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Showing papers on "Pitching moment published in 1973"


01 May 1973
TL;DR: In this article, a new method has been developed for calculating the pressure distribution and aerodynamic characteristics of wing-body-tail combinations in subsonic and supersonic potential flow.
Abstract: A new method has been developed for calculating the pressure distribution and aerodynamic characteristics of wing-body-tail combinations in subsonic and supersonic potential flow. A computer program has been developed to perform the numerical calculations. The configuration surface is subdivided into a large number of panels, each of which contains an aerodynamic singularity distribution. A constant source distribution is used on the body panels, and a vortex distribution having a linear variation in the streamwise direction is used on the wing and tail panels. The normal components of velocity induced at specified control points by each singularity distribution are calculated and make up the coefficients of a system of linear equations relating the strengths of the singularities to the magnitude of the normal velocities. The singularity strengths which satisfy the boundary condition of tangential flow at the control points for a given Mach number and angle of attack are determined by solving this system of equations using an interactive procedure. Once the singularity strengths are known, the pressure coefficients are calculated, and the forces and moments acting on the configuration determined by numerical integration.

103 citations


01 Dec 1973
TL;DR: In this paper, wind-tunnel tests have been conducted to determine the low-speed two-dimensional aerodynamic characteristics of a 17-percent-thick airfoil designed for general aviation applications (GA(W)-1).
Abstract: Wind-tunnel tests have been conducted to determine the low-speed two-dimensional aerodynamic characteristics of a 17-percent-thick airfoil designed for general aviation applications (GA(W)-1). The results were compared with predictions based on a theoretical method for calculating the viscous flow about the airfoil. The tests were conducted over a Mach number range from 0.10 to 0.28. Reynolds numbers based on airfoil chord varied from 2.0 million to 20.0 million. Maximum section lift coefficients greater than 2.0 were obtained and section lift-drag ratio at a lift coefficient of 1.0 (climb condition) varied from about 65 to 85 as the Reynolds number increased from about 2.0 million to 6.0 million.

90 citations


Journal ArticleDOI
TL;DR: In this paper, the span-wise aerodynamic loading of a wing with minimum induced drag was derived for prescribed lift and root-bending moment, in the case that the lift and its moment of inertia about the longitudinal axis of the aircraft are given.
Abstract: The spanwise aerodynamic loading of wings having minimum induced drag is derived for prescribed lift and root-bending moment. THis problem is an alternative to Prandtl's solution for the case that the lift and its moment of inertia about the longitudinal axis of the aircraft are given.

32 citations


01 Apr 1973
TL;DR: In this article, an engineering-type method is presented for estimating normal-force, axial-force and pitching-moment coefficients for slender bodies of circular and non-circular cross section alone and with lifting surfaces.
Abstract: An engineering-type method is presented for estimating normal-force, axial-force, and pitching-moment coefficients for slender bodies of circular and noncircular cross section alone and with lifting surfaces. Static aerodynamic characteristics computed by the method are shown to agree closely with experimental results for slender bodies of circular and elliptic cross section and for winged-circular and winged-elliptic cones. However, the present experimental results used for comparison with the method are limited to angles of attack only up to about 20 deg and Mach numbers from 2 to 4.

23 citations


Patent
08 Mar 1973
TL;DR: In this paper, a deployment mechanism employing multiple, nonproportional, four-bar linkages for guiding the movement of an airfoil control surface is proposed, where the linkages are connected at one of their ends to the control surface and at their other end to an air foil, and the connection points of each linkage define a set of lines which are skewed at progressively greater angles with respect to a reference line.
Abstract: A deployment mechanism employing multiple, nonproportional, four-bar linkages for guiding the movement of an airfoil control surface. The linkages are connected at one of their ends to the airfoil control surface and at their other end to an airfoil. The connection points of each linkage to the airfoil define a set of lines which are skewed at progressively greater angles with respect to a reference line. The mechanism when activated guides the extension of the control surface into an overlapping position with the airfoil. The overlap forms a nozzle between the airfoil control surface and the airfoil surface. The amount of overlap at any point is a constant percentage of the airfoil chord length at that same point.

