Showing papers on "Pitching moment published in 1994"

2-D and 3-D oscillating wing aerodynamics for a range of angles of attack including stall

Abstract: A comprehensive experimental investigation of the pressure distribution over a semispan wing undergoing pitching motions representative of a helicopter rotor blade was conducted Testing the wing in the nonrotating condition isolates the three-dimensional (3-D) blade aerodynamic and dynamic stall characteristics from the complications of the rotor blade environment The test has generated a very complete, detailed, and accurate body of data These data include static and dynamic pressure distributions, surface flow visualizations, two-dimensional (2-D) airfoil data from the same model and installation, and important supporting blockage and wall pressure distributions This body of data is sufficiently comprehensive and accurate that it can be used for the validation of rotor blade aerodynamic models over a broad range of the important parameters including 3-D dynamic stall This data report presents all the cycle-averaged lift, drag, and pitching moment coefficient data versus angle of attack obtained from the instantaneous pressure data for the 3-D wing and the 2-D airfoil Also presented are examples of the following: cycle-to-cycle variations occurring for incipient or lightly stalled conditions; 3-D surface flow visualizations; supporting blockage and wall pressure distributions; and underlying detailed pressure results

An active flexible wing multi-disciplinary design optimization method

Abstract: This paper discusses Active Flexible Wing (AFW) technology and describes how it differs from conventional wing design. The benefits of AFW are briefly described. Two design procedures which aid in the design and optimization of an AFW wing are described in detail. The first procedure is for the design of an AFW control system on an existing wing. This procedure optimizes control surface positions to maximize air vehicle maneuverability, without exceeding structural limits. The second procedure is for the design of a new wing using AFW technology. This procedure simultaneously couples aerodynamic, structural, and external load designs. The process optimizes a wing structure and control surface positions for minimum weight and drag, while satisfying structural constraints. = buckling constraints = bending moment = drag of case i = hinge moment = roll moment of load case i = lift of load case i = pitching moment of load case i = roll rate = torsion moment = twist and camber variables (e.g., wing jig shape design) = structural design variables = flutter constraints = structural weight = control surface positions and air vehicle flight angle design variables = roll axis inertia = roll acceleration = stress constraints of load case i {a} = vector of rigid aerodynamic panel deflections {!} = lift vector on aerodynamic panels * Project Engineer, Advanced Aircraft Member AIAA [ A ] = aerodynamic panel lift due to alpha influence coefficient matrix [ B ] = aerodynamic to structural transformation matrix [ K ] = global stiffness matrix [ d ~ / dtk] = derivative of the global structural stiffness matrix with respect to structural design variables [ S I ] =structural flexibility matrix (in units of deflection per force) on aerodynamic model [ S A ] = structural flexibility matrix (in units of rotation per force) on aerodynamic model

Navier-Stokes simulations of Orbiter aerodynamic characteristics including pitch trim and bodyflap

Abstract: An analysis of the longitudinal aerodynamics of the shuttle orbiter in the hypersonic flight regime is made through the use of computational fluid dynamics. Particular attention is given to establishing the cause of the 'pitching moment anomaly,' which occurred on the orbiter's first flight, and to computing the aerodynamics of a complete orbiter configuration at flight conditions. Data from ground-based facilities as well as orbiter flight data are used to validate the computed results. Analysis shows that the pitching moment anomaly is a real-gas chemistry effect that was not simulated in ground-based facilities, which used air as a test gas. Computed flight aerodynamics for the orbiter are within 5% of the measured flight values and trim bodyflap deflections are predicted to within 10%.

Aerodymics of the Shuttle Orbiter at high altitudes

Abstract: The high-altitude/high-Knudsen number aerodynamics of the Shuttle Orbiter are computed from Low-Earth Orbit down to 100 km using three-dimensional direct simulation Monte Carlo and free molecule codes. Results are compared with Blanchard's latest Shuttle aerodynamic model, which is based on in-flight accelerometer measurements, and bridging formula models. Good comparison is observed, except for the normal force and pitching moment coefficients. The present results were obtained for a generic Shuttle geometry configuration corresponding to a zero deflection for all control surfaces.

