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Showing papers on "Pitching moment published in 1971"


01 Feb 1971
TL;DR: Computer program for estimating subsonic aerodynamic characteristics of various aerodynamic configurations is presented and examples and typical running times of various types of configurations are provided.
Abstract: Computer program for estimating subsonic aerodynamic characteristics of various aerodynamic configurations is presented. Program represents lifting planforms with vortex-lattice. Specific aerodynamic characteristics to be determined are described. Examples and typical running times of various types of configurations are provided.

126 citations


Journal ArticleDOI
TL;DR: In this article, the authors show that strong nonlinearities in the pitching moment occur with angle of attack, which tend to decrease stability with increasing angle for small nose bluntness and increase stability with an increasing angle.
Abstract: Free-flight tests of 12.5° half-angle cones with bluntness ratios (nose radius to base radius) from 0 to 0.5 have been conducted at Mach numbers 17 to 18. Conventional wind-tunnel tests were also conducted on 10° half-angle cones with bluntness ratios from 0 to 0.25 at Mach 10.6. Results from both of these tests are compared to recent theoretical calculations as well as to existing theories and data. The results show that the strong dependence of initial momentcurve slope on nose bluntness known to occur for Mach numbers near 10 is found to be even more pronounced at M = 17. This effect is accurately predicted by recent theoretical calculations using method of characteristics both at zero angle of attack and angles up to 5°. Other theoretical results are also shown which give good agreement, including Ericsson's theory based on the embedded Newtonian flow concept and Clay and Walchner's empirical correlation. The present paper shows, in addition, that strong nonlinearities in the pitching moment occur with angle of attack which tend to decrease stability with increasing angle for small nose bluntness and increase stability with increasing angle for large nose bluntness. Experimental results and some comparisons with theory are also presented for normal force and center-of-pressure location.

9 citations


ReportDOI
01 Jul 1971
TL;DR: In this paper, the authors describe a combined theoretical-experimental program which has been conducted with the aim of developing a computer program to predict three-degree-of-freedom trajectoreis of stores when dropped from fighter-bomber type aircraft at speeds up to the critical speed.
Abstract: : This report is the final technical report which describes a combined theoretical-experimental program which has been conducted with the aim of developing a computer program to predict three-degree-of-freedom trajectoreis of stores when dropped from fighter-bomber type aircraft at speeds up to the critical speed. Both single store and multiple store installations are treated. The report first describes the mathematical models used to represent the various aircraft components. Then the calculation of the flow field, accounting for primary interference effects, using these models is described followed by the method of calculating the normal force and pitching moment from this flow field. A method of accounting for additional interference between the wing, pylon, and store is next presented.

7 citations




01 Feb 1971
TL;DR: Stability and response characteristics of directly controlled rigid rotors at high advance ratios and correlation of mathematical model with wind tunnel test data were analyzed in this article, showing that the model is robust to wind tunnel tests.
Abstract: Stability and response characteristics of directly controlled rigid rotors at high advance ratios and correlation of mathematical model with wind tunnel test data

6 citations


Book ChapterDOI
01 Jan 1971
TL;DR: In this article, the authors investigate the magnitude and characteristics of the time dependent aerodynamic forces produced by an aircraft passing over the trailing vortices of a preceding one, where the pilot of the second aircraft is assumed to have sufficient control power to maintain his aircraft in level flight, but in practice, this condition is likely to be violated when the aircraft is close to one of the leading aircraft, particularly when the latter happens to be a large transport plane.
Abstract: When an aircraft flies across the wake of a preceding one, it is subjected to changing airloads and moments induced by the trailing vortices of the first aircraft. The purpose of this paper is to investigate the magnitude and characteristics of the time dependent aerodynamic forces so produced. Both aircraft are assumed to be in horizontal flight but the direction of flight of the second aircraft is assumed to be inclined at a small angle to the trailing vortices of the first aircraft. The airloads then will change relatively slowly with time and may be estimated with reasonable accuracy by quasi-steady aerodynamic theory without taking Wagner growth of lift effects into account. In the present study, the pilot of the second aircraft is assumed to have sufficient control power to maintain his aircraft in level flight. However, in practice, this condition is likely to be violated when the aircraft is close to one of the trailing vortices of the leading aircraft, particularly when the latter happens to be a large transport plane. In such circumstances the following aircraft could stall and would, in any case, be subjected to big changes of lift and rolling moment as indicated by the results presented. In the development of the theory it is assumed that the trailing vortices are a chord length or more below the following aircraft and that the vorticity distribution over its wings will have negligible effect on the trailing vortices themselves. The pair of vortices will give rise to an upwash distribution over the wings of an approaching aircraft which must be balanced by an equal and opposite velocity distribution induced by the vorticity distribution created over the wing. The effects of the induced velocity components along the span and in the direction of flight are neglected. The problem then is one of finding the appropriate vorticity distribution at each stage of the aircraft’s flight over the trailing vortices. This can be done approximately by using a modified lifting line theory or, more accurately, by lifting surface theory. To illustrate the methods of analysis employed, calculations were done for an aircraft with rectangular wings, but only the airloads on the wings were determined. Values of the lift, rolling moment and pitching moment coefficients, corresponding to vortex inclinations of 0, l0, 20, 30, and 60° were obtained for a rectangular wing of aspect ratio 6 at different times during its passage over the trailing vortices.

