ALBERT

Description: No abstract available

Description: The presence of radomes and instruments that are sensitive to water films or ice formations in the nose section of all-weather aircraft and missiles necessitates a knowledge of the droplet impingement characteristics of bodies of revolution. Because it is possible to approximate many of these bodies with an ellipsoid of revolution, droplet trajectories about an ellipsoid of revolution with a fineness ratio of 10 were computed for incompressible axisymmetric air flow. From the computed droplet trajectories, the following impingement characteristics of the ellipsoid surface were obtained and are presented in terms of dimensionless parameters: (1) total rate of water impingement, (2) extent of droplet impingement zone, and (3) local rate of water impingement. These impingement characteristics are compared briefly with those previously reported for an ellipsoid of revolution with a fineness ratio of 5.

Description: An investigation at a free-stream Mach number of 2.02 was made to determine the effects of a propulsive jet on a wing surface located in the vicinity of a choked convergent nozzle. Static-pressure surveys were made on a flat surface that was located in the vicinity of the propulsive jet. The nozzle was operated over a range of exit pressure ratios at different fixed vertical distances from the flat surface. Within the scope of this investigation, it was found that shock waves, formed in the external flow because of the presence of the propulsive jet, impinged on the flat surface and greatly altered the pressure distribution. An integration of this pressure distribution, with the location of the propulsive jet exit varied from 1.450 propulsive-jet exit diameters to 3.392 propulsive-jet exit diameters below the wing, resulted in an incremental lift for all jet locations that was equal to the gross thrust at an exit pressure ratio of 2.86. This incremental lift increased with increase in exit pressure ratio, but not so rapidly as the thrust increased, and was approximately constant at any given exit pressure ratio.

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: An investigation was made in the N.A.C.A. 7- by 10- foot wind tunnel to determine the aerodynamic section characteristics of an N. A. C. A. 23012 airfoil with a single main slotted flap equipped successively with auxiliary flaps of the plain, split, and slotted types. A test installation mas used in which an airfoil of 7-foot span was mounted vertically between the upper and the lower sides of the closed test section so that two-dimensional flow was approximated. On the basis of maximum lift coefficient, low drag at moderate and high lift coefficients, and high drag at high lift coefficients, the optimum combination of the arrangements was found to be the double slotted flap . All the auxiliary flaps tested, however, increased the magnitudes of the pitching moments over those of the main slotted flap alone.

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: The effects of several wing leading-edge camber and deflected-tip modifications on the force and moment characteristics of a 1/20-scale model of the Convair F-102 airplane have been determined at Mach numbers from 0.60 t o 1.14 for angles of attack up to 14 deg. in the Langley 8-foot transonic tunnel. The effects of elevator deflections from 0 deg. to -10 deg. were also obtained for a configuration incorporating favorable leading- edge and tip modifications. Leading-edge modifications which had a small amount of constant-chord camber obtained by vertically adjusting the thickness distribution over the forward (3.9 percent of the mean aerodynamic chord) portion of the wing were ineffective in reducing the drag at lifting conditions at transonic speeds. Leading edges with relatively large cambers designed to support nearly elliptical span load distributions at lift coefficients of 0.15 and 0.22 near a Mach number of 1.0 produced substantial reductions in drag at most lift coefficients.

Description: An investigation has been made in the Langley low-turbulence pressure tunnel of the aerodynamic characteristics of the NACA 0012, 64(sub 2)-015, and 64(sub 3)-018 airfoil sections. Data were obtained at Mach numbers from 0.3 to that for tunnel choke, at angles of attack from -2deg to 30deg, and with the surface. of each airfoil smooth-and with roughness applied at the leading edge.The Reynolds numbers of the tests ranged from 0.8 x 10(exp 6) to 4.4 x 10(exp 6). The results are presented as variations of lift, drag, and quarter-chord pitching-moment coefficients with Mach number.

Description: The effects of deflecting full-span, constant-chord, leading-edge flaps, having either round or sharp leading edges, upon the lift, drag,. and pitching moment characteristics of a model of an interceptor-type aircraft have been determined experimentally at subsonic and supersonic speeds. Results indicate that the variations of lift with angle of attack and of pitching moment with lift were unaffected by either the shape of the flap leading edge or flap deflection. Deflection of the flaps having either a round or sharp leading edge increased the drag at zero lift at both subsonic and supersonic speeds. In spite of the increase in the drag at zero lift, however, deflection of the flaps increased the maximum lift-drag ratio at subsonic speeds and had no deleterious effect at supersonic speeds.