Ancillary aerodynamic structures for an unmanned aerial vehicle having ducted, coaxial counter-rotating rotors

Abstract: An unmanned aerial vehicle (UAV) (100) having a toroidal fuselage (120) and a rotor assembly (170) including counter-rotating rotors coaxially mounted with respect to the toroidal fuselage incorporates ancillary aerodynamic structures (18) having a cambered airfoil profile to provide a nose-down pitching moment to counteract the nose-up pitching moment generated by airflow over the toroidal fuselage during forward translational flight of the UAV. The ancillary aerodynamic structures are symmetrically mounted in combination with the lateral sides of the toroidal fuselage so that the centers of lift are located aftwardly of the fuselage axis of the toroidal fuselage in forward translational flight modes. In a first embodiment, the ancillary aerodynamic structures (18) are fixedly mounted in combination with the toroidal fuselage (10) at a predetermined angle of incidence. In a second embodiment, the ancillary aerodynamic structures (19) are rotatably mounted in combination with the toroidal fuselage (10') to provide variable incidence ancillary aerodynamic structures for the UAV.

Pitch-up characteristics for hsct class planforms: survey and estimation

Abstract: Low aspect ratio, highly-swept cranked delta and arrow wing planforms are often proposed for high-speed civil transports. These wing planforms offer low supersonic drag without suffering greatly from low liftldrag ratios in low-speed flight. They can, however, suffer from pitch-up at modest angles of attack (as low as 5' angle of attack) during low-speed flight due to leading edge vortex influence, flow separation and vortex breakdown. This paper describes an investigation conducted to study past research on the longitudinal aerodynamic characteristics of highlyswept cranked wing planforms and a new method to estimate pitch-up. The survey of past research placed emphasis on 1) understanding the problem of pitch-up, and 2) ascertaining the effects of leading and trailing edge flaps. The estimation method uses a vortex lattice method to calculate the inviscid flow solution. Then, the results are adjusted to account for flow separation on the outboard wing section by imposing a limit on the equivalent 2-D sectional lift coefficient. The method offers a means of making low cost estimates of the non-linear pitching moment characteristics of slender, cranked arrow wing configurations. Numerous comparisons with data are included.

Three-dimensional flow calculations for a projectile with standard and dome bases

Abstract: Test firings of the 155-mm XM825 artillery projectile have shown that its flight performance was affected by configurational changes to the base cavity. This was an unexpected result, and a clear understanding of why these changes affected the flight behavior did not exist. A computational study has been made for the two different base-cavity configurations which were flight tested. Flowfield computations have been performed at 0.8

Dynamic-Stall and Structural-Modeling Effects on Helicopter Blade Stability with Experimental Correlation

Abstract: The effects of blade and root-flexure elasticity and dynamic stall on the stability of hingeless rotor blades are investigated. The dynamic stall description is based on the ONERA models of lift, drag, and pitching moment. The structural analysis is based on three blade models that range from a rigid flap-lag model to two elastic flap-lag-torsion models, which differ in representing root-flexure elasticity. The predictions are correlated with the measured lag damping of an experimental isolated three-blade rotor; the correlation covers rotor operations from near-zero-thrust conditions in hover to highly stalled, high-thrust conditions in foward flight. That correlation shows sensitivity of lag-damping predictions to structural refinements in blade and root-flexure modeling. Moreover, this sensitivity increases with increasing control pitch angle and advance ratio. For high-advance-ratio and high-thrust conditions, inclusion of dynamic stall generally improves the correlation.

