Showing papers on "Pitching moment published in 1995"

Natural-laminar-flow airfoil for general-aviation applications

Abstract: A natural-Iamina r-flow airfoil, the NLF(1)-0115, has been recently designed for general-aviation aircraft at the NASA Langley Research Center. During the design of this airfoil, special emphasis was placed on experiences and observations gleaned from other successful general-aviation airfoils. For example, the flight lift-coefficient range is the same as that of the turbulent-flow NACA 23015 airfoil. Also, although beneficial for reducing drag and producing high lift, the NLF(1)-0115 airfoil avoids the use of aft loading, which can lead to large stick forces if utilized on portions of the wing having ailerons. Furthermore, not using aft loading eliminates the concern that the high pitching-moment coefficient generated by such airfoils can result in large trim drag if cruise flaps are not employed. The NASA NLF(1)-0115 airfoil has a thickness of 15% chord. It is designed primarily for general-aviation aircraft with wing loadings of 720-960 N/m2 (15-20 lb/ft2). Low-profile drag as a result of laminar flow is obtained over the range from c, = 0.1 and R = 9 x 106 (the cruise condition) to c, = 0.6 and R = 4 x 106 (the climb condition). While this airfoil can be used with flaps, it is designed to achieve a c,,max of 1.5 at R = 2.6 x 10 6 without flaps. The zero-lift pitching moment is held to c,H,0 = -0.055. The hinge moment for a 20% chord aileron is fixed at a value equal to that of the NACA 632-215 airfoil, CH = — 0.0022. The loss in cAmax due to leading-edge roughness at R = 2.6 x 10 6 is 11% as compared with 14% for the NACA 23015.

Effects of grit roughness and pitch oscillations on the S809 airfoil

Abstract: An S809 airfoil model was tested in The Ohio State University Aeronautical and Astronautical Research Laboratory (OSU/AARL) 3{times}5 subsonic wind tunnel (3{times}5) under steady flow and stationary model conditions, and also with the model undergoing pitch oscillations. To study the possible extent of performance loss due to surface roughness, a standard grit pattern (LEGR) was developed to simulate leading edge contamination. After baseline cases were completed, the LEGR was applied for both steady state and model pitch oscillation cases. The Reynolds numbers for steady state conditions were 0.75, 1, 1.25, and 1.5 million, while the angle of attack ranged from {minus}20, to +40 {degrees}. With the model undergoing pitch oscillations, data were acquired at Reynolds numbers of 0.75, 1, 1.25, and 1.4 million, at frequencies of 0.6, 1.2, and 1.8 Hz. Two sine wave forcing functions were used; {plus_minus} 5.5{degrees} and {plus_minus} 10{degrees}, at mean angles of attack of 8{degrees}, 14{degrees}, and 20{degrees}. For purposes herein, any reference to unsteady conditions means the model was in pitch oscillation about the quarter chord. In general, the unsteady maximum lift coefficient was from 4% to 86% higher than the steady state maximum lift coefficient, and variation in the quarter chord pitching moment coefficient magnitude was from {minus}83% to 195% relative to steady state values at high angles of attack. These findings indicate the importance of considering the unsteady flow behavior occurring in wind turbine operation to obtain accurate load estimates.

