ALBERT

Description: An investigation was made in the Langley stability tunnel to study the influence of number of fins, fin shrouding, and fin aspect ratio on the spin instability of mortar-shell tail surfaces. It was found that the 12-fin tails tested spun less rapidly throughout the angle-of-yaw range than did the 6-fin tails and that fin shrouding reduced the spin encountered by a large amount.

Description: A study of many crash deceleration records suggested a simplified model of a crash deceleration pulse, which incorporates the essential properties of the pulse. The model pulse is considered to be composed of a base pulse on which are superimposed one or more secondary pulses of shorter duration. The results of a mathematical analysis of the seat-passenger deceleration in response to the airplane deceleration pulse are provided. On the basis of this information, presented as working charts, the maximum deceleration loads experienced by the seat and passenger in response to the airplane deceleration pulse can be computed. This maximum seat-passenger deceleration is found to depend on the natural frequency of the seat containing the passenger, considered as a mass-spring system. A method is presented that shows how to arrive at a combination of seat strength, natural frequency, and ability to absorb energy in deformation beyond the elastic limit that will allow the seat to serve without failure during an airplane deceleration pulse taken as the design requirement.

Description: The addition of fuselage volume, concentrated on top of the forward portion of the fuselage, for the purpose of delaying the drag-rise Mach number of subsonic airplanes at lifting conditions is investigated. The additions have been designed on the basis of the area rule and other important considerations to provide greater practicability of application compared with shapings previously investigated. The addition delayed the drag-rise Mach number by an increment of approximately 0.03 for a configuration having a wing with moderate thickness and 35 deg of sweepback at a lift coefficient of 0.3. A lesser delay was obtained for a configuration with a thicker wing. The additions increase the nonlinearities of the variations of pitching moment with lift,

Description: A flight investigation was conducted to determine the effects of an inlet modification and rocket-rack extension on the longitudinal trim and low-lift drag of the Douglas F5D-1 airplane. The investigation was conducted with a 0.125-scale rocket-boosted model which was flight tested at the Langley Pilotless Aircraft Research Station at Wallops Island, Va. Results indicate that the combined effects of the modified inlet and fully extended rocket racks on the trim lift coefficient and trim angle of attack were small between Mach numbers of 0.94 and 1.57. Between Mach numbers of 1.10 and 1.57 there was an average increase in drag coefficient of about o,005 for the model with modified inlet and extended rocket racks. The change in drag coefficient due to the inlet modification alone is small between Mach numbers of 1.59 and 1.64

Description: The basic aerodynamic characteristics of a 0.04956-scale model of the Convair TF-102A airplane with controls undeflected have been determined at Mach numbers from 0.60 to 1.135 for angles of attack up to approximately 22 deg in the Langley 8-foot transonic tunnel. In addition, comparisons have been made with data obtained from a previous investigation of a 0.04956-scale model of the Convair F-102A airplane. The results indicated the TF-102A airplane was longitudinally stable for all conditions tested. An increase in lift-curve slope from 0.045 to 0.059 and an 11-percent rearward shift in aerodynamic-center location occurred with increases in Mach number from 0.60 to approximately 1.05. The zero-lift drag coefficient for the TF-102A airplane increased 145 percent between the Mach numbers of 0.85 and 1.075; the maximum lift-drag ratio decreased from 9.5 at a Mach number of 0.60 to 5.0 at Mach numbers above 1.025. There was little difference in the lift and pitching-moment characteristics and drag due to life between the TF-102A and F-102A configurations. However, as compared with the F-102A airplane, the zero-lift drag-rise Mach number for the TF-102A was reduced by at least 0.06, the zero-lift peak wave drag was increased 50 percent, and the maximum lift-drag ratio was reduced as much as 20 percent.