Kinematics and aerodynamics of the velocity vector roll

Abstract: The velocity-vector roll is defined as an angular rotation of an airplane about its instantaneous velocity vector, constrained to be performed at constant angle of attack (AOA), no sideslip, and constant velocity. Consideration of the aerodynamic force equations leads to requirements for body-axis yawing and pitching rotations that must be present to satisfy these constraints. Here, the body-axis rotations and the constraints are used in the moment equations to determine the aerodynamic moments required to perform the velocity-vector roll. The total aerodynamic moments, represented in the reference body-axis coordinate system, are then analyzed to determine the conditions under which their maxima occur. It is shown, for representative tactical airplanes, that the conditions for maximum pitching moment are strongly a function of the orientation of the airplane, occurring at about 90 deg of bank in a level trajectory. Maximum required pitching moment occurs at peak roll rate and is achieved at an AOA in excess of 45 deg. The conditions for maximum rolling moment depend on the value of the roll mode time constant. For a small time constant (fast response) the maximum rolling moment occurs at maximum roll acceleration and zero AOA, largely independent of airplane orientation; for a large time constant, maximum required rolling moment occurs at maximum roll rate, at maximum AOA, and at 180 deg of bank in level flight. The maximum yawing moment occurs at maximum roll acceleration and maximum AOA and is largely independent of airplane orientation. Results are compared with those obtained using conventional assumptions of zero pitch and yaw rates and show significant improvement, especially in the prediction of maximum-pitching-moment requirements.

Adaptive computations of flow around a delta wing with vortex breakdown

Abstract: An adaptive unstructured mesh solution method for the three-dimensional Euler equations was used to simulate the flow around a sharp edged delta wing. Emphasis was on the breakdown of the leading edge vortex at high angle of attack. Large values of entropy, which indicate vortical regions of the flow, specified the region in which adaptation was performed. The aerodynamic normal force coefficients show excellent agreement with wind tunnel data measured by Jarrah, and demonstrate the importance of adaptation in obtaining an accurate solution. The pitching moment coefficient and the location of vortex breakdown are compared with experimental data measured by Hummel and Srinivasan, showing good agreement in cases in which vortex breakdown is located over the wing.

Aerodynamic characteristics of strake vortex flaps on a strake-wing configuration

Abstract: The effects of strake vortex flaps (SVFs), determined experimentally, on the aerodynamic characteristics of a strake-wing configuration are presented. SVFs may improve cruise performance over that of a planar strake by partially unloading the strake and generating a thrust component. The magnitude of the nose-up pitching moment may also be reduced by unloading the strakes. The use of SVFs as lateral control devices is also investigated. The results indicate that cruise performance is improved for all vortex flap deflection angles compared to planar strakes, with minimal or no concomitant penalty at high lift coefficients. Positive pitching moment is also substantially reduced. Differentially deflected strakes appear to be capable of generating significant rolling moments at high angles of attack.

Wind Tunnel Results of Pneumatic Forebody Vortex Control Using Rectangular Slots a Chined Forebody

Abstract: A subsonic wind tunnel investigation of pneumatic vortex flow control on a chined forebody using slots was accomplished at a dynamic pressure of 50 psf resulting in a R(n)/ft of 1.3 x 10(exp 6). Data were acquired from angles of attack ranging from -4deg to +34deg at side slips of +0.4deg and +10.4deg. The test article used in this study was the 10% scale Fighter Lift and Control (FLAC) advanced diamond winged, vee-tailed fighter configuration. Three different slot blowing concepts were evaluated; outward, downward, and tangential with ail blowing accomplished asymmetrically. The results of three different mass flows (0.067, 0.13, and 0.26 lbm/s; C(sub mu)'s of less than or equal to 0.006, 0.011. and 0.022 respectively) were analyzed and reported. Test data are presented on the effects of mass flows, slot lengths and positions and blowing concepts on yawing moment and side force generation. Results from this study indicate that the outward and downward blowing slots developed yawing moment and side force increments in the direction opposite of the blowing side while the tangential blowing slots generated yawing moment and side force increments in the direction towards the blowing side. The outward and downward blowing slots typically produced positive pitching moment increments while the tangential blowing slots typically generated negative pitching moment increments. The slot blowing nearest the forebody apex was most effective at generating the largest increments and as the slot was moved aft or increased in length, its effectiveness at generating forces and moments diminished.

