Showing papers on "Pitching moment published in 1992"

Multipoint inverse airfoil design method based on conformal mapping

Abstract: An exact method of multipoint inverse airfoil design for incompressible flow is presented. Multipoint design is handled by dividing the airfoil into a number of desired segments. For each segment the velocity distribution is prescribed together with an angle of attack at which the prescribed velocity distribution is to be achieved. In this manner, multipoint design objectives can be taken into account in the initial specification of the velocity distribution. In order for the multipoint inverse airfoil design problem to be well posed, three integral constraints and several conditions arise which must be satisfied. Further restrictions are imposed if the airfoil is to have a specified pitching moment, thickness ratio, or other constraints. The system of equations is solved partly as a linear system and partly through multidimensional Newton iteration. Since the velocity distribution is prescribed about the circle in the angular coordinate, specification of the velocity in terms of arc length is handled through the multidimensional Newton iteration as well. The current formulation sets the stage for a more general multipoint inverse airfoil design method in which it will be possible to specify the velocity distribution, some boundary-layer development, or the surface geometry along a segment.

Unsteady flow past an airfoil pitching at a constant rate

Abstract: The unsteady flow past a NACA 0012 airfoil that is undertaking a constant-rate pitching up motion is investigated experimentally by the PIDV technique in a water towing tank. The Reynolds number is 5000, based upon the airfoil's chord and the free-stream velocity. The airfoil is pitching impulsively from 0 to 30 deg. with a dimensionless pitch rate alpha of 0.131. Instantaneous velocity and associated vorticity data have been acquired over the entire flow field. The primary vortex dominates the flow behavior after it separates from the leading edge of the airfoil. Complete stall emerges after this vortex detaches from the airfoil and triggers the shedding of a counter-rotating vortex near the trailing edge. A parallel computational study using the discrete vortex, random walk approximation has also been conducted. In general, the computational results agree very well with the experiment.

Generalized multipoint inverse airfoil design

Abstract: In a rather general sense, inverse airfoil design can be taken to mean the problem of specifying a desired set of airfoil characteristics, such as the airfoil maximum thickness ratio, pitching moment, part of the velocity distribution, or boundary-layer development. From thie information, the corresponding airfoil shape is determined. We present a method that approaches the design problem from this perspective. In particular, the airfoil is divided into segments along which, together with the design conditions, either the velocity distribution or boundary-layer development may be prescribed

Calculation of real-gas effects on blunt-body trim angles

Abstract: The effect of vibrational excitation and dissociation at high temperatures on the trim angle of attack of a blunt lifting body is calculated for a nonequilibrium flow regime in air using a CFD technique. The vibrational-electronic temperature and the species densities are calculated assuming the flow to be in a nonequilibrium state. The forebody flow of a two-dimensional blunt body of the shape of the Apollo Command Module at a finite angle of attack is calculated. The results show that the pitching moment around a reference point is larger and the trim angle of attack is smaller for a reacting gas than for a perfect gas. The calculated shift in the trim angle due to the real-gas effect is of the same order as that seen during the Apollo flights.

Experimental Study of a Low Reynolds Number Tandem Airfoil Configuration

Abstract: Experiments were concluded with two identical Wortmann FX63-137 airfoils in closely coupled tandem configurations at a Reynolds number of 8.5 x 10 4. For the data presented here, the values of the stagger and gap were 1.5 and 0, respectively. The decalage angles were 0 and ±10 deg. Direct measurement of lift, drag, and 1/4-chord pitching moment, as well as static pressure distributions, were acquired for each airfoil. Flow visualization using kerosene smoke was performed to complement the experimental data. The total drag reduction and lift increase resulted in a significant increase in the lift-to-drag ratio for a number of configurations. Nomenclature Aw = wing area Cd = section drag coefficient, D/(Awqx) Ci = section lift coefficient, L/(AwqJ) Cm = section 1/4-chord moment coefficient, Cp = pressure coefficient, l-q/qx c = chord length D = drag force G = gap, \y\lc L = lift force M = 1/4-chord pitching moment q = dynamic pressure, \pU2 Rc = Reynolds number based on chord, Uxc/v St = stagger, x/c (positive when upstream airfoil is above the downstream airfoil) U = flow velocity x = distance in streamwise direction y - distance normal to streamwise directions a = angle of attack 8 = decalage, aua - ada