21 citations


24 Oct 1973
TL;DR: In this paper, a method initially proposed by Bryson is extended to include asymmetric shedding, which employs the impulsive flow analogy, and models each wake vortex using a single-point vortex.
Abstract: : A method initially proposed by Bryson is extended to include asymmetric shedding. This method employs the impulsive flow analogy, and models each wake vortex using a single-point vortex. Free parameters inherent in the problem formulation are determined empirically. Normal force, pitching moment and yawing force coefficients are predicted for slender bodies with a nose fineness ratio greater than four and at a Mach number less than 0.9. (Modified author abstract)

14 citations


Journal ArticleDOI
TL;DR: In this paper, the authors considered the application of complex coordinates to the study of the dynamic characteristics of the tip-pathplane of a helicopter rotor and provided a convenient and natural framework for investigation of the response characteristics of fully articulated and hingeless rotors.
Abstract: The application of complex coordinates to the study of the dynamic characteristics of the tip-pathplane of a helicopter rotor is considered. The coordinates describing rotor blade motions can be transformed from individual blade flapping angles to linear combinations of flapping angles. Two of the transformed coordinates can be interpreted as describing the tilting motion of the tip-path-plane. In a stationary reference frame, these two new coordinates are coupled. However, by defining a set of complex coordinates, for near hovering flight these two coupled differential equations can be combined into a single second-order differential equation describing the tilting motion of the tip-path-plane. This formulation provides a convenient and natural framework for investigation of the response characteristics of fully articulated and hingeless rotors. Considerable insight into the influence of various physical parameters on the behavior of the tip-pathrplane can be gained. The approach is illustrated by consideration of the transient and frequency response characteristics of the tip-pathplane and the influence of flapping feedback. An extension of the root locus method is described which makes the investigation of flapping feedback convenient. Nomenclature! a - rotor blade lift curve slope b = number of blades A(s) = transfer function c = rotor blade chord Ci = rotor hub rolling moment coefficient, positive for right roll, Ci = L/pKR2(VR)2R Cm — rotor hub pitching moment coefficient, positive nose up, Cm = M/pirR2(QR)2R Cmz = complex rotor hub moment coefficient, Cmz— Cm + iCi /i = rotor blade flapping moment of inertia j = constant of proportionali ty between complex inflow and complex rotor aerodynamic hub moment coefficient KH = gain parameter associated with integral flapping feedback Kp,e = parameters associated with proportional flapping feedback mz — dimensionless aerodynamic moment acting on rotor blade due to forward speed. Effect of cyclic pitch, angular rates and flapping not included p = rotor blade natural frequency in flapping divided by rotor rpm, also roll rate nondimensionalized by rotor rpm, positive roll right q = pitch rate nondimensionalized by rotor rpm, positive nose up

13 citations


01 Nov 1973
TL;DR: In this article, an exploratory investigation has been made at Mach numbers from 0.40 to 0.95 to determine the effects on lift, drag, and pitching moment of blowing a jet exhaust over the upper surface of a 50 deg swept leading-edge wing.
Abstract: An exploratory investigation has been made at Mach numbers from 0.40 to 0.95 to determine the effects on lift, drag, and pitching moment of blowing a jet exhaust over the upper surface of a 50 deg swept leading-edge wing. Also investigated were the effects of varying the longitudinal and vertical location of the nozzle exit on the induced effects of jet blowing.

9 citations


01 Oct 1973
TL;DR: In this article, the authors used the Maple-Synge theory to establish the analytical form of the aerodynamic forces and moments acting on a four-finned, curved finned vehicle and investigated the effects of the curved- finned aerodynamics on the stability of motion.
Abstract: : Maple-Synge theory is used to establish the analytical form of the aerodynamic forces and moments acting on a four-finned, curved-finned vehicle. The linearized equations of motion are solved, and the effects of the curved- finned aerodynamics on the stability of motion are investigated. It is shown that small and moderate values of yawing moment due to angle-of-attack, characteristic of curved-finned configurations, can have significant effects on both transient and steady-state stability.

