Showing papers on "Pitching moment published in 1980"

Upper surface blown powered lift system for aircraft

Abstract: In an aircraft having a pair of engines mounted forwardly of the wing and discharging jet exhaust over the wing for augmented lift, a trailing edge flap is positioned directly behind each of the engines. Each flap has a stowed position for the cruise mode, a rearwardly and downwardly extending position for maximum deflection of the jet exhaust downwardly for STOL operation, and various intermediate positions. The upper aerodynamic surface of each flap is curved in a manner that the radius of curvature increases along the chord length of the flap from the forward knee of the flap to the trailing edge of the flap. Each flap is moved from its stowed position, through intermediate positions to its fully deployed position in a manner that the upper aerodynamic surface of each flap forms a smooth and continuous extension of the fixed rear portion of the upper aerodynamic surface of the wing. Vortex generators are provided on the upper wing surfaces forward of flaps to generate vortices in the jet exhaust that travels over the upper surfaces of the flaps. This arrangement results in improved lift characteristics, reduced pitching moment with the flap in its deployed position, and improved operating characteristics.

Effect of Interblade Phase Angle and Incidence Angle on Cascade Pitching Stability

Abstract: A comprehensive test program was performed at low subsonic velocity on a linear cascade of airfoils oscillating in pitch about their midchords for incidence angles up to 10 deg, reduced frequencies up to 0.193, and over a range of interblade phase angles from σ = −60 deg to +60 deg. The test conditions represent significant changes in blade loading and dimensionless frequency, and the range of interblade phase angle includes those values usually encountered in actual turbomachines. The measured pressure time histories over the airfoil chord were used to calculate the stability parameters of the system including the unsteady pitching moment coefficient and the aerodynamic damping parameter. For the range of parameters tested it was found that the interblade phase angle is the most important parameter affecting the stability of oscillating cascaded airfoils. The system was unstable for most positive values of σ over the entire range of loading and frequency. This was similar in behavior (but not in magnitude) to the predictions of available potential flow cascade theories and suppports the observation that blade stall need not be present for torsional “stalled” flutter to occur. System stability for negative values of σ was more dependent on loading and frequency and, conformed more closely with the observed behavior of stalled flutter. Specifically, for σ < 0 deg stability increased with frequency and decreased with loading. A preliminary evaluation of the pressure time histories shows that a second harmonic behavior renders the 1.2 percent chord station ineffective in contributing to the blade damping. Under these circumstances it is surmised that the induced damping is associated mainly with the first harmonic component of the pressure response at the 6.2 percent chord station.

Low-speed aerodynamic characteristics of a 17-percent-thick medium speed airfoil designed for general aviation applications

Abstract: Wind tunnel tests were conducted to determine the low speed two dimensional aerodynamic characteristics of a 17 percent thick medium speed airfoil (MS(1)-0317) designed for general aviation applications. The results were compared with data for the 17 percent thick low speed airfoil (LS(1)-0417) and the 13 percent thick medium speed airfoil (MS(1)-0313). Theoretical predictions of the drag rise characteristics of this airfoil are also provided. The tests were conducted in the Langley low turbulence pressure tunnel over a Mach number range from 0.10 to 0.32, a chord Reynolds number range from 2 million to 12 million, and an angle of attack range from about -8 to 20 deg.

A system for aerodynamic design and analysis of supersonic aircraft. Part 1: General description and theoretical development

Abstract: An integrated system of computer programs was developed for the design and analysis of supersonic configurations. The system uses linearized theory methods for the calculation of surface pressures and supersonic area rule concepts in combination with linearized theory for calculation of aerodynamic force coefficients.

Airplane all-moving airfoil with moment reducing apex

Abstract: Disclosed is an airplane all-moving airfoil having a moment reducing apex for facilitating control of the airfoil. In a preferred embodiment, the apex protrudes forward from the leading edge of an all-moving horizontal stabilizer and operates through its aerodynamic effect on the stabilizer to reduce the moment required to maintain and vary the stabilizer position. Counter-rotating airflow vortices produced by the apex reduce the rearward displacement of the center of pressure on the stabilizer as the stabilizer is deflected into an increasing angle of attack. As a result, lighter weight hydraulic stabilizer and elevator actuating mechanisms can be employed. In a preferred embodiment an aeroelastically flexible apex is employed to enhance moment reduction at high angles of attack.

