Showing papers on "Pitching moment published in 1972"

Aerodynamic effects of leading-edge serrations on a two-dimensional airfoil

Abstract: An investigation was conducted to determine the flow field and aerodynamic effects of leading-edge serrations on a two-dimensional airfoil at a Mach number of 0.13. The model was a NACA 66-012 airfoil section with a 0.76 m (30 in.) chord, 1.02 m (40 in.) span, and floor and end plates. It was mounted in the Ames 7- by 10-Foot Wind Tunnel. Serrated brass strips of various sizes and shapes were attached to the model in the region of the leading edge. Force and moment data, and photographs of tuft patterns and of oil flow patterns are presented. Results indicated that the smaller serrations, when properly placed on the airfoil, created vortices that increased maximum lift and angle of attack for maximum lift. The drag of the airfoil was not increased by these serrations at airfoil angles of attack near zero and was decreased at large angles of attack. Important parameters were serration size, position on the airfoil, and spacing between serrations.

Investigation of Airfoil Dynamic Stall and Its Influence on Helicopter Control Loads

Abstract: : Measurements were made of the unsteady normal force and pitching moment on an NACA 0012 airfoil model oscillated both sinusoidally and nonsinusoidally over a range of incidence angles, including a substantial penetration into stall. The sinusoidal normal force and pitching moment data were reduced and tabulated as functions of the angle of attack, the angular velocity parameter, and the angular acceleration parameter. This generalized form of the data was used to reconstruct the measured sinusoidal aerodynamic response of the model airfoil with excellent results. Additional correlations were made using nonsinusoidal pitch schedules which included periodic ramp changes in angle of attack and a flexured angular blade response to a one-per- rev sinusoidal incidence angle change typical of that for a helicopter blade. The agreement between predicted and measured normal force and moment loops was very good for the ramp motion.

Numerical method for design of minimum-drag supersonic wing camber with constraints on pitching moment and surface deformation

Abstract: A numerical method, based on linearized theory, for designing minimum-drag supersonic wing camber surfaces of arbitrary planform for a given lift, with options for constraining the pitching moment and/or the surface deformation at the trailing edge of the root chord and for selecting any desired combination of eight specified wing-loading distributions to be employed in the optimization procedure is presented. Two examples are given to illustrate applications of the method. The results indicate that relatively small drag penalties are incurred in designing wings to be self-trimming and to have a reasonable camber surface.

Low-speed aerodynamic characteristics of an airfoil optimized for maximum lift coefficient

Abstract: An investigation has been conducted in the Langley low-turbulence pressure tunnel to determine the two-dimensional characteristics of an airfoil optimized for maximum lift coefficient. The design maximum lift coefficient was 2.1 at a Reynolds number of 9.7 million. The airfoil with a smooth surface and with surface roughness was tested at angles of attack from 6 deg to 26 deg, Reynolds numbers (based on airfoil chord) from 2.0 million to 12.9 million, and Mach numbers from 0.10 to 0.35. The experimental results are compared with values predicted by theory. The experimental pressure distributions observed at angles of attack up to at least 12 deg were similar to the theoretical values except for a slight increase in the experimental upper-surface pressure coefficients forward of 26 percent chord and a more severe gradient just behind the minimum-pressure-coefficient location. The maximum lift coefficients were measured with the model surface smooth and, depending on test conditions, varied from 1.5 to 1.6 whereas the design value was 2.1.

A study of the variable impedance surface concept as a means for reducing noise from jet interaction with deployed lift-augmenting flaps

Abstract: The feasibility of quieting the externally-blown-flap (EBF) noise sources which are due to interaction of jet exhaust flow with deployed flaps was demonstrated on a 1/15-scale 3-flap EBF model. Sound field characteristics were measured and noise reduction fundamentals were reviewed in terms of source models. Test of the 1/15-scale model showed broadband noise reductions of up to 20 dB resulting from combination of variable impedance flap treatment and mesh grids placed in the jet flow upstream of the flaps. Steady-state lift, drag, and pitching moment were measured with and without noise reduction treatment.

The Static and Magnus Aerodynamic Characteristics of the M823 Research Store Equipped with Fixed and Freely Spinning Stabilizers

Abstract: : The report is the first in a two-part series of technical reports on the dynamics and aerodynamics of free-fall stores using freely spinning stabilizers Presented are the results of recent wind-tunnel measurements of the normal force, pitching moment and Magnus force and moment on the M823 Research Store in transonic flow Comparisons are made between configurations equipped with fixed and freely spinning stabilizers, with regard to sign and relative magnitude of the Magnus force and moment A method is presented and applied whereby the Magnus force and moment are corrected for flow angularity

Wind-tunnel measurements of aerodynamic load distribution on an NASA supercritical-wing research airplane configuration

Abstract: Wind tunnel tests have been conducted on a research airplane model with an NASA supercritical wing to define the general character of the flow over the wing and to aid in structural design of the full scale airplane. Pressure measurements were made at Mach numbers from 0.25 to 1.30 for sideslip angles from -2.50 deg to 2.50 deg over a moderate range of angles of attack and dynamic pressures. Except for representative figures, the results are presented in tabular form without detailed analysis.

