ALBERT

Description: Comparison of transition locations for an open-nose cone, a conventional sharp cone, and a hollow cylinder showed that transition locations on the open-nose cone and the hollow cylinder were identical but differed greatly from those on the sharp cone. This is believed to be caused by the essentially two-dimensional character of leading edge of the open-nose cone. Bluntness effects on the open-nose cone observed on the hollow cylinder. Transition 2.2 times the sharp-cone transition distance by blunting the tip.

Description: An all-internal conical compression inlet with annular bleed at the throat was investigated at Mach 5.0 and zero angle of attack. The minimum contraction ratio of the supersonic diffuser, coincident with a mass-flow ratio of 1.0, was determined to be 0.084 as compared with the isentropic contraction ratio of 0.04 at Mach 5.0. The over-all inlet performance was very sensitive to the amount of annular bleed at the throat because of the extensive boundary layer. For example, the critical recovery varied from 41 percent with 6-percent bleed to 59 percent with 25-percent bleed. Decreasing the spacing between the supersonic and subsonic diffusers increased the critical mass-flow ratio but reduced the range of subcritical mass-flow regulation. A constant-area section was required ahead of the subsonic diffuser in order to obtain reasonable performance. An inlet-engine net-thrust analysis indicated that the optimum performance occurred with from 20- to 25-percent bleed, depending on how the bypassed air was handled.

Description: A supplementary investigation has been conducted in the langley 20-foot free-spinning tunnel on a l/24-scale model of the Grumman F11F-1 airplane to determine the spin and recovery characteristics with alternate nose configurations, the production version and the elongated APS-67 version, with and without empty and full wing tanks. When spins were obtained with either alternate nose configuration, they were oscillatory and recovery characteristics were considered unsatisfactory on the basis of the fact that very slow recoveries were indicated to be possible. The simultaneous extension of canards near the nose of the model with rudder reversal was effective in rapidly terminating the spin. The addition of empty wing tanks had little effect on the developed spin and recovery characteristics. The model did not spin erect with full wing tanks. For optimum recovery from inverted spins, the rudder should be reversed to 22O against the spin and simultaneously the flaperons should be moved with the developed spin; the stick should be held at or moved to full forward longitudinally. The minimum size parachute required to insure satisfactory recoveries in an emergency was found to be 12 feet in diameter (laid out flat) with a drag coefficient of 0.64 (based on the laid-out-flat diameter) and a towline length of 32 feet.

Description: Incipient spin characteristics have been investigated on a l/35-scale dynamic model of the Convair F-10% airplane. The model was launched by a catapult apparatus into free flight with various control settings, and the motions obtained were photographed. The model was ballasted for the combat loading. All tests were made with the speed brakes and landing gear retracted, and engine effects were not simulated. The results of the investigation indicated that the model would enter motions apparently simulating entry phases of spins when the elevators were deflected full up. Deflecting the rudder had little effect on the direction of the motion obtained, but when ailerons were deflected the model always rotated in a direction opposite to the aileron setting (that is, the model entered a right spin with the stick to the left). The ailerons were very influential in initiating spin entry, and the pilot should avoid, as far as possible, the use of ailerons in low-speed flight.

Description: An investigation has been made in the Langley 8-foot transonic tunnels on the aerodynamic characteristics of a 0.15-scale model of the North American Aviation 255-inch fin-stabilized external store over a maximum Mach number range of 0.60 to 1.2 and on the effects of mounting lugs, of fin orientation, of fin aspect ratio, and of fixed-transition. The Reynolds number (based on a body length of 37.50 inches) varied from 9.8 x 10(exp 6) to 13.1 x 10(exp 6). The results indicate that the static margin of the finned store at low lift coefficients was only 9 percent of body length at subsonic Mach numbers and was reduced to zero at a Mach number of 1.0, Increasing the fin aspect ratio from 1.82 to 2.41 increased the subsonic static margin to 18 percent and provided a minimum margin of 9 percent near a Mach number of l.O. Store mounting lugs or fin orientation had only small effects on the aerodynamic characteristics of the basic store.

Description: An investigation has been made in the Langley Unitary Plan wind tunnel at Mach numbers of 1.60, 1.80, and 2.00 to determine the aerodynamic characteristics of a 0.03-scale model of the Avro CF-105 airplane. The investigation included the determination of the static longitudinal and lateral stability, the control and the hinge-moment characteristics of the elevator, rudder, and aileron, as well as the vertical-tail-load characteristics. Although the data are presented without analysis, a limited inspection of the longitudinal control results indicates a loss in maximum lift-drag ratio due to trimming of about 1.8 because of the large static margin. A reduction in static margin would be expected to improve the trim lift-drag ratio but would also reduce the directional stability. With the existing static margin, the configuration is directionally unstable at angles of attack above about 6 deg or 8 deg.

