Showing papers on "Pitching moment published in 2000"

Low Reynolds Number Aerodynamics of Low-Aspect-Ratio, Thin/Flat/Cambered-Plate Wings

Abstract: The design of micro aerial vehicles requires a better understanding of the aerodynamics of small low-aspect-ratio wings An experimental investigation has focused on measuring the lift, drag, and pitching moment about the quarter chord on a series of thin flat plates and cambered plates at chord Reynolds numbers varying between 60,000 and 200,000 Results show that the cambered plates offer better aerodynamic characteristics and performance It also appears that the trailing-edge geometry of the wings and the turbulence intensity in the wind tunnel do not have a strong effect on the lift and drag for thin wings at low Reynolds numbers Moreover, the results did not show the presence of any hysteresis, which is usually observed with thick airfoils/wings

Aerodynamic Measurements at Low Raynolds Numbers for Fixed Wing Micro-Air Vehicles

Abstract: : A description of the micro-air vehicle (MAV) concept and design requirements is presented These vehicles are very small and therefore operate at chord Reynolds numbers below 200,000 where very little data is available on the performance of lifting surfaces, ie, airfoils and low aspect-ratio wings This paper presents the results of a continuing study of the methods that can be used to obtain reliable force and moment data on thin wings in wind and water tunnels To this end, a new platform force and moment balance, similar to an already existing balance, was designed and built to perform lift, drag and moment measurements at low Reynolds numbers Balance characteristics and validation data are presented Results show a good agreement between published data and data obtained with the new balance Results for lilt, drag and pitching moment about the quarter chord with the existing aerodynamic balance on a series of thin flat plates and cambered plates at low Reynolds numbers are presented They show that the cambered plates offer better aerodynamic characteristics and performance Moreover, it appears that the trailing-edge geometry of the wings and the turbulence intensity up to about 1% in the wind tunnel do not have a strong effect on the lilt and drag for thin wings at low Reynolds numbers However, the presence of two endplates for two-dimensional tests and one endplate for the semi-infinite tests appears to have an undesirable influence on the lift characteristics at low Reynolds numbers

Aerodynamic and Aeroacoustic Optimization of Rotorcraft Airfoils via a Parallel Genetic Algorithm

Abstract: A parallel genetic algorithm (GA) methodology was developed to generate a family of two-dimensional airfoil designs that address rotorcraft aerodynamic and aeroacoustic concerns The GA operated on 20 design variables, whichconstitutedthecontrolpointsforasplinerepresentingtheairfoilsurfaceTheGAtookadvantageofavailable computer resources by operating in either serial mode, where the GA and function evaluations were run on the same processor or “ manager/worker” parallel mode, where the GA runs on the manager processor and function evaluations areconducted independently on separate workerprocessors The multiple objectives of this work were to minimizethedrag and overall noiseof the airfoil Constraintswereplaced on liftcoefe cient, moment coefe cient, andboundary-layerconvergenceTheaerodynamicanalysiscodeXFOILprovidedpressureandsheardistributions in addition to liftand drag predictions Theaeroacousticanalysis code, WOPWOP, provided thicknessand loading noise predictions The airfoils comprising the resulting Pareto-optimal set exhibited favorable performance when compared with typical rotorcraft airfoils under identical design conditions using the same analysis routines The relationship between the quality of results and the analyses used in the optimization is also discussed The new airfoil shapes could provide starting points for further investigation

Method of reducing a nose-up pitching moment in a ducted rotor unmanned aerial vehicle

Abstract: A method for reducing a nose-up pitching moment in an unmanned aerial vehicle (10) during forward flight. The unmanned aerial vehicle includes counter-rotating rotor assemblies (38, 40) that are mounted within a duct (18). Each rotor assembly (38, 40) includes a plurality of rotor blades. The method involves adjusting the rotor blades to have substantially zero pitch. Then rotating the rotor asemblies (38, 40) to produce a virtual plane (62) across the duct (18). The virtual plane (62) is operative for substantially deflecting air (70) passing over the fuselage (44) away from the duct (18). In one embodiment of the invention, the method involves the further step of obstructing at least a portion of the bottom of the duct (18) to inhibit air (70) that is flowing across the bottom of the duct from passing into the duct (18).