Description: Four solid-propellant rocket engines of nominal 1000-pound-thrust were tested for starting characteristics at pressure altitudes ranging from 112,500 to 123,000 feet and at a temperature of -75 F. All engines ignited and operated successfully. Average chamber pressures ranged from 1060 to ll90 pounds per square inch absolute with action times from 1.51 to 1.64 seconds and ignition delays from 0.070 t o approximately 0.088 second. The chamber pressures and action times were near the specifications, but the ignition delay was almost twice the specified value of 0.040 second.

Description: An investigation has been completed in the Langley 20-foot free-spinning tunnel on a l/18-scale model of the Ryan X-13 airplane to determine its spin, recovery, and tumbling characteristics, and to determine the minimum altitude from which a belly landing could be made in case of power failure in hovering flight. Model spin tests were conducted with and without simulated engine rotation. Tests without simulated engine rotation indicated two types of spins: one, a slightly oscillatory flat spin; and the other, a violently oscillatory spin. Tests with simulated engine rotation indicated that spins to the left were fast rotating and steep and those to the right were slow rotating and flat. The optimum technique for recovery is reversal of the rudder to against the spin and simultaneous movement of the ailerons to full with the spin followed by movement of the elevators to neutral after the spin rotation ceases. Tumbling tests made on the model indicated that although the Ryan X-13 airplane will not tumble in the ordinary sense (end-over-end pitching motion), it may instead tend to enter a wild gyrating'motion. Tests made to simulate power failure in hovering flight by dropping the model indicated that the model entered what appeared to be a right spin. An attempt should be made to stop this motion immediately by moving the rudder to oppose the rotation (left pedal), moving the ailerons to with the spin (stick right), and moving the stick forward after the spin rotation ceases to obtain flying speed for pullout. The minimum altitude required for a belly landing in case of power failure in hovering flight was indicated to be about 4,200 feet.

Description: The results of the NACA tests to determine the tensile strength of some structural sheet materials heated to failure at temperature rates from 0.2 deg. F to 100 deg F per second under constant load conditions are reviewed . Yield and rupture stresses obtained under rapid-heating conditions are compared with the results of conventional elevated-temperature tensile tests. The relation between rapid-heating tests, short-time creep tests, and conventional creep tests is discussed . The application of a phenomenological theory for calculating rapid-heating curves is shown. Methods are given for predicting yield and rupture stresses and temperatures from master curves and temperature-rate parameters

Description: An investigation was conducted on a 35 deg swept-wing fighter airplane to determine the effects of several blunt-trailing-edge modifications to the wing and tail on the high-speed stability and control characteristics and tracking performance. The results indicated significant improvement in the pitch-up characteristics for the blunt-aileron configuration at Mach numbers around 0.90. As a result of increased effectiveness of the blunt-trailing-edge aileron, the roll-off, customarily experienced with the unmodified airplane in wings-level flight between Mach numbers of about 0.9 and 1.0 was eliminated, The results also indicated that the increased effectiveness of the blunt aileron more than offset the large associated aileron hinge moment, resulting in significant improvement in the rolling performance at Mach numbers between 0.85 and 1.0. It appeared from these results that the tracking performance with the blunt-aileron configuration in the pitch-up and buffeting flight region at high Mach numbers was considerably improved over that of the unmodified airplane; however, the tracking errors of 8 to 15 mils were definitely unsatisfactory. A drag increment of about O.OOl5 due to the blunt ailerons was noted at Mach numbers to about 0.85. The drag increment was 0 at Mach numbers above 0.90.

Description: Measurement of average skin-friction coefficients have been made on six rocket-powered free-flight models by using the boundary-layer rake technique. The model configuration was the NACA RM-10, a 12.2-fineness-ratio parabolic body of revolution with a flat base. Measurements were made over a Mach number range from 1 to 3.7, a Reynolds number range 40 x 10(exp 6) to 170 x 10(exp 6) based on length to the measurement station, and with aerodynamic heating conditions varying from strong skin heating to strong skin cooling. The measurements show the same trends over the test ranges as Van Driest's theory for turbulent boundary layer on a flat plate. The measured values are approximately 7 percent higher than the values of the flat-plate theory. A comparison which takes into account the differences in Reynolds number is made between the present results and skin-friction measurements obtained on NACA RM-10 scale models in the Langley 4- by 4-foot supersonic pressure tunnel, the Lewis 8- by 6-foot supersonic tunnel, and the Langley 9-inch supersonic tunnel. Good agreement is shown at all but the lowest tunnel Reynolds number conditions. A simple empirical equation is developed which represents the measurements over the range of the tests.