Dynamic Stall Control for Advanced Rotorcraft Application

Abstract: Advanced concepts designed to improve the lift, drag, and pitching moment characteristics of rotor blades have been investigated for the purpose of enhancing rotor maneuver capability. The advantages and disadvantages of these concepts have been evaluated using both computational and experimental means. The concepts that were considered in this study included a leading-edge slat, a deformable leading-edge, and upper-surface blowing. The results show the potential of these concepts for substantially improving the performance of a rotor. HE next generation of rotorcraft will be required to operate at much higher performance levels than in the past, particularly in the areas of nap-of-the-ea rth (NOE), deep-penetration operations, and air-to-air combat. These new requirements will require highly maneuverable, agile, and survivable rotorcraft, far exceeding the capabilities of those in the current inventory. The objectives of this project include an increase in the maneuverability/agility capability of the helicopter and a reduction in the acoustic detection range. The single most important element of the rotorcraft for meeting these requirements is the rotor itself, since it is the primary source of lift, control, and speed. At the same time, the rotor is also a major source of acoustically detectable radiation. Among the many factors affecting rotorcraft performance, the aerodynamic characteristics of the rotor system are the most important and are the main subject of this paper. The maneuvering capability of a rotorcraft can be improved by re- ducing or suppressing the vibratory loads on the rotor blades caused by aerodynamic separation and stall. This would have the effect of expanding the stall-limiting boundary of the rotor and thereby increase the available load factor in all flight regimes. The con- ventional way to obtain higher lift is to increase the blade area, however, this usually results in a heavier rotor that is also less ef- ficient. With regard to compressibili ty effects and acoustic radia- tion, improvements have been obtained by sweeping, tapering, and thinning the tip region of the rotor blade. As a result, numerous families of airfoils and planform shapes have evolved that offer bet- ter advancing-blade characteristics. However, improvements on the retreating-blade side have not been as impressive. One reason for this imbalance may be that design codes are available for treating blades at low angles of attack and high Mach number (characteris- tic of the advancing side), whereas the design strategy has had to depend heavily on costly empirical studies for blades at high angles of attack and having some amount of separation (characteristic of the retreating side). Increasing the tip speed of the rotor to achieve a maneuvering ad- vantage may produce a dangerous condition with regard to acoustic detection. Rapid advancements in passive acoustic sensor arrays and advanced signal processing technologies pose a serious threat to the mission effectiveness of Army helicopters. Since the rotor blade generates acoustic radiations that can be easily detected and identified, airfoil and planform shapes must be carefully optimized to reduce the detection range of the rotorcraft. The requirements for improved maneuverability and reduced sus- ceptibility will clearly demand a substantial growth in the technolo- gies for addressing rotor aerodynamics. New control techniques must be considered, both passive and active, and these must be ac- companied by a more thorough physical understanding of these flow phenomena along with substantially improved prediction capabili- ties. To meet these requirements, computational and experimental efforts have been initiated to evaluate the effectiveness of various concepts. At present these concepts include airfoils with slats and slots, airfoils that deform, and airfoils with flow energizers. Description of Experiment and Computational Fluid Dynamics Code

Numerical investigation of dynamic stall of an oscillating wing

Abstract: The unsteady three-dimensional flowfield over an oscillating wing is investigated with the numerical solution of the compressible, time-dependent, Reynolds-averaged Navier-Stokes equations. Spatial discretization is performed with a third-order accurate, upwind-biased, vertex-based, finite volume scheme. An alternative direction implicit, iterative scheme is used for the time integration. The high Reynolds number turbulent flow behavior is modeled with a one-equation turbulence model. The effect of subiterations, time step and grid density on the accuracy of the computed solutions is investigated. It is found that scaling of the time step with the angular velocity of the motion produces accurate solutions at a reduced computational cost. The computational domain over an aspect ratio 5 wing with rounded tip and NACA-0015 airfoil sections is discretized with a single-block grid. The light stall flowfield over the wing oscillating in a subsonic freestream with a mean angle of attack of 11 deg and an amplitude of 4.2 deg is computed. The structure of the separated, unsteady flowfield is investigated and comparisons with available experimental data are performed.

Identification of an Unsteady Aerodynamic Stall Model from Flight Test Data.

Abstract: An unsteady aerodynamic model for high lift including flow separation and stall is presented. The model provides four parameters and two time constants, which can be identified from suitable flight test data in order t o describe specific stall characteristics. Using this model, lift, drag, and pitching moment are a function of an internal state, namely the position of the flow separation point on the wing surface. This position is formulated as a function of the angle of attack, the angle of attack rate, and the time (first order differential equation). The identifiability of the model parameters and the plausibility of the model structure is discussed using flight test data of C-160 and VFW-614 ATTAS aircraft. Validation plots demonstrate the model accuracy and it can be seen clearly that there are considerable unsteady effects a t high angles of attack caused by flow separation, which cannot be described properly using conventional flight mechanic models.