Aerodynamic and aeroacoustic optimization of rotorcraft airfoils via a parallel genetic algorithm

Abstract: A parallel genetic algorithm (GA) methodology was developed to generate a family of two-dimensional airfoil designs that address rotorcraft aerodynamic and aeroacoustic concerns. The GA operated on 20 design variables, which constituted the control points for a spline representing the airfoil surface. The GA took advantage of available computer resources by operating in either serial mode, where the GA and function evaluations were run on the same processor or manager/worker parallel mode, where the GA runs on the manager processor and function evaluations are conducted independently on separate worker processors. The multiple objectives of this work were to minimize the drag and overall noise of the airfoil. Constraints were placed on lift coefficient, moment coefficient, and boundary-layer convergence. The aerodynamic analysis code XFOIL provided pressure and shear distributions in addition to lift and drag predictions. The aeroacoustic analysis code, WOPWOP, provided thickness and loading noise predictions. The airfoils comprising the resulting Pareto-optimal set exhibited favorable performance when compared with typical rotorcraft airfoils under identical design conditions using the same analysis routines

A backstepping design for flight path angle control

Abstract: A nonlinear approach to flight path angle control is presented. Using backstepping, a globally stabilizing control law is derived. Although the nonlinear nature of the lift force is considered, the pitching moment to be produced is only linear in the measured states. Thus, the resulting control law is much simpler than if feedback linearization had been used. The free parameters that spring from the backstepping design are used to achieve a desired linear behavior around the operating point.

Feedback Control of a Space Vehicle With Unactuated Fuel Slosh Dynamics

Abstract: We develop a mathematical model that describes the accelerating flight of a spacecraft in a fixed plane. The spacecraft is represented as a rigid body and fuel slosh dynamics are included using a common pendulum model. The control inputs are defined by a transverse body fixed force and a pitching moment about the center of mass of the spacecraft; the slosh dynamics are assumed to be unactuated. The model is placed in the form of a nonlinear control system that allows for the study of planar vehicle maneuvers. We develop a nonlinear feedback controller that stabilizes a relative equilibrium corresponding to suppression of the transverse, pitch, and slosh dynamics. This controller is a significant extension of what has been done previously, since it simultaneousl y controls both the rigid vehicle motion and the fuel slosh dynamics.

Robust Vortex Control of a Delta Wing by Distributed Microelectromechanical-Systems Actuators

Abstract: Micromachined actuators have been used successfully to control leading-edge vortices of a delta wing by manipulating the thin boundary layer before flow separation. In an earlier work, we demonstrated that small disturbances generated by these microactuators could alter large-scale vortex structures and consequently generate appreciable aerodynamic moments along all three axes for flight control. In the current study, we explored the possibility of independently controlling these moments. Instead of using a linearly distributed array of microactuators covering the entire leading edge as done in the previous study, we applied a shorter array of actuators located on either the forward or the rear half-section of the leading edge. Both one- and two-sided control configurations have also been investigated. Data showed that the pitching moment could be generated independently by appropriate actuation of the microactuators. To understand the interaction between the microactuators and leading-edge vortices, we conducted surface pressure distribution, direct force measurements, and flow visualization experiments. We investigated the effects of microactuators on the vortex structure, especially vortex core location

Aerodynamic forces and flow structures of an airfoil in some unsteady motions at small Reynolds number

Abstract: The aerodynamic forces and flow structures of a NACA 0012 airfoil in some unsteady motions at small Reynolds number (Re=100) are studied by numerically solving the Navier-Stokes equations. These motions include airfoil acceleration and deceleration from one translational speed to another and rapidly pitching up in constant freestream (equivalent to pitching up during translational motion at constant speed). It is shown that at small Reynolds number (Re=100), when the airfoil is performing fast acceleration or deceleration from one speed to another, a large aerodynamic force can be generated during and for a time period after the acceleration or deceleration; a large aerodynamic force can also be generated when the airfoil is performing a fast pitching motion in a constant freestream. In these fast unsteady motions, an airfoil in low Re flow can produce a large aerodynamic force as effective as in large Re flow, or the effect of unsteady motion dominates the Reynolds number effect. During the fast unsteady motion of the airfoil, new layers of strong vorticity are formed near the upper and lower surfaces of the airfoil under the previously existing thick vorticity layers, and it is the generation and motion of the new vorticity layers that is mainly responsible for the generation of the large aerodynamic force; the large-scale structure and movement of the newly produced vorticity layers are similar to that of high Re flow.