Description: Operation of the original engine configuration disclosed a severe compressor stall problem at high altitude, which was largely attributed to a radial flow distortion entering the high-pressure compressor. Engine modifications for eliminating or alleviating the stall problem were investigated. These included use of variable high-pressure compressor inlet guide vanes, increased turbine-stator areas, and minor alterations in both the low- and high-pressure compressor rotors.

Description: A combined analytical and experimental determination is made of the coupled natural frequencies and mode shapes in the longitudinal plane of symmetry for a dynamic model of a single-rotor helicopter. The analytical phase is worked out on the basis of a seven-degree-of-freedom system combining elastic deflections of the rotor blades, rotor shaft, pylon, and fuselage. The calculated coupled frequencies are first compared with calculated uncoupled frequencies to show the general effects of coupling and then with measured coupled frequencies to determine the extent to which the coupled frequencies can be calculated. The coupled mode shapes are also calculated and were observed visually with stroboscopic lights during the tests. A comparison of the coupled and uncoupled natural frequencies shows that significant differences exist between these frequencies for some of the modes. Good agreement is obtained between the measured and calculated values for the coupled natural frequencies and mode shapes. The results show that the coupled natural frequencies and mode shapes can be determined by the analytical procedure presented herein with sufficient accuracy if the mass and stiffness distributions of the various components of the helicopter are known.

Description: A study is made of aerodynamic performance and static stability and control at hypersonic speeds. In a first part of the study, the effect of interference lift is investigated by tests of asymmetric models having conical fuselages and arrow plan-form wings. The fuselage of the asymmetric model is located entirely beneath the wing and has a semicircular cross section. The fuselage of the symmetric model was centrally located and has a circular cross section. Results are obtained for Mach numbers from 3 to 12 in part by application of the hypersonic similarity rule. These results show a maximum effect of interference on lift-drag ratio occurring at Mach number of 5, the Mach number at which the asymmetric model was designed to exploit favorable lift interference. At this Mach number, the asymmetric model is indicated to have a lift-drag ratio 11 percent higher than the symmetric model and 15 percent higher than the asymmetric model when inverted. These differences decrease to a few percent at a Mach number of 12. In the course of this part of the study, the accuracy to the hypersonic similarity rule applied to wing-body combinations is demonstrated with experimental results. These results indicate that the rule may prove useful for determining the aerodynamic characteristics of slender configurations at Mach numbers higher than those for which test equipment is really available. In a second part of the study, the aerodynamic performance and static stability and control characteristics of a hypersonic glider are investigated in somewhat greater detail. Results for Mach numbers from 3 to 18 for performance and 0.6 to 12 for stability and control are obtained by standard text techniques, by application of the hypersonic stability rule, and/or by use of helium as a test medium. Lift-drag ratios of about 5 for Mach numbers up to 18 are shown to be obtainable. The glider studied is shown to have acceptable longitudinal and directional stability characteristics through the range of Mach numbers studied. Some roll instability (negative effective dihedral) is found at Mach numbers near 12.

Description: An investigation was performed in the Langley Unitary Plan wind tunnel to determine the aerodynamic characteristics of a model of a 450 swept-wing fighter airplane, and to determine the loads on attached stores and detached missiles in the presence of the model. Also included was a determination of aileron-spoiler effectiveness, aileron hinge moments, and the effects of wing modifications on model aerodynamic characteristics. Tests were performed at Mach numbers of 1.57, 1.87, 2.16, and 2.53. The Reynolds numbers for the tests, based on the mean aerodynamic chord of the wing, varied from about 0.9 x 10(exp 6) to 5 x 10(exp 6). The results are presented with minimum analysis.

Description: Results have been obtained from an investigation in the Langley Unitary Plan wind tunnel at Mach numbers from 2.5 to 3.5 of a canard-type configuration designed for supersonic cruise flight. Tests extended over an angle-of-attack range from about -4 deg to 11 deg and an angle-of-sideslip range from -4 deg to 6 deg. For the present tests, the results indicate that forebody deflection was an efficient means of providing a sizable positive pitching-moment shift with little or no increase in drag. The test configuration had a trimmed lift-drag ratio of approximately 6.0 at Mach numbers near 3.0 and at a Reynolds number of 2.52 x 10(exp 6). The configuration was both longitudinally and directionally stable. The lift-drag ratios are believed to be somewhat low inasmuch as the models used for the present tests had large-grain-size transition strips fixed to the various surfaces and these strips added wave drag. Also, the model boundary- layer diverter is oversized with respect to a full - scale configuration and therefore contributes additional drag.