A simple analytical aerodynamic model of Langley Winged-Cone Aerospace Plane concept

Abstract: A simple three DOF analytical aerodynamic model of the Langley Winged-Coned Aerospace Plane concept is presented in a form suitable for simulation, trajectory optimization, and guidance and control studies The analytical model is especially suitable for methods based on variational calculus Analytical expressions are presented for lift, drag, and pitching moment coefficients from subsonic to hypersonic Mach numbers and angles of attack up to +/- 20 deg This analytical model has break points at Mach numbers of 10, 14, 40, and 60 Across these Mach number break points, the lift, drag, and pitching moment coefficients are made continuous but their derivatives are not There are no break points in angle of attack The effect of control surface deflection is not considered The present analytical model compares well with the APAS calculations and wind tunnel test data for most angles of attack and Mach numbers

Application of two procedures for dual-point design of transonic airfoils

Abstract: Two dual-point design procedures were developed to reduce the objective function of a baseline airfoil at two design points. The first procedure to develop a redesigned airfoil used a weighted average of the shapes of two intermediate airfoils redesigned at each of the two design points. The second procedure used a weighted average of two pressure distributions obtained from an intermediate airfoil redesigned at each of the two design points. Each procedure was used to design a new airfoil with reduced wave drag at the cruise condition without increasing the wave drag or pitching moment at the climb condition. Two cycles of the airfoil shape-averaging procedure successfully designed a new airfoil that reduced the objective function and satisfied the constraints. One cycle of the target (desired) pressure-averaging procedure was used to design two new airfoils that reduced the objective function and came close to satisfying the constraints.

Unsteady boundary-layer separation on a pitching airfoil

Abstract: This research IS a theoretical and computational study of unsteady boundary-layer separation from a two-diinensional thin airfoil immersed in a uniform flow stream when the' angle of attack is varied as a function of time. The flow is considered to be high speed in the sense that the Reynolds number is large; limiting situations corresponding to the case Re -+ ~ are considered. As the angle of attack is suddenly changed beyond a critical angle, unsteady processes initiate in the boundary layer in the leading edge region of the airfoil that eventually lead to a phenomen.on known as dynamic stall. As these processes develop, a recirculating eddy is first formed in the boundary layer, and soon the flow near the top surface of the airfoil is focused into a region that progressively narrows in the streamwise direction. This leads to tlie eventual development of a separation singularity in the solution of the boundary-layer equations and a strong viscous-inviscid interaction wherein the boundary layer erupts Into the outer inviscid flow. Computatio'nal results are obtained for the angle of attack increasing linearly with time, and these illustrate the events leading up to the ejection of bou~dary-Iayer vorticity into the upstream external flow at a finite time. Separation develops as a thin 'I "spike" extending from' the upper surface of the airfoil on the upstream side of the recirculating zone. The time required to reach separation, starting from an altitude of zero incidence, is evaluated as a function of pitching rate. The effect of reversing the pitching rate is considered; it is found that under certain circumstances, leading edge separation can be delayed.

Propulsion-induced aerodynamic effects measured with a full-scale STOVL model

Abstract: Propulsion-induced aerodynamic interference effects are presented for a full-scale, powered, STOVL fighter aircraft model. The ejector-lift/vectored thrust configuration, designated the E-7A, was tested at low speed in the 40by 80and 80by 120-ft wind tunnels of the National Full-Scale Aerodynamics Complex located at NASA Ames Research Center. Aerodynamic effects on vehicle lift, drag, and pitching moment are presented over a range of effective velocities for simultaneous operation of all lifting jets. The jet/airframe interactions for separate operation of the lifting ejector system and vector able ventral nozzle are also presented. Ejector and engine inlet momentum effects were fully simulated in these full-scale, powered tests. The jet thrust vector angle of the ventral nozzle was varied to simulate transition from hover to wingborne flight modes. When the lifting ejector system and ventral nozzle are operated simultaneously, the induced effects on lift decrease as the thrust vector angle of the ventral nozzle approaches the horizontal. A negative increment in drag is produced over a narrow portion of the transition speed range when the ejectors and ventral nozzle are operated together. Aerodynamic induced effects of the ejector system measured at full-scale compare well with the small-scale data. Changes in lift and pitching moment due to ventral nozzle operation are smaller at full scale.