Numerical Simulation of Transient Jet-Interaction Phenomenology in a Supersonic Freestream

Abstract: The objective is to evaluate the transient effects of a reaction control jet on the aerodynamic performance of a genericinterceptormissileoperating at supersonice ightconditions.Three-dimensionalcomputationsofthehighly turbulent e owe eld produced by a pulsed, supersonic, lateral-jet control thruster interacting with the supersonic freestream and missile boundary layer of a generic interceptor missile are evaluated at different altitudes and thruster conditions. A generic missile interceptor cone guration consisting of a long, slender body containing e xed dorsal and tail e ns is simulated. Parametric computational e uid dynamic solutions are obtained at altitude conditions corresponding to 19.7 and 35.1 km for 1 ) steady-state conditions with the lateral control jet turned off, 2) steady-state conditions with the lateral control jet turned on, 3 ) transient jet startup conditions, and 4 ) transient jet shutdown conditions. A thermally and calorically perfect gas with a specie c heat ratio equal to 1.4 was assumed for both the Mach number 5 freestream and Mach number 3 lateral jet. Vehicle forces and moments are assessed from each solution by integrating the surface pressures and viscous shear stresses computed on the missile surfaces. These results are used to determine the ine uence of the jet-interaction effects on the transient aerodynamicperformanceofthemissile.Theanalysispredictsstrongtransientine uencesfortheintegratednormal force and pitching moment.

A non-linear aeroelastic model for the study of flapping-wing flight

Abstract: A = element cross-sectional area AR = aspect ratio a = mass-proportional damping constant b = stiffness-proportional damping constant Cd = drag coefficient Cdf = skin-friction drag coefficient Cmac = airfoil moment coefficient about its aerodynamic center Cn = normal force coefficient c = wing segment chord length Dc = drag due to camber Df = friction drag E = modulus of elasticity Fy = total chordwise force F'(k), G'(k) = terms for modified Theodorsen function G = shear modulus of elasticity gr = acceleration due to gravity h = total plunging displacement h = elastic component of plunging displacement hQ imposed displacement / = moment of inertia J = polar moment of inertia k = reduced frequency based on | c L = total lift M = total twisting moment acting on a wing segment

A New Aerodynamic Approach to Advanced Automobile Basic Shapes

Abstract: Aerodynamic basic shapes are generally intended as non wheeled bodies moving at a small distance from the ground, effective and suitable for automobile applicationsThe shape is furthermore designed to comply with requirements other than aerodynamic accomodating occupants, luggage and mechanical parts within as small as possible overall dimensions However, even though the basic body drag coefficient can be as low as 005, the addition of wheels may increase the body drag, by two to three times The new approach starts from the definition of aerodynamic criteria such as total lift close to zero, the pitching moment sign and value consistent with road holding and stability, a reduced sensitivity to side wind, gradual variation of the cross sections, etc Then, the presence of the wheels is taken into account in order to reduce their aerodynamic interference with the body, and to manage the wake mechanisms in order to recover the kinetic energy of the flow without fitting the body with a solid diffuser: in fact, this would increase the car length without contributing much to the usable space Theoretical and experimental work leading to a new shape particularly short in the rear part and able to improve both the accomodation of passengers in the rear seats and the visibility outside at present a major drawback of actual cars are shown together with aerodynamic drag benefits Other advantages resulting from the application of the method are finally discussed

Hypersonic Flow Control Using Upstream Focused Energy Deposition

Abstract: A numerical study of centerline and off-centerline power deposition at a point upstream of a two-dimensional blunt body at Mach 6.5 at 30 km altitude are presented. The full Navier-Stokes equations are used. Wave drag, lift, and pitching moment are presented as a function of amount of power absorbed in the flow and absorption point location. It is shown that wave drag is considerably reduced. Modifications to the pressure distribution in the flow field due to the injected energy create lift and a pitching moment when the injection is off-centerline. This flow control concept may lead to effective ways to improve the performance and to stabilize and control hypersonic vehicles.