9 citations


01 May 1973
TL;DR: In this paper, a wind-tunnel investigation was conducted with a vectored-thrust V/STOL fighter configuration, which was equipped with two nacelle-mounted V2F jet simulators and one lift-jet simulator.
Abstract: A wind-tunnel investigation was conducted with a vectored-thrust V/STOL fighter configuration. The model was equipped with two nacelle-mounted vectored-thrust jet simulators and one lift-jet simulator. The vectored-thrust jet could be tested at two alternate longitudinal positions and three nozzle deflection angles. The vectored-thrust configuration with the rear nozzles showed an increase in lift and a decrease in pitching moment when compared with the forward nozzles. The rear nozzles also improve stall characteristics.

7 citations


01 Sep 1973
TL;DR: In this paper, a steady aerodynamic analysis was performed on a hypothetical RPV as it approached the mothership for docking, and the correct trend in the variation of lift and static pitch moment of the RPV was predicted using a horseshoe vortex model to represent the mother ship.
Abstract: : A potential approach for recovery of remotely piloted vehicles (RPV) is the in-flight docking of the RPV with a mothership Some design concepts are presented for in-flight recovery systems A steady aerodynamic analysis was performed on a hypothetical RPV as it approached the mothership for docking The correct trend in the variation of lift and static pitch moment of the RPV was predicted using a horseshoe vortex model to represent the mothership

01 Aug 1973
TL;DR: In this article, the effects of flap span and wing aspect ratio on the static longitudinal aerodynamic characteristics and chordwise and spanwise pressure distributions on the wing and trailing-edge flap of a straight-wing STOL model having an externally blown jet flap without vertical and horizontal tail surfaces were investigated.
Abstract: An investigation has been conducted to determine the effects of flap span and wing aspect ratio on the static longitudinal aerodynamic characteristics and chordwise and spanwise pressure distributions on the wing and trailing-edge flap of a straight-wing STOL model having an externally blown jet flap without vertical and horizontal tail surfaces. The force tests were made over an angle-of-attack range for several thrust coefficients and two flap deflections. The pressure data are presented as tabulated and plotted chordwise pressure-distribution coefficients for angles of attack of 1 and 16. Pressure-distribution measurements were made at several spanwise stations.

01 Feb 1973
TL;DR: A review of a number of research investigations as a part of a joint NASA/Army rotorcraft project is presented in this paper, which is directed toward achieving a better understanding of rotor unsteady airfoils.
Abstract: The basic unsteady aerodynamic environment of the rotary wing is summarized. Some of the observed trends in the state of the art are discussed. Some of the research needs that will require attention are reported. A review of a number of research investigations as a part of a joint NASA/Army rotorcraft project is presented. The research is directed toward achieving a better understanding of rotor unsteady airfoils. The investigations include: (1) rotor maneuver loads; (2) level flight and maneuver wake prediction; (3) tip-vortex flow; (4) blade-vortex interactions; (5) dynamic stall; (6) transient Mach number air loads; and (7) development of variable geometry rotors.

01 Jul 1973
TL;DR: In this article, a 10-deg. circular cone with various spherical and conical (45 degrees) nose bluntness of 1.7, 10% and 25% was investigated in ARL's Mach-14 wind tunnel.
Abstract: : A 10-deg. circular cone with various spherical and conical (45 degrees) nose bluntnesses of 1.7%, 10% and 25% was investigated in ARL's Mach-14 wind tunnel. Test results confirm that the static and dynamic stability coefficients are not equal in pitch and in yaw for nonzero angles of attack if the pitching moment becomes a nonlinear function of angle of attack due to nose flunting. The inequality of the in plane and out of plane stability derivatives was found at small angles of attack which are only fractions of the cone half angle. (Author)