5 citations


01 Oct 1971
TL;DR: Aerodynamic characteristics of HL-10 lifting body and aerodynamic loads on center fin and control surfaces at subsonic, transonic, and supersonic speeds were analyzed in this article.
Abstract: Aerodynamic characteristics of HL-10 lifting body and aerodynamic loads on center fin and control surfaces at subsonic, transonic, and supersonic speeds

4 citations



Journal ArticleDOI
TL;DR: In this article, a method for theoretically predicting forces and moments on aircraft stores in supersonic flow is investigated, where linear theory is used to predict the flowfield due to a jet fighter-bomber type aircraft, representing aircraft wing, nose, pylons, and inlets.
Abstract: A method for theoretically predicting forces and moments on aircraft stores in supersonic flow is investigated. Linear theory is used to predict the flowfield due to a jet fighter-bomber type aircraft, representing aircraft wing, nose, pylons, and inlets. The interference loading is integrated over the store length by considering crossflow effects and buoyancy effects. The method is computerized. Theoretical pitching and yawing moment calculations for a store under an F-4C aircraft at Mach 1.2 are compared with wind-tunnel data. The results show reasonably good agreement, with the exception that finite shock effects shift the experimental data axially forward of the linear theory prediction.

1 citations


01 Nov 1971
TL;DR: In this article, a clean configuration missile with and without N-vanes was photographed with a high speed camera and the angular data obtained were fit using the WOBBLE computer program to obtain the coefficients of pitching moment, pitch damping moment, and Magnus moment.
Abstract: : Subsonic one-degree-of-freedom and three-degrees-of-freedom wind tunnel tests were performed to determine whether N-vanes attached to a spinning, statically stable missile would eliminate the Magnus instabilities which can occur at high spin rates. The clean configuration missile that was tested had a large precession limit cycle of about 35 degrees amplitude. With N-vanes at + 15 degrees, it was possible to eliminate the precession limit cycle. With the N-vanes at -15 degrees, the precession limit cycle was eliminated, and a nutation limit cycle of 25 degrees was created. The motion of the missile with and without N-vanes was photographed with a high speed camera. The angular data obtained were fit using the WOBBLE computer program to obtain the coefficients of pitching moment, pitch damping moment, and Magnus moment.


22 Jan 1971
TL;DR: In this paper, a digital computer program was developed to compute the lift, pitching moment, and flap hinge moment on a two-dimensional hydrofoil with a trailing edge flap operating near a free surface with waves.
Abstract: : A discussion is given of the development of a digital computer program to compute hydrofoil loads and some aspect of hydrofoil control. The program computes the lift, pitching moment, and flap hinge moment on a two- dimensional hydrofoil with a trailing edge flap operating near a free surface with waves. The computational approach involves the numerical solution of an integral equation relating an upwash distribution to a kernel function and pressure distribution. The pressure distribution is expanded in a truncated Glauert series; the integration is carried out numerically using a Gaussian quadrature; and the coefficients of the Glauert series are evaluated by a minimum error collocation method. The control problem investigated involves the positioning of a pivoted hydrofoil by means of a servo-controlled trim tab. When the foil is pivoted at its quarter chord point and control is implemented solely by means of a servo tab, the system is virtually uncontrollable. However, by pivoting the foil off the quarter chord point or by augmenting the servo tab with a servo attached directly to the foil, the system can be controlled.