New Rotor Airfoil Design Procedure for Unsteady Flow Control

Abstract: The present paper describes two different experiments performed in a transonic wind tunnel facility at DLR-Goettingen. The first experiment was conducted in order to study compressible vortices behind a cylinder and investigating the feasibility of combining two different measuring techniques: the Background Oriented Schlieren (BOS) technique and the Particle Image Velocimetry (PIV) which allow respectively to measure both density and velocity fields. The second experiment described in the present paper is done in the same wind tunnel facility where a new test section has been developed to investigate the unsteady flow about oscillating models under dynamic stall conditions. Dynamic stall is characterized by the development, movement and shedding of one or more concentrated vortices on the blade upper surface, the hysteresis loops of lift-, drag- and pitching moment are highly influenced by these vortices. To understand the very complicated unsteady flow involved, a detailed knowledge of the instantaneous flow fields is of crucial importance. With the application of the described measuring techniques it is expected to gain more insight into the problem. In recent years numerical codes based on the time-accurate solution of the Reynolds-Averaged Navier-Stokes equations (RANS) have been developed. Results from these codes are ready for comparison with experimental data. A section of the present paper is dedicated to the comparison of numerical with corresponding experimental data.

The influence of mid-chord battle damage on the aerodynamic characterstics of two-dimensional wings

Abstract: This paper briefly considers the method of simulating gunfire damage to a wing and outlines the key basic assumptions used in modelling. The results of qualitative and quantitative investigations into the aerodynamic characteristics of a wing damaged at quarter chord are then presented. The results are discussed in terms of flow mechanisms, changes to surface pressure distributions and increments in lift, drag and pitching moment coefficients. For the damaged wing, the influence on force and moment coefficients was attributed to flow through the damage. This through flow was driven by the pressure differential between the upper and lower wing surfaces, and took one of two forms. The first form was a ‘weak-jet’ which formed an attached wake and resulted in small changes in force and moment coefficients. The second form resulted from either increased incidence, or damage size. This was the ‘strong-jet’, where through flow penetrated into the freestream flow, resulting in separation of the oncoming surface flow, and the development of a larger separated wake with reverse flow. The effect on force and moment coefficients was significant. The paper also compares the structure of the damage through flow with previously published results for jets in crossflows. Many similarities in the flow features were identified, although there were significant differences in the surface pressure distributions for the two cases.

Model Validation of Wake-Vortex / Aircraft Encounters

Abstract: Wake-vortex effects on an 10% scale model of the B737-100 aircraft are calculated using both strip theory and vortex-lattice methods. The results are then compared to data taken in the 30ft x 60ft wind tunnel at NASA Langley Research Center (LaRC). The accuracy of the models for a reduced geometry, such with the horizontal stabilizer and the vertical tail removed, is also investigated. Using a 10% error in the circulation strength and comparing the model's results with the experiment illustrates the sensitivity of the models to the vortex circulation strength. It was determined that both strip theory and the vortex lattice method give accurate results when all the geometrical information is used. When the horizontal stabilizer and vertical tail were removed there were difficulties modeling the sideforce coefficient and pitching moment. With the removal of only the vertical tail unacceptable errors occurred when modeling the sideforce coefficient and yawing moment. Lift could not be accurately modeled with either the full geometry or the reduced geometry.

The Aerodynamic Stability of a Le Mans Prototype Race Car Under Off-Design Pitch Conditions

Abstract: The current generation of sports racing cars such as those competing under the Le Mans “LM”P and “LM”GTP regulations are particularly sensitive to the pitch of the vehicle. This is a consequence of the low ground clearances that must be adopted to maximise the benefits that can be gained from ground effect and of the very large floor plan area of these cars. To achieve optimum cornering and straight line performance the suspension characteristics are often tuned to the aerodynamic forces in order to reduce the pitch and hence the drag of the vehicle at high speeds whilst retaining relatively high downforce when cornering. A series of accidents at the 1999 Le Mans 24-hour race have highlighted the potential instability of these vehicles which resulted in the catastrophic ‘take-off’ of one of the “LM”GTP cars during the race and others during qualifying and the pre-race ‘warm-up’. The data presented here have been extracted from a detailed experimental study of a typical “LM”GTP car under design and off-design pitch conditions including extreme cases of nose-up pitching moment to assess the onset of instability i.e rotation leading to take-off. Additional data are presented to demonstrate the influence of possible regulation changes upon these parameters.