Ultra-soft, ultra-elastic airfoils

Abstract: A novel rotating aerodynamic toy comprising a ultra-elastic gel airfoil which is suitable for launch in light or heavy wind conditions and capable of performing various aerodynamic effects including climb, stall, return, straight-line flight and other aerobatics. The ultra-elastic properties of the airfoil allow the airfoil to transform its aerodynamic profile at launch and while in flight.

Fourier functional analysis for unsteady aerodynamic modeling

Abstract: A method based on Fourier analysis is developed to analyze the force and moment data obtained in large amplitude forced oscillation tests at high angles of attack. The aerodynamic models for normal force, lift, drag, and pitching moment coefficients are built up from a set of aerodynamic responses to harmonic motions at different frequencies. Based on the aerodynamic models of harmonic data, the indicial responses are formed. The final expressions for the models involve time integrals of the indicial type advocated by Tobak and Schiff. Results from linear two- and three-dimensional unsteady aerodynamic theories as well as test data for a 70-degree delta wing are used to verify the models. It is shown that the present modeling method is accurate in producing the aerodynamic responses to harmonic motions and the ramp type motions. The model also produces correct trend for a 70-degree delta wing in harmonic motion with different mean angles-of-attack. However, the current model cannot be used to extrapolate data to higher angles-of-attack than that of the harmonic motions which form the aerodynamic model. For linear ramp motions, a special method is used to calculate the corresponding frequency and phase angle at a given time. The calculated results from modeling show a higher lift peak for linear ramp motion than for harmonic ramp motion. The current model also shows reasonably good results for the lift responses at different angles of attack.

Transonic aeroelasticity analysis using state-space unsteady aerodynamic modeling

Abstract: An aeroelastic analysis is conducted on a two-degree-of-freedom airfoil in transonic flow using a generalized state-space approximation for the unsteady aerodynamics. The aerodynamic representation is validated against computational fluid dynamic solutions for angle of attack oscillations up to Mach numbers of 0.875 and at reduced frequencies up to 1.0. Despite the inherent nonlinear nature of transonic flow, it is shown that a linear finite-state model with as few as eight states can provide a good approximation to the unsteady lift and moment behavior if appropriate allowance is made for Mach number effects on the airfoil's static lift curve slope and mean aerodynamic center. It is shown how the aerodynamic representation can be coupled to the structural equations of a typical airfoil section with bending and torsional degrees of freedom. The stability of the resulting aeroelastic system is determined by eigenanalysis. This aeroelastic analysis is shown to be in excellent agreement with calculations performed using more sophisticated unsteady aerodynamic theories.

Navier-Stokes methods to predict circulation control airfoil performance

Abstract: : The predictive capability of the two-dimensional compressible mass- averaged Navier-Stokes equations was investigated for a typical circulation control air-foil. The governing equations were solved using the implicit approximate factorization algorithm of Beam-Warming with the turbulence model of Baldwin-Lomax. To account for the unique characteristics of circulation control airfoils, an empirical turbulence model correction due to Bradshaw was used. This thesis is unique in that the predictive capability of the computational method is explored by examining the importance of the empirical Bradshaw curvature correction constant on the computed results. Using a generic value of the curvature constant at various blowing coefficient levels, the computational method was able to accurately predict airfoil pitching moment and lift curve slope due to blowing. Predicted levels of airfoil lift coefficient, although reasonable, were found to be consistently low compared with experiment due to the generic curvature constant providing premature jet detachment from the Coanda surface. Computed and measured airfoil drag results followed the same trends, but lack of overall drag coefficient agreement was disappointing. Lift coefficient was found to be quite sensitive, pitching moment not sensitive, and drag coefficient moderately sensitive to the value of the curvature constant used. For the highest blowing coefficient case considered, the value of curvature constant required for the computational lift coefficient to match the experimental lift coefficient was also determined.

