ALBERT

Beschreibung: research

Beschreibung: An drei Orten im nordwestdeutschen Flachlande konnten zwei Interstadiale der Weichseleiszeit festgestellt werden. Vermutlich gehören die Torfe oberhalb des Interglazials von Eversen bei Rotenburg/Hann. ebenfalls einem dieser Interstadiale an (Wolff & Schröder 1940). Die stratigraphischen Verhältnisse lassen eine eindeutige Eingliederung in die Weichselvereisung zu. In allen Fällen liegen die beiden Torfbänke oberhalb des letzten Interglazials und sind in Sanden der letzten Eiszeit eingeschlossen. Von einer Parallelisierung mit den Stadien der Weichselvereisung soll vorläufig Abstand genommen werden. Durch die Pollenanalyse ergab sich eine Unterteilung in eine Birken- und in eine Kiefernphase. Das Klima war subarktisch-ozeanisch. Die Wiedervereisung nach dem letzten Interglazial und den beiden Interstadialen wurde durch ein feucht-kühles, niederschlagsreiches (vermutlich schneereiches) Klima eingeleitet. Beide Interstadiale haben mit der Alleröd-Oszillation gemeinsame Züge, lassen sich jedoch klar von ihr trennen. Ob eine Unterscheidung der beiden Interstadiale mit Hilfe der Pollenanalyse vorgenommen werden kann, ließen die bisherigen Untersuchungen nicht erkennen und muß daher den weiteren Arbeiten vorbehalten bleiben.

Beschreibung: research

Beschreibung: research

Beschreibung: research

Beschreibung: research

Beschreibung: research

Beschreibung: research

Beschreibung: research

Beschreibung: research

Beschreibung: research

Beschreibung: research

Beschreibung: research

Beschreibung: research

Beschreibung: research

Beschreibung: In der Nachbarschaft von Flüssen und größeren Bächen nimmt der Löß sehr häufig eine sandige Beschaffenheit an, in vielen Gebieten ist sogar die Sandkomponente in Form von verschieden starken Schichten (Sandbänder) eingeschaltet. Die vorliegende Arbeit beschäftigt sich in der Hauptsache mit der Lagerung der Sandbänder im vertikalen Profil und mit den Veränderungen in horizontaler Richtung. Auf Grund der allgemeinen Ausbildung und der Körnungsunterschiede des Lößes konnten westlich Königgrätz 4 Faziesbereiche unterschieden werden und zwar die Flugsand-, Bänder-, Übergangs- und Normalfazies. Die 4 Fazieszonen folgen von O nach W aufeinander. Die speziellen Untersuchungen an den Sandbändern führten zu folgenden Ergebnissen: 1. Die Abnahme des prozentualen Anteils der Gesamtmächtigkeit der Sandbänder am Löß erfolgt linear mit dem Logarithmus der Entfernung vom Auswehungsgebiet. 2. Die durchschnittliche Anzahl der Sandbänder je Meter Löß ändert sich mit der Zunahme der Entfernung vom Ursprungsgebiet, wie im einzelnen aus Abb. 10 hervorgeht. 3. Neben der Gesamtmächtigkeit und der Anzahl der Sandbänder wurde auch die Häufigkeit der verschiedenen Mächtigkeitsstufen ermittelt und graphisch aufgetragen (Abb. 11). Dabei ergab sich, daß in unmittelbarer Nähe des Auswehungsgebietes neben wenigen 9 und 10 cm mächtigen Bändern die 2 und 2,5 cm starken überwiegen. In einer Entfernung von etwa 300 m vom Flußufer dominieren die 1 und 1,5 cm starken. Weiter nach W nehmen aber nach und nach die dickeren Lagen wieder zu, so daß am Ende der Sandbänderfazies die 2 und 2,5 cm starken die häufigsten sind. Aus den mitgeteilten Ergebnissen und aus zahlreichen Körnungsanalysen wird geschlossen, daß sowohl die eingelagerten Sandschichten als auch der gößte Teil des Lößes äolischen Ursprungs sind. Sandlagen und Löß wurden durch östliche Winde aus dem breiten Tal der Elbe ausgeweht und auf die benachbarten Höhen transportiert. Dabei gelangte der in Form von Bändern abgelagerte Sand bis zu einer Entfernung von ca. 700 m. Die Kornfeinheit des Lößes nimmt nach W hin immer mehr zu. Weitere Beobachtungen an dem Bystrice- und Cidlina-Tal im W des Untersuchungsgebietes bestätigen die Annahme von der Auswehung des Lößes und der Sandschichten aus den Tälern durch Winde aus östlicher Richtung. Daß die allgemeine Lößverbreitung auch von der Breite der Täler abhängig ist, zeigt ein Blick auf die Karte (Abb. 1).