Wing performance in moderate rain

Abstract: Emphasis is placed on the correlation of surface-film behavior including rivulet formulation with measured values of lift, drag, and moment at angles of attack up to stall. Four regions of surface flow are identified: 1) the droplet-impact, 2) film-convection, 3) rivulet-formation, and 4) droplet-convection regions. The extent that each of these regions covers the airfoil surface changes with incidence and correlates with changes in aerodynamicforce coefficients. Additionally, results quantify effects of the use of boundary-laye r trips for linking flight and wind-tunnel models in rain, and show that surface water phenomena affect laminar-to-turbulent transition in a manner that is inconsistent with the use of transition fixing to increase the effective test Reynolds number. Nomenclature ^wing = wing planform area Cd = drag coefficient, DRAG/^wing C, = lift coefficient, LIFT/^wing Cm — moment coefficient, PITCHING MOMENT/^wing c = chord length Re = Reynolds number based on chord length and freestream velocity a = angle of attack AQ = drag coefficient increment, (Crf)wet - (Cd)dry AC/% = percent change in lift coefficient, 100 x [(Q wet ~ (Q^MQ^ ACm .= moment coefficient increment, (Cw)wet - (Cm)dry

A Longitudinal Aerodynamic Data Repeatability Study for a Commercial

Abstract: A high Reynolds number investigation of a commercial transport model was conducted in the National Transonic Facility (NTF) at Langley Research Center. This investigation was part of a cooperative effort to test a 0.03-scale model of a Boeing 767 airplane in the NTF over a Mach number range of 0.70 to 0.86 and a Reynolds number range of 2.38 to 40.0 x 10\super{6} based on the mean aerodynamic chord. One of several specific objectives of the current investigation was to evaluate the level of data repeatability attainable in the NTF. Data repeatability studies were performed at a Mach number of 0.80 with Reynolds numbers of 2.38, 4.45, and 40.0 x 10\super{6} and also at a Mach number of 0.70 with a Reynolds number of 40.0 x 10\super{6}. Many test procedures and data corrections are addressed in this report, but the data presented do not include corrections for wall interference, model support interference, or model aeroelastic effects. Application of corrections for these three effects would not affect the results of this study because the corrections are systematic in nature and are more appropriately classified as sources of bias error. The repeatability of the longitudinal stability-axis force and moment data has been assessed. Coefficients of lift, drag, and pitching moment are shown to repeat well within the pretest goals of \pm0.005, \pm0.0001, and \pm0.001, respectively, at a 95-percent confidence level over both short- and near-term periods.

A longitudinal aerodynamic data repeatability study for a commercial transport model test in the National Transonic Facility

Abstract: A high Reynolds number investigation of a commercial transport model was conducted in the National Transonic Facility (NTF) at Langley Research Center. This investigation was part of a cooperative effort to test a 0.03-scale model of a Boeing 767 airplane in the NTF over a Mach number range of 0.70 to 0.86 and a Reynolds number range of 2.38 to 40.0 x 10(exp 6) based on the mean aerodynamic chord. One of several specific objectives of the current investigation was to evaluate the level of data repeatability attainable in the NTF. Data repeatability studies were performed at a Mach number of 0.80 with Reynolds numbers of 2.38, 4.45, and 40.0 x 10(exp 6) and also at a Mach number of 0.70 with a Reynolds number of 40.0 x 10(exp 6). Many test procedures and data corrections are addressed in this report, but the data presented do not include corrections for wall interference, model support interference, or model aeroelastic effects. Application of corrections for these three effects would not affect the results of this study because the corrections are systematic in nature and are more appropriately classified as sources of bias error. The repeatability of the longitudinal stability-axis force and moment data has been accessed. Coefficients of lift, drag, and pitching moment are shown to repeat well within the pretest goals of plus or minus 0.005, plus or minus 0.0001, and plus or minus 0.001, respectively, at a 95-percent confidence level over both short- and near-term periods.