Rotary-wing blade of rotary-wing aircraft

Abstract: A rotary-wing blade of a rotary-wing aircraft has an airfoil section which has an asymmetrical region of positive camber from its leading edge to its approximately 30-percent chord length point and an essentially symmetrical airfoil region from its 30-percent chord point to approximately its 90-percent chord length point. The region aft of the 90-percent chord point can be optionally curved or reflexed upward or downward to a limited degree. This blade has a large lift coefficient and a small pitching moment and is readily adaptable to meet various design conditions of lift coefficient and pitching moment.

Computational investigation of predicted store loads in mutual interference flow fields

Abstract: Steady-state numerical solutions of the Euler equations for the flow field about a wing/pylon/finned store configuration at a Mach number of 0.95 have been obtained for several store locations and attitudes. The objectives of the study were to gain insight into requirements for future computational trajectory prediction methods, to compare computationally predicted loads and pressures to measured data, and to investigate a mutual interference correction to a semi-empirical load prediction method. To meet these objectives, computational fluid dynamics (CFD) solutions were used to predict loads with the store placed at its carriage position and at 0.25, 0.50, 1 .O, and 4.0 store body diameters below the carriage position. An additional solution was obtained with the store in a position determined by a wind tunnel trajectory simulation test. Load predictions were also obtained using the Influence Function Method (IFM) for these positions. The CFD-predicted pressure distributions for the store at the carriage and trajectory positions agreed well with the measured test data, and the CFD-predicted loads on the store in these two cases agreed fairly well with the test loads. Conversely, the loads predicted by the basic IFM were in poor agreement for all cases with the measured loads and loads predicted by the CFD calculations. However, when properly applied to the IFM, the mutual interference loads correction provided a reasonable approximation to the CFD-predicted loads. Over the course of this investigation, it was found that grid density, geometric accuracy, and viscosity requirements for CFD trajectory predictions are extremely dependent on the physical properties of the store of interest, the method of release, and the portion of the trajectory over which the calculations will be performed. Nomenclature Influence coefficients used by the Influence Function Method to calculate normal and side forces. Influence coefficients used by the Influence Function Method to calculate yawingand pitching-moment coefficients. Rolling-moment coefficient, (rolling moment)/(Q*SD). Positive in the positive @ sense (see Fig. 1 ). Pitching-moment coefficient taken about the c.g., (pitching moment)/(Q*S*D). Positive in the positive 8 sense (see Fig.1). Normal-force coefficient, (normal force)/ (QnS). Positive in the positive Z direction (see Fig. 1 ). Yawing-moment coefficient taken about the c.g., (yawing moment)/(Q*S*D). Positive in the positive 6 sense (see Fig. 1). Pressure coefficient. Side force coefficient, (side force)/(Q*S). Positive in the positive Y direction (see Fig. 1). Local chord of fin on model store. Store center of gravity located 2.79 in. (model scale) axially from the store nose. Store model diameter, 1 .O .in. Distance over which the double delta correction is applied to the IFM-predicted loads, store diameters. Store model length, 5.941 in. Free-stream Mach number. Free-stream dynamic pressure, 1/2pV2. T h e research reported herem was performed by the Arnold Engineering Development Center (AEDC), A r Force Systems Command Work and analysis for thts research were done by personnel of Calspan Corportat~onJAEDC Operations, operatmg contractor for the AEDC aerospace fl~ght dynam~cs fac l~t~es Further reproduction 1s author~zed to satlsfy needs of the U S Government

Effects of trailing-edge flap on buffet characteristics of a supercritical airfoil

