Showing papers on "Pitching moment published in 1990"

Mechanics of Forward Flight in Bumblebees: II. QUASI-STEADY LIFT AND POWER REQUIREMENTS

Abstract: This paper examines the aerodynamics and power requirements of forward flight in bumblebees. Measurements weremade of the steady-state lift and drag forces acting on bumblebee wings and bodies. The aerodynamic force and pitching moment balances for bumblebees previously filmed in free flight were calculated. A detailed aerodynamic analysis was used to show that quasi-steady aerodynamic mechanisms are inadequate to explain even fast forward flight. Calculations of the mechanical power requirements of forward flight show that the power required to fly is independent of airspeed over a range from hovering flight to an airspeed of 4.5 ms −1

State-space representation of unsteady airfoil behavior

Abstract: A method is presented to model the unsteady lift, pitching moment, and drag acting on a two-dimensional airfoil operating under attached-flow conditions in a compressible flow. Starting from suitable generalizations and approximations to aerodynamic indicial functions, the unsteady airloads due to an artibrary forcing are represented in a state-space (differential equation) form. This model is in a form compatible with the aeroelastic analyses of both fixed-wing and rotary-wing systems

Dynamic stall of a constant-rate pitching airfoil

Abstract: A computational study is presented for the dynamic stall of an airfoil that is pitched at a constant rate from zero incidence to a high angle of attack. The unsteady flow is simulated employing the mass-averaged NavierStokes equations and an algebraic turbulent eddy viscosity model. The approach is first validated by comparison of computed and experimental results for a pitching airfoil at low freestream Mach numbers. The computed dynamic stall events, as well as the computed effects of pitch rate and axis location, are found in qualitative agreement with experimental observations. The effect of compressibility on dynamic stall is investigated. As the freestream Mach number increases, the appearance of a supersonic region provides—through the shock/boundarylayer interaction—an additional mechanism in the dynamic stall process. The main effects of compressibility are found to be 1) a change from trailing-edge stall to leading-edge stall and 2) a reduction in the stall delay and in the attained maximum lift.

All wing aircraft

Abstract: An all-wing aircraft is disclosed that has novel foreplane and depending aftplane, a center wing section and outer wing panel flying surfaces which cooperate aerodynamically to eliminate the need for conventional fuselage and tail structures. The foreplanes are strategically located to create a positive pitching moment which is sufficient to significantly reduce elevator forces and to balance the negative pitching moment induced by the outer wing panels and the downwardly extending aftplane units that provide static and dynamic pitch and yaw stability. At the same time, the foreplane structures serve as an unobstructed means to mount engines forward on the airframe to established a forward empty center of gravity. Additionally the aftplane structure serves as a means to mount main landing gears, elevators, ruddervators and to provide structural means for interconnecting outboard wing sections of the airplane. The all-wing design is of a nature permitting adoption of the principles thereof in either a multi-engine airplane or single-engine airplane. In both instances, a center wing section, foreplane and aftplane structures and outer wing sections cooperate to provide stability in all three axes of movement of the airplane while decreasing fuel burn by virtue of an improved lift to drag ratio and empty weight reduction by removal of the conventional fuselage and tail structures.

Numerical Computations of Transonic Critical Aerodynamic Behavior

Abstract: The determination of aerodynamic coefficients by shell designers is a critical step in the development of any new projectile design. Of particular interest is the determination of the aerodynamic coefficients at transonic speeds. It is in this speed regime that the critical aerodynamic behavior occurs and a rapid change in the aerodynamic coefficients is observed. Three-dimensional, transonic, flowfield computations over projectiles have been made using an implicit, approximately factored, partially flux-split algorithm. A composite grid scheme has been used to provide the increased grid resolution needed for accurate numerical simulation of three-dimensional transonic flows. Details of the asymmetrically located shock waves on the projectiles have been determined. Computed surface pressures have been compared with experimental data and are found to be in good agreement. The pitching moment coefficient, determined from the computed flowfields, shows the critical aerodynamic behavior observed in free flights. I. Introduction T HE flight of projectiles covers a wide range of speeds. The accurate prediction of projectile aerodynamics at these speeds is of significant importance in the early design stage of a projectile. The critical aerodynamic behavior occurs in the transonic speed regime, 0.9

Unsteady aerodynamic characteristics of a fighter model undergoing large-amplitude pitching motions at high angles of attack

Abstract: The effects of harmonic or constant-rate-ramp pitching motions (giving angles of attack from 0 to 75 deg) on the aerodynamic performance of a fighter-aircraft model with highly swept leading-edge extensions are investigated experimentally in the NASA Langley 12-ft low-speed wind tunnel. The model configuration and experimental setup are described, and the results of force and moment measurements and flow visualizations are presented graphically and discussed in detail. Large force overshoots and hysteresis are observed and attributed to lags in vortical-flow development and breakup. The motion variables have a strong influence on the persistence of dynamic effects, which are found to affect pitch-rate capability more than flight-path turning performance.