Experimental and Computational Analysis of Shuttle Orbiter Hypersonic Trim Anomaly

Abstract: During the high-Mach-number, high-altitude portion of the first entry of the Shuttle Orbiter, the vehicle exhibited a nose-up pitching moment relative to preflight prediction of approximately Delta Cm = 0.03. This trim anomaly has been postulated to be due to compressibility, viscous, and/or real-gas (lowered specific heat ratio gamma) effects on basic body pitching moment, body-flap effectiveness, or both. In order to assess the relative contribution of each of these effects, an experimental study was undertaken to examine the effects of Mach number, Reynolds number, and ratio of specific heats. Complementary computational solutions were obtained for wind-tunnel and flight conditions. The primary cause of the anomaly was determined to be lower pressures on the aft windward surface of the Orbiter than deduced from hypersonic wind-tunnel tests with ideal- or near-ideal-gas test flow. The lower pressure levels are a result of the lowering of the flowfield gamma due to high-temperature effects. This phenomenon was accurately simulated in a hypersonic wind tunnel using a heavy gas, which provided a lower, gamma, and was correctly predicted by Navier-Stokes computations using nonequilibrium chemistry.

System and method for modeling the flow performance features of an object

Abstract: The method and apparatus includes a neural network for generating a model of an object in a wind tunnel from performance data on the object. The network is trained from test input signals (e.g., leading edge flap position, trailing edge flap position, angle of attack, and other geometric configurations, and power settings) and test output signals (e.g., lift, drag, pitching moment, or other performance features). In one embodiment, the neural network training method employs a modified Levenberg-Marquardt optimization technique. The model can be generated "real time" as wind tunnel testing proceeds. Once trained, the model is used to estimate performance features associated with the aircraft given geometric configuration and/or power setting input. The invention can also be applied in other similar static flow modeling applications in aerodynamics, hydrodynamics, fluid dynamics, and other such disciplines. For example, the static testing of cars, sails, and foils, propellers, keels, rudders, turbines, fins, and the like, in a wind tunnel, water trough, or other flowing medium.

An evaluation of aerodynamic prediction methods applied to the XB-70 for use in high speed aircraft stability and control system design

Abstract: A key consideration in the development of flight control systems early in the design stage is the availability of aerodynamic information for different candidate configurations. Aerodynamic estimation methods must be available to provide the connection between the configuration geometry and its stability and control characteristics. The accuracy of current estimation methods is studied in this paper through application of DATCOM and APAS to available XB-70 wind tunnel and flight test data. The study was carried out for the subsonic approach condition and three supersonic conditions. Tables and charts are presented to provide a quantitative assessment of the accuracy of the predictions, and areas requiring improvement are identified. Results show that APAS and DATCOM predictions are good for most lateral/directional stability and control derivatives. Estimations for the pitching moment slope, yaw damping and yawing moment due to flap deflection derivatives are only fair. The most difficult derivative to predict is the rolling moment due to sideslip.

Analysis of flowfields over missile configurations at subsonic speeds

Abstract: Flowfields over missile configurations at subsonic speeds and high angle of attack are computed with a NavierStokes flow solver using overset grids. The accuracy of the computed solutions is first validated for flows at high incidence over a fuselage-wing and a fuselage-canard-wing configuration. The effects of grid density and turbulent versus laminar solutions are assessed by comparison with detailed experimental surface pressures. Development of vortex breakdown over the wing is predicted in accordance with the experiment. Delay of vortex breakdown over the wing caused by the presence of the canard was also captured by the numerical solution. The computed surface pressures are in good agreement with the experiment. Solutions for a complete missile configuration with fins and tails at 45-deg roll are also obtained for subsonic flow at high incidence. The computed normal force and pitching moment are in good agreement with available measurements. The effect of the fin deflection on the development of the vortical flowfield is investigated. It is found that the flowfield in the fin and gap regions can be of primary importance to the overall development of the forebody flowfield.