Abstract: The buffet characteristics of a 16% thickness-to-chord ratio supercritical airfoil were investigated in the Institute for Aerospace Research (IAR) High Reynolds Number Two-Dimensional Test Facility. The trailingedge flap dimension was 13.5% chord and it was deflected at various angles to study the effect of modifying the downstream pressure on controlling flow separation over the airfoil. The unsteady normal force was measured and the buffet boundary was determined from the divergence of the fluctuating normal force. The investigation was conducted quite deep into the buffet regime. Spectral analyses of the normal force were carried out and the frequencies of shock-wave oscillations were measured. They were found to be Mach-number dependent and varied between 50-80 Hz for M = 0.612 to 0.792. The effects of varying the flap angles on the shock-wave position and drag of the airfoil were also investigated. Results for an off-design Mach number of 0.612 were given in some detail.

Multigrid Euler calculations over complete aircraft

Abstract: A three-dimensional Euler code is applied to a series of configurations of increasing complexity. Comparisons are made to experiments, or to other computations when the former is not available. The method uses the multigrid approach on sets of equally spaced Cartesian grid cells. A unique and robust implementation of the body surface boundary condition on grid cells not aligned with the surface provides accurate results on relatively coarse grids. All computational results are compared to experimental data with good agreement exhibited over a wide range of flight conditions. The solution accuracy is assessed for the Onera M-6 wing, with errors due to grid resolution of less than 1% being achieved for the lift and pitching moment coefficients.

Research in robust control for hypersonic vehicles

Abstract: During the first reporting period research concentrated on finishing the modeling work required for a representative model of a scramjet propulsion system for hypersonic vehicles. An existing hypersonic propulsion code was adjusted to the winged-cone configuration. In this process the complete force and moment calculation was revised. The advantageous feature of the code to account for angle of attack variations was then used to compute the thrust, lift, and pitching moment contributions of the propulsion system not only for various Mach numbers and fuel equivalence ratios, but also for different angles of attack.

Helicopter rotor blade with improved performance characteristics

Abstract: In a helicopter rotor blade having a root end for attachment to a rotor head, a central portion of constant chord dimension and a tip portion having a chord dimension greater than that of the central portion, an outboard region of the central portion has a negative basic pitching moment coefficient, an inboard region of the central portion has a basic pitching moment coefficient more positive than that of the outboard region and the tip portion has a positive basic pitching moment coefficient.

New Methods for Predicting Nonlinear Lift, Center of Pressure, and Pitching Moment on Missile Configurations.

Abstract: : A new semiempirical method has been developed to predict normal force, pitching moment, and center of pressure on missile configurations up to angles of attack of 30 deg . The new method is based on linear theory and slender body techniques at low angle of attack and uses wind tunnel data to derive nonlinear angle-of-attack corrections as angle of attack increases. The new improved theories include body alone, wing alone, and body-wing and wing-body interference. While the new theory is databased, simple analytical formulas are derived that allow general use of the techniques. Comparison of the new theory to the linearized approaches used in the former NSWCDD aeroprediction code shows significant reductions in errors of aerodynamics dynamics above about 5 deg to 10 deg angle of attack. Limited comparisons of the new theory to other state-of-the-art engineering codes show the new theory to be as good as or better than anything known to be available computing planar aerodynamics up to 30 deg angle of attack. nonlinear lift, pitching moment, center of pressure. slender body theory, linear theory.

Low-speed longitudinal and lateral-directional aerodynamic characteristics of the X-31 configuration

Abstract: An experimental investigation of a 19 pct. scale model of the X-31 configuration was completed in the Langley 14 x 22 Foot Subsonic Tunnel. This study was performed to determine the static low speed aerodynamic characteristics of the basic configuration over a large range of angle of attack and sideslip and to study the effects of strakes, leading-edge extensions (wing-body strakes), nose booms, speed-brake deployment, and inlet configurations. The ultimate purpose was to optimize the configuration for high angle of attack and maneuvering-flight conditions. The model was tested at angles of attack from -5 to 67 deg and at sideslip angles from -16 to 16 deg for speeds up to 190 knots (dynamic pressure of 120 psf).