Description: Transfer functions descriptive of the response of most engine variables were determined from transient data that were obtained from approximate step inputs in fuel flow and in exhaust-nozzle area. The speed responses of both spools to fuel flow and to turbine-inlet temperature appeared as identical first-order lags. Response to exhaust-nozzle area was characterized by a first-order lag response of the outer-spool speed, accompanied by virtually no change in inner-spool speed.

Description: The static lateral- and directional-stability characteristics of a high-speed fighter-type airplane, obtained from wind-tunnel tests of a model, are presented. The model consisted of a thin, unswept wing of aspect ratio 2.3 and taper ratio 0.385, a body, and a horizontal tail mounted in a high position on a vertical tail. Rolling-moment, yawing moment, and cross-wind-force coefficients are presented for a range of sideslip angles of -5 deg. to +5 deg, for Mach numbers of 0.90, 1.45, and 1.90. Data are presented which show the effects on the lateral and directional stability of: (1) component parts of the complete model, (2) modification of the empennage so as to provide different heights of the horizontal tail above the wing plane, (3) angle of attack, and (4) dihedral of the wing.

Description: An investigation to determine the altitude performance of the J57-P-1 turbojet engine and components was conducted at the NACA Lewis altitude wind tunnel. Data were obtained over a corrected inboard rotor speed range from 56 to 106 percent of rated speed, with intercompressor bleeds both open and closed, at altitudes from 15,000 to 50,000 feet and at a flight Mach number of 0.81. The corresponding range of Reynolds number indices was from 0.858 to 0.213. All data presented were obtained with a fixed-area exhaust nozzle sized according to the manufacturer's specification. Over-all engine performance parameters are presented as functions of inboard rotor speed corrected on the basis of engine inlet temperature. Component parameters are presented as functions of their respective corrected rotor speeds. A tabulation of all performance data is included in addition to the graphical presentation. Corrected net thrust is unusually sensitive to changes in corrected inboard rotor speed in the high speed region. A change of 1 percent in speed, at sated speed, produced a change of 6 percent in corrected net thrust . At rated engine speed, increasing the altitude from 15,000 to 50,000 feet at a constant flight Mach number of 0.81 increased the specific fuel consumption 13 percent but did not affect corrected net thrust.

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 made of a 1/10-scale dynamically similar model of the North American F-86 airplane to study its behavior when ditched. The model was landed in calm water at the Langley tank no. 2 monorail. Various landing attitudes, speeds, and conditions of damage were simulated. The behavior of the model was determined from visual observations, acceleration records, and motion-picture records of the ditchings. Data are presented in tabular form, sequence photographs, and time-history acceleration curves. From the results of the investigation it was concluded that the airplane should be ditched at the nose-high, 14 deg attitude to avoid the violent dive which occurs at the 4 deg attitude. The flaps and leading-edge slats should be fully extended to obtain the lowest possible landing speed. The wing tanks should be jettisoned to avoid the undesirable behavior which occurs with the tanks attached. In a calm-water ditching under these conditions the airplane will run smoothly for about 600 feet. Maximum longitudinal and vertical decelerations of about 3g will be encountered.

Description: An investigation is being conducted to determine the dynamic stability and control characteristics of a 0.13-scale flying model of Convair XFY-1 vertically rising airplane. This paper presents the results of flight and force tests to determine the stability and control characteristics of the model in vertical descent and landings in still air. The tests indicated that landings, including vertical descent from altitudes representing up to 400 feet for the full-scale airplane and at rates of descent up to 15 or 20 feet per second (full scale), can be performed satisfactorily. Sustained vertical descent in still air probably will be more difficult to perform because of large random trim changes that become greater as the descent velocity is increased. A slight steady head wind or cross wind might be sufficient to eliminate the random trim changes.