Navier-Stokes predictions of pitch damping for finned projectiles using steady coning motion

Abstract: Previous theoretical investigations have proposed that the side force and moment acting on a body of revolution in steady coning motion could be related to the pitch-damping force and moment. In the current research effort, this approach has been applied to produce the first known Navier-Stokes predictions of the pitch damping for finned projectiles. The flow field about finned kinetic energy projectiles in steady coning motion has been successfully computed using a parabolized Navier-Stokes computational approach. The computations make use of a rotating coordinate frame in order to solve the steady flow equations. From the computed flow field, the side moment due to coning motion is used to determine the pitchdamping coefficient. The computational predictions of the slope of the side moment coefficient with coning rate normalized by the sine of the angle of attack have been compared with pitch damping coefficients determined from range firings for two finned projectile configurations. The predictions show good agreement with the range data. This computational approach provides a significant predictive capability for the design of kinetic energy projectiles whose terminal ballistic performance can be degraded by moderate levels of yaw at the target. Nomenclature a00 freestream speed of sound c m pitching moment coefficient c m , slope of the pitching moment coefficient with angle of attack Cm, + Cm, pitch damping moment coefficient c n side moment coefficient * Senior Member, AIAA t Associate Fellow, AIAA Cn, slope of the side moment coefficient mith angle of attack c n slope of the side moment coefficient with coning rate c n , Magnus moment coefficient Chra slope of the normal force coefficient with angle of attack CN, + CN, pitch damping force coefficient slope of the side force coefficient with coning rate Magnus force coefficient projectile diameter total energy per unit volume flux vectors in transformed coordinates source term in Navier-Stokes eqs. jacobian characteristic length, typically D freestream Mach number pressure, as used in N-S eqs. spin rate, as used roll equations freestream static pressure Reynolds number, amp, Dip, distance downrange center of gravity shift, calibers viscous flux vector reference area of projectile, 7 r ~ ~ / 4 time velocity components in x,y,z directions freest ream velocity Cartesian coordinates w.r.t. body axial location of body center of gravity Note: Force coefficients are scaled, F / ; ~ , a& M& Srej ; Moment coefficients are scaled, M / f p, a& M& DS,,~

Analysis of cone-derived waveriders by hypersonic small-disturbance theory

Abstract: Hypersonic small-disturbance theory is used to analyze slender waveriders derived from axisymmetric flows past circular cones. Viscous effects are accounted for by means of laminar boundary-layer theory. The shape of the waverider is determined by specifying the upper freestream surface, from which the lower compression surface can then be obtained. When this is done, the lift, drag, and pitching moment are determined in terms of quadratures over the shock layer in the base plane. They are functions of freestream Mach number, freestream Reynolds number based on the length of the waverider, some measure of slenderness, and other parameters relating to the shape of the waverider. The functional relationships can be cast in terms of similarity laws. The lift-to-drag ratio is determined for a wide range of shapes and parameters.

The aerodynamic interference between tanker and receiver aircraft during air-to-air refuelling

Abstract: Wind tunnel data from a tanker wing and receiver aircraft model at varying vertical separation have been compared with theoretical results. In the aerodynamic model the tanker wing is represented by a horseshoe vortex while the aerodynamic loads on the receiver are determined by the vortex lattice method and lifting-line theory, although an approximate method is used to determine the side force on the fin. In the longitudinal case data were obtained for low, mid and high tailplane positions and, with the exception of the pitching moment results, fairly good agreement is obtained between theory and experiment. The relatively small differences are due mainly to the wind tunnel boundary interference effect which could not be quantified for the pitching moment measurements. The lateral aerodynamic interference was determined by banking the tanker wing and displacing it sideways and by yawing the receiver model. Fairly good agreement is obtained between the theory and experiment for the most significant terms which are the rolling moments due to bank and sideways displacement. The effect of the sidewash due to the tanker wake on the receiver in yaw is found to be relatively insignificant. Over the range of bank, yaw and sideways displacements tested the results are almost linear.

