ALBERT

Description: A comprehensive discussion of the various factors affecting the determination of stability and control derivatives from flight data is presented based on the experience of the NASA High-Speed Flight Station. Factors relating to test techniques, determination of mass characteristics, instrumentation, and methods of analysis are discussed. For most longitudinal-stability-derivative analyses simple equations utilizing period and damping have been found to be as satisfactory as more comprehensive methods. The graphical time-vector method has been the basis of lateral-derivative analysis, although simple approximate methods can be useful If applied with caution. Control effectiveness has been generally obtained by relating the peak acceleration to the rapid control input, and consideration must be given to aerodynamic contributions if reasonable accuracy is to be realized.. Because of the many factors involved In the determination of stability derivatives, It is believed that the primary stability and control derivatives are probably accurate to within 10 to 25 percent, depending upon the specific derivative. Static-stability derivatives at low angle of attack show the greatest accuracy.

Description: The aerodynamic characteristics of several noncircular two-dimensional cylinders with axes normal to the stream at various flow incidences (analogous to angles of attack of a two-dimensional airfoil and obtained by rotating the cylinders about their axes) for a range of Reynolds numbers have been determined from low-speed wind-tunnel tests. The results indicate that these parameters have rather large effects on the drag and side force developed on these cylinders. The side force is especially critical and very often undergoes a change in sign with a change in Reynolds number. Since the flow incidences correspond to combined angles of attack and sideslip in the crossflow plane of three-dimensional bodies, these two-dimensional results appear to have strong implications with regard to directional stability of fuselages at high angles of attack. These implications, along with those associated with the spin-recovery characteristics of aircraft, are briefly discussed.

Description: The influence of the deflected flow caused by the fuselage (especially by unsymmetrical attitudes) on the lift and the rolling moment due to sideslip has been discussed for infinitely long fuselages with circular and elliptical cross section. The aim of this work is to add rectangular cross sections and, primarily, to give a principle by which one can get practically usable contours through simple conformal mapping. In a few examples, the velocity field in the wing region and the induced flow produced are calculated and are compared with corresponding results from elliptical and strictly rectangular cross sections.

Description: The mutual influences of compression shocks and friction boundary layers were investigated by means of high speed wind tunnels.Schlieren optics provided a clear picture of the flow phenomena and were used for determining the location of the compression shocks, measurement of shock angles, and also for Mach angles. Pressure measurement and humidity measurements were also taken into consideration.Results along with a mathematical model are described.

Description: The longitudinal aerodynamic characteristics of a wing-body-horizontal-tail configuration designed for efficient performance at transonic speeds has been investigated at Mach numbers from 0.80 to 1.03 in the Langley 16-foot transonic tunnel. The effect of adding an outboard leading-edge chord-extension to the highly tapered 45 deg. swept wing was also obtained. The average Reynolds number for this investigation was 6.7 x 10(exp 6) based on the wing mean aerodynamic chord. The relatively low tail placement as well as the addition of a chord-extension achieved some alleviation of the pitchup tendencies of the wing-fuselage configuration. The maximum trimmed lift-drag ratio was 16.5 up to a Mach number of 0.9, with the moment center located at the quarter-chord point of the mean aerodynamic chord. For the untrimmed case, the maximum lift-drag ratio was approximately 19.5 up to a Mach number of 0.9.

Description: The motion of different bodies imersed in liquid or gaseous media is accompanied by characteristic sound which is excited by the formation of unstable surfaces of separation behind the body, usually disintegrating into a system of discrete vortices(such as the Karman vortex street due to the flow about an infintely long rod, etc.).In the noise from fans,pumps,and similar machtnery, vortexnQif3eI?Yequently predominates. The purpose of this work is to elucidate certain questions of the dependence ofthis sound upon the aerodynamic parameters and the tip speed of the rotating rods,or blades. Although scme material is given below,insufficientto calculate the first rough approximation to the solution of this question,such as the mechanics of vortex formation,never the less certain conclusions maybe found of practical application for the reduction of noise from rotating blades.