Description: A supplementary investigation was conducted in the Langley 20-foot free-spinning tunnel on a 1/24-scale model of the Grumman F9F-6 airplane. The primary purpose of the investigation was to reevaluate the spin-recovery characteristics of the airplane in view of the fact that the ailerons had been eliminated from the flaperon-aileron lateral control system of the airplane. A spin-tunnel investigation on a model of the earlier version of the F9F-6 airplane had indicated that use of ailerons with the spin (stick right in a right spin) was essential to insure recovery. The results indicate that with.ailerons eliminated, it may be difficult to obtain an erect developed spin but if a fully developed spin is obtained on the airplane, recovery therefrom may be difficult or impossible. Flaperon deflection should have little effect on spins or recoveries.

Description: An experimental investigation has been conducted to determine the stability and control characteristics of a 0.13-scale free-flight model of the Convair XFY-1 airplane during take-offs and landings in steady winds. The tests indicated that take-offs in headwinds up to at least 20 knots (full scale) will be fairly easy to perform although the airplane may be blown downstream as much as 3 spans before a trim condition can be established. The distance that the airplane will be blown down-stream can be reduced by restraining the upwind landing gear until the instant of take-off. The tests also indicated that spot landings in headwinds up to at least 30 knots (full scale) and in crosswinds up to at least 20 knots (full scale) can be accomplished with reasonable accuracy although, during the landing approach, there will probably be an undesirable nosing-up tendency caused by ground effect and by the change in angle of attack resulting from vertical descent. Some form of arresting gear will probably be required to prevent the airplane from rolling downwind or tipping over after contact. This rolling and tipping can be prevented by a snubbing line attached to the tip of the upwind' wing or tail or by an arresting gear consisting of a wire mesh on the ground and hooks on the landing gear to engage the mesh.

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 effect on drag of positioning symmetrically mounted Douglas Aircraft Company, Inc. stores in pairs on a parabolic fuselage of fineness ratio 10.0 has been determined by flight tests of rocket-propelled, zero-lift models through a range of Mach number from 0.9 to 1.8. The stores were mounted in half-submerged positions and on pylons and were tested in three longitudinal locations on the fuselage with the forward position being located at the maximum diameter of the fuselage. The effects on drag of removing the half-submerged stores or extending them outward on pylons also was investigated by tests of models with half-submerged-store cavities on the fuselage. Two pylons differing in airfoil section and thickness were tested at the forward position of the stores on the fuselage with cavities. The half-submerged stores gave the smallest drag increments, which were approximately equal regardless of their respective longitudinal locations. Removing the half-submerged stores to expose the cavities increased the drag increments from two to three times. For the pylon-mounted stores, the store in the midposition had less drag than in the forward or rear positions at supersonic speeds. Adding the half-submerged-store cavities to the pylon-mounted-store configurations reduced the drag at the rear position between Mach numbers 0.95 and 1.50 and increased the drag at the midposition throughout the speed range. Changing from the 6-percent-thick flat pylon to the 10-percent-thick airfoil pylon increased the total drag slightly above Mach number 1.10. Good agreement was obtain& between the experimental and theoretical interference drag coefficients for the pylon-mounted stores (without fuselage cavities} in the three longitudinal locations tested at Mach numbers 1.2 and 1.5.

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: 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. 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 crossflow drag coefficient along the body and the effect of the wing downwash field on the afterbody loading.

Description: Preliminary information on the complex subject of the fatigue strength of fabricated structural members for aircraft is presented in the test results obtained on several different types of airship girders subjected to axial tension and compression in a resonance fatigue machine. A description of this machine as well as numerous photographs of the fatigue failures are given. There is also presented an extended bibliography on the subject of fatigue strength.

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: A flight test has been conducted to determine the longitudinal stability and control,characteristics of a 0.133-scale model of the Consolidated Vultee XFY-1 airplane without propellers for the Mach number range between 0.73 and 1.19.

Description: An investigation was conducted in the Langley high-speed 7- by 10-foot tunnel to determine effects of modifications to a bomb model (particularly with regard to drag) when mounted on a wing-fuselage model and tested at Mach numbers from 0.70 to 1.10. In addition, the static longitudinal stability characteristics of several configurations of a larger scale model of the bomb alone were obtained over a Mach number range from 0.50 to 0.95. The results obtained for the wing-fuselage-bomb model indicate that large reductions in installation drag were obtained for the wing-fuselage-bomb model when the flat nose of the basic bomb was replaced by rounded or pointed noses of various calibers. Shortening the mounting pylon gave further decreases in the installation drag. The tests of the bomb alone indicated that only the flat-nose configurations were stable over the greater part of the Mach number range. Nose-shape modifications which improved the drag also caused the bombs to become unstable at low angles of attack. The stability of the low-drag bomb configurations could be improved by lengthening the cylindrical portion of the body behind the center of gravity.

