing rock was investigated using flow visualization in a water tunnel. Models with 70deg, 75-deg, 80-deg, and 85-deg sweeps were tested in a free-to-roll rig and a forced oscillation rig. Wing rock was observed to occur on various wings in the absence of asymmetric vortex liftoff, vortex breakdown, and static hysteresis. These phenomena are therefore not the necessary ingredients for wing rock. The presence of these phenomena, however, can have strong influence on characteristics such as the amplitude and frequency of wing rock.

Multidisciplinary optimization of aeroservoelastic systems using reduced-size models

Abstract: Efficient analytical and computational tools for simultaneous optimal design of the structural and control components of aeroservoelastic systems are presented. The optimization objective is to achieve aircraft performance requirements and sufficient flutter and control stability margins with a minimal weight penalty and without violating the design constraints. Analytical sensitivity derivatives facilitate an efficient optimization process which allows a relatively large number of design variables. Standard finite element and unsteady aerodynamic routines are used to construct a modal data base. Minimum State aerodynamic approximations and dynamic residualization methods are used to construct a high accuracy, low order aeroservoelastic model. Sensitivity derivatives of flutter dynamic pressure, control stability margins and control effectiveness with respect to structural and control design variables are presented. The performance requirements are utilized by equality constraints which affect the sensitivity derivatives. A gradient-based optimization algorithm is used to minimize an overall cost function. A realistic numerical example of a composite wing with four controls is used to demonstrate the modeling technique, the optimization process, and their accuracy and efficiency.

Twenty-five years of aerodynamic research with infrared imaging

Abstract: A review is presented of the commercial introduction of IR imaging systems in the midsixties that has opened the possibilities to visualize viscous interactions between a body and the surrounding airflow by mapping the surface temperature distributions on configurations of interest. The capability of IR imaging systems to produce in real-time thermograms, which can be interpreted both locally and globally, makes them useful for heat transfer and skin-friction aerodynamic studies. Attention is given to IR systems and data processing, supersonic and hypersonic studies, Space Shuttle flight experiments, subsonic and transonic studies, and propulsion studies.

Optimum cruise lift coefficient in initial design of jet aircraft

Abstract: shows the average lift during the ramp-down motion is the dominant factor in the decline of the total average lift at high amax. In the analysis in Ref. 1 a dynamic lift augmentation of AL = 1.5 was assumed, and this resulted in an ALV with a near 20% improvement in straight-and-level range over the conventional baseline vehicle. In Fig. 3c, an augmentation of AL = 1.5 was achieved at the highest pitch rate of K = 0.2 and a0 = 0 deg and for both cases of K = 0.05 and 0.1 for starting angles of aQ = 10 deg. The asssumption of Ref. 1 concerning the existence of augmentation parameters near 1.5 and above would appear then to have some justification. Any benefits that might be gained from augmented lift must be weighed against the penalty levied by the increase in dynamic drag loading. Figure 4 shows the total time-average drag coefficients for each test motion. Figure 4 clearly shows that decreasing amax is advantageous from the standpoint of average drag reduction. From Fig. 3c at the pitch rates of K = 0.05 and 0.1 and starting angles of 10 deg, a value of amax = 35 deg results in nearly optimum lift augmentation with respective AL values of 1.5 and 1.63. The corresponding average drag coefficient is seen in Fig. 4 to be near 0.7, which by some standards is large. Notice, however, that the average drag decreases rapidly with decreasing amax, and for amax = 25 deg and pitch rates of K = 0.05 and 0.1, the average drag coefficient has dropped to 0.3 and 0.24, respectively. At the same time, the lift augmentation remains well above unity and, from Fig. 3c, has values of 1.38 and 1.34, respectively. Thus for the rate of K = 0.1, decreasing amax from 35 to 25 deg results in a drop in the average lift of 18% (though still maintaining significant lift augmentation), whereas the average drag decreases over 60%. In the present study the pitch rate for the ramp-up and ramp-down motions was the same. For a stopping angle of amax = 25 deg the data of Fig. 3a indicate that the average lift during ramp up generally increases with pitch rate, whereas in Fig. 3b the average lift during ramp down decreases with pitch rate. There may then be some advantage in ramping up at a high rate followed by ramping down at a lower rate. In the motions studied in Ref. 1 the rate during pull-up was lower than that for pitch-down. Maintaining acceptable drag loading may be a limiting condition for defining airfoil motions for the purpose of utilizing augmented lift.

Navier-Stokes calculations of inboard stall delay due to rotation

Abstract: Momentum blade element methods, which have historically been used to predict rotor blade thrust, underpredict the power that is required for highly twisted rotors at high thrust levels. During a test of the V-22 tilt rotor blade conducted at the NASA Ames Research Center Outdoor Aerodynamic Research Facility (OARF), measurements of the thrust as a function of collective angle showed that the thrust levels achieved were higher than anticipated and drag levels at the high thrust levels were much lower than predicted. Because the maximum thrusting capability of the rotor in hover usually determines the maximum pay load of the vehicle, it is important to be able to accurately predict high thrust performance characteristics. A three-dimensional Navier-Stokes method was used to investigate the flow phenomenon that causes the increase in lift coefficients and reduction in drag coefficients obtained near the stall of a rotating blade. Comparison of the computational results with the experimental data indicates that this method can be used to predict thrust levels. The Navier-Stokes results indicate that flow separation near the hub is curtailed for a rotating blade compared with the momentum blade element predictions.

Forebody flow control on a full-scale F/A-18 aircraft

Abstract: A full-scale F/A-18 was tested in the 80- by 120-Foot Wind Tunnel at NASA Ames Research Center to measure the effectiveness of pneumatic forebody vortex control devices. By altering the forebody vortex flow, yaw control can be maintained to angles of attack greater than 50 deg. Two forebody vortex control devices were tested: a discrete circular jet and a tangential slot. The tests were conducted for angles of attack between 25 and 50 deg, and angles of sideslip from 0 to +/- 15 deg. The Reynolds number based on wing mean aerodynamic chord ranged from 4.5 x 10 exp 6 to 12.0 x 10 exp 6. The time-averaged side forces and yawing moments, along with both time-averaged and time-dependent pressures on the forebody of the aircraft are presented here for various configurations. Of particular interest was the results that the tangential slot blowing had a greater effect on the yawing moment than the discrete circular jet. Additionally, it was found that blowing very close to the radome apex was not as effective as blowing slightly farther aft on the radome, and that a 16-inch slot was more effective than either an 8- or 48-inch long slot.