Correlation between geometric similarity of ice shapes and the resulting aerodynamic performance degradation - A preliminary investigation using WIND

Abstract: Aerodynamic performance calculations were performed using WIND on ten experimental ice shapes and the corresponding ten ice shapes predicted by LEWICE 2.0. The resulting data for lift coefficient and drag coefficient are presented. The difference in aerodynamic results between the experimental ice shapes and the LEWICE ice shapes were compared to the quantitative difference in ice shape geometry presented in an earlier report. Correlations were generated to determine the geometric features which have the most effect on performance degradation. Results show that maximum lift and stall angle can be correlated to the upper horn angle and the leading edge minimum thickness. Drag coefficient can be correlated to the upper horn angle and the frequency-weighted average of the Fourier coefficients. Pitching moment correlated with the upper horn angle and to a much lesser extent to the upper and lower horn thicknesses.

Comparison of Strain Gage and Fiber Optic Sensors On A Sting Balance In A Supersonic Wind Tunnel

Abstract: Force and moment balances have proved to be essential in the measurement and calculation of aerodynamic properties during wind tunnel testing. With the recent advancements of technology, new fiber optic sensors have been designed to replace the conventional foil strain gage sensors commonly found on balances, thereby offering several distinct advantages. The use of fiber optic sensors on a balance brings with it some potential advantages over conventional strain gage balances including increased resolution and accuracy, insensitivity to electromagnetic interference, and the capability of use at high temperatures. By using the fiber optic sensors, some of the limitations of the conventional balance can be overcome, leading to a better overall balance design. This thesis considers an initial trial application of new fiber optic sensors on a conventional, six-component sting balance while retaining the original foil strain gage sensors for comparison. Tests were conducted with a blunt, 10o half-angle cone model in the Virginia Tech 9x9 inch Supersonic Wind Tunnel at Mach 2.4 with a total pressure of 48 psia and ambient total temperature of 25.3oC. Results showed a close comparison between the foil strain gages and the fiber optic sensor measurements, which were set up to measure the normal force and pitching moment on the blunt cone model. A Finite Element Model (FEM) of the sting balance was produced in order to determine the best locations for the fiber optic sensors on the sting balance. Computational Fluid Dynamics (CFD) was also used in order to predict and compare the results acquired from all of the sensors.

Wind tunnel testing of performance degradation of ice contaminated airfoils

Abstract: An experimental investigation was conducted to study the aerodynamic effect of simulated supercooled large droplet ice accretion on a MS(1) 0313 aircraft airfoil by means of wind tunnel testing. The airfoil was equipped with an aileron with an axis of movement at 78 % of airfoil chord. Ice accretion was simulated by a strip of a quadrantal section with front edge perpendicular to the surface of the airfoil. Two strips of different dimensions were examined, of the height of 1.33 % of chord and 2.25 % of chord, respectively. Every strip was positioned at several positions on the upper surface of the airfoil, and its influence on the basic aerodynamic characteristics of the airfoil was evaluated, i. e. the lift, drag, moment and first of all the hinge moment of the aileron. The ice accretion strip was positioned from 5 % to 45 % of chord, the deflections of aileron were up to 20 degrees down and 30 degrees up and the angle of attack was changed from 0 up to the angle of maximum positive and maximum negative lift coefficient. Wind tunnel tests were performed in the VZLU 3 m diameter low - speed wind tunnel at a Reynolds number of 2.10 6 .

Phugoid Approximation for Conventional Airplanes

Abstract: An improved closed-form approximation for phugoid motion in conventional airplanes is presented. Although several closed-form approximations for phugoid motion are currently available and widely used, none of these approximations accurately predict all of the fundamental characteristics of phugoid motion. The new approximation accounts for changes in angle of attack as well as the effects of pitch stability and pitch damping. The total phugoid damping is shown to depend on pitch damping aswell as aircraftdrag. In addition, thissolution pointsout another important contribution to phugoid damping called phase damping. It is shown that the phase-damping contribution to the real component of the phugoid eigenvalue is always positive and tends to reduce the total phugoid damping. Under certain conditions this phase damping can cause the phugoid mode to become divergent. Nomenclature Aw = planform area of the wing CD = total drag coefe cient CDp = parasitic drag coefe cient CD,a = change in drag coefe cient with angle of attack CL = lift coefe cient CL,a = change in lift coefe cient with angle of attack CM = pitching moment coefe cient CM,a = change in pitching moment coefe cient with angle of attack CM,$ = change in pitching moment coefe cient with dimensionless pitching rate ¯ c = mean chord length e = Oswald efe ciency factor FT = thrust force g = acceleration of gravity Iyy = pitching moment of inertia in body-e xed coordinates m = aircraft mass RA = aspect ratio Rd = phugoid pitch-damping ratio Rg = dimensionless gravitational acceleration RM = dimensionless change in pitching moment with axial velocity RM,a = dimensionless change in pitching moment with angle of attack RM,$ = dimensionless change in pitching moment with pitching rate Rp = phugoid phase-divergence ratio Rs = phugoid stability ratio Rx = dimensionless change in axial force with axial velocity Rxa = complex amplitude Rxc = complex coefe cient Rxp = complex phase Rx,a = dimensionless change in axial force with angle of attack Rz = dimensionless change in normal force with axial velocity Rza = complex amplitude Rzc = complex coefe cient Rzp = complex phase Rz,a = dimensionless change in normal force with angle of attack