Description: The factors which influence the selection of landing approach speeds are discussed from the pilot's point of view. Concepts were developed and data were obtained during a landing approach flight investigation of a large number of jet airplane configurations which included straight-wing, swept-wing, and delta-wing airplanes as well as several applications of boundary-layer control. Since the fundamental limitation to further reductions in approach speed on most configurations appeared to be associated with the reduction in the pilot's ability to control flight path angle and airspeed, this problem forms the basis of the report. A simplified equation is presented showing the basic parameters which govern the flight path angle and airspeed changes, and pilot control techniques are discussed in relation to this equation. Attention is given to several independent aerodynamic characteristics which do not affect the flight path angle or airspeed directly but which determine to a large extent the effort and attention required of the pilot in controlling these factors during the approach. These include stall characteristics, stability about all axes, and changes in trim due to thrust adjustments. The report considers the relationship between piloting technique and all of the factors previously mentioned. A piloting technique which was found to be highly desirable for control of high-performance airplanes is described and the pilot's attitudes toward low-speed flight which bear heavily on the selection of landing approach speeds under operational conditions are discussed.

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: An investigation has been made to determine the aerodynamic characteristics in pitch at a Mach number of 6.8 of hypersonic missile configurations with cruciform trailing-edge flaps and with all-movable control surfaces. The flaps were tested on a configuration having low-aspect-ratio cruciform fins with an apex angle of 5 deg the all-movable controls were mounted at the 46.7-percent body station on a configuration having a 10 deg flared afterbody. The tests were made through an angle-of-attack range of -2 deg to 20 deg at zero sideslip in the Langley 11-inch hypersonic tunnel. The results indicated that the all-movable controls on the flared afterbody model should be capable of producing much larger values of trim lift and of normal acceleration than the trailing-edge -flap configuration. The flared -after body configuration had considerably higher drag than the cruciform-fin model but only slightly lower values of lift drag ratio.

Description: A supplementary investigation to determine the effect of external fuel tanks on the spin and recovery characteristics of a l/28-scale model of the North American FJ-4 airplane has been conducted in the Langley 20-foot free-spinning tunnel. The model had been extensively tested previously (NACA Research Memorandum SL38A29) and therefore only brief tests were made to evaluate the effect of tank installation. Erect spin tests of the model indicate that flat-type spins-are more prevalent with 200-gallon external fuel tanks than with tanks not installed. The recovery technique determined for spins without tanks, rudder reversal to full against the spin accompanied by simultaneous movement of ailerons to full with the spin, is recommended for spins encountered with external tanks installed. If inverted spins are encountered with external tanks installed, the tanks should be jettisoned and recovery attempted by rudder reversal to full against the spin with ailerons maintained at neutral.

Description: The evaluation of the altitude operational characteristics was part of the over-all investigation of the early developmental Iroquois engine. Engine steady-state windmilling characteristics were evaluated over a range of flight Mach numbers from 0.48 to 1.72 at altitudes of 35,000 and 50,000 feet. Engine altitude ignition limits were obtained over a range of flight Mach numbers from 0.5 to 1.5 with the standard engine ignition system and also with an oxygen boost system. A short investigation of high-speed altitude reignition following combustor blowout was conducted.

Description: An experimental investigation has been made to determine the dynamic stability and control characteristics of a 1/5-scale flying model of a jet-powered vertical-attitude VTOL research airplane in hovering and transition flight. The model was powered with either a hydrogen peroxide rocket motor or a compressed-air jet exhausting through an ejector tube to simulate the turbojet engine of the airplane. The gyroscopic effects of the engine were simulated by a flywheel driven by compressed-air jets. In hovering flight the model was controlled by jet-reaction controls which consisted of a swiveling nozzle on the main jet and a movable nozzle on each wing tip; and in forward flight the model was controlled by elevons and a rudder. If the gyroscopic effects of the jet engine were not represented, the model could be flown satisfactorily in hovering flight without any automatic stabilization devices. When the gyroscopic effects of the jet engine were represented, however, the model could not be controlled without the aid of artificial stabilizing devices because of the gyroscopic coupling of the yawing and pitching motions. The use of pitch and yaw dampers made these motions completely stable and the model could then be controlled very easily. In the transition flight tests, which were performed only with the automatic pitch and yaw dampers operating, it was found that the transition was very easy to perform either with or without the engine gyroscopic effects simulated, although the model had a tendency to fly in a rolled and sideslipped attitude at angles of attack between approximately 25 and 45 deg because of static directional instability in this range.

Description: Additional wind-tunnel tests were made of a 1/8-scale model of the Republic XP-91 airplane to determine its characteristics with various modifications. The modifications included a revised conventional tail, revised rocket arrangement, drooped wing tips, and revised landing gear and doors. Tests were also made to determine the effectiveness of the control surfaces of the model with the conventional tail and the effect of changing wing incidence and tail length. The revised rocket arrangement provided a considerable increase in the static directional stability contributed by the vee tail at small angles of yaw. The conventional tail provided a greater static directional stability than the vee tail without increasing the rolling moment due to sideslip. The rolling moment die to sideslip was considerable reduced by either drooped wing tips or open main landing-gear doors. The reduction in rolling moment due to sideslip resulting from the drooped tips was less with the landing-gear doors open than with the doors closed. A change in wing incidence from 0 degrees to 6 degrees reduced the elevator angle required for balance by approximately 6 degrees.