Measured forces and moments on a delta wing during pitch-up

Abstract: A series of low-speed wind tunnel tests on a 70-deg, sharp, leading-edge delta wing undergoing ramp pitching motion of high amplitude were performed to investigate the aerodynamic forces and moments. Forces and moments were obtained from a six-component interanl balance. Large amplitude oscillatory motion was produced by sinusoidally oscillating the model over a range of reduced frequencies. Ramp motion was produced by pitching the model through a half cycle of sinusoidal motion at a root chord Reynolds number of 1.54 million. The effect of ramp and oscillatory motions on the forces and moments are almost identical at matched pitch rates. Pitch rate had strong effect on the magnitude of the aerodynamic forces and moments. Upon completion of the model motion, some time is required for the forces and moments to decay to their static values. This convergence of the dynamic values to the static ones was a function of the pitch rate.

Subsonic and Transonic Low-Reynolds-Number Airfoils with Reduced Pitching Moments

Abstract: A subsonic and a transonic airfoil are presented for application in a high-altitude long-endurance aircraft and a very-high-altitude aircraft, respectively. The subsonic airfoil is designed for a lift coefficient c(l) = 1.4 at a chord Reynolds number Re = 700,000 and a very low Mach number. The transonic airfoil is designed for c(l) = 1.0 at Re = 500,000 and a transonic Mach number M = 0.7. Both airfoils are developed to perform as well or better than previously designed airfoils. However, the present airfoils are developed for a constrained pitching moment to reduce aircraft trim drag and to relieve, to some extent, the torsional loads in the typically high-aspect-ratio wings. The beneficial effects of a cruise flap and of boundary-layer transition control on the off-design performance characteristics are illustrated.

Analytical aerodynamic model of a high alpha research vehicle wind-tunnel model

Abstract: A 6 DOF analytical aerodynamic model of a high alpha research vehicle is derived. The derivation is based on wind-tunnel model data valid in the altitude-Mach flight envelope centered at 15,000 ft altitude and 0.6 Mach number with Mach range between 0.3 and 0.9. The analytical models of the aerodynamics coefficients are nonlinear functions of alpha with all control variable and other states fixed. Interpolation is required between the parameterized nonlinear functions. The lift and pitching moment coefficients have unsteady flow parts due to the time range of change of angle-of-attack (alpha dot). The analytical models are plotted and compared with their corresponding wind-tunnel data. Piloted simulated maneuvers of the wind-tunnel model are used to evaluate the analytical model. The maneuvers considered are pitch-ups, 360 degree loaded and unloaded rolls, turn reversals, split S's, and level turns. The evaluation finds that (1) the analytical model is a good representation at Mach 0.6, (2) the longitudinal part is good for the Mach range 0.3 to 0.9, and (3) the lateral part is good for Mach numbers between 0.6 and 0.9. The computer simulations show that the storage requirement of the analytical model is about one tenth that of the wind-tunnel model and it runs twice as fast.

Experimental and predicted aerodynamic characteristics of a proposed Assured Crew Return Vehicle (ACRV) lifting-body configuration at Mach 6 and 10

Abstract: The effects of Mach number, Reynolds number, and ratio of specific heats on the aerodynamic characteristics of a proposed Assured Crew Return Vehicle (ACRV) lifting-body configuration were examined for a range of angles of attack from -5 deg to 50 deg. Predictions made with a Langley-developed, three-dimensional Navier Stokes solver known as LAURA, which was exercised as an Euler solver for the present study, are compared with the experimental results. Unlike the Shuttle Orbiter, which experienced a significant nose-up increment in pitching moment with decreasing specific heat ratio (i.e., real gas effects), the aerodynamic characteristics of this lifting-body configuration are insensitive to changes in specific heat ratio. The maximum trimmed lift-to-drag ratio achieved was about 1.5. Predicted inviscid values of aerodynamic coefficients were generally in good agreement with measurement.

Stability of floating and submerged blocks

Abstract: The rotational stability of floating and submerged rectangular blocks is described. The limit of stability is reached when the underturning moment acting on the block is equal to the maximum hydrostatic righting moment. The hydrostatic righting moment is derived and a convenient expression for its maximum is presented in nondimensional form. A moment coefficient is defined that relates the underturning moment at the limit of stability to the moment produced by the product of the dynamic pressure of the flow and the plan area of the block. An exponential function of the ratio of block thickness to flow depth is postulated as a general expression for the moment coefficient. The parameters of this function are related to the block geometry by analyzing the existing experimental data. The limit of rotational stability for rectangular blocks can then be described in terms of a densimetric Froude number based on block thickness.