Description: An investigation has been conducted in the Langley 8-foot transonic tunnel to determine the effects of several fuselage modifications on the transonic drag-rise characteristics of a 1/20-scale model of the Convair F-102 airplane. Tests covered an angle-of-attack range from 0deg to about 10deg and a Mach number range from 0.60 to 1.14. Results indicated that the transonic drag rise .for the basic F-102 airplane could be substantially reduced by extending the fuselage after-body approximately 8 percent of the fuselage length. Tests of other bodies indicated that a shorter (4-percent) afterbody extension may have a similar effect on the drag rise. Further improvement of the axial cross-sectional-area distribution of the 8-percent extended configuration through the addition of fuselage volume resulted in additional reductions in the drag rise at a Mach number of 1.0 and caused no or only slight drag penalties at the higher Mach numbers. The results of the present tests generally substantiate the area-rule concept with respect to the prediction of the transonic drag rise through the use of an equivalent-area body of revolution for a practical delta-wing airplane configuration.

Description: An experimental investigation has been conducted to determine the dynamic stability and control characteristics in hovering and transition flight of a 0.13-scale flying model of the Convair XFY-1 vertically rising airplane with the lower vertical tail removed. The purpose of the tests was to obtain a general indication of the behavior of a vertically rising airplane of the same general type as the XFY-1 but without a lower vertical tail in order to simplify power-off belly landings in an emergency. The model was flown satisfactorily in hovering flight and in the transition from hovering to normal unstalled forward flight (angle of attack approximately 30deg). From an angle of attack of about 30 down to the lowest angle of attack covered in the flight tests (approximately 15deg) the model became progressively more difficult to control. These control difficulties were attributed partly to a lightly damped Dutch roll oscillation and partly to the fact that the control deflections required for hovering and transition flight were too great for smooth flight at high speeds. In the low-angle-of-attack range not covered in the flight tests, force tests have indicated very low static directional stability which would probably result in poor flight characteristics. It appears, therefore, that the attainment of satisfactory directional stability, at angles of attack less than 10deg, rather than in the hovering and transition ranges of flight is the critical factor in the design of the vertical tail for such a configuration.

Description: A low-speed wind-tunnel investigation has been made of some aspects of the aerodynamic problems associated with the use of air-to-air missiles when carried externally on aircraft. Measurements of the forces and moments on a missile model for a range of positions under the mid-semispan location of a 45deg sweptback wing indicated longitudinal and lateral forces with regard to both carriage and release of the missiles. Surveys of the characteristics of the flow field in the region likely to be traversed by the missiles showed abrupt gradients in both flow angularity and in local dynamic pressure. Through the use of aerodynamic data on the isolated missile and the measured flow-field characteristics, the longitudinal forces and moments acting on the missile while in the presence of the wing-fuselage combination could be estimated with fair accuracy. Although the lateral forces and moments predicted were qualitatively correct, there existed some large discrepancies in absolute magnitude.

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: Experimental results showing the static longitudinal-stability and control characteristics of a model of a fighter airplane employing a low-aspect-ratio unswept wing and an all-movable horizontal tail are presented. The investigation was made over a Mach number range from 0.60 to 0.90 and from 1.35 to 1.90 at a constant Reynolds number of 2.40 million, based on the wing mean aerodynamic chord. Because of the location of the horizontal tail at the tip of the vertical tail, interference was noted between the vertical tail and the horizontal tail and between the wing and the horizontal tail. This interference produced a positive pitching-moment coefficient at zero lift throughout the Mach number range of the tests, reduced the change in stability with increasing lift coefficient of the wing at moderate lift coefficients in the subsonic speed range, and reduced the stability at low lift coefficients at high supersonic speeds. The lift and pitching-moment effectiveness of the all movable tail was unaffected by the interference effects and was constant throughout the lift-coefficient range of the tests at each Mach number except 1.90.