Interpretation of experimental high-alpha aerodynamics. Implications for flight prediction

Abstract: Nomenclature b = wingspan, or largest dimension C — aerodynamic coefficient; with no superscript, in body axes system Cij = dCi/d(jl/2V), i = /, m, n\ j = q, q', d, a(I = c)'J =p,j8, (/ = b) Cik = dCf/dk, i = l,m,n\k = a, /3, a Cl = rolling moment coefficient Cm = pitching moment coefficient C,, = yawing moment coefficient Cp = static pressure coefficient CY = side force coefficient c = mean aerodynamic chord d — body maximum diameter h = height of test section h() = model length / = generalized reference length M^ = freestream Mach number /?, px = local static pressure /?, q, r = body-axes angular velocities q^ = freestream dynamic pressure Re, Reynolds number based on /, / c, b, or d S = reference area V, V^ = freestream velocity w = minimum dimension of test section X, Y, Z = inertial axes system Xf, Yf; Zf = flight coordinate system, Fig. 2 x, _y , z = body axes system a, /3 = angles of attack and sideslip A = increment or amplitude A = inclination of rotation axis, Fig. 6 = coning rate, parameter fib/2V or £1 a) = reduced circular frequency, coll(2V), where / = c or b, as appropriate

Effects of Sonic Line Transition on Aerothermodynamics of the Mars Pathfinder Probe

Abstract: Flowfield solutions over the Mars Pathfinder Probe spanning the trajectory through the Martian atmosphere at angles of attack from 0 to 11 degrees are obtained. Aerodynamic coefficients derived from these solutions reveal two regions where the derivative of pitching moment with respect to angle of attack is positive at small angles of attack. The behavior is associated with the transition of the sonic line location between the blunted nose and the windside shoulder of the 70 degree half-angle cone in a gas with a low effective ratio of specific heats. The transition first occurs as the shock layer gas chemistry evolves from highly nonequilibrium to near equilibrium, above approximately 6.5 km/s and 40 km altitude, causing the effective specific heat ratio to decrease. The transition next occurs in an equilibrium flow regime as velocities decrease through 3.5 km/s and the specific heat ratio increases again with decreasing ent halpy. The effects of the expansion over the shoulder into the wake are more strongly felt on the fustrum when the sonic line sits on the shoulder. The transition also produces a counter-intuitive trend in which windside heating levels decrease with increasing angle of attack resulting from an increase in t he effective radius of curvature. Six-degree-of-freedom trajectory analyses utilizing the computed aerodynamic coe fficients predict a moderate, 3 to 4 degree increase in total angle of attack as the probe, spinning at approximately 2 revolutions per minute, passes through these regions.

Unsteady flow control on rotor airfoils

Abstract: Airfoil designs for helicopter rotors are based on compromises: On the advancing side the blade encounters high speed transonic flows including moving shock waves, on the retreating side, however, the high incidence of the blade may lead to dynamic stall with strong hysteresis effects in force and moment loops. It is the aim of the present numerical investigation to systematically develop airfoils which are changing their shape during operation. With specially designed dynamically deforming airfoil sections the dynamic stall characteristics can successfully be influenced. Rather small airfoil modifications show already considerable benefits with respect to a reduction of the negative (nose-down) pitching moment peak as well as with the avoidance of negative aerodynamic damping.

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

Body spoiler for yaw control of a supersonic airplane

Abstract: A body spoiler rotatably mounted within a recess formed within opposite sides of the forebody of a supersonic air vehicle and configured to conform with the fuselage of the vehicle. A sensor positioned close to the inlet of each engine powering the flight of the vehicle measures the pressure of the fluid entering the engine. If an engine on one side of the vehicle malfunctions, the pressure sensor sends a deploy signal to the body spoiler mounted on the same side of the vehicle. An actuator receives the deploy signal and rotates the spoiler through an angle which is dependent upon the speed of the vehicle. When deployed, the spoiler generates a shock wave which in turn produces pressure along the forebody thereby creating a yaw moment to substantially counterbalance the yaw moment generated by the malfunctioning engine. A plurality of spoilers can also be pivotably mounted along and about the periphery of the fuselage and used as an emergency aid in generating other forces and moments along and about the three main axes of the vehicle. In addition to providing yaw moment control, spoilers deployed on the crown or keel of the vehicle at supersonic speeds could be used in dive recovery by generating drag and a nose up pitching moment.