Beschreibung: research

Beschreibung: research

Beschreibung: research

Beschreibung: research

Beschreibung: Pressure-distribution measurements have been made on the fus elage of the Bell X- 1 research airplane. Data are presented for angles of attack from 2 deg. to 8 deg. during pull-ups at Mach numbers of about 0.78, 0.85, 0.88, and 1.02. The results of the investigation indicated that a large portion of the load carried by the fuselage was in the vicinity of the wing and may be attributed to wing-to-fuselage carryover. The presence of the wing from the 41 to 60 percent fuselage stations influenced the fuselage pressures from about 30 to 65 percent fuselage length at Mach numbers of approximat ely 0.78, 0.85, and 0.88, and from about 35 to 80 percent fuselage length at a Mach number of approximately 1.02. The fuselage contributed about 20 percent of the total airplane normal-force coefficient. The center of pressure of the fuselage load throughout the tests was located from 41 to 51 percent fuselage length, which corresponds to the forward half of the wing root-chord location.

Beschreibung: A method has been proposed for predicting the effect of a rapid blade-pitch increase on the thrust and induced-velocity response of a helicopter rotor. General equations have been derived for the ensuing motion of the helicopter. These equations yield time histories of thrust, induced velocity, and helicopter vertical velocity for given rates of blade-pitch-angle changes and given rotor-angular-velocity time histories. The results of the method have been compared with experimental results obtained with a rotor mounted on the Langley helicopter test tower. The calculated and experimental results are in good agreement, although, in general, the calculated thrust-coefficient overshoots are about 10 percent greater than those obtained experimentally.

Beschreibung: A cascade of 65-(12)10 compressor blades was tested at one geometric setting over a range of inlet Mach number from 0.12 to 0.89. Two groups of data are presented and compared: the first from the cascade operating conventionally with no boundary-layer control, and the second with the boundary layer controlled by a combination of upstream slot suction and porous-wall suction at the blade tips. A criterion for two-dimensionality was used to specify the degree of boundary-layer control by suction to be applied. The data are presented and an analysis is made to show the effect of Mach number on turning angle, blade wake, pressure distribution about the blade profile and static-pressure rise. The influence of boundary-layer control on these parameters as well as on the secondary losses is illustrated. A system of correlating the measured static-pressure rise through the cascade with the theoretical isentropic values is presented which gives good agreement with the data. The pressure distribution about the blade profile for an inlet Mach number of 0.21 is corrected with the Prandtl-Glauert, Karman-Tsien, and vector-mean velocity - contraction coefficient compressibility correction factors to inlet Mach numbers of 0.6 and 0.7. The resulting curves are compared with the experimental pressure distributions for inlet Mach numbers of 0.6 and 0.7 so that the validity of applying the three corrections can be evaluated.

Beschreibung: The empirical relation between the induced velocity, thrust, and rate of vertical descent of a helicopter rotor was calculated from wind tunnel force tests on four model rotors by the application of blade-element theory to the measured values of the thrust, torque, blade angle, and equivalent free-stream rate of descent. The model tests covered the useful range of C(sub t)/sigma(sub e) (where C(sub t) is the thrust coefficient and sigma(sub e) is the effective solidity) and the range of vertical descent from hovering to descent velocities slightly greater than those for autorotation. The three bladed models, each of which had an effective solidity of 0.05 and NACA 0015 blade airfoil sections, were as follows: (1) constant-chord, untwisted blades of 3-ft radius; (2) untwisted blades of 3-ft radius having a 3/1 taper; (3) constant-chord blades of 3-ft radius having a linear twist of 12 degrees (washout) from axis of rotation to tip; and (4) constant-chord, untwisted blades of 2-ft radius. Because of the incorporation of a correction for blade dynamic twist and the use of a method of measuring the approximate equivalent free-stream velocity, it is believed that the data obtained from this program are more applicable to free-flight calculations than the data from previous model tests.