Dual strain gage balance system for measuring light loads

Abstract: A dual strain gage balance system for measuring normal and axial forces and pitching moment of a metric airfoil model imparted by aerodynamic loads applied to the airfoil model during wind tunnel testing includes a pair of non-metric panels being rigidly connected to and extending towards each other from opposite sides of the wind tunnel, and a pair of strain gage balances, each connected to one of the non-metric panels and to one of the opposite ends of the metric airfoil model for mounting the metric airfoil model between the pair of non-metric panels. Each strain gage balance has a first measuring section for mounting a first strain gage bridge for measuring normal force and pitching moment and a second measuring section for mounting a second strain gage bridge for measuring axial force.

Supersonic aerodynamic characteristics of a Mach 3 high-speed civil transport configuration

Abstract: The NASA-Langley AST31 high-speed civil transport (HSCT) configuration, a blended wing/fuselage concept scaled for 250-passenger carriage over 6500 nmi, has been subjected to both CFD and wind tunnel tests; the wind tunnel tests were conducted in the Mach 1.6-3.6 range. The inability of the linear theory method to accurately predict stability levels as well as nonlinear pitching moment characteristics demonstrates the need to undertake high-order analyses as early in the design process as possible. An HSCT laminar flow capability experimental study was conducted in order to simulate 50-percent laminar flow conditions, and thereby ascertain the difference in drag relative to fully turbulent conditions.

Pitch control trimming system for canard design aircraft

Abstract: A pitch control trimming system for a canard design aircraft has the ability to reposition the center of gravity of the aircraft along the longitudinal axis. The design permits trimming of the aircraft without external trim tabs and permits flaps to be used on a canard design aircraft without a nose down or negative pitching moment occurring. The aircraft has a substantial mass associated with the aircraft preferably the power plant including engine and propeller, which is movable substantially along the longitudinal axis of the aircraft, and includes an actuator to move the mass forward and aft on the longitudinal axis to reposition the center of gravity of the aircraft.

An in-flight investigation of ground effect on a forward-swept wing airplane

Abstract: A limited flight experiment was conducted to document the ground effect characteristics of the X-29A research aircraft. This vehicle has an aerodynamic platform which includes a forward-swept wing and close-coupled, variable incidence canard. The flight-test program obtained results for errors in the air data measurement and for incremental normal force and pitching moment caused by ground effect. Correlations with wind-tunnel and computational analyses were made. The results are discussed with respect to the dynamic nature of the flight measurements, similar data from other configurations, and pilot comments. The ground effect results are necessary to obtain an accurate interpretation of the vehicle's landing characteristics. The flight data can also be used in the development of many modern aircraft systems such as autoland and piloted simulation.

The aerodynamic design of the oblique flying wing supersonic transport

Abstract: The aerodynamic design of a supersonic oblique flying wing is strongly influenced by the requirement that passengers must be accommodated inside the wing. It was revealed that thick oblique wings of very high sweep angle can be efficient at supersonic speeds when transonic normal Mach numbers are allowed on the upper surface of the wing. The goals were motivated by the ability to design a maximum thickness, minimum size oblique flying wing. A 2-D Navier-Stokes solver was used to design airfoils up to 16 percent thickness with specified lift, drag and pitching moment. A new method was developed to calculate the required pressure distribution on the wing based on the airfoil loading, normal Mach number distribution and theoretical knowledge of the minimum drag of oblique configurations at supersonic speeds. The wing mean surface for this pressure distribution was calculated using an inverse potential flow solver. The lift to drag ratio of this wing was significantly higher than that of a comparable delta wing for cruise speeds up to Mach 2.

Subsonic sting interference on the aerodynamic characteristics of a family of slanted-base ogive-cylinders

Abstract: Support interference free drag, lift, and pitching moment measurements on a range of slanted base ogive cylinders were made using the NASA Langley 13 inch magnetic suspension and balance system. Typical test Mach numbers were in the range 0.04 to 0.2. Drag results are shown to be in broad agreement with previous tests with this configuration. Measurements were repeated with a dummy sting support installed in the wind tunnel. Significant support interferences were found at all test conditions and are quantified. Further comparison is made between interference free base pressures, obtained using remote telemetry, and sting cavity pressures.