Local bridging to predict aerodynamic coefficients in hypersonic, rarefied flow

Abstract: A computational method is given for the prediction of local pressure and viscous shear stress on windward surfaces of convex, axisymmetric or quasi-axisymmetric, hypersonic bodies in the transitional, rarefied flow regime. Overall aerodynamic forces and moments are computed by integration of the local quantities. The method is based on a correlation of local pressure and shear stress computed by the direct simulation Monte Carlo (DSMC) numerical technique for cold-wall, real-gas conditions and some supplemental data from low-density, hypersonic wind tunnels. Two-dimensional shapes and leeward surfaces are not included in the scope of the method as it is presented here. Results are compared with DSMC and viscous shock layer computations for both local and overall coefficients. Also included are sphere and blunt cone drag obtained from both computation and experiment, as well as lift and pitching moment coefficients for the NASA Aeroassist Flight Experiment vehicle at various angles of attack. It appears that the method presented here will prove adequate for relatively bluff bodies approximately defined by wetted length-to-nose radius ratios of less than 10.

Finite wing aerodynamics with simulated glaze ice

Abstract: The effect of a simulated glaze ice accretion on the aerodynamic performance of a three-dimensional wing is studied experimentally. The model used for these tests was a semi-span wing of effective aspect ratio five, mounted from the sidewall of the UIUC subsonic wind tunnel. The model has an NACA 0012 airfoil section on a rectangular, untwisted planform with interchangeable leading edges to allow for testing both the baseline and the iced wing geometry. A three-component sidewall balance was used to measure lift, drag and pitching moment on the clean and iced model. A four-beam two-color fiberoptic laser Doppler velocimeter (LDV) was used to map the flowfield along several spanwise cuts on the model. Preliminary results from LDV scans, which will be the bulk of this paper, are presented following the force balance measurement results. Initial comparison of LDV surveys compare favorably with inviscid theory results and 2D split hot-film measurements near the model surface.

The Effect of Yaw Cards on the Pitching and Yawing Motion of Symmetric Projectiles

Abstract: : This report presents an improved method for analysis of the effect of yaw cards on the determination of the pitching moment coefficient of symmetric missiles. Several comparisons of the present method axe made with spark photography range results obtained for the same projectiles, and a proper treatment of yaw-card data is shown to significantly improve the agreement between the two techniques.

Lift augmentation on a delta wing via leading edge fences and the Gurney flap

Abstract: Wind tunnel tests were conducted on two devices for the purpose of lift augmentation on a 60 deg delta wing at low speed. Lift, drag, pitching moment, and surface pressures were measured. Detailed flow visualization was also obtained. Both the leading edge fence and the Gurney flap are shown to increase lift. The fences and flap shift the lift curve as much as 5 deg and 10 deg, respectively. The fences aid in trapping vortices on the upper surface, thereby increasing suction. The Gurney flap improves circulation at the trailing edge. The individual influences of both devices are roughly additive, creating high lift gain. However, the lower lift to drag ratio and the precipitation of vortex burst caused by the fences, and the nose down pitching moment created by the flap are also significant factors.

The Effect of Vortex Generators on Crossflow Separation on a Submarine in a Turning Maneuver

Abstract: : The effect of vortex generators on the crossflow separation of a 688 class submarine in a turning maneuver was studied. The vortex generators are located on the top and bottom centerline of the submarine. The intent of the vortex generators is to improve turning performance by changing the inherent hydrodynamic forces incurred from crossflow separation in such a maneuver. The number, size, axial location, and local angle of attack of the vortex generators are studied. Oil flow visualization was used as the primary diagnostic in determining the effectiveness of various vortex generator configurations. Force and moment measurements were taken to relate the fluid dynamics to changes in force and moment distribution. The vortex generators were found to be very effective in delaying crossflow separation. Separation was delayed by as much as 35 on the bottom of the submarine model. With the addition of vortex generators, the normal force was reduced up to 33%, the axial force was increased up to 233%, and the yaw moment was increased up to 150%. Also, the vertical force increased up to 150%, the pitching moment increased up to 40%, and the roll moment was unaffected.