Description: An investigation has been conducted to determine the static stability and control and damping in roll and yaw of a 0.13-scale model of the Convair XFY-1 airplane with propellers off from 0 deg to 90 deg angle of attack. The tests showed that a slightly unstable pitch-up tendency occurred simultaneously with a break in the normal-force curve in the angle-of-attack range from about 27 deg to 36 deg. The top vertical tail contributed positive values of static directional stability and effective dihedral up to an angle of attack of about 35 deg. The bottom tail contributed positive values of static directional stability but negative values of effective dihedral throughout the angle-of-attack range. Effectiveness of the control surfaces decreased to very low values at the high angles of attack, The model had positive damping in yaw and damping in roll about the body axes over the angle-of-attack range but the damping in yaw decreased to about zero at 90 deg angle of attack.

Description: Altitude performance characteristics of the J65-B3 turbojet engine and its components were obtained at engine-inlet conditions corresponding to Reynolds number indices from 0.2 to 0.8 over a range of corrected engine speeds from 70 to 110 percent of rated speed. Engine operational limits up to an altitude of 75,000 feet together with ignition and windmilling characteristics were also obtained. The engine and component data are presented both in graphical and in tabulated form. The operational characteristics are presented in graphical form.

Description: An investigation was conducted in the Langley 20-foot free-spinning tunnel on a 1/23-scale model of the McDonnell F3H-1N airplane. The effects of control settings and movements upon the erect and inverted spin and recovery characteristics of the model were determined for the clean condition. Spin-recovery parachute tests were also performed. The results indicated that erect spins obtained on the airplane for the take-off or combat loadings should be satisfactorily terminated if full rudder reversal is accompanied by moving the ailerons to full with the spin (stick full right in a right spin). The spins obtained should be oscillatory in pitch, roll, and yaw. Recoveries from inverted spins should be satisfactory by full reversal of the rudder. A 16.7-foot- diameter tail parachute with a towline length of 30 feet and a drag coefficient of 0.734 should be adequate for emergency recovery from demonstration spins.

Description: The lift, pitching-moment, and drag characteristics of a missile configuration having a body of fineness ratio 9.33 and a cruciform triangular wing and tail of aspect ratio 4 were measured at a Mach number of 1.99 and a Reynolds number of 6.0 million, based on the body length. The tests were performed through an angle-of-attack range of -5 deg to 28 deg to investigate the effects on the aerodynamic characteristics of roll angle, wing-tail interdigitation, wing deflection, and interference among the components (body, wing, and tail). Theoretical lift and moment characteristics of the configuration and its components were calculated by the use of existing theoretical methods which have been modified for application to high angles of attack, and these characteristics are compared with experiment. The lift and drag characteristics of all combinations of the body, wing, and tail were independent of roll angle throughout the angle-of-attack range. The pitching-moment characteristics of the body-wing and body-wing- tail combinations, however, were influenced significantly by the roll angle at large angles of attack (greater than 10 deg). A roll from 0 deg (one pair of wing panels horizontal) to 45 deg caused a forward shift in the center of pressure which was of the same magnitude for both of these combinations, indicating that this shift originated from body-wing interference effects. A favorable lift - interference effect (lift of the combination greater than the sum of the lifts of the components) and a rearward shift in the center of pressure from a position corresponding to that for the components occurred at small angles of attack when the body was combined with either the exposed wing or tail surfaces. These lift and center-of-pressure interference effects were gradually reduced to zero as the angle of attack was increased to large values. The effect of wing-tail interference, which influenced primarily the pitching-moment characteristics, is dependent on the distance between the wing trailing vortex wake and the tail surfaces and thus was a function of angle of attack, angle of roll, and wing- tail interdigitation. Although the configuration at zero roll with the wing and tail in line exhibited the least center-of-pressure travel, the configuration with the wing and tail interdigitated had the least change in wing- tail interference over the angle - of-attack range. The lift effectiveness of the variable-incidence wing was reduced by more than 70 percent as a result of an increase in the combined angle of attack and wing incidence from 0 deg to 40 deg center dot The wing- tail interference (effective downwash at the tail) due to wing deflection was nearly zero as a result of a region of negative vorticity shed from the inboard portion of the wing. The lift characteristics of the configuration and its components were satisfactorily predicted by the calculated results, but the pitching moments at large angles of attack were not because of the influence of factors for which no adequate theory is available, such as the variation of the cross flow drag coefficient along the body and the effect of the wing downwash field on the after body loading.