Upper surface blown powered lift system

Abstract: In an aircraft having a pair of engines mounted forwardly of the wing and discharging jet exhaust over the wing for augmented lift, a trailing edge flap is positioned directly behind each of the engines. Each flap has a stowed position for the cruise mode, a rearwardly and downwardly extending position for maximum deflection of the jet exhaust downwardly for STOL operation, and various intermediate positions. The upper aerodynamic surface of each flap is curved in a manner that the radius of curvature is a constant and uniform minimum at a forward portion of the flap and substantially greater at a rear portion of the flap. Each flap is moved from its stowed position, through intermediate positions to its fully deployed position in a manner that the upper aerodynamic surface of each flap forms a smooth and continuous extension of the fixed rear portion of the upper aerodynamic surface of the wing. Vortex generators are provided on the upper wing surfaces forward of flaps to generate vortices in the jet exhaust that travels over the upper surfaces of the flaps. This arrangement results in improved lift characteristics, reduced pitching moment with the flap in its deployed position, and improved operating characteristics.

Supersonic Wind Tunnel Measurements of Static and Magnus Aerodynamic Coefficients for Projectile Shapes with Tangent and Secant Ogive Noses

Abstract: : Wind tunnel tests have been conducted for a series of projectile configurations. The shapes tested include a three caliber secant ogive nose and three caliber tangent ogive nose configuration. The overall model length was fixed at 6 calibers for all configurations tested. Static and Magnus aerodynamic coefficient data were obtained at Mach numbers 2.0, 3.0 and 4.0 for angles of attack up to 10.0 deg. The data are presented in graphical form along with tabulations of the summary data. The tangent ogive nose produced a slight increase in the Magnus moment coefficient when compared to similar secant ogive configurations. The boattail configuration, when compared to the cylinder shapes, was found to increase the Magnus moment; however, this effect was shown to decrease with increasing Mach number.

Flow Patterns and Aerodynamic Characteristics of a Wing-Strake Configuration

Abstract: Surface oil flow visualizations, force tests, and pressure measurements were conducted at low, transonic, and supersonic speeds. Four flow patterns on a wing with strake at low speed have been found. The flow on the wing upper surface is affected and controlled by the formation, development, and breakdown of the strake vortices. The differences in flow patterns are reflected in the force and moment results. The lift increment results from the effect of the strake vortex not only on the inner panel but also on the outer panel. The nonlinear variations of the pitching moment are discussed. The lift increment is decreased with an increase in Mach number at transonic speeds, primarily because the flow over the wing without strake changes from leading edge separation to leading edge attached flow with shock-induced separation. An increase in the lift-drag ratio is due to the lift increase at low speeds and the drag decrease at supersonic speeds.

Arc wing aircraft

Abstract: A lightweight, man carrying aircraft including a skeletal fuselage assembly in combination with a primary arc-shaped lifting airfoil having variable camber, incidence angle, and pitching moment and a secondary stabilizing airfoil of a tubular, ring-like configuration. The arc-shaped primary airfoil is superior to traditional low speed airfoil forms due to its high aerodynamic efficiency, inherent design simplicity and strength. The ring-tail assembly, offering marginal aerodynamic lift, contributes significantly to the in-flight stability and safety of the aircraft. Both airfoil members are independently controllable and constructed of a fabric or like skin material which is fitted over a wing assembly and tensioned into an operable airfoil form by the aerodynamic forces of the air during flight. Typically, the wing assemblies are supported and maintained by a lightweight, tubular fuselage structure to which is affixed the propulsion system, control mechanisms and ground support assemblies.

Calculation of Laminar Separation Bubbles and Their u oc.- a f Effect on Airfoil Performance 2 & o o y

Abstract: The "bursting" of leading edge laminar separation bubbles that can occur on moderately thick airfoil sections is the cause of sudden large increases in drag, decrease of lift, and change in pitching moment associated with abrupt stall at high angles of attack. This paper describes a semiempirical theory for the description and calculation of the development and bursting of laminar separation bubbles. The semiempirical theory uses ia experimental correlation to relate the length of the laminar free shear layer to the freestream turbulence and a modified version of Hortons method to calculate the properties and reattachment position of the turbulent shear layer. The method is used to predict the development of laminar separation bubbles for NACA airfoils. Comparison of these results with NACA experiments shows good agreement.