ReportDOI
01 Feb 1973
TL;DR: An in-flight simulation to investigate minimum longitudinal stability for large delta-wing transports in landing approach and touchdown (including ground effect) was conducted using the USAF/Calspan Total In-Flight Simulator (TIFS) airplane as discussed by the authors.
Abstract: : TRANSPORT PLANES), APPROACH, FLIGHT SIMULATORS, STABILITY, INSTRUMENT FLIGHT, PILOTS, REACTION(PSYCHOLOGY), AVIATION SAFETY, GROUND EFFECT*LONGITUDINAL STABILITY, TOUCHDOWN, EVALUATIONAn in-flight simulation to investigate minimum longitudinal stability for large delta-wing transports in landing approach and touchdown (including ground effect) was conducted using the USAF/Calspan Total In-Flight Simulator (TIFS) airplane. Aerodynamic, inertial and control data for this class of airplane were obtained from a prototype Concorde package supplied by the FAA. The simulation program involved the examination of 20 configurations by four evaluation pilots. The configurations evaluated were based upon a systematic variation of the longitudinal stability characteristics for this class of airplane. These variations were designed to examine the influence of pitch stiffness, backsideness, pitch damping and nonlinear pitching moment effects on pilot acceptability of minimum longitudinal stability for the landing approach task. A total of 61 evaluations was performed.

Journal ArticleDOI
TL;DR: In this article, the effect of noise in the data on the accuracy of the extracted information was investigated and the relationship between noise level and the performance of the evaluated aerodynamic derivatives of the model was presented.
Abstract: The dynamic testing of a model in the University of Virginia cold magnetic balance wind-tunnel facility is expected to consist of measurements of the balance forces and moments, and the observation of the essentially six degree of freedom motion of the model. The aerodynamic derivatives of the model are to be evaluated from these observations. The basic feasibility of extracting aerodynamic information from the observation of a model which is executing transient, complex, multi-degree of freedom motion is demonstrated. It is considered significant that, though the problem treated here involves only linear aerodynamics, the methods used are capable of handling a very large class of aerodynamic nonlinearities. The basic considerations include the effect of noise in the data on the accuracy of the extracted information. Relationships between noise level and the accuracy of the evaluated aerodynamic derivatives are presented.

01 Feb 1973
TL;DR: In this article, an analysis of numerical methods for extracting aerodynamic coefficients from dynamic test data has been conducted and the emphasis of the analysis is on the effects that random measurement errors in the data and random disturbances in the system have on the accuracy with which the coefficients for linear and nonlinear systems can be determined.
Abstract: : An analysis of numerical methods for extracting aerodynamic coefficients from dynamic test data has been conducted The emphasis of the analysis is on the effects that random measurement errors in the data and random disturbances in the system have on the accuracy with which the coefficients for linear and nonlinear systems can be determined Both deterministic and stochastic methods for extracting the coefficients and determining their uncertainties are considered The deterministic technique considered, due to Chapman and Kirk, provides excellent estimates of both linear and nonlinear static pitching moment coefficients for the range of measurement efforts and system noise considered Somewhat less accurate estimates of pitch damping coefficients are obtained The stochastic approach considered demonstrates the feasibility of using an extended Kalman filter, with a parameter augmented state vector, for determining the values of the aerodynamic coefficients and their uncertainties from noisy dynamic test data Parameter estimates obtained from the extended filter compare favorably with previously obtained results using deterministic techniques Estimates of the parameter uncertainties provided by the filter are generally superior to those obtained with deterministic techniques particularly when system noise has corrupted the data

01 May 1973
TL;DR: In this paper, the authors describe a lifting surface theory for calculating the aerodynamic characteristics of jet-flapped wings based on a finite-element scheme -the method of elementary vortex distribution or the EVD method, which is represented by a set of overlapped elementary vortex distributions.
Abstract: : The report describes a lifting surface theory for calculating the aerodynamic characteristics of jet-flapped wings. Based on a finite-element scheme - the method of Elementary Vortex Distribution or the EVD method, the wing and jet sheet are represented by a set of overlapped elementary vortex distributions. A solution is obtained by satisfying a set of mixed-type boundary conditions on both the wing and jet sheet. The EVD method, as described, provides the following: spanwise and chordwise loading; spanwise variation of induced drag; a capability to investigate the effects of part span flaps, part span blowing, rolling, yawing, pitching, and sideslip; and total lift and induced drag (momentum method), pitching moment, yawing and rolling moments, and side force. (Modified author abstract)