Development of Theoretical Method for Two-Dimensional Multi-Element Airfoil Analysis and Design. Part I: Viscous-Flow Analysis Method.

Abstract: : A method has been developed for the analysis of arbitrary multi-element airfoils in viscous flow The viscous solution is obtained through an inviscid analysis of an equivalent system defined from viscous considerations An iterative procedure has been formulated to implement this analysis The inviscid solution is obtained through a distributed singularity method A finite-difference method is used to determine boundary-layer characteristics Methods are included to predict laminar-flow bubbles and separation and transition points The equivalent airfoil is defined for airfoils with attached flow as well as for airfoils with flow separation The validity of the method is established through comparison of the predicted results with experimental data for several single- and multi-element airfoils The comparisons show good agreement for lift coefficient and maximum lift coefficient and fair agreement for drag and pitching moment coefficients Details of the computer program developed to implement this method are described, including input and output details, FORTRAN source deck listing, and a sample problem (Author)

An Experimental Study of Certain Aerodynamic Effects on Erosion

Abstract: : A preliminary set of data was obtained from a unique erosion tester, designed such that the aerodynamics over the specimen are an integral part of the test parameters. The data obtained from this test system supports a hypothesis that a pitching moment on the longer sand particles could rotate them in a manner that could influence the erosion rate. It was concluded that more work is necessary in order to substantiate the above hypothesis. (Author)

Evaluation of Techniques for Predicting Static Aeroelastic Effects on Flexible Aircraft

Abstract: An experimental evaluation of analytical techniques for predicting the longitudinal stability characteristics of a large flexible aircraft is presented. Analytical methods based on both the modal approach and stiffness influence coefficients are used to predict the aerodynamic characteristics of a flexible airplane. These methods are then applied to a flexibly scaled model of a supersonic transport configuration. Comparisons between wind-tunnel data, the modal approach, and calculations based on stiffness influence coefficients are presented over the Mach number range from M = 0.6-2.7. The results of this study show good agreement for most cases between analyses and experiment. Nomenclature a.c. = aerodynamic center position Aa.c. = shift in aerodynamic center position due to flexibility, positive forward b — wing span c = local chord measured streamwise c = reference chord CL = lift coefficient, lift/gS CLx = lift-curve slope, 8CL/dot at a = 0° CL | a =o = lift coefficient at a = 0° CLq = lift coefficient due to pitching, 8CL/d(9c/2V) CL5e = elevator effectiveness in lift, dCL/88e

Effect of Lagging Pitching Moment on Re-Entry Vehicle Dynamic Stability

Abstract: Recent work has shown that the effect of ablation on re-entry vehicle (R/V) dynamic stability may be significant. Waterfall has explained flight test anomalies by postulating a dynamic model in terms of an ablative pitching moment derivative Cmfla and associated time lag r and by assuming that the resulting behavior was representable in the classical two-frequency form. The problem has been treated from a similar point of view by Ericsson. The present work is concerned with establishing the conditions under which the two-frequency assumption is valid and examing the effects of large phase lags. Exact solutions of the difference-differential equation which results from the inclusion of the term with retarded argument have been generated for the special case of constant dynamic pressure for a range of pertinent aerodynamic parameters. A significant result is that even for large time lags (T ~ 1 sec) the solution retains its classical two-component form. The solutions are compared to the approximate technique of Waterfall and a convenient means of generating the solutions is presented.

The effects of configuration changes on the aerodynamic characteristics of a full-scale mockup of a light twin engine airplane

Abstract: Wind tunnel tests of a full-scale model of a light twin engine aircraft were conducted The angle of attack was varied from minus 4 degrees to plus 20 degrees The sideslip range was plus or minus 8 degrees Thrust coefficients were 0, 020, and 044 Tests were made with various nacelle configurations, modes of propeller rotation, orientation of the thrust axis, and airfoil section at Reynolds numbers of 296 times one million amd 205 times one million

A four-level technique for estimation of tactical missile aerodynamic parameters

Abstract: A four-level technique is described for estimation of aerodynamic moment coefficient M α , normal acceleration force coefficient Z α , and velocity V . These parameters are associated with a simplified model of pitch plane dynamics for a tactical missile. Contrary to past studies where aerodynamic parameters were estimated from measurements which are very difficult to obtain (e.g., angle of attack), it is shown herein how estimates of the above parameters can be realized from measurements which are more easily obtained. Specifically, parameters are estimated from measurements of rate, normal acceleration, and gimbal angle deflection.