Description: This paper presents the results of an investigation of the dynamic stability and controllability of a model which approximately represents the Lockheed XFV-1 airplane to a 1/8 scale. The investigation consisted of hovering flights in still air at a considerable height above the ground, hovering flights very close to the ground, vertical take-offs and landings, flights through the transition range from hovering to normal forward flight, and sideways translational flights. The model could be flown smoothly and easily in hovering flight despite the fact that the uncontrolled pitching and yawing motions were unstable oscillations. There was a noticeable reduction in the controllability of the model when hovered very close to the ground but take-offs could be made easily and landings on a g,ven spot could be made accurately in spite of this adverse ground effect. Flights through the transition range from hovering to normal forward flight could be performed fairly easily. The model seemed to have stability of angle of attack and angle of roll over most of the transition range. The yawing motion was divergent in the very high angle-of-attack range but could be controlled easily. At the lower angles of attack, the model seemed to become stable in yaw. In sideways flight there was an increasingly strong tendency to diverge in roll as the speed was increased and finally, at a speed of about 25 knots (full scale), the model rolled off despite efforts of the pilot to control it.

Description: An investigation has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 1.41 to determine the static stability and control and drag characteristics of a l/l5-scale model of the Grunman F9F-9 airplane. The effects of alternate fuselage shapes, wing camber, wing fences, and fuselage dive brakes on the aerodynamic characteristics were also investigated. These tests were made at a Reynolds number of 1.96 x l0 (exp 6) based on the wing mean aerodynamic chord of 0.545 foot. The basic configuration had a static margin of stability of 38.4 percent of the mean aerodynamic chord and a minimum drag coefficient of 0.049. For the maximum horizontal tail deflection investigated (-l0 deg), the maximum trim lift coefficient was 0.338. The basic configuration had positive static lateral stability at zero angle of attack and positive directional control throughout the angle-of-attack range investigated up to ll deg.

Description: The longitudinal stability and control characteristics of a 1/30-scale model of the Republic XF-103 airplane were investigated in the Langley 8-foot transonic tunnel. The effect of speed brakes located at the end of the fuselage was also investigated. The main part of the investigation was made with internal flow in the model, but some data were obtained with no internal flow. The longitudinal stability and control at transonic-speeds appeared satisfactory. The transonic drag rise was small. The speed brakes had no adverse effects on longitudinal stability.

Description: An investigation to determine the steady-state and surge characteristics of the J57-P-1 two-spool turbojet engine with various inlet air-flow distortions was conducted in the altitude wind tunnel at the NACA Lewis laboratory. Along with a uniform inlet total-pressure distribution, one circumferential and three radial pressure distortions were investigated. Data were obtained over a complete range of compressor speeds both with and without intercompressor air bleed at a flight Mach number of 0.8 and at altitudes of 35,000 and 50,000 feet. Total-pressure distortions of the magnitudes investigated had very little effect on the steady-state operating line for either the outer or inner compressor. The small radial distortions investigated also had engine over that obtained with the uniform inlet pressure distribution. The circumferential distortion, however, raised the minimum speed at which the engine could operate without encountering surge when the intercompressor bleeds were closed. This increase in minimum speed resulted in a substantial reduction in the operable speed range accompanied by a reduction in the altitude operating limit.

Description: The performance and operational characteristics of the J71-A2 turbojet-engine afterburner were investigated for a range of altitudes from 23,000 to 60,000 feet at a flight Mach number of 0,9 and at flight Mach numbers of 0.6, 0.9, and 1.0 at an altitude of 45,000 feet. The combustion performance and altitude operational limits, as well as the altitude starting characteristics have been determined.

Description: Tests have been made in the N.A.C.A. 8-foot high-speed wind tunnel of the drag caused by roughness on the surface of an airfoil of N.A.C.A. 23012 section and 5-foot chord. The tests were made at speeds from 80 t o 500 miles per hour at lift coefficients from 0 to 0.30. For conditions corresponding to high-speed flight, the increase in the drag was 30 percent of the profile drag of the smooth airfoil for the roughness produced by spray painting and 63 percent for the roughness produced. by 0.0037-inch carborundum grains. About one-half the drag increase was caused by the roughness on the forward one-fourth of the airfoil. Sandpapering the painted surface with No. 400 sandpaper made it sufficiently smooth that the drag was no greater than when the surface was polished. In the lower part of the range investigated the drag due to roughness increased rapidly with Reynolds Number.