Flow simulation of NAL experimental supersonic airplane/booster separation using overset unstructured grids

Abstract: The overset unstructured grid method developed for multiple-body problems is applied to a flow simulation about an experimental supersonic airplane separation from a rocket booster. An unstructured grid around the rocket booster is overset on the stationary grid around the airplane and moves with time to simulate the separation process. Detailed components of the rocket booster are faithfully reproduced by the unstructured grid. This capability of the unstructured grid reduces the number of required overset grids and significantly simplifies the overset procedure. The computed result of the airplane/booster separation clearly simulates the complex reflection patterns of shock waves between two bodies during the separation process. The computed lift and pitching moment coefficients are compared with the wind tunnel results. The computational accuracy of the aerodynamic coefficients is improved by including the detailed components of the airplane and rocket.

Flow Simulation of NAL Experimental Supersonic Airplane/Booster Separation Using Overset Unstructured Grids.

Abstract: The use of overset concept for the unstructured grid method is relatively unexplored. However, the overset approach can extend the applicability of the unstructured grid method for real engineering problems without much efforts in code developments. Multiple moving-body problem is one of those applications. In this paper, the overset unstructured grid method is applied to a flow simulation about an experimental supersonic airplane separation from a rocket booster. Two unstructured grids, each of which covers the airplane and the rocket booster respectively, are used for the simulation. The grid around the rocket booster moves with time in the stationary grid about the airplane. The computed result clearly simulates the shock wave patterns between two bodies. Comparisons with the experimental results show good agreements in the lift and pitching moment coefficients of the airplane and booster during the separation process.

Aerodynamics of the 120-mm m831a1 projectile: analysis of free-flight experimental data

Abstract: The 120-mm M831A1 projectile is a low-cost training projectile used by U.S. armor troops. For the last several years, program managers have received feedback from the user that in some cases, M831A1 impact performance did not appear consistent with the current M831A1 computer correction factor (CCF). Based on this informantion, a low-scale but in-depth experimental analysis of the round was conducted to assess its aeroballistic qualities and hopefully identify any potential issues that could affect accuracy performance. The fiveround experiment was conducted by the Army Research Laboratory (ARL) at the Transonic Experimental Facility (TEF), Aberdeen Proving Ground, Maryland. Yaw magnitudes displayed variability, and several shots had at least moderate levels. The source of the yaw levels measured was the launch dynamics, and a detailed study of inbore dynamics is in progress. Most shots exhibited a "stepping" motion in plots of total yaw vs. range. This phenomenon is the result of a trim believed to be caused by an aerodynamic asymmetry. A source of the trim has not yet been isolated. Accurate freeflight drag and pitching moment coefficients were