Description: Tests were conducted to find the effects of compressibility on the longitudinal stability and control of a 1/7-scale semispan model of the Northrop YB-49 airplane. Lift, drag, pitching moment, and elevon hinge moments were measured and are presented in graphical form. The results show that, due to a loss of lift on the outboard portion of the wing, the longitudinal static stability decreased rapidly as the Mach numbers increased above 0.735 the model experienced a climbing moment at positive lift coefficients. Also, a longitudinal-control effectiveness began to decrease at a Mach number of about 0.725

Description: A diamond wing and body combination was designed to have an area distribution which would result in near optimum zero-lift wave-drag coefficients at a Mach number of 1.00, and decreasing wave-drag coefficient with increasing Mach number up to near sonic leading-edge conditions for the wing. The airfoil section were computed by varying their shape along with the body radii (blending process) to match the selected area distribution and the given plan form. The exposed wing section had an average maximum thickness of about 3 percent of the local chords, and the maximum thickness of the center-line chord was 5.49 percent. The wing had an aspect ratio of 2 and a leading-edge sweep of 45 deg. Test data were obtained throughout the Mach number range from 0.20 to 3.50 at Reynolds numbers based on the mean aerodynamic chord of roughly 6,000,000 to 9,000,000. The zero-lift wave-drag coefficients of the diamond model satisfied the design objectives and were equal to the low values for the Mach number 1.00 equivalent body up to the limit of the transonic tests. From the peak drag coefficient near M = 1.00 there was a gradual decrease in wave-drag coefficient up to M = 1.20. Above sonic leading-edge conditions of the wing there was a rise in the wave-drag coefficient which was attributed in part to the body contouring as well as to the wing geometry. The diamond model had good lift characteristics, in spite of the prediction from low-aspect-ratio theory that the rear half of the diamond wing would carry little lift. The experimental lift-curve slope obtained at supersonic speeds were equal to or greater than the values predicted by linear theory. Similarly the other basic aerodynamic parameters, aerodynamic center position, and maximum lift-drag ratios were satisfactorily predicted at supersonic speeds.

Description: An investigation of a model of a standard size body in combination with a representative 45 deg swept-wing-fuselage model has been conducted in the Langley 8-foot transonic pressure tunnel over a Mach number range from 0.80 to 1.43. The body, with a fineness ratio of 8.5, was tested with and without fins, and was pylon-mounted beneath the fuselage or wing. Force measurements were obtained on the wing-fuselage model with and without the body, for an angle-of-attack range from -2 deg to approximately 12 deg and an angle-of-sideslip range from -8 deg to 8 deg. In addition, body loads were measured over the same angle-of-attack and angle-of-sideslip range. The Reynolds number for the investigation, based on the wing mean aerodynamic chord, varied from 1.85 x 10(exp 6) to 2.85 x 10(exp 6). The addition of the body beneath the fuselage or the wing increased the drag coefficient of the complete model over the Mach number range tested. On the basis of the drag increase per body, the under-fuselage position was the more favorable. Furthermore, the bodies tended to increase the lateral stability of the complete model. The variation of body loads with angle of attack for the unfinned bodies was generally small and linear over the Mach number range tested with the addition of fins causing large increases in the rates of change of normal-force coefficient and nose-down pitching-moment coefficient. The variation of body side-force coefficient with sideslip for the unfinned body beneath the fuselage was at least twice as large as the variation of this load for the unfinned body beneath the wing. The addition of fins to the body beneath either the fuselage or the wing approximately doubled the rate of change of body side-force coefficient with sideslip. Furthermore, the variation of body side-force coefficient with sideslip for the body beneath the wing was at least twice as large as the variation of this load with angle of attack.

Description: An investigation was made of the effects of body shape on the drag of a 45 deg sweptback-wing-body combination at Mach numbers from 0.90 to 1.43. Both the expansion and compression fields induced by body indentation were swept back as the stream Mach number increased from 0.94. The line of zero pressure change was generally tangent to the Mach lines associated with the local velocities over the wing and body. The strength of the induced pressure fields over the wing were attenuated with spanwise distance and the major effects were limited to the inboard 60 percent of the wing semispan. Asymmetrical body indentation tended to increase the lift on the forward portion of the wing and reduce the lift on the rearward portion. This redistribution of lift had a favorable effect on the wave drag due to lift. Symmetrical body indentation reduced the drag loading near the wing-body juncture at all Mach numbers. The reduction in drag loading increased in spanwise extent as the Mach number increased and the line of zero induced pressure became more nearly aligned with the line of maximum wing thickness. Calculations of the wave drag due to thickness, the wave drag due to lift, and the vortex drag of the basic and symmetrical M = 1.2 body and wing combinations at an angle of attack of 0 deg predicted the effects of indentation within 11 percent of the wing-basic-body drag throughout the Mach number range from 1.0 to 1.43. Calculations of the wave drag due to thickness, the wave drag due to lift, and the vortex drag for the basic, symmetrical M = 1.2, and asymmetrical M = 1.4 body and wing combinations predicted the total pressure drag to within 8 percent of the experimental value at M = 1.43.