Description: Tests have been made in the NACA 8-foot high-speed wind tunnel of the drag caused by four types of lap joint. The tests were made on an airfoil of NACA 23012 section and 5-foot chord and covered in a range of speeds from 80 to 500 miles per hour and lift coefficients from 0 to 0.30. The increases in profile drag caused by representative arrangements of laps varied from 4 to 9%. When there were protruding rivet heads on the surface, the addition of laps increased the drag only slightly. Laps on the forward part of a wing increased the drag considerably more than those farther back.

Description: Tests were made in the NACA 8-foot high speed wind tunnel of a metal-covered, riveted, 'service' wing of average workmanship to determine the aerodynamic effects of the manufacturing irregularities incident to shop fabrication. The wing was of 5-foot chord and of NACA 23012 section and was tested in the low-lift range at speeds from 90 to 450 miles per hour corresponding to Reynolds numbers from 4,000,000 to 18,000,000. At a cruising condition the drag of the service wing was 46% higher than the drag of a smooth airfoil, whereas the drag of an accurately constructed airfoil having the same arrangement of 3/32-inch brazier-head rivets and lap joints showed a 29% increase. The difference, or 17% of the smooth-wing drag, is apparently the drag caused by the manufacturing irregularities: sheet waviness, departures from true profile, and imperfect laps. the service wing, for one condition at least, showed a drag increase due to compressibility at a lower air speed than did the more accurate airfoil.

Description: Tests have been conducted in the NACA 8-foot high-speed wind tunnel to determine the effect of exposed rivet heads and spot welds on wing drag. Most of the tests were made with an airfoil of 5-foot chord. The air speed was varied from 80 to 500 miles per hour and the lift coefficient from 0 to 0.30. The increases in the drag of the 5-foot airfoil varied from 6%, due to countersunk rivets, to 27%, due to 3/32-inch brazier-head rivets, with the rivets in a representative arrangement. The drag increases caused by protruding rivet heads were roughly proportional to the height of the heads. With the front row of rivets well forward, changes in spanwise pitch had negligible effects on drag unless the pitch was more than 2.5% of the chord. Data are presented for evaluating the drag reduction attained by removing rivets from the forward part of the wing surface; for example, it is shown that over 70% of the rivet drag is caused by the rivets on the forward 30% of the airfoil in a typical case.

Description: Wing ducts for liquid-cooled engine radiators have been investigated in the N.A.C.A. full-scale wind tunnel on a large model airplane. Th e tests were made to determine the relative merits of several types of duct and radiator installations for an airplane of a particular des ign. In the test program the principal duct dimensions were system atically varied, and the results are therefore somewhat applicable to the general problems of wing duct design, although they should be co nsidered as preliminary and only indicative of the inherent possibil ities.

Description: An investigation was made in the NACA 5-foot vertical wind tunnel of a large variety of duct inlets and outlets to obtain information relative to their design for the cooling or the ventilation systems on aircraft. Most of the tests were of openings in a flat plate but, in order to determine the best locations and the effects of interference, a few tests were made of openings in an airfoil. The best inlet location for a system not including a blower was found to be at the forward stagnation point; for one including a blower, the best location was found to be in the region of lowest total head, probably in the boundary layer near the trailing edge. Design recommendations are given, and it is shown that correct design demands a knowledge of the external flow and of the internal requirements in addition to that obtained from the results of the wind tunnel tests.

Description: The systematic investigation of wing cooling ducts at the NACA laboratory has been continued with tests in the full-scale wind tunnel on ducts of finite span. These results extend the previous investigation on section characteristics of ducts to higher Reynolds numbers and indicate the losses due to the duct ends. The data include comparisons between ducts completely within the ring and the conventional underslung ducts. Methods of flow regulation were studied and data were obtained for a wide range of internal duct resistance. The results show satisfactory correlation between the finite span and the previously measured section characteristics obtained with full-span ducts. The effects of the various design parameters on the duct characteristics are discussed. The cooling power required for the internal duct installation is shown to be only a small percentage of the engine power.