Aerodynamic test and analysis of wrap-around fins at supersonic Mach numbers utilizing design of experiments

Abstract: Cxb =body axial force coefficient at zero angle of The aerodynamics of wrap-around fins (WAFs) are influenced by various factors, including Mach number, Icngth-to-diameter (LD) ratio, fin sweep angle, fin root chord length, and fin thickness. This paper presents the rcsults of an investigation into the effects of these factors on WAFs at supersonic Mach numbers. This effort was conducted to furthcr define WAF flight characteristics. The design of experiments (DOE) test approach was used to detcrmine the feasibility of DOE in extracting aerodynamic coefficients from spark range testing, as well as to provide a measure of cost savings. The results of the free flight data reduction and predictions arc summarized and compared to the DOE model and to engineering code predictions. These results show the effect of each factor on the tin axial force coefficient, fin normal force, and total pitching moment derivative coefficients. In addition, an analysis was completed for thc roll and side force. The results of this analysis show good agrcement with the DOE modcl and engineering prediction codes. This study also shows that DOE methodology is an efficient way to predict aerodynamic coefficients for spark range testing. -i

A method for the determination of the effect of propeller slipstream on static longitudinal stability and control of multi-engined aircraft

Abstract: The method makes possible the determination of slipstream eftects on wing lift-off pitching moment and flow cond itions at the horizontal tail

A computational study of the aerodynamics of a body immersed in a supersonic wake

Abstract: A computational investigation of a body submerged within the supersonic wake was performed to determine the aerodynamic characteristics and possible flight trajectory of the trailing body. The effects of two leading body geometries, wake profile perturbations, laminar versus turbulent wakes, and various placements of the trailing body were examined. At large separation distances, which placed the trailing body downstream of the wake neck and coaxial with the incoming wake, a toroidal vortex was formed at the nose of the trailing body. The size of the toroidal vortex was dependent on the relative strengths of the reattachment pressure and the stagnation pressure behind the shock at the centerline. When the trailing body was radially displaced, the toroidal vortex was replaced with a horseshoe vortex which diminished in size with increasing displacement. Drag coefficients for the trailing body with coaxial alignment increased with increasing separation distance for both leading body geometries. However, they were highly dependent on these leading geometries. For off-axis alignments, the drag increased with exposure to freestream conditions. The variation of lift coefficient with asymmetry indicated that the trailing body will tend to move towards the wake centerline, but the pitching moment suggested that an angle of attack would develop.

Flutter stability of a detuned cascade in subsonic compressible flow

Abstract: A mathematical model is developed to predict the unsteady aerodynamics of a detuned two-dimensional flat plate cascade in subsonic compressible flow. Aerodynamic detuning is introduced by nonuniform circumferential spacing and chordwise offset. Combined aerodynamic-structural detuning is accomplished by replacing alternate airfoils with splitter blades. A torsion mode stability analysis that considers aerodynamic and combined aerodynamic-structural detuning is developed by combining the unsteady aerodynamic model with a single degreeof-freedom structural model. The effect of these detuning techniques on flutter stability is then demonstrated by applying this model to a baseline unstable 12-bladed rotor and detuned variations of this rotor. This study demonstrates that detuning is a viable passive flutter control technique. Nomenclature A = cascade A airfoil index a — freestream speed of sound B - cascade B airfoil index CL = lift coefficient CM - moment coefficient C, = ratio of chord length of cascade B to cascade A CA = chord length of cascade A CB = chord length of cascade B ea = elastic axis location k - reduced frequency a>c/W n = airfoil index os - chordwise offset of cascade B relative to cascade A

Aerodynamic Characteristics of Aircraft Under Side Gust

Abstract: Wind-tunnel tests were conducted to investigate the aerodynamic characteristics of an aircraft when exposed to side gust resulting from microburst windshear. External airstream was supplied through the test section side wall such that it was normal to the flight pathline and parallel to the ground. The effect of side gust was studied in terms of the model performance in undisturbed flight compared to its response during side flow. The model aerodynamic characteristics were examined at several sideslip angles as well as during straight flight. The wind-tunnel results of the straight-flight mode indicated that as side gust was introduced, there was a slight but consistent drop in lift, considerable accumulation of drag, increase in nose-up pitching moment, and constant lateral drift. The loss of lift became more appreciable at negative sideslip angles. In remedy, this situation was associated with a minimal amount of added-on drag. It was demonstrated that such loss of lift could be minimized by yawing the aircraft in the opposite direction, i.e., as the gust direction became closer to being head wind. However, this configuration of positive sideslip yielded higher accumulation of drag. Severe nose-up pitching moment was experienced during side gust exposure at negative sideslip orientations. It is found that while the trim condition was displaced, the stability margin was retained unaffected.