Calculation of Real-Gas Effects on Airfoil Aerodynamic Characteristics

Abstract: The effects of high temperature thermochemical phenomena on the aerodynamic characteristics at hypersonic speeds are calculated for two-dimensional airfoils in air. The calculations are performed on an airfoil similar to that used for the Space Shuttle Orbiter, and ellipses of thickness ratios varying between 5 and 15 percent. For the airfoil, one flight condition is considered. For the ellipses, the calculations are carried out over a range of chord lengths, flight velocities, flight altitudes, and angles of attack. It is shown that the lift and drag coefficients are consistently reduced by the thermochemical phenomena, and that the behavior can be represented by a specific heat ratio value less than 1.4. The center of pressure shifts forward due to the thermochemical phenomena, but its extent is sensitively affected by the geometry and angle of attack and cannot be represented by a fixed specific heat ratio. The calculated results are in qualitative agreement with the data obtained during the entry flights of the Space Shuttle vehicle.

An airfoil for general aviation applications

Abstract: A new airfoil, the NLF(1)-0115, has been recently designed at the NASA Langley Research Center for use in general-aviation applications. During the development 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 having large amounts of lift, the NLF(1)-0115 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 drags if cruise flaps are not employed. The NASA NLF(1)-0115 has a thickness of 15 percent. It is designed primarily for general-aviation aircraft with wing loadings of 718 to 958 N/sq m (15 to 20 lb/sq ft). Low profile drag as a result of laminar flow is obtained over the range from c sub l = 0.1 and R = 9x10(exp 6) (the cruise condition) to c sub l = 0.6 and R = 4 x 10(exp 6) (the climb condition). While this airfoil can be used with flaps, it is designed to achieve c(sub l, max) = 1.5 at R = 2.6 x 10(exp 6) without flaps. The zero-lift pitching moment is held at c sub m sub o = 0.055. The hinge moment for a .20c aileron is fixed at a value equal to that of the NACA 63 sub 2-215 airfoil, c sub h = 0.00216. The loss in c (sub l, max) due to leading edge roughness, rain, or insects at R = 2.6 x 10 (exp 6) is 11 percent as compared with 14 percent for the NACA 23015.

Unsteady surface pressure distributions on a delta wing undergoing large amplitude pitching motions

Abstract: Wind tunnel experiments were performed on a 70-deg-sweep delta wing to determine the effect of a sinusoidal pitching motion on the pressure field on the suction side of the wing. Pressure taps were placed from 35-90 percent of the chord, at 60 percent of the local semi-span. Pressure coefficients were measured as functions of Reynolds number and pitch rate. The surface pressure distribution was seen to vary at the same frequency as the pitching frequency, though distortion due to the vortex breakdown was observed. Comparing the upstroke (angle of attack increasing) and downstroke (angle of attack decreasing) pressures for a specific angle of attack, a time lag in the pressure distribution was observed. The downstroke pressures were slightly larger at the forward chord locations. Vortex breakdown was seen to have the most significant effect at the 40-45-percent chord location, where an increase in local pressure was apparent, as well as a distortion of the periodic pressure fluctuation.

Rotor-blade for helicopter, main-rotor for helicopter and helicopter

Abstract: PURPOSE: To reduce aerodynamic pitching moments by providing blade droop for changing the distribution of a bound vortex in a blade tip edge so as to reduce the aerodynamic pitching moments of a rotor blade during forward flight. CONSTITUTION: A main rotor blade 11 has a front edge 12 and a rear edge 13 for defining a chord of blade, a root end 14 and a retracted blade tip edge 30, rotated about an axis 16 and has a feathering axis 18 at the position of about 1/4 as short as a point of the chord of blade. During forward flight, a bound vortex 22 generates local lift load distribution 31 caused by air current 24 widthwise the blade and generates one pitching moment for each rotation. Thus, a droop plane is defined by a droop line 40 on a blade tip part 35 along the retracted blade tip edge 36. Thus, the aerodynamic pitching moments can be reduced or removed.

Wing effects on asymmetric vortex formation for a ship-launched missile

Abstract: Wind tunnel tests were performed on a vertically-launched surface-to-air missile model to investigate the effects of low-aspect-ratio strakes and wings on the side forces and yawing moments induced by nose-generated asymmetric vortices at high angles of attack. Force and moment data were collected for angles of attack from 0 to 90 deg at a Reynolds number, based on model diameter, of 1.1 x lo5, and at a Mach number of 0.1. In a second series of tests at 50 deg angle of attack, flowfield measurements were made at two crossplane locations for vortex positions and strengths. Three body configurations were tested. Induced side forces and yawing moments were found to be maintained with the addition of wings and strakes. At the onset of the asymmetric vortices, yawing moments were of opposite sign to the induced side forces, indicating the dominance of afterbody side forces. However, as side forces increased with angle of attack, the forebody vortices quickly became dominant. Velocity vector plots, total pressure loss contours, and vorticity contours all proved useful in characterizing the lee flowfield. Though the lee flowfield for the two winged configurations (0 and 45 deg roll angle) differed greatly, both maintained identical levels of yawing moment coefficient .