Airfoil Modification Effects on Subsonic and Transonic Pressure Distributions and Performance for the EA-6B Airplane

Abstract: Longitudinal characteristics and wing-section pressure distributions are compared for the EA-6B airplane with and without airfoil modifications. The airfoil modifications were designed to increase low-speed maximum lift for maneuvering, while having a minimal effect on transonic performance. Section contour changes were confined to the leading-edge slat and trailing-edge flap regions of the wing. Experimental data are analyzed from tests in the Langley 16-Foot Transonic Tunnel on the baseline and two modified wing-fuselage configurations with the slats and flaps in their retracted positions. Wing modification effects on subsonic and transonic performance are seen in wing-section pressure distributions of the various configurations at similar lift coefficients. The modified-wing configurations produced maximum lift coefficients which exceeded those of the baseline configuration at low-speed Mach numbers (0.300 and 0.400). This benefit was related to the behavior of the wing upper surface leading-edge suction peak and the behavior of the trailing-edge pressure. At transonic Mach numbers (0.725 to 0.900), the wing modifications produced a somewhat stronger nose-down pitching moment, a slightly higher drag at low-lift levels, and a lower drag at higher lift levels.

Separation characteristics of generic stores from lee side of an inclined flat plate at Mach 6

Abstract: An experimental investigation was conducted to determine the aerodynamic characteristics of a store as it was separated from the lee side of a flat plate inclined at 15 deg to the free-stream flow at Mach 6. Two store models were tested: a cone cylinder and a roof delta. Force and moment data were obtained for both stores as they were moved in 0.5-in. increments away from the flat plate lee-side separated flow region into the free-stream flow while the store angle of attack was held constant at either 0 deg or 15 deg. The results indicate that both stores had adverse separation characteristics (i.e., negative normal force and pitching moment) at an angle of attack of 0 deg, and the cone cylinder had favorable separation characteristics (i.e., positive normal force and pitching moment) at an angle of attack of 15 deg. At an angle of attack of 15 deg, the separation characteristics of the roof delta are indeterminate at small separation distances and favorable at greater separation distances. These characteristics are the result of the local flow inclination relative to the stores as they traversed through the flat plate lee-side flow field. In addition to plotted data, force and moment data are tabulated and schlieren photographs of the stores and flat plate are presented.

Trailing Vortex Effects on Large Receiver Aircraft

Abstract: The aerodynamic forces and moments on the Hercules receiver aircraft, due to its horizontal and vertical position and bank, yaw, and pitch attitudes in the wake of the KC10 tanker aircraft, are assessed relative to the receiver's aerodynamic characteristics in free air. Large changes in lift, drag, and pitching moment are predicted near the tanker wake centerline. As the receiver is displaced sideways towards the tanker wingtip vortices it experiences large side force and yawing moment and particularly high rolling moment. The most significant term due to the receiver attitude is the rolling moment due to bank. b C»t CL C. ^'V. cm '" H ,

Rotor blade for a rotary-wing aircraft

Abstract: A rotor blade for a rotary-wing aircraft, of an airfoil section has a basic blade thickness at 12% chord length, a portion of an increased blade thickness between the leading edge and a point corresponding to about 90% chord length, a maximum blade thickness at a position shifted backward, a curvature distribution concentrated on the middle portion thereof, and a portion contiguous with the trailing edge and cambered. The drag coefficient of the airfoil section is small when the Mach number is about 0.6 and the lift coefficient is about 0.6, and the airfoil section has a large maximum lift coefficient and a large zero-lift drag divergence Mach number, and generates a small pitching moment.