Euler and Navier-Stokes solutions for supersonic flow around a complex missile

Abstract: The numerical solution of the Euler and Navier-Stokes system of equations and experimental data for the supersonic flow around a complex missile configuration are compared. The Euler and Navier-Stokes equations are solved with the DLR finite-volume method CEVCATS using multiblock mesh structure and multigrid acceleration technique. The numerical solution of the flow equations for this missile with eight fins has provided a flow with a rich vortex structure and manifold three-dimensional effects. The aerodynamic coefficients for normal-force and pitching moment obtained with the Euler solution are in good agreement with experimental values. Therefore, the position of the aerodynamic center is wellpredicted by the inviscid solution. However for the flow close to the surface a Navier-Stokes solution is required to obtain agreement with the experimental results.

Influence of cavitation on blade characteristics

Abstract: An experimental study of flow around a blade with a modified NACA 4418 profile was conducted in a water tunnel that also enables control of the cavitation conditions within it. Pressure, lift force, drag force and pitching moment acting on the blade were measured for different blade angles and cavitation numbers, respectively. Relationships between these parameters were elaborated and some of them are presented here in dimensionless form. The analysis of results confirmed that cavitation changes the pressure distribution significantly. As a consequence, lift force and pitching moment are reduced, and the drag force is increased. When the cavitation cloud covers one side of the blade and the flow becomes more and more vaporous, the drag force also begins to decrease. The cavity length is increased by increasing the blade angle and by decreasing the cavitation number.

高速走行車両の縦運動に及ぼす空気力の影響 : 第1報,空力特性の定式化

Abstract: The objective of this study is to describe aerodynamic effects on vehicle dynamics. This paper concentrates on the mathematical formulations and the experimental certification of the longitudinal single-degree-of-freedom motion. The lift and pitching moment coefficients of a scaled car model are measured with the deviation of the body attitude. Quasi-steady aerodynamic modelling reduces the formulation to a linearized system. The transient responses simulating bouncing motion under high-speed conditions are calculated and found to be in good agreement with the results obtained in the wind tunnel tests.

Low-speed longitudinal aerodynamic characteristics of a flat-plate planform model of an advanced fighter configuration

Abstract: A flat-plate wind tunnel model of an advanced fighter configuration was tested in the NASA LaRC Subsonic Basic Research Tunnel and the 16- by 24-inch Water Tunnel. The test objectives were to obtain and evaluate the low-speed longitudinal aerodynamic characteristics of a candidate configuration for the integration of several new innovative wing designs. The flat plate test allowed for the initial evaluation of the candidate planform and was designated as the baseline planform for the innovative wing design study. Low-speed longitudinal aerodynamic data were obtained over a range of freestream dynamic pressures from 7.5 psf to 30 psf (M = 0.07 to M = 0.14) and angles-of-attack from 0 to 40 deg. The aerodynamic data are presented in coefficient form for the lift, induced drag, and pitching moment. Flow-visualization results obtained were photographs of the flow pattern over the flat plate model in the water tunnel for angles-of-attack from 10 to 40 deg. The force and moment coefficients and the flow-visualization photographs showed the linear and nonlinear aerodynamic characteristics due to attached flow and vortical flow over the flat plate model. Comparison between experiment and linear theory showed good agreement for the lift and induced drag; however, the agreement was poor for the pitching moment.