Nacelle-airframe interference at low supersonic Mach numbers.

Abstract: The aerodynamic interference between the propulsion system and airframe for a low supersonic transport with wing-mounted nacelles is examined. Both a flowfield analysis and the equivalent body approach were used to predict the interference lift, drag, and pitching moment as functions of nacelle size, shape, and position. The results indicate that the interference lift and pitching moment, as well as drag, must be included in the analysis to properly assess the interference effects. In addition, the performance of the basic wing was found to play an important role in determining the effectiveness of the interference lift in reducing the net installation drag. Based on a conservative prediction, the interference effects can reduce the installed propulsion system drag to 40% of the isolated drag of the nacelles. Furthermore, including the interference effects in the optimization of the engine cycle from a thermodynamic and weight standpoint can result in a considerable reduction in the net propulsion system weight fraction (fuel plus engines) while increasing the optimum engine bypass ratio of a typical transport vehicle.

Aerodynamic characteristics of 2 NASA supercritical airfoils with different maximum thickness

Abstract: Transonic wind tunnel tests were conducted at Mach numbers from 0.60 to 0.81 to determine the aerodynamic characteristics of two NASA supercritical airfoils. The airfoils had maximum thicknesses of 10 and 11 percent of the chord. Normal forces and pitching moments acting on the airfoils were determined from surface static pressure measurements. Drag forces acting on the airfoils were derived from vertical variations of the total and static pressures measured across the wake. For the thinner airfoil, stall begins at approximately 0.1 higher normal-force coefficient at the higher test Mach numbers, and the drag divergence Mach number at a normal-force coefficient of 0.7 was 0.01 higher.

Blunt Cones at Angle of Attack in Supersonic Nonuniform Freestreams

Abstract: A = velocity defect parameter [see Eq. (1)] Ab = local cone cross-sectional area B = wake width parameter [see Eq. (1)] CD = drag coefficient, drag/qnAb CL = lift coefficient, \ift/qnAb CMQ = nose pitching moment coefficient, nose momQnt/qnAbL (positive nose up) CLa = lift coefficient slope, dCJdx CM* = pitching moment coefficient slope, dCMQldoL Cp = pressure coefficient, (p — pn)lqn L = reference length equal to nose radius qn = dynamic pressure based on wake centerline conditions, ' PnVn/2 S' = nondimensional surface distance measured from the apex of a sharp cone, S/L § = nondimensional surface distance measured from the forward stagnation point on a blunt cone, S/L

Effects of wing-pivot location and forewing configuration on the low-speed aerodynamic characteristics of a variable-sweep airplane model

Abstract: An investigation has been made to determine the effects of the location of the wing pivot and geometry of the forewing on the static longitudinal aerodynamic characteristics at subsonic speeds of a model representing a variable-sweep supersonic fighter airplane. Results indicate that as the wing-pivot location moves aft and outboard, the change in static margin due to wing sweep is reduced. Increasing the forewing area resulted in a forward shift of the aerodynamic center as well as a slight reduction in the aerodynamic-center variation due to wing sweep.

Wind-tunnel tests of a single stage to orbit space shuttle at Mach numbers of 2.60, 3.85, and 4.64

Abstract: Wind tunnel tests were conducted to determine the static aerodynamic characteristics of a single-stage-to-orbit space shuttle model at Mach numbers of 2.60, 3.85, and 4.64. Test parameters included various payloads for the launch configurations and various corner radii for the entry configurations. The test results for the launch configurations generally indicated that a decrease in payload size resulted in a marked increase in axial force. Decreasing the corner radii of the entry configuration led to large increases in axial force. The results also indicated that the presence of an engine door on the entry configuration caused a measurable positive increment in pitching moment. Changing the entry configuration afterbody geometry from a biconic shape to a conic shape increased the normal-force and stability level.

Static Aerodynamic Stability Characteristics of a Body of Revolution with Wraparound Fins at Mach Numbers from 0.5 to 1.3

Abstract: : Wayne ;Craft,J. C. ;RD-TM-72-14AMCDR-1009*Guided missiles, *Fins, Aerodynamic stability, Aerodynamic forces, Pitch(Motion), Roll, Yaw, Angle of attackExperimental aerodynamic investigations were conducted in the NASA Langley Unitary Plan wind tunnel for several wraparound fins with straight and with swept leading edge mounted on bodies of revolution. Mach number ranged from 1.6 to 2.86 while angle of attack was varied from -6 degrees to 6 degrees at 0 degrees sideslip angle. Six component aerodynamic force and moment data were recorded on the complete missile configuration while simultaneously recording the aerodynamic normal force and bending moments acting on each fin. The results of these tests are presented herein in plotted form. The number designation given the test is UPWT 980. (Author)