Aerodynamic characteristics of ribbon stabilized grenades

Abstract: An extensive experimental program has been conducted to determine the aerodynamic characteristics of grenade ribbon stabilizers. While data has been acquired from vertical and horizontal wind tunnel tests, free-flight drop tests, and freeflight gun tests, this paper presents only the data from the horizontal wind tunnel test. During this test, both static and dynamic free-yaw data were obtained at speeds ranging from 90 to 180 feet/second. The test obtained axial force, side force, yawing moment and pitching moment while the grenade was free to yaw, or while statically fixed at a given yaw angle. Analysis of the static and dynamic forces and moments indicates that there are four types of ribbon induced oscillatory motion. These types are presented, as is the zero-yaw drag as a function of ribbon length and ribbon width. NOMENCLATURE Svmbols: AoA Angle-of-attack, degrees. AF Axial force, Ibs. CM Pitching moment coefficient, = PM/QSD. CY Side force coefficient, = SF/QS. CLN Yawing moment coefficient, = YM/QSD. CAF Axial force coefficient, = AF/QS. CD Drag COeffiCient (derived from CU. CAF & Theta) CL Lift COeffiCient (derived from Cy, CAF & Theta) D Reference length, = 1.515 inches. PM Pitching moment, in-lbs. S Reference area, sq. in., = nD2/4. SF Side force, Ibs. YM Yawing moment, in-lbs. Theta Yaw-angle, degrees. * AIAA Member-Aerospace Engineer This paper is declared a work of the U.S. Government and is not subject to copyright protection in the United States. C. Wayne Dahlke* David C. Purinton* Dynetics, Inc Huntsville, AL FORWARD An effort has been undertaken to reduce the hazardous dud rate of rocket-dispensed grenades to less than 1 percent. As a part of this effort, the current study was initiated as a low cost solution to this problem. The objective being to eliminate side impact, the largest single cause of hazardous duds. Following dispense the grenades experience large oscillatory motion until the ribbon stabilizer is deployed. After the stabilizer is deployed, the grenade motion quickly damps out and slows to terminal velocity. Once the grenade has reached terminal velocity it should not experience any oscillatory motion in excess of 20 degrees. Previous wind tunnel tests of ribbon stabilized grenades may be classified into three types of tests. These are static force and moment tests, dynamic “free-yaw” tests, and vertical flight tests. Static force and moment data presented by Dahlke”‘*’ were collected utilizing a pyramidal balance to obtain the grenade axial force, side force andyawing moment. Dynamic “free-yaw’ tests presented by Dahlket” I measured the grenades yaw angle as a function of time and from that data the dynamic derivatives (CMa, CMq) were calculated. Analysis of the static data alone would lead the analyst to believe that the current ribbon provides adequate stability. Furthermore, incorporation of the static data into a 6-DOF simulation would no doubt indicate that grenade oscillations imparted during dispense quickly damp out and the remainder of the grenade trajectory is at relatively small angles-of-attack. However, data taken during the “free-yaw” tests and observations made during the vertical wind tunnel tests indicate that this is not the case. This discrepancy may be attributed to the ribbon dynamics. While the flapping ribbon does provide the drag needed to stabilize the grenade body (and arm the fuze), the flapping ribbon also acts to destabilize the grenade. The task of designing a test that truly captures all the flight dynamics of the grenade is not currently possible. A “perfect” test would require a breakthrough in a miniaturized

Drag, lift and pitching moment increments due to wind tunnel wall constraint: extension to three dimensions

Abstract: Wind-tunnel-induced flow distortion around a test model causes drag, lift and pitching moment increments that are absent in free air. Previous two-dimensional predictions of these effects are extended to three dimensions and means for evaluating them are discussed. A new extension to Maskell's blockage prediction method, designated "Maskell III", is introduced that estimates the lift increment. Other boundary correction approaches include procedures based on the use of wall pressure measurements and the application of pressure signature and two-variable algorithms. These studies are supported by dedicated experiments on a family of flat plate wings, carried out at the NRC. Initial applications of the new correction methods to these data have produced encouraging results. Nomenclature b model span. B tunnel width. c wing mean chord. cS spacing of solid blockage source-sink pair. C tunnel cross sectional area, B x H. CD, CL, CM wind-axis drag, lift and pitching moment coefficients. ∆ CD,∆ CL,∆ CM

Aerodynamic force measuring apparatus and method

Abstract: PROBLEM TO BE SOLVED: To provide an aerodynamic force measuring apparatus and method which enable measuring of a non-.steady aerodynamic force at a higher accuracy with a load cell of a relatively small capacity in any case where an aerodynamic force is large or is where the level of the aerodynamic force is very small in the measurement of the aerodynamic force working on a model arranged in an air flow space with the load cell. SOLUTION: A model arranged within an air flow space such as air tunnel is mounted on a frame and when the model is vibrated by an exciting means through the frame to measure the aerodynamic force working on the model by an air flow, the model is suspended through an elastic support member from the frame to support the weight of the model. A lower part of the model is linked to the load cell to make the aerodynamic force work mainly on the load cell so that the position of a link part is changed between a drive link mechanism and the frame. This allows the frame and the model to select vertical movement or rotary motion through the exciting means and the drive link mechanism.