Comment on 'Effect of thrust/speed dependence on long-period dynamics in supersonic flight'

Abstract: Introduction R EFERENCE 1 discusses effects of speed-dependent variations in thrust on the long-period motions of aircraft in supersonic flight. Its principal contribution is the demonstration of effects that a complex pair of zeros have on speed-tothrottle feedback. It is concluded that there are fundamental differences between thrust/speed effects in subsonic and supersonic flight. It also is stated that "thrust variations with speed (or Mach number) have practically no effect on the phugoid in supersonic flight," then that "the phugoid is not influenced at all in supersonic flight." The purpose of this Comment is to show that thrust/speed effects are indeed small but not zero in supersonic flight and that this is a kinematic effect of increasing speed rather than an aerodynamic effect of Mach number. In the process, a simplified analytical model is offered for further study. Furthermore, it is noted that pitching moment/speed sensitivity due to thrust-axis offset is likely to have important long-period effect even when the direct thrust/speed effects are small.

Wind tunnel investigation and analysis of the sm701 airfoil

Abstract: A wind tururel test was performed on a two-dimensional model of the SM701 airfoil designed for use on World Class gliders. The test covered a range of Reynolds Number conditions from one million to 2.5 million. Aerodynamic forces and moments were measured with an extemal balance. Wake-rake measurements of the two dimensional drag were also made. Flow visualization techniques provided information on transition from laminar to turbulent flow. Post stall conditions were examined for both positive and negative angles of attack. Llft, drag, and pitching moment were analyzed and comparisons made with numerical predictions. The model was designed, constructed, and the test conducted by students at Texas A&M University.

Comparison of pitch rate history effects on dynamic stall

Abstract: Dynamic stall of an airfoil is a classic case of forced unsteady separated flow. Flow separation is brought about by large incidences introduced by the large amplitude unsteady pitching motion of an airfoil. One of the parameters that affects the dynamic stall process is the history of the unsteady motion. In addition, the problem is complicated by the effects of compressibility that rapidly appear over the airfoil even at low Mach numbers at moderately high angles of attack. Consequently, it is of interest to know the effects of pitch rate history on the dynamic stall process. This abstract compares the results of a flow visualization study of the problem with two different pitch rate histories, namely, oscillating airfoil motion and a linear change in the angle of attack due to a transient pitching motion.

Study of the Leading-Edge Vortex Dynamics in the Unsteady Flow Over an Airfoil

Abstract: : The two-year project to study the dynamics of the leading-edge vortex (LEV) over a pitching airfoil under conditions of dynamic stall, was started in January 1990. Several accomplishments have been made during these two years. The most significant of these are (1) the construction of a special water channel suitable for the study of dynamic stall over a pitching airfoil, (2) the measurement of surface pressure distributions over the airfoil under several key operating conditions, and (3) development of the techniques of Particle Image Velocimetry (PIV) and its use in the measurement of instantaneous velocity and vorticity field in the two-dimensional flow around the airfoil. Some of these data which are the first of their kind have been used to understand the physics of unsteady vorticity dynamics over a pitching airfoil. These data are being made available to other investigators for use as database in validating their numerical models. Unsteady Aerodynamics, Dynamic Stall, Supermaneuverability, Vortex Dynamics.

Generalized multipoint inverse arifoil design

Abstract: In a rather general sense, inverse airfoil design can be taken to mean the problem of specifying a desired set of airfoil characteristics, such as the airfoil maximum thickness ratio, pitching moment, part of the velocity distribution, or boundary-layer development. From thie information, the corresponding airfoil shape is determined. We present a method that approaches the design problem from this perspective. In particular, the airfoil is divided into segments along which, together with the design conditions, either the velocity distribution or boundary-layer development may be prescribed