Beschreibung: The autorotative performance of an assumed helicopter was studied to determine the effect of inoperative jet units located at the rotor-blade tip on the helicopter rate of descent. For a representative ramjet design, the effect of the jet drag is to increase the minimum rate of descent of the helicopter from about 1,OO feet per minute to 3,700 feet per minute when the rotor is operating at a tip speed of approximately 600 feet per second. The effect is less if the rotor operates at lower tip speeds, but the rotor kinetic energy and the stall margin available for the landing maneuver are then reduced. Power-off rates of descent of pulse-jet helicopters would be expected to be less than those of ramjet. helicopters because pulse jets of current design appear to have greater ratios of net power-on thrust to power-off, drag than currently designed rain jets. Iii order to obtain greater accuracy in studies of autorotative performance, calculations in'volving high power-off rates of descent should include the weight-supporting effect of the fuselage parasite-drag force and the fact that the rotor thrust does not equal the weight of the helicopter.

Beschreibung: 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.

Beschreibung: The damping in roll and rolling effectiveness of two models of a missile having cruciform, triangular, interdigitated wings and tails have been determined through a Mach number range of 0.8 to 1.8 by utilizing rocket-propelled test vehicles. Results indicate that the damping in roll was relatively constant over the Mach umber range investigated. The rolling effectiveness was essentially constant at low supersonic speeds and increased with increasing mach numbers in excess of 1.4 over the Mach number range investigated. Aeroelastic effects increase the rolling-effectiveness parameters pb/2V divided by delta and decrease both the rolling-moment coefficient due to wing deflection and the damping-in-roll coefficient.

Beschreibung: 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.

Beschreibung: A tank investigation has been conducted on a 1/8-size powered dynamic model of the Grumman JRF-5 airplane equipped with twin hydro-skis. The results of tests using two types of skis are presented: one had vertical sides joining the top surface to the chine; the other had the top surface faired to the chine to eliminate the vertical sides. Both configurations had satisfactory longitudinal stability although the model had a slightly greater stable elevator range available when the skis without the vertical sides were attached. Free model tests indicated no instability present when one ski emerged before the other. Considerable excess thrust was available at all speeds with either type of skis. A hump gross load-resistance ratio of 3.37 was obtained with the skis with the vertical sides and 3.53 with the other skis. Landing behavior in smooth water with yaw up to 15deg and roll up to 15deg in opposite directions was satisfactory with either type of skis.

Beschreibung: An investigation was made to determine the static longitudinal and lateral stability and control characteristics of a l/6-scale model of the revised Republic XF-84H airplane with and without the propeller operating. The model had a 40deg swept wing of aspect ratio 3.45 and was equipped with a thin, three-blade supersonic-type propeller. Modifications incorporated in the revised model included a raised horizontal tail, increased rudder size, wing fences at 65 percent semispan, and a modified wing leading edge outboard of the fences. The test results for flap-retracted and flap-deflected conditions indicated that the revised configuration should be satisfactory for most normal flight conditions provided the angle of attack does not exceed the angle for pitch-up. An abrupt pitch-up tendency of the model was evident for the zero thrust condition above approximately 15' angle of attack. Although the effects of power were destabilizing, power-on longitudinal stability was satisfactory through the angle-of-attack range for which the model was stable with zero thrust. Above the angle of attack for pitch-up, an uncontrollable left roll-off tendency would be expected with power on and slats retracted. Projection of wing slats or use of leading-edge chord-extensions with only the left extension drooped were found beneficial in controlling the roll-off tendency with power on; however the most effective means found was projection of only the left slat.

