ALBERT

Description: The hazards of lightning strokes to aircraft fuel tanks have been investigated in artificial-lightning-generation facilities specifically constructed to duplicate closely the natural lightning discharges to air craft determined through flight research programs and analysis of lightning-damaged aircraft over a period of many years. Explosion studies were made in an environmental explosion chamber using small fuel tanks under various simulated flight conditions. The results showed that there is a primary hazard whenever there is direct puncture of the fuel-tank wall, whereas the ignition of fuel by hot spots on tank walls due to lightning strikes is unlikely. Punctures of fuel-tank walls by artificial-lightning discharges produced explosions of the fuel in the mixture range from excessively lean to rich mixtures. None of the aluminum alloys, 0.081 inch thick or over, were punctured by the laboratory discharges representative of natural-lightning discharges to aircraft; however, reliance on this wall thickness for complete protection would not be justified, because occasional strokes are known to be of greater magnitude and because statistics reveal variations in the damage pattern. Data gathered by the Lightning and Transients Research Institute on lightning strokes to aircraft show that 90 percent of the strokes recorded have occurred in the temperature range of -10 to +10 C, where many of the jet fuels are flammable but where aviation gasoline is overrich. Also, 10 percent of the strokes recorded have been to the wings, which are the principal fuel-storage areas for modern aircraft. Thus, there is a hazard, particularly for jet fuels. Certain protective measures are indicated by the studies to date, such as the use of lightning diverter rods, thickening of the wing skin in areas near the most probable stroke paths, and the use of fuel-tank liners in critical areas.

Description: The technical memorandum briefly summarizes the growth of interest in aeroelastic phenomena as aircraft speed increased and wing designs changed for faster aircraft. Different types of aircraft vibrations are then introduced, and the mathematical basis for the theory behind them is described. Special attention is given to static oscillations, wing flutter, and the flutter of skin panels. The last section of the memorandum deals with the prevention of flutter by design specifications.

Description: An investigation of the low-subsonic stability and control characteristics of a model of a flat-bottom hypersonic boost-glide configuration having 78 deg sweep of the leading edge has been made in the Langley full-scale tunnel. The model was flown over an angle-of-attack range from 10 to 35 deg. Static and dynamic force tests were made in the Langley free-flight tunnel. The investigation showed that the longitudinal stability and control characteristics were generally satisfactory with neutral or positive static longitudinal stability. The addition of artificial pitch damping resulted in satisfactory longitudinal characteristics being obtained with large amounts of static instability. The most rearward center-of-gravity position for which sustained flights could be made either with or without pitch damper corresponded to the calculated maneuver point. The lateral stability and control characteristics were satisfactory up to about 15 deg angle of attack. The damping of the Dutch roll oscillation decreased with increasing angle of attack; the oscillation was about neutrally stable at 20 deg angle of attack and unstable at angles of attack of about 25 deg and above. Artificial damping in roll greatly improved the lateral characteristics and resulted in flights being made up to 35 deg angle of attack.

Description: Acceleration, airspeed, and altitude data obtained with an NACA VGH recorder from a four-engine commercial transport airplane operating over a northwestern United States-Alaska route were evaluated to determine the magnitude and frequency of occurrence of gust and maneuver accelerations., operating airspeeds, and gust velocities. The results obtained were then compared with the results previously reported in NACA Technical Note 3475 for two similar airplanes operating over transcontinental routes in the United States. No large variations in the gust experience for the three operations were noted. The results indicate that the gust-load experience of the present operation closely approximated that of the central transcontinental route in the United States with which it is compared and showed differences of about 4 to 1 when compared with that of the southern transcontinental route in the United States. In general, accelerations due to gusts occurred much more frequently than those due to operational maneuvers. At a measured normal-acceleration increment of 0.5g, accelerations due to gusts occurred roughly 35 times more frequently than those due to operational maneuvers.