Forebody-mounted jet nozzles for yaw control on a generic fighter configuration with a chined forebody

Abstract: A wind tunnel investigation was conducted in the NASA Langley 12-Foot Low-Speed wind tunnel to determine the yaw-control effectiveness of jet nozzle blowing on a chined forebody. The results indicated that jet nozzle blowing can be used on a chined forebody to generate yawing moment levels up to twice the value of maximum rudder control. In addition, this yawing moment increment was controllable at angles of attack up to and beyond maximum lift and at moderate sideslip angles. The nozzle pointing angle and forebody cross-section were shown to have a pronounced impact on the effectiveness of nozzle blowing. Although the effect on pitching moment was small, nozzle blowing on a chined forebody produced substantially large rolling moments in the region of maximum lift. The rolling moment was proverse, however, and could be used to coordinate wind-axis roll maneuvers at high angles of attack.

Validation of a multipoint approach for modeling spin aerodynamics

Abstract: during the spin. The rolling-moment and yawing-moment coefficient variations were relatively small, and thus, the steadyspin assumption is reasonable for these quantities. In contrast is the large amplitude of the pitching moment coefficient trace. The maximum and minimum values of the free-spin results are approximately equal to the rotary balance coefficient ±90%. Clearly, the steady-state results do not capture the true nature of the moment coefficient in this case. This trend would obviously be more exaggerated if the spin became more oscillatory.

Sailplane Glide Performance and Control Using Fixed and Articulating Winglets

Abstract: An experimental study was conducted to investigate the effects of controllable articulating winglets on glide performance and yawing moments of high performance sailplanes. Testing was conducted in the Texas A&M University 7 x 10 foot Low Speed Wind Tunnel using a full-scale model of the outboard 5.6 feet of a 15 meter class high performance sailplane wing. Different wing tip configurations could be easily mounted to the wing model. A winglet was designed in which the cant and toe angles as well as a rudder on the winglet could be adjusted to a range of positions. Cant angles used in the investigation consisted of 5, 25, and 40 degrees measured from the vertical axis. Toe-out angles ranged from 0 to 22.5 degrees. A rudder on the winglet was used to study the effects of changing the camber of the winglet airfoil on wing performance and wing yawing moments. Rudder deflections consisted of-10, 0, and 10 degrees. Test results for a fixed geometry winglet and a standard wing tip are presented to show the general behavior of winglets on sailplane wings, and the effects of boundary-layer turbulators on the winglets are also presented. By tripping the laminar boundary-layer to turbulent before laminar separation occurs, the wing performance was increased at low Reynolds numbers. The effects on the lift and drag, yawing moment, pitching moment, and wing root bending moment of the model are presented. Oil flows were used on the wing model with the fixed geometry winglet and the standard wing tip to visualize flow directions and areas of boundary layer transition. A cant angle of 25 degrees and a toe-out angle of 2.5 degrees provided an optimal increase in wing performance for the cant and toe angles tested. Maximum performance was obtained when the winglet rudder remained in the neutral position of zero degrees. By varying the cant, toe, and rudder angles from their optimized positions, wing performance decreases. Although the winglet rudder proved to be more effective in increasing the yawing moment compared to varying the cant and toe angles, the amount of increased yawing moment was insignificant when compared to that produced by the vertical tail. A rudder on the winglet was determined to be ineffective for providing additional yaw control.

Interpretation of Experimental High-Alpha Aerodynamics—Implications for Flight Prediction

Abstract: Nomenclature b = wingspan, or largest dimension C — aerodynamic coefficient; with no superscript, in body axes system Cij = dCi/d(jl/2V), i = /, m, n\ j = q, q', d, a(I = c)'J =p,j8, (/ = b) Cik = dCf/dk, i = l,m,n\k = a, /3, a Cl = rolling moment coefficient Cm = pitching moment coefficient C,, = yawing moment coefficient Cp = static pressure coefficient CY = side force coefficient c = mean aerodynamic chord d — body maximum diameter h = height of test section h() = model length / = generalized reference length M^ = freestream Mach number /?, px = local static pressure /?, q, r = body-axes angular velocities q^ = freestream dynamic pressure Re, Reynolds number based on /, / c, b, or d S = reference area V, V^ = freestream velocity w = minimum dimension of test section X, Y, Z = inertial axes system Xf, Yf; Zf = flight coordinate system, Fig. 2 x, _y , z = body axes system a, /3 = angles of attack and sideslip A = increment or amplitude A = inclination of rotation axis, Fig. 6 = coning rate, parameter fib/2V or £1 a) = reduced circular frequency, coll(2V), where / = c or b, as appropriate