The delta wing in oscillatory gusts

Abstract: This report describes systematic aerodynamic lift and pitching moment measurements on two sharp edged delta wings of aspect ratio 1 and 2 in oscillatory vertical gusts of varying frequency parameter and gust amplitude. The effects of the nonlinearity associated with the separated flow region are explored and the influence of the potential flow and the vortex induced lift on the measured oscillating forces are discussed with implications that may be of value for developing or validating theoretical models.

Low-speed aerodynamic characteristics of a 14-percent-thick NASA phase 2 supercritical airfoil designed for a lift coefficient of 0.7

Abstract: The low speed aerodynamic characteristics of a 14 percent thick supercritical airfoil are documented. The wind tunnel test was conducted in the Low Turbulence Pressure Tunnel. The effects of varying chord Reynolds number from 2,000,000 to 18,000,000 at a Mach number of 0.15 and the effects of varying Mach number from 0.10 to 0.32 at a Reynolds number of 6,000,000 are included.

Comparison of linear analysis with experiment for a single-track vehicle in a steady turn

Abstract: The work of Fu (1965) was extended to determine whether a linear analysis which includes the overturning and aligning moments of the pneumatic tire will yield improved accuracy in the prediction of the equilibrium turning behavior of a single-track vehicle Comparisons between theory and experiment showed that it is necessary to account for small changes in tire properties caused by the shift in tire normal loads resulting from aerodynamic drag, lift, and pitching moment, in order to make a reasonably accurate estimate of the understeer gradient prevailing at different speeds An examination of the reasons for the inability to make better predictions of the steering torque required to hold a cycle in a steady turn showed that this prediction requires very precise data defining the geometry and inertial properties of the front fork and wheel assembly as well as highly accurate data defining the mechanical properties of the pneumatic tire Measurement of intake manifold pressure showed, as suspected, that single-track vehicles, in contrast to double-track vehicles, do not require a significant increase in propulsive power in order to maintain speed on a turn of decreasing radius

Flutter Analysis of a Two-Dimensional and Two-Degree-of-Freedom Supercritical Airfoil in Small-Disturbance Unsteady Transonic Flow.

Abstract: : Flutter analysis is performed for a MBB A-3 supercritical airfoil at transonic Mach numbers. Two degrees of freedom, pitching and plunging, are considered. The unsteady aerodynamic data are obtained by using two separate transonic aerodynamic computational codes: (1) LTRAN2 based on the time integration method and (2) STRANS2 and UTRANS2 based on the harmonic analysis method. The steady aerodynamic results are shown in the form of upper and lower surface pressure curves. The unsteady aerodynamic coefficients are obtained and tabulated for various values of low reduced frequency by pitching the MBB A-3 airfoil about the quarter chord axis. For the case of zero mean angle of attack, the coefficients obtained by the two computer codes are directly compared and discussed. At design Mach number of 0.765 and zero mean angle of attack, flutter results are obtained by both methods and compared. They are presented as plots of flutter speed and the corresponding reduced frequency vs. one of the four parameters: airfoil - air mass ratio; position of mass center; position of elastic axis; and plunge-to-pitch frequency ratio.

Computational parametric study of the aerodynamics of spinning slender bodies at supersonic speeds

Abstract: : Three dimensional finite-difference flow field computation techniques have been employed to generate a parametric aerodynamic study at supersonic speeds. Computations for viscous turbulent and inviscid flow have been performed for cone-cylinder, secant-ogive-cylinder, and tangent-ogive-cylinder bodies for a Mach number range of 1.75 or = M or = 5. The aerodynamic coefficients computed are pitching moment, normal force, center of pressure, Magnus moment, Magnus force, Magnus center of pressure, form drag, viscous drag, roll damping and pitch damping. All aerodynamic coefficients are computed in a conceptually exact manner. The only empirical input is that required for turbulence modeling. Computed results are compared to experimental data from free flight aerodynamic ranges and wind tunnels in order to validate the computational techniques. parametric comparisons illustrate the effects of body configuration and Mach number for the ten aerodynamic coefficients. The results for Magnus and pitch damping are of particular interest.