Beschreibung: Free-flight tests have been made to determine the zero-lift drag of several configurations of the XAAM-N-2 pilotless aircraft. Base-pressure measurements were also obtained for some of the configurations. The results show that increasing the wing-thickness ratio from 4 to 6 percent increased the wing drag by about 100 percent at M = 1.3 and by about 30 percent at M = 1.8. Increasing the nose fineness ratio from 5.00 to 6.25 reduced the drag coefficient of the wingless models a maximum of about 0.030 (10 percent) at M = 2.0. A corresponding change in nose shape for the winged models decreased the drag coefficient by about 0.05 in the Mach number range from 1.1 to 1.4; at Mach numbers greater than 1.6 no measurable reduction in drag coefficient was obtained. The drag of the present Sparrow fuselage is less than that of a parabolic fuselage which could contain the same equipment.

Beschreibung: A limited investigation of a 1/24-scale dynamically similar model of the Navy Bureau of Aeronautics DR-77 design was conducted in Langley tank no. 2 to determine the calm-water take-off and the rough-water landing characteristics of the design with particular regard to the take-off resistance and the landing accelerations. During the take-off tests, resistance, trim, and rise were measured and photographs were taken to study spray. During the landing tests, motion-picture records and normal-acceleration records were obtained. A ratio of gross load to maximum resistance of 3.2 was obtained with a 30 deg. dead-rise hydro-ski installation. The maximum normal accelerations obtained with a 30 deg. dead-rise hydro-ski installation were of the order of 8g to log in waves 8 feet high (full scale). A yawing instability that occurred just prior to hydro-ski emergence was improved by adding an afterbody extension, but adding the extension reduced the ratio of gross load to maximum resistance to 2.9.

Beschreibung: The stator-blade angles in the twelfth through fifteenth stages of a 16-stage axial-flow compressor were increased 3O. The over-all performance of this modified compressor is compared to the performance of the compressor with original blade angles. The matching characteristics of the modified compressor and a two-stage turbine were obtained and compared to those of the compressor with original blade angles and the same turbine.

Beschreibung: 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.

Beschreibung: 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.

Beschreibung: The study of the hydrodynamic properties of planing bottom of flying boats and seaplane floats is at the present time based exclusively on the curves of towing tests conducted in tanks. In order to provide a rational basis for the test procedure in tanks and practical design data, a theoretical study must be made of the flow at the step and relations derived that show not only qualitatively but quantitatively the inter-relations of the various factors involved. The general solution of the problem of the development of hydrodynamic forces during the motion of the seaplane float or flying boat is very difficult for it is necessary to give a three-dimensional solution, which does not always permit reducing the analysis to the form of workable computation formulas. On the other had, the problem is complicated by the fact that the object of the analysis is concerned with two fluid mediums, namely, air and water, which have a surface of density discontinuity between them. The theoretical and experimental investigations on the hydrodynamics of a ship cannot be completely carried over to the design of floats and flying-boat hulls, because of the difference in the shape of the contour lines of the bodies, and, because of the entirely different flow conditions from the hydrodynamic viewpoint.

Beschreibung: An impulse-momentum method for determining impact conditions for landing gears in eccentric landings is presented. The analysis is primarily concerned with the determination of contact velocities for impacts subsequent to initial touchdown in eccentric landings and with the determination of the effective mass acting on each landing gear. These parameters determine the energy-absorption requirements for the landing gear and, in conjunction with the particular characteristics of the landing gear, govern the magnitude of the ground loads. Changes in airplane angular and linear velocities and the magnitude of landing-gear vertical, drag, and side impulses resulting from a landing impact are determined by means of impulse-momentum relationships without the necessity for considering detailed force-time variations. The effective mass acting on each gear is also determined from the calculated landing-gear impulses. General equations applicable to any type of eccentric landing are written and solutions are obtained for the particular cases of an impact on one gear, a simultaneous impact on any two gears, and a symmetrical impact. In addition a solution is presented for a simplified two-degree-of-freedom system which allows rapid qualitative evaluation of the effects of certain principal parameters. The general analysis permits evaluation of the importance of such initial conditions at ground contact as vertical, horizontal, and side drift velocities, wing lift, roll and pitch angles, and rolling and pitching velocities, as well as the effects of such factors as landing gear location, airplane inertia, landing-gear length, energy-absorption efficiency, and wheel angular inertia on the severity of landing impacts. -A brief supplementary study which permits a limited evaluation of variable aerodynamic effects neglected in the analysis is presented in the appendix. Application of the analysis indicates that landing-gear impacts in eccentric landings can be appreciably more severe than impacts in symmetrical landings with the same sinking speed. The results also indicate the effects of landing-gear location, airplane inertia, initial wing lift, side drift velocity, attitude, and initial rolling velocity on the severity of both initial and subsequent landing-gear impacts. A comparison of the severity of impacts on auxiliary gears for tricycle and quadricycle configurations is also presented.