Description: An analysis is made of wing deflection and streamwise twist measurements in rough-air flight of a large flexible swept-wing bomber. Random-process techniques are employed in analyzing the data in order to describe the magnitude and characteristics of the wing deflection and twist responses to rough air. Power spectra and frequency-response functions for the wing deflection and twist responses at several spanwise stations are presented. The frequency-response functions describe direct and absolute response characteristics to turbulence and provide a convenient basis for assessing analytic calculation techniques. The wing deformations in rough air are compared with the expected deformations for quasi-static loadings of the same magnitude, and the amplifications are determined. The results obtained indicate that generally the deflections are amplified by a small amount, while the streamwise twists are amplified by factors of the order of 2.0. The magnitudes of both the deflection velocities and the twist angles are shown to have significant effects on the local angles of attack at the various stations and provide the main source of aerodynamic loading, particularly at frequencies in the vicinity of the first wing-vibration mode.

Description: Analysis of the vortex model proposed by Kriebel, Seidel, and Schwind shows this representation of rotating stall satisfies, at least approximately, the requirements at the cascade. Cascade-parameter-variation effects on rotating stall were studied in a circular cascade and single-stage compressor. Modification of the single-stage compressor stopped the rotating-stall pattern and permitted observation of the pressure and velocity distribution around the annulus. Closer observation might be possible with proper flow-visualization techniques, such as a water pump.

Description: In an attempt to find an aerodynamic means of counteracting the transonic trim change of a fighter airplane, lower surface spoilers were tested on a 0.055-scale wind-tunnel model. The Mach number range of the tests was 0.8 to 1.2 at Reynolds numbers of approximately 4 million. Although the spoilers produced a moderate decrease in the trim change at low altitudes, they also produced a large increase in drag. Pressure-distribution tests with external fuel tanks showed large pressure changes on the lower surface of the wing due to the tanks.

Description: The Levy method which deals with an idealized structure was used to obtain the natural modes and frequencies of a large-scale built-up 45 deg. delta wing. The results from this approach, both with and without the effects of transverse shear, were compared with the results obtained experimentally and also with those calculated by the Stein-Sanders method. From these comparisons it was concluded that the method as proposed by Levy gives excellent results for thin-skin delta wings, provided that corrections are made for the effect of transverse shear.

Description: A cambered and twisted triangular wing of aspect ratio 2 in combination with a cambered body was investigated experimentally to determine the effectiveness of the camber in reducing the drag due to lift at trim at supersonic speeds. Four arrangements were tested comprising all combinations of a symmetrical and a cambered wing with a symmetrical and a cambered body. The camber shape investigated was derived by linearized lifting surface theory for triangular wings with sonic leading edges and satisfied the requirement that the wing be trimmed at the design Mach number and lift coefficient. The experimental results for the cambered wing and cambered body showed that the drag coefficient at trim was always greater, at the same lift coefficient, than that for the untrimmed symmetrical wing and body. The trim lift coefficient was positive and decreased with increasing Mach number. At the design Mach number of 2.24, the trim lift coefficient was somewhat lower and the drag coefficient was higher than values predicted by linearized lifting surface theory for the wing alone. A comparison of the trim lift-drag ratio of the cambered wing and cambered body with values obtained by trimming the symmetrical wing and symmetrical body either with a canard or a trailing-edge flap showed that, at approximately the design Mach number the cambered configuration developed a somewhat higher value than the trailing-edge flap configuration but a lower value than the canard configuration.

Description: The maximum Mach number and altitude capabilities of the Bell X-2 research airplane were achieved during a program conducted by the U.S. Air Force with Bell Aircraft Corp. providing operational support and the National Aeronautics and Space Administration providing instrumentation and advisory engineering assistance. A maximum geometric altitude of 126,200 feet was attained at a static pressure of 9.4 pounds per square foot and a dynamic pressure of 19.1 pounds per square foot. During the last flight of the airplane, a maximum Mach number of 3.20 was reached. The directionally divergent maneuver which terminated the final high Mach number flight was precipitated by the loss in directional stability that resulted from increasing the angle of attack. The yawing moment from the lateral control was sufficient to initiate the divergence and also to cause,, indirectly, rolling moments that were greater than the aileron capabilities of the airplane. The ensuing violent motions-resulting from inertial roll coupling caused the loss of the aircraft.