Aerodynamic Analysis of a Fighter Aircraft with a Higher Order Paneling Method

Abstract: : This report presents results from analyses of steady and unsteady flows about several configurations involving the wing of the F-5 fighter airplane. The analyses were performed using higher order panel methods and the configurations analyzed included the following: the clean wing (both in an unbounded and in a wind tunnel wall bounded atmosphere), the wing with an external missile store mounted at the wing tip, and the wing with an external missile store mounted on a pylon at the lower surface of the wing. The flow mach number ranged from 0.6 to 1.35 in steady flow and from 0.6 to 0.95 in unsteady flow. Each steady flow case is analyzed at three angles of attack (0.5 deg, 0.0 deg, 0.5 deg) while each unsteady flow case consisted of unsteady pitch oscillation about zero angle of attack. The reduced frequency of the oscillation was in the range from 0.2498 to 0.3955. The computed results include chordwise pressure distributions along wing sections at eight spanwise locations from 18.1 to 97.7 per cent semispan. The results also include the coefficients of lift and pitching moment for each complete configuration as well as the coefficients of aerodynamic force and couple arising from the pressure on only the surface of the missile store.

Propulsive aerodynamics of an advanced nozzle/forward swept wing aircraft configuration

Abstract: Nonaxisymmetric exhaust nozzles can offer advanced aircraft configurations improved performance through the propulsive-aerodynamic interaction of the jet exhaust and the wing flowfield. An experimental investigation of this interaction was conducted at Mach numbers from 0.3 to 0.9 on a half-span wing-body configuration with vectoring (0, 15, 30 deg) nonaxisymmetric exhaust nozzles incorporated into the trailing edge of a forward swept wing. Results indicate that at all Mach numbers, the aerodynamic lift and drag coefficients increase while the pitching moment coefficient becomes more negative with increasing nozzle vectoring and nozzle flow. Thrust vectoring improves the aircraft drag polar especially at higher angles of attack.

Influence of Pitching Moment Characteristics on Departure and Uncoordinated Roll Reversal Boundaries for Fighter Configurations.

Abstract: : A study is reported which generated design charts and developed associated boundaries for identifying departure and uncoordinated roll-reversal flight characteristics as a function of aerodynamic rolling and yawing moment coefficients typical of fighter airplanes for various pitching moment characteristics. This information should be valuable for specification, design and evaluation purposes. The investigation utilized a large angle, six-degree-of-freedom digital computer program to simulate the motions of a fighter performing a severe air combat maneuver. The results showed that the developed uncoordinated roll reversal boundaries may be applied to any fighter configuration and that the developed departure boundaries are applicable to fighter configurations exhibiting static pitch stability. Unaugmented airframes having pitch instability impose more stringent requirements on the rolling and yawing moment coefficients to avoid departure susceptibility. A simple angle-of-attack feedback augmentation system can markedly improve the departure resistance of both stable and unstable airframes given sufficient control authority. (Author)

An analysis method for multi-component airfoils in separated flow

Abstract: The multi-component airfoil program (Langley-MCARF) for attached flow is modified to accept the free vortex sheet separation-flow model program (Analytical Methods, Inc.-CLMAX). The viscous effects are incorporated into the calculation by representing the boundary layer displacement thickness with an appropriate source distribution. The separation flow model incorporated into MCARF was applied to single component airfoils. Calculated pressure distributions for angles of attack up to the stall are in close agreement with experimental measurements. Even at higher angles of attack beyond the stall, correct trends of separation, decrease in lift coefficients, and increase in pitching moment coefficients are predicted.

Life and pitching moment characteristics of stores determined from flight drop tests

Abstract: An accurate estimation of the aerodynamic characteristics of externally carried stores, like bombs, drop tanks, missiles in the presence of the interference flowfield of the aircraft is essential for prediciton of reliable separation trajectories. The currently available theoretical methods are restricted to steady-state conditions of the flowfield around the model. Wind-tunnel methods such as the captive model, drop model and flow field survey techniques can be used to generate aerodynamic data for pre-flight simulation studies. But they are incapable of simulating the pronounced dynamic conditions of the stores encountered during the free fall which give rise to considerable aerodynamic forces and moments of noncirculator y origin. The present paper describes a simple parameter identification procedure employed to estimate the lift and pitching moment coefficients of a family of droppable fuel tanks using photographic data obtained from full-scale drop tests. Results are presented for the stores showing the combined effect of store incidence and pitch rate.