Beschreibung: A wind-tunnel investigation has been conducted to determine the stability and control characteristics of a full-size model of the Hughes MX-904 missile. Aerodynamic characteristics of the complete model through moderate ranges of angles of attack and yaw, with an additional test made through an angle of attack of 180 degrees, are presented. The effects of horizontal tail deflection are also included.

Beschreibung: A free-flight 0.12-scale rocket-boosted model of the North American MX-770 (X-10) missile has been tested in flight by the Pilotless Aircraft Research Division of the Langley Aeronautical Laboratory. Drag, longitudinal stability, and duct performance data were obtained at Mach numbers from 0.8 to 1.7 covering a Reynolds number range of about 9 x 10(exp 6) to 24 x 10(exp 6) based on wing mean aerodynamic chord. The lift-curve slope, static stability, and damping-in-pitch derivatives showed similar variations with Mach number, the parameters increasing from subsonic values in the transonic region and decreasing in the supersonic region. The variations were for the most part fairly smooth. The aerodynamic center of the configuration shifted rearward in the transonic region and moved forward gradually in the supersonic region. The pitching effectiveness of the canard control surfaces was maintained throughout the flight speed range, the supersonic values being somewhat greater than the subsonic. Trim values of angle of attack and lift coefficient changed abruptly in the transonic region, the change being associated with variations in the out-of-trim pitching moment, control effectiveness, and aerodynamic-center travel in this speed range. Duct total-pressure recovery decreased with increase in free-stream Mach number and the values were somewhat less than normal-shock recovery. Minimum drag data indicated a supersonic drag coefficient about twice the subsonic drag coefficient and a drag-rise Mach number of approximately 0.90. Base drag was small subsonically but was about 25 percent of the minimum drag of the configuration supersonically.

Beschreibung: An investigation of a 1/24- scale dynamically similar model of the Douglas C-124 airplane was made to determine the ditching characteristics and proper technique for ditching the airplane. Various conditions of damage, landing attitude, flap setting, and speed were investigated. The behavior of the model was determined from visual observations, motion- picture records, and time-history deceleration records. The results of the investigation are presented in table form, photographs, and curves. It was concluded on the basis of results from model tests with scale-strength bottoms (equivalent to 1150 pounds per square foot, full scale) that the airplane should be ditched at a medium nose-high landing attitude (near 7deg) with flaps full down. The airplane will probably make a smooth run with considerable damage resulting to the fuselage bottom just forward of the wing, but it is not likely that the water inflow will be overwhelming to personnel provided they are not in the belly compartment. Longitudinal decelerations in calm water will be about 2 1/2g and the landing run will be about four fuselage lengths.

Beschreibung: 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.

Beschreibung: An application of airfoil design methods was used to design series of related turbine-blade profiles to satisfy the conditions of inlet flow angle and turning angle encountered in the usual range of turbine operation. A series of blade profiles applicable to most turbine blading requirements and a secondary series with particular reference to impulse conditions were designed. Five blade sections from these series ranging in mean-line turning angles from 63 deg. to 120 deg. were tested in low-speed cascade tunnels. From low-speed test results optimum blade angles of attack were selected at each test condition. The induced angle and the deviation angle of the flow were determined from the low-speed data. If these angles are known for the solidity and inlet angle of an application, the necessary camber is specified. A method of predicting high-speed pressure distributions from low-speed cascade test results is presented to extend the usefulness of the low-speed data. Sample high-speed tests of two of the five blade sections were made at Mach numbers up to the critical value. The results indicated satisfactory flow conditions in all of the blade passages tested.