Experimental and theoretical aerodynamic characteristics of a high-lift semispan wing model

Abstract: Experimental and theoretical aerodynamic characteristics were compared for a high-lift, semispan wing configuration that incorporated a slightly modified version of the NASA Advanced Laminar Flow Control airfoil section. The experimental investigation was conducted in the Langley 14- by 22-Foot Subsonic Tunnel at chord Reynolds numbers of 2.36 and 3.33 million. A two-dimensional airfoil code and a three-dimensional panel code were used to obtain aerodynamic predictions. Two-dimensional data were corrected for three-dimensional effects. Comparisons between predicted and measured values were made for the cruise configuration and for various high-lift configurations. Both codes predicted lift and pitching moment coefficients that agreed well with experiment for the cruise configuration. These parameters were overpredicted for all high-lift configurations. Drag coefficient was underpredicted for all cases. Corrected two-dimensional pressure distributions typically agreed well with experiment, while the panel code overpredicted the leading-edge suction peak on the wing. One important feature missing from both of these codes was a capability for separated flow analysis. The major cause of disparity between the measured data and predictions presented herein was attributed to separated flow conditions.

Dynamic ground-effect measurements on the F-15 STOL and Maneuver Technology Demonstrator (S/MTD) configuration

Abstract: A moving-model ground-effect testing method was used to study the influence of rate-of-descent on the aerodynamic characteristics for the F-15 STOL and Maneuver Technology Demonstrator (S/MTD) configuration for both the approach and roll-out phases of landing. The approach phase was modeled for three rates of descent, and the results were compared to the predictions from the F-15 S/MTD simulation data base (prediction based on data obtained in a wind tunnel with zero rate of descent). This comparison showed significant differences due both to the rate of descent in the moving-model test and to the presence of the ground boundary layer in the wind tunnel test. Relative to the simulation data base predictions, the moving-model test showed substantially less lift increase in ground effect, less nose-down pitching moment, and less increase in drag. These differences became more prominent at the larger thrust vector angles. Over the small range of rates of descent tested using the moving-model technique, the effect of rate of descent on longitudinal aerodynamics was relatively constant. The results of this investigation indicate no safety-of-flight problems with the lower jets vectored up to 80 deg on approach. The results also indicate that this configuration could employ a nozzle concept using lower reverser vector angles up to 110 deg on approach if a no-flare approach procedure were adopted and if inlet reingestion does not pose a problem.

Design of a three-component wall-mounted balance

Abstract: The design and evaluation of a three-component, wall-mounted pyramidal balance for a small wind tunnel is discussed. The balance was designed to measure lift, drag, pitching moment, and angle of attack. The specific design of each component and mathematical models used to design the balance are covered. Balance evaluation consisted of calibration, tare, and interaction analysis.

Prediction of forces and moments for flight vehicle control effectors. Part 1: Validation of methods for predicting hypersonic vehicle controls forces and moments

Abstract: Many types of hypersonic aircraft configurations are currently being studied for feasibility of future development. Since the control of the hypersonic configurations throughout the speed range has a major impact on acceptable designs, it must be considered in the conceptual design stage. The ability of the aerodynamic analysis methods contained in an industry standard conceptual design system, APAS II, to estimate the forces and moments generated through control surface deflections from low subsonic to high hypersonic speeds is considered. Predicted control forces and moments generated by various control effectors are compared with previously published wind tunnel and flight test data for three configurations: the North American X-15, the Space Shuttle Orbiter, and a hypersonic research airplane concept. Qualitative summaries of the results are given for each longitudinal force and moment and each control derivative in the various speed ranges. Results show that all predictions of longitudinal stability and control derivatives are acceptable for use at the conceptual design stage. Results for most lateral/directional control derivatives are acceptable for conceptual design purposes; however, predictions at supersonic Mach numbers for the change in yawing moment due to aileron deflection and the change in rolling moment due to rudder deflection are found to be unacceptable. Including shielding effects in the analysis is shown to have little effect on lift and pitching moment predictions while improving drag predictions.