Subsonic static and dynamic stability characteristics of a NASP configuration

Abstract: Wind-tunnel tests of a National Aero-Space Plane configuration were conducted in the NASA Langley Research Center (LaRC) 7- by 10-Foot High Speed Tunnel. The model used is a LaRC designed blended body configuration. Static and dynamic stability characteristics were measured at Mach numbers of 0.3, 0.6, and 0.8. In addition to tests of the baseline configuration, component buildup tests with a canard surface and a body flap were conducted. The baseline configuration showed positive static stability except at the higher angles of attack at 0.8 Mach number. The baseline configuration has positive damping about all three axes. There was generally good agreement between the in-phase dynamic parameters and the corresponding static data. Also included are comparisons of the experimental damping parameters with results from the engineering predictive code Aerodynamic Preliminary Analysis System (APAS). The APAS damping predictions are good for the roll mode, and only fair for the pitch and yaw modes. Nomenclature Static longitudinal data are referred to the stability axis system and the static lateral stability data are referred to the body axis system aligned with water line zero (WL 0) as shown in Fig. 1. The dynamic stability data presented are referred to the body axis system inclined 4.25 deg relative to WL 0. Both the static and dynamic stability balances and stings were offset 4.25 deg from the model centerline to avoid any alteration of the upper aft surface of the model. The origin of the axes was located to correspond to the moment reference position shown in Fig. 2. The model reference length for the pitching moment coefficients is the body reference length of 33.6 in. (see Fig. 2). For the yawing and rolling moment coefficients the reference length is the overall wing span of 12.00 in. The reference area of 161.04 in.2 is the theoretical wing planform area (including elevens) with the wing leading edges projected to the centerline of the vehicle. The area of the body flap at the rear of the body and the canard are not included in the reference area. In the following Nomenclature list a dot over a quantity indicates a first derivative with respect to time. b

Effects on non-linearities on aircraft poststall motion

Abstract: The poststall maneuverability controlled by thrust vectoring has become one of the important aspects of new fighter development projects. In simplified case, the motion of aircraft can be described by 6DOF nonlinear system. The lecture deals with the longitudinal motion of poststall maneuverable aircraft. The investigation made about the effects of non-linearities in aerodynamic coefficients having considerable non-linearities and hysteresisis an the poststall motions. There were used some different models of aerodynamic coefficients. The results of investigation have shown that the poststall domain of vectored aircraft can be divided into five different pHs in field of thrust - pitch vector angle, and the chaotic motions of aircraft can be found at the different frequencies of thrust deflection. There were defined an unstable right domain with an unstable oscillation and a field of overpulling at poststall motion. The certain frequency chaotic attractors were got at frequencies of Oxitation between the 0.15 and 0.65 rad/sec. The pitching moment derivatives had the big influence on the chaotic motions, while the lift coefficient derivatives bad the reasonable effects, only.

Sensitivity of lag-damping correlations to structural and aerodynamic approximations of isolated experimental rotors in forward flight

Abstract: The predictions of regressive lag-mode damping levels are correlated with the database of an isolated, soft-inplane, three-blade rotor operated untrimmed. The database was generated at the Army Aeroflightdynamics Directorate at Ames. The correlation covers a broad range of data, from near-zero thrust conditions in hover to high-thrust and highly stalled conditions in forward flight with advance ratio as high as 0.55 and shaft angle as high as 20 degrees. In the experimental rotor, the airfoil or blade portion has essentially uniform mass and stiffness distributions, but the root flexure has highly nonuniform mass and stiffness distributions. Accordingly, the structural approximations refer to four models of root-flexure-blade assembly. They range from a rigid flap-lag model to three elastic flap-lag-torsion models, which differ in modeling the root flexure. The three models of root-flexure are: three root springs in which the bending-torsion couplings are fully accounted for; a finite-length beam element with some average mass and stiffness distributions such that the fundamental frequencies match those of the experimental model; and accurate modal representation in which the actual mass and stiffness distributions of the experimental root-flexure-blade assembly are used in calculating the nonrotating mode shapes. The four models of root-flexure-blade assembly are referred to as the rigid flap-lag model, spring model, modified model and modal model. For each of these four models of the root-flexure-blade assembly, the predictions are based on the following five aerodynamic theories: ear theory, which accounts for large angle-of-attack and reverse-flow effects on lift, and has constant drag and pitching moment; quasisteady stall theory, which includes quasisteady stall lift, drag and pitching moment characteristics of the airfoil section, dynamics stall theory, which uses the ONERA dynamic stall models of lift, drag and pitching moment; dynamic wake theory, which is based on a finite-state three-dimensional wake model and includes all wake effects including both shed and trailing vorticity; and dynamics and wake theory, which combines both dynamic stall theory and dynamic wake theory and is a relatively complete aerodynamic representation.