Beschreibung: A ditching investigation of a model of the Convair-Liner airplane was made to observe the behavior and determine the safest procedure for making an emergency water landing. The ditching model was designed and constructed by the National Advisory Committee for Aeronautics. Design information on the airplane was furnished by the Consolidated Vultee Aircraft Corporation. A three-view drawing of the airplane is shown. The investigation was made in calm water at the Langley tank no. 2 monorail.

Beschreibung: An investigation has been made in the Langley gust tunnel with two identical airplane models approximating 1/40-scale models of the B-29, coupled in tandem with a boom so that the individual centers of gravity were equidistant from the single coupling joint at the tail of the lead airplane. Time histories of the boom joint load were obtained as the models were flown through a gust. The results indicate that on a similar configuration involving airplanes the size of B-29 airplanes a load on the boom joint of 10,000 to 14,000 pounds could be induced by encountering a gust of 50 feet per second and having a gradient distance of 17 chords, at a forward speed of 380 feet per second and that the total load is extremely sensitive to the steadiness of flight that can be maintained with or without a gust. It is felt that the results are probably satisfactory to show order of magnitude, but it does not appear possible that a precise determination of the joint load that would be applicable to the full-scale airplanes can be obtained by gust-tunnel tests.

Beschreibung: Flight tests have been conducted on rocket-propelled models of an airplane configuration incorporating a sweptback wing with inverse taper to investigate the drag, stability, and control characteristics at transonic and supersonic speeds. The models were tested with a conventional tail arrangement in the Mach number range from 0.55 to 1.2. In addition to the various aerodynamic parameters obtained, the flying qualities were computed for a full-scale airplane with the center-of-gravity location at 18 percent of the mean aerodynamic chord. Also, included in this investigation are drag measurements made on relatively simple fixed-control models tested with both conventional and V-tail arrangements.

Beschreibung: An elementary type of analysis has been used to determine the amount of wing tip that must be severed to produce irrevocable loss of control of a B-29 airplane. The remaining inboard structure of the Boeing B-29 wing has then been analyzed and curves are presented for the estimated reduction in structural strength due to four general types of damage produced by rod-type warhead fragments. The curves indicate the extent of structural damage required to produce a kill of the aircraft within 10 seconds.

Beschreibung: A systematic research program is being carried out in the Langley high-speed 7- by 10-foot tunnel to determine the aerodynamic characteristics of various arrangements of the component parts of research-type airplane models, including some complete model configurations. Data are being obtained on characteristics in pitch, sideslip, and during steady roll at Mach numbers from 0.40 to about 0.95. This paper presents results which show the effect of taper ratio on the aerodynamic characteristics in sideslip of wing-fuselage combinations having wings with a sweep of 45 degrees at the quarter-chord line, an aspect ratio of 4, and a NACA 65A006 airfoil section.

Beschreibung: 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.

Beschreibung: An experimental investigation of the variation of aileron rolling effectiveness and total drag with Mach number has been made using 1/6-scale rocket-propelled models of the Bell MX-776. Three models having constant-chordwise-thickness full-span aileron at approximate deflections of 2 deg, 5 deg, and 15 deg have been flown. Positive control effectiveness over the Mach number range between approximately 0.5 and 1.2 was obtained from the models and no indication of reversal of effectiveness was encountered. The ratio of tip helix angle to aileron deflection indicated a decrease in proportional rolling effectiveness with increasing deflections in the Mach number range from approximately 0.7 to 1.0. A drag rise of about 125 percent in the transonic region between Mach numbers of 0.85 and 1.02 followed by a gradual decrease at higher speeds was revealed.