Description: An investigation of the 1XP excitation of inclined single-rotation propellers has indicated a new concept for determining propeller shaft forces and moments of an inclined propeller. This report presents preliminary results, in particular to the counterrotating propeller.

Description: An investigation of a 1/14-scale dynamically similar model of a panto-base version of the Chase C-123 airplane was conducted to evaluate the hydrodynamic characteristics of the airplane. The resistance, longitudinal stability, and spray patterns during take-off and general behavior in calm- and rough-water landings were determined. Brief calm-water tests were made to compare the initial vertical impact accelerations of the model with and without hydro-skis. Take-off stability was satisfactory for calm-water operation. A ratio of gross load to maximum resistance of 3,6 was obtained. Heavy spray reached the propellers only during ski emergence. The landing behavior in calm water and in waves 3 feet by 150 feet (full scale) was satisfactory for a normal range of trim angles. Initial impacts in calmwater landings resulted in vertical accelerations of about 2 1/2 with the hydro-skis installed and about 4g with the hydro-skis removed,

Description: Missions for which a rocket interceptor is suited and the effect of rocket-engine performance on interceptor performance are discussed. Flight missions for interceptors having rocket and turbojet engines are compared, and circumstances under which a combination of rocket and turbojet may be advantageous are discussed.

Description: Additional results on the static longitudinal and lateral stability characteristics of a 0.05-scale model of the Convair F2Y-1 water-based fighter airplane were obtained in the Langley high-speed 7- by 10-foot tunnel over a Mach number range of 0.50 to 0.92. The maximum angle-of-attack range (obtained at the lower Mach numbers) was from -2 degrees to 25 degrees. The sideslip-angle range investigated was from -4 degrees to 12 degrees. The investigation included effects of various arrangements of wing fences, leading-edge chord-extensions, and leading-edge notches. Various fuselage fences, spoilers, and a dive brake also were investigated. From overall considerations of lift, drag, and pitching moments, it appears that there were two modifications somewhat superior to any of the others investigated: One was a configuration that employed a full-chord fence and a partial-chord fence located at 0.63 semispan and 0.55 semispan, respectively. The second was a leading-edge chord-extension that extended from 0.68 semispan to 0.85 semispan in combination with a leading-edge notch located at 0.68 semispan. With plus or minus 10 degrees aileron, the estimated wing-tip helix angle was reduced from 0.125 at a Mach number of 0.50 to 0.088 at a Mach number of 0.92, with corresponding rates of roll of 4.0 and 5.2 radians per second. The upper aft fuselage dive brake, when deflected 30 degrees and 60 degrees, reduced the rudder effectiveness about 10 to 20 percent and about 35 to 50 percent, respectively.

