ALBERT

Description: A wind-tunnel investigation was made at low speed in the Langley stability tunnel in order to determine the effects of fuselage nose length and a canopy on the oscillatory yawing derivatives of a complete swept-wing model configuration. The changes in nose length caused the fuselage fineness ratio to vary from 6.67 to 9.18. Data were obtained at various frequencies and amplitudes for angles of attack from 0 deg. to about 32 deg. Static lateral and longitudinal stability data are also presented.

Description: Results of an investigation of the static longitudinal stability and control characteristics of an aspect-ratio-3.1, unswept wing configuration equipped with an aspect-ratio-4, unswept horizontal tail are presented without analysis for the Mach number range from 0.70 to 2.22. The hinge line of the all-movable horizontal tail was in the extended wing chord plane, 1.66 wing mean aerodynamic chords behind the reference center of moments. The ratio of the area of the exposed horizontal-tail panels to the total area of the wing was 13.3 percent and the ratio of the total areas was 19.9 percent. Data are presented at angles of attack ranging"from -6 deg to +18 deg for the horizontal tail set at angles ranging from +5 deg to -20 deg and for the tail removed.

Description: An analytical approach is presented which is applicable to the optimization of homing navigation guidance systems which are forced to operate in the presence of radar noise. The two primary objectives are to establish theoretical minimum miss distance performance and a method of synthesizing the optimum control system. The factors considered are: (1) target evasive maneuver, (2) radar glint noise, (3) missile maneuverability, and (4) the inherent time-varying character of the kinematics. Two aspects of the problem are considered. In the first, consideration is given only to minimization of the miss distance. The solution given cannot be achieved in practice because the required accelerations are too large. In the second, results are extended to the practical case where the limited acceleration capabilities of the missile are considered by placing a realistic restriction on the mean-square acceleration so that system operation is confined to the linear range. Although the exact analytical solution of the latter problem does not appear feasible, approximate solutions utilizing time-varying control systems can be found. One of these solutions - a range multiplication type control system - is studied in detail. It is shown that the minimum obtainable miss distance with a realistic restriction on acceleration is close to the absolute minimum for unlimited missile maneuverability. Furthermore, it is shown that there is an equivalence in performance between the homing and beam-rider type guidance systems. Consideration is given to the effect of changes in target acceleration, noise magnitude, and missile acceleration on the minimum miss distance.

Description: A flight investigation was made to obtain experimental information on the handling qualities of a normal-acceleration type of automatic longitudinal control system. The control system was installed in a subsonic fighter-type airplane. In hands-off (stick-free) flight the normal-acceleration control system attempted to regulate the normal acceleration to a constant value which is dependent on the automatic-control-system trim setting. In maneuvering flight a given pilot's stick deflection produced a proportional change in normal acceleration, the change in acceleration being independent of flight condition. A small side-located controller stick was used by the pilot to introduce signals into the automatic control system. In the flight program emphasis was placed on the acceleration-limiting capabilities of the control system. The handling qualities were investigated in maneuvers such as slow and rapid pull-ups and turns and also in flight operations such as cruising, stalls, landings, aerobatics, and air-to-air tracking. Good acceleration limiting was obtained with the normal-acceleration control system by limiting the magnitude of the input signal that the pilot could introduce into the control system. The same values of control-system gain settings could be used from an acceleration-limiting stand-point at both 10,000 and 30,000 feet for the complete speed range of the airplane. The response characteristics of the airplane-control system combination were also satisfactory at both high and low altitude with these same values of control-system gain setting. In the pilot's opinion, the normal-acceleration control system provided good stability and control characteristics in flight operations such as cruising, stalls, landings, aerobatics, and air-to-air tracking.

Description: An investigation has been made utilizing a three-blade, 10-foot- diameter, supersonic-ty-pe propeller to determine propeller flutter characteristics. The particular flutter characteristics of interest were (1) the effect of stall flutter on a propeller operating in positive and negative thrust, (2) the effect of stall flutter on a propeller operating with the thrust axis inclined, and (3) the variation of vibratory blade shear stresses as the stall flutter boundary is penetrated and exceeded. Thrust and power measurements were made for all test conditions. Wake and inflow surveys were made when appropriate, to define the thrust and torque distributions and the magnitude of the inflow velocity. Stress measurements were made simultaneously to obtain the propeller flutter and bending response. It was found when operating both in the positive and negative thrust regions that, for most cases after the onset of flutter, the magnitude of the flutter stresses at first increased rapidly with section blade angle P, after which further increases in 0 resulted in only a moderate increase or a reduction in stress. Thrust-axis inclination up to the limit of the tests (angle of attack of 15 deg and dynamic pressure of 40 psf) appeared to have no effect on stall flutter. The stall flutter stresses were found to be directly associated with the section thrust characteristics of the blades. The onset of flutter was found to occur simultaneously with the divergence of the section thrust variation with blade angle from linearity for stations outboard of the blade 0.8-radius station. The maximum flutter stresses appeared to be a function of the maximum section thrust obtained at or in the vicinity of the blade 0.8-radius station. In an attempt to correlate two-dimensional airfoil data with three-dimensional data to predict the stall angle of attack (divergence of the section thrust) of the blade sections, it was found that no consistent correlation could be obtained. Also, a knowledge of the inflow conditions appeared to be insufficient to account for differences in airfoil characteristics between the two-dimensional and the three-dimensional cases.

Description: An investigation to determine the low-speed rolling, yawing, and sideslipping derivatives of a 1/7-scale model which was used to represent the original configuration and a modified configuration of the North American X-15 airplane has been conducted in the Langley free-flight tunnel. The original model was modified to approximately represent the final airplane configuration by reducing the size of the fuselage side fairings and changing the vertical-tail arrangement. The effects of various tail arrangements were determined for both configurations and the effect of small forebody strakes was determined for the modified configuration only.