Beschreibung: As part of a program to determine the feasibility of using a fighter airplane as a parasite in combination with a Consolidated Vultee RB-36 for long-range reconnaissance missions (project FICON), an experimental investigation has been made in the Langley free-flight tunnel to determine the dynamic stability and control characteristics of a 1/17.5-scale model of a Chance Vought F7U-3 airplane in several tow configurations. The investigation consisted of flight tests in which the model was towed from a strut in the tunnel by a towline and by a direct coupling which provided complete angular freedom. The tests with the direct coupling also included a study of the effect of spring restraint in roll in order to simulate approximately the proposed full-scale arrangement in which the only freedom is that permitted by the flexibility of the launching and retrieving trapeze carried by the-bomber. For the tow configurations in which a towline was used (15 and 38 feet full scale), the model had a very unstable lateral oscillation which could not be controlled. The stability was also unsatisfactory for the tow configuration in Which the model was coupled directly to the strut with complete angular freedom. When spring restraint in roll was added, however, the stability was satisfactory. The use of the yaw damper which increased the damping in yaw to about six times the normal value of the model appeared to have no appreciable effect on the lateral oscillations in the towline configurations, but produced a slight improvement in the case of the direct coupling configurations. The longitudinal stability was satisfactory for those cases in which the lateral stability was good enough to permit study of longitudinal motions.

Beschreibung: An investigation of a 1/24-scale dynamically similar model of the Boeing B-47 airplane was made to determine the ditching characteristics and proper ditching technique for the airplane. Various conditions of damage, landing attitude, flap setting, and speed were investigated. The behavior of the model was determined from visual observations, motion-picture records, and time-history deceleration records. The results of the investigation are presented in table form, photographs, and curves. The airplane should be ditched at the lowest speed and highest attitude consistent with adequate control; the flaps should be full down. The airplane will probably make a deep but fairly smooth run. The fuselage bottom will be damaged and partially filled with water; consequently, crew members should be assigned ditching stations near an exit in the upper or forward part of the fuselage. The nacelles may be expected to be torn away from the wing. In calm water the maximum decelerations will be about 3g and the landing run will be about 6 fuselage lengths.

Beschreibung: At the request of the Air Materiel Command, an investigation was made in the Langley free-flight tunnel to determine the longitudinal stability and control characteristics of models coupled together in a tandem configuration for aerial refueling similar to one proposed by the Douglas Aircraft Company, Inc. Static force tests were made with 1/20-scale models of the B-29 and F-80 airplanes to determine the effects of rigidly coupling the airplanes together. The Douglas configuration differs from the rigid configuration tested in that it provides for some freedom in pitch and vertical displacement. The force tests showed that, for the bomber alone, the aerodynamic center was 0.21 mean aerodynamic chord behind the center of gravity (stable) but that for the tandem configuration with rigid coupling the aerodynamic center was 0.28 mean aerodynamic chord forward of the center of gravity of the combination (unstable). This reduction in stability was caused by the downwash of the bomber on the fighter. The pitching moment produced by elevator deflection of the bomber was reduced approximately 50 percent by addition of the fighter. Some recent flight tests made in the free-flight tunnel on models in a similar tandem configuration indicated that, with a hinged coupling permitting freedom in pitch, the stability of the combination was better than that obtained with a rigid coupling and was about the same as that for the bomber alone.

Beschreibung: At the request of the Air Materiel Command an investigation was made in the Langley free-flight tunnel to determine the static longitudinal stability and control characteristics of models coupled together in a tandem configuration proposed by All American Airways, Inc. Force tests were made using 1/20-scale models of B-29 end F-80 airplanes to determine the effects of coupling the fighter to the tail of the bomber. The results of the investigation showed that for the bomber alone the aerodynamic center was 0.21 mean aerodynamic chord behind the center of gravity (stable) but that for the tandem configuration the aerodynamic center was 0.09 mean aerodynamic chord forward of the center of gravity, of the combination (unstable). The elevator effectiveness of the bomber was reduced approximately 50 percent by addition of the fighter. Some recent flight tests made in the free-flight tunnel with models simulating the proposed configuration indicate that the reduction in stability may be minimized by incorporating a hinged coupling permitting freedom in pitch.