Description: Free-flight tests in the transonic speed range utilizing rocketpropelled models have been made on three pairs of 0.11-scale North American F-100 airplane wings having an aspect ratio of 3.47, a taper ratio of 0.308, 45 degree sweepback at the quarter-chord line, and thickness ratios of 31 and 5 percent to investigate the possibility of flutte r. Data from tests of two other rocket-propelled models which accidentally fluttered during a drag investigation of the North American F-100 airplane are also presented. The first set of wings (5 percent thick) was tested on a model which was disturbed in pitch by a moving tail and reached a maximum Mach number of 0.85. The wings encountered mild oscillations near the first - bending frequency at high lift coefficients. The second set of wings 9 percent thick was tested up to a maximum Mach number of 0.95 at (2) angles of attack provided by small rocket motors installed in the nose of the model. No oscillations resembling flutter were encountered during the coasting flight between separation from the booster and sustainer firing (Mach numbers from 0.86 to 0.82) or during the sustainer firing at accelerations of about 8g up to the maximum Mach number of the test (0.95). The third set of wings was similar to the first set and was tested up to a maximum Mach number of 1.24. A mild flutter at frequencies near the first-bending frequency of the wings was encountered between a Mach number of 1.15 and a Mach number of 1.06 during both accelerating and coasting flight. The two drag models, which were 0.ll-scale models of the North American F-100 airplane configuration, reached a maximum Mach number of 1.77. The wings of these models had bending and torsional frequencies which were 40 and 89 percent, respectively, of the calculated scaled frequencies of the full-scale 7-percent-thick wing. Both models experienced flutter of the same type as that experienced-by the third set of wings.

Description: Data were obtained in an altitude test chamber for a range of altitudes from 20,000 to 58,000 feet at a flight Mach number of 0.9, and for several flight Mach numbers at an altitude of 45,000 feet. Data approximating sea-level operation are also included. Engine component performance data are presented in addition to windmilling, exhaust-nozzle, and ejector performance.

Description: An investigation was made of the take-off characteristics of a 1/10-scale dynamic model of the Convair XF2Y-1 airplane. This airplane is a water-based, jet-propelled, delta-wing fighter incorporating a hydro-ski landing gear. Tests were made with the original configuration, with the beaching wheels removed, and with the wheels installed and fairings added in front of the wheels. Each configuration was tested at weight and balance conditions simulating 17,000 pounds gross weight with the moment due t o 7,600 pounds of thrust, 17,300 pounds gross weight with a 9,500-pound thrust condition, and 23,000 pounds gross weight with a 9,300-pound thrust condition. Constant-speed runs were made at various elevon settings and vertical ski-strut positions; and trim, rise, and resistance were measured. Accelerated runs were made with controlled elevons and scale shock struts which could be extended as desired, and the longitudinal stability and spray characteristics were observed and photographed.

Description: An investigation was conducted in the Langley 19-foot pressure tunnel on a 0.3-scale model of the Republic RF-84F airplane to determine modifications which would eliminate the pitch-up that occurred near maximum lift during flight tests of the airplane. The effects of high-lift and stall-control devices, horizontal tail locations, external stores, and various inlets on the longitudinal characteristics of the model were investigated. For the most part, these tests were conducted at a Reynolds number of 9.0 x 10(exp 6) and a Mach number of 0.19. The results indicated that from the standpoint of stability the inlets should possess blunted side bodies. The horizontal tail located at either the highest or lowest position investigated improved the stability of the model. Three configurations were found for the model equipped with the production tail which eliminated the pitch-up through the lift range up to the maximum lift and provided a stable static margin which did not vary more than 15% of the mean aerodynamic chord through the lift range up to 85% of maximum lift. The three configurations are as follows: the production wing-fuselage-tail combination with an inlet similar to the production inlet but smaller in plan form in conjunction with either (1) a wing fence located at 65% of the win semispan or (2) an 11.7% chord leading-edge extension extending from 65.8 to 95.8% of the wing semispan and (3) the production wing-fuselage-tail combination with the production inlet and an 11.7% chord leading-edge extension extending from 70.8 to 95.8% of the wing semispan.