01 Dec 1973
TL;DR: In this paper, an experimental investigation was conducted in a 9- by 7-foot supersonic wind tunnel to determine the effect of plume-induced flow separation and aspiration effects due to operation of both the orbiter and the solid rocket motors on a 0.019-scale model of the space shuttle vehicle.
Abstract: An experimental investigation was conducted in a 9- by 7-foot supersonic wind tunnel to determine the effect of plume-induced flow separation and aspiration effects due to operation of both the orbiter and the solid rocket motors on a 0.019-scale model of the launch configuration of the space shuttle vehicle. Longitudinal and lateral-directional stability data were obtained at Mach numbers of 1.6, 2.0, and 2.2 with and without the engines operating. The plumes exiting from the engines were simulated by a cold gas jet supplied by an auxiliary 200 atmosphere air supply system, and by solid body plume simulators. Comparisons of the aerodynamic effects produced by these two simulation procedures are presented. The data indicate that the parameters most significantly affected by the jet plumes are the pitching moment, the elevon control effectiveness, the axial force, and the orbiter wing loads.

R. T. Medan1
01 Sep 1973
TL;DR: In this paper, a computer program that provides the geometry and boundary conditions appropriate for an analysis of a lifting, thin wing with control surfaces in linearized, subsonic, steady flow is presented.
Abstract: A computer program that provides the geometry and boundary conditions appropriate for an analysis of a lifting, thin wing with control surfaces in linearized, subsonic, steady flow is presented. The kernel function method lifting surface theory is applied. The data which is generated by the program is stored on disk files or tapes for later use by programs which calculate an influence matrix, plot the wing planform, and evaluate the loads on the wing. In addition to processing data for subsequent use in a lifting surface analysis, the program is useful for computing area and mean geometric chords of the wing and control surfaces.

01 Jan 1973
TL;DR: An exploratory wind tunnel test has been conducted at Mach numbers from 0.60 to 0.80 to investigate the effects of nozzle geometry and upper surface blowing on the aerodynamic characteristics of a 14-percent-thick airfoil as discussed by the authors.
Abstract: An exploratory wind tunnel test has been conducted at Mach numbers from 0.60 to 0.80 to investigate the effects of nozzle geometry and upper surface blowing on the aerodynamic characteristics of a 14-percent-thick airfoil. Measured data included lift, drag, pitching moments, surface pressures, and nozzle thrust.

01 Jun 1973
TL;DR: In this article, a low speed wind tunnel test was conducted to assess the effects of the larger JT8D refan nacelles on the stability and control characteristics of the DC-9-30, with emphasis on the deep stall regime.
Abstract: A low speed wind tunnel test was conducted to assess the effects of the larger JT8D refan nacelles on the stability and control characteristics of the DC-9-30, with emphasis on the deep stall regime. Deep stall pitching moment and elevator hinge moment data, and low angle of attack tail-on and tail-off pitching moment data are presented. The refan nacelle was tested in conjunction with various pylons of reduced span relative to the production DC-9-30 pylon. Also, a horizontal tail that was larger than the production tail was tested. The data show that the refan installation has a small detrimental effect on the DC-9-30 deep stall recovery capability, that recovery characteristics are essentially independent of pylon span, and that the larger horizontal tail significantly increases recovery margins. The deep stall characteristics with the refan installation, within the range of pylon spans tested, are acceptable with no additional design changes anticipated.

Journal ArticleDOI
TL;DR: In this paper, the power spectral technique has been extended to show the effect of aerodynamic nonlinearities on the normal acceleration response of a rigid aircraft in the cruise configuration, and the resulting changes from the linear root-mean-square values of normal acceleration were only 3 to 5½% for a root-means-square vertical gust velocity of 20 m/s.
Abstract: The power spectral technique has been extended to show the effect of aerodynamic non-linearities on the normal acceleration response of a rigid aircraft in the cruise configuration. Non-linearities in the normal force and pitching moment variations with incidence have been considered. The resulting changes from the linear root-mean-square values of normal acceleration were only 3 to 5½% for a root-mean-square vertical gust velocity of 20 m/s.