Physical aspects of stall hysteresis on an airfoil with slotted flap

Abstract: This article reports details of stall hysteresis in wind tunnel tests of the GA(Wp2 airfoil with 25% slotted flap. The data included force, pitching moment, surface pressure, and flow visualization at a Mach number of 0.13 and Reynolds number of 2.2x106, close to the flight range of general aviation airplanes. Stall hysteresis on the flapped-airfoil configurations was discussed in parallel to that encountered on single-element airfoils at relatively low Reynolds numbers. Attention was given to explore the slot flow effects and associated boundary layer flow events on flap setting optimized for high Cha. The study reveals that the stall hysteresis on the two-element airfoil is primarily controlled by the flap flow field, whose suction pressures are leveled at post-stall a. Suppressed flat suction pressures on the flap are not “unsuppressed” within the a limits of hysteresis loop.

Aerodynamic Characteristics of Rotor Airfoils as Affected by Simulated Ballistic Damage.

Abstract: : Tests were made using a 2-D insert in a subsonic wind tunnel to examine the effects of simulated ballistic damage on the aerodynamic characteristics of ten helicopter rotor blade sections. Two undamaged baseline blade sections, comprised of SC1095 and SC1095R8 airfoils, were tested and then modified with different simulated ballistic damage configurations. These comprised of a circular hole with the surrounding skin removed, fore and aft circular holes, and an aft wedge-shaped hole. One blade section was subjected to actual ballistic damage near the trailing edge. The sectional lift, drag and pitching moment were measured at positive and negative angles of attack at Reynolds numbers of one, two, and three million. Pressure measurements were also made for two configurations fitted with pressure taps. Additional tests were conducted over a full 360 degree range in angle of attack for a Reynolds number of one million. The measurements were complemented by oil flow visualization on the blade sections. The simulated damage caused large disturbances in the flowfield near and downwind of the damaged regions. Generally, flow separation was initiated at the upstream leading edge of the damage, followed by a growth in separation, both in span and intensity, with increasing angle of attack. The aerodynamic characteristics were significantly degraded, with up to a 60 percent reduction in lift-curve-slope, a loss of maximum lift capability of nearly 30 percent, and a significant decrease in the lift-to-drag ratio due to drag increases of up to nearly 340 percent at low angles of attack.

Application of an Euler Code on a modern combat aircraft configuration

Abstract: An Euler Code is applied on a modern combat aircraft configuration to study the effect of deflections of leading edge high lift devices and control surfaces on the aerodynamic coefficients. The calculations for deflection of Leading Edge Vortex. CONtroller (LEVCON) and Elevons are done at Mach No. = 0.7, and angle of attack = 10° and for deflection of leading edge slats at Mach No. = 0.95, angle of attack = 7.5°. It is found that the incremental change in the aerodynamic coefficients — lift, induced drag, and the pitching moment due to the deflections of LEVCON, slats and elevon are predicted reasonably accurate both quantatively and qualitatively.

Balances for the measurement of multiple components of force in flows of a millisecond duration

Abstract: Paper reports a new balance for the measurement of three components of force — lift, drag and pitching moment — in impulsively started flows which have a duration of about one millisecond. The basics of the design of the balance are presented and results of tests on a 15° semi-angle cone set at incidence in the T4 shock tunnel are compared with predictions. These results indicate that the prototype balance performs well for a 1.9 kg, 220 mm long model. Also presented are results from initial bench tests of another application of the deconvolution force balance to the measurement of thrust produced by a 2D scramjet nozzle.