A dynamic analysis of the motion of a low-wing general aviation aircraft about its calculated equilibrium flat spin mode

Abstract: The properties of the flat spin mode of a general aviation configuration have been studied through analysis of rotary balance data, numerical simulation, and analytical study of the equilibrium state. The equilibrium state is predicted well from rotary balance data. The variations of yawing moment and pitching moment as functions of sideslip have been shown to be of great importance in obtaining accurate modeling. These dependencies are not presently available with sufficient accuracy from previous tests or theories. The stability of the flat spin mode has been examined extensively using numerical linearization, classical perturbation methods, and reduced order modeling. The stability exhibited by the time histories and the eigenvalue analyses is shown to be strongly dependent on certain static cross derivatives and more so on the dynamic derivatives. Explicit stability criteria are obtained from the reduced order models.

Prediction of nacelle aerodynamic interference effects at low supersonic Mach numbers

Abstract: The accuracy of analytical predictions of nacelle aerodynamic interference effects at low supersonic speeds are studied by means of test versus theory comparisons. Comparisons shown include: (1) isolated wing body lift, drag, and pitching moments; (2) isolated nacelle drag and pressure distributions; (3) nacelle interference shock wave patterns and pressure distributions on the wing lower surface; (4) nacelle interference effects on wing body lift, drag, and pitching moments; and (5) total installed nacelle interference effects on lift, drag, and pitching moment. The comparisons also illustrate effects of nacelle location, nacelle spillage, angle of attack, and Mach numbers on the aerodynamic interference. The initial results seem to indicate that the methods can satisfactorily predict lift, drag, pitching moment, and pressure distributions of installed engine nacelles at low supersonic Mach numbers with mass flow ratios from 0.7 to 1.0 for configurations typical of efficient supersonic cruise airplanes.

On the hydrodynamic forces for shallow draft ships in shallow water (3rd report). on the pitch hydrodynamic forces of a circular disk

Abstract: The hydrodynamic forces of pitch mode on a circular disk as a shallow draft ship in shallow water are investigated. The boundary value problem is formulated by the use of the concept of the surface distributed sources so that integral equations for the source densities are obtained. In the case of long waves, the problem is solved analytically. The numerical solution of the integral equation is found. The added moment of inertia, wave damping factor, wave exciting moment, radiation pressures, wave exciting pressures and motions are calculated. The corresponding experiments are carried out and the results of the numerical calculation are in good agreement with those of the experiments. (Author)

Unsteady Pressure Distributions on Airfoils in Cascade.

Abstract: : The results of a study of unsteady pressure distributions in a two-dimensional cascade of blades caused by spatial inflow velocity variations are presented. An existing incompressible, inviscid theory which employs a simplified vortex model in conjunction with the assumptions of thin airfoil theory has been used by Henderson and Bruce to derive expressions for the unsteady response, which includes the cascade unsteady lift and pitching moment. An alternative way to obtain these unsteady response parameters is to establish the expression for the unsteady pressure distribution. The unsteady lift and pitching moment are calculated by direct numerical integration over the unsteady pressure difference across the airfoil chord. Comparison of the computed theoretical results using these two approaches shows satisfactory agreement except when the wavelength of the velocity variations approaches the cascade blade spacing. Good agreement is also observed between the existing measured and predicted data. The effects of design parameters of a cascade, such as space-chord ratio, maximum blade camber, and mean incidence angle, on the unsteady response are presented and discussed. (Author)

Large-scale wind-tunnel tests of inverting flaps on a STOL utility aircraft model

Abstract: A unique inverting flap system was investigated on a large scale deflected slipstream model in the Ames 40 by 80 foot wind tunnel. The subject tests utilized 33% chord double-slotted flaps on a low aspect ratio wing that was fully immersed in the propeller slipstream. Evaluation of the flap effectiveness is aided by comparisons with the results of tests of other flap systems on the same twin propeller, twin tail boom STOL utility aircraft mode. No extreme or abrupt force or moment increments were encountered when the flaps were deflected through a wide range, corresponding to the complete retraction/extension spectrum. The lift and descent capability of the inverting flaps compared very favorably with that of the other flap systems that have been tested on this model, including some with much greater mechanical complexity. As expected, the flaps caused large nose down, pitching moment increments at the high lift settings; however, the trimmed characteristics are still competitive with those obtained from the more complicated flap systems. It is believed that these flaps may have promising potential application to the design of relatively simple STOL utility aircraft with improved performance capabilities. In addition, they may merit consideration as retrofits to existing aircraft with less effective flap systems.