Description: An investigation was made in the NACA 7- by 10-foot wind tunnel of a large-chord wing model with a duct to house a simulated radiator suitable for a liquid-cooled engine. The duct was expanded to reduce the radiator losses, and the installation of the duct and radiator was made entirely within the wing to reduce form and interference drag. The tests were made using a two-dimensional flow set-up with a full-span duct and radiator. Section aerodynamic characteristics of the basic airfoil are given and also curves showing the characteristics of the various duct-radiator combinations. An expression for efficiency, the primary criterion of merit of any duct, and the effect of the several design parameters of the duct-radiator arrangement are discussed. The problem of throttling is considered and a discussion of the power required for cooling is included. It was found that radiators could be mounted in the wing and efficiently pass enough air for cooling with duct outlets located at any point from 0.25c to 0.70c from the wing leading edge on the upper surface. The duct-inlet position was found to be critical and, for maximum efficiency, had to be at the stagnation point of the airfoil and to change with flight attitude. The flow could be efficiently throttled only by a simultaneous variation of duct inlet and outlet sizes and of inlet position. It was desirable to round both inlet and outlet lips. With certain arrangements of duct, the power required for cooling at high speed was a very low percentage of the engine power.

Description: The torsional deflection of the blades of three full-scale duralumin propellers operating under various loading conditions was measured by a light-beam method. Angular bending deflections were also obtained as an incidental part of the study. The deflection measurements showed that the usual present-day type of propeller blades twisted but a negligible amount under ordinary flight conditions. A maximum deflection of about 1/10th of a degree was found at V/nD of 0.3 and a smaller deflection at higher values of V/nD for the station at 0.70 radius. These deflections are much smaller than would be expected from earlier tests, but the light-beam method is considered to be much more accurate than the direct-reading transit method used in the previous tests.

Description: At the request of the Bureau of Aeronautics, Department of the Navy, National Advisory Committee for Aeronautics has conducted a preliminary investigation at high subsonic speeds of the static longitudinal and lateral stability characteristics of a 0.05-scale model of the Convair F2Y-1 water-based fighter airplane. The tests covered a Mach number range from 0.5 to 0.94 and corresponding Reynolds numbers, based on the wing mean aerodynamic chord, from 3.3 x 10(exp 6) to 4.3 x 10(exp 6). The maximum angle-of-attack range (obtained at the lower Mach numbers) was from -2 degrees to 25 degrees. Sideslip angles from -4 degrees to 12 degrees also were investigated. The investigation included effects of various arrangements of wing fences and of rocket packages.

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: A 1/5-scale, rocket-propelled model of the Convair F-102 configuration was tested in free flight to determine zero-lift drag at Mach numbers up to 1.34 and at Reynolds numbers comparable to those of the full-scale airplane. This large-scale model corresponded to the prototype airplane and had air flow through the duct. Additional zero-lift drag tests involved a series of small equivalent bodies of revolution which were launched by means of a helium gun. The several small-scale models tested corresponded to: the basic configuration, the 1/5-scale rocket-propelled model configuration, a 2-foot (full-scale) fuselage-extension configuration, and a 7-foot (full-scale) fuselage-extension configuration. Models designed to correspond to the area distribution at a Mach number of 1.0 were flown for each of these 'shapes and, in addition, models designed to correspond to the area distribution at a Mach number of 1.2 were flown for the 1/5-scale rocket-propelled model and the 7-foot-fuselage-extension configuration. The value of external pressure drag coefficient (including base drag) obtained from the large-scale rocket model was 0.0190 at a Mach number of 1..05 and the corresponding values from the equivalent-body tests varied from 0.0183 for the rocket-propelled model shape to 0.0137 for the 7-foot-fuselage-extension configuration. From the results of tests of equivalent bodies designed to correspond to the area distribution at a Mach number of 1.0, it is evident that the small changes in shape incorporated in the basic and 2-foot-fuselage-extension configurations from that of the rocket-propelled model configuration will provide no significant change in pressure drag. On the other hand, the data from the 7-foot-fuselage-extension model indicate a substantial reduction in pressure drag at transonic speeds.

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: The first four stages were found to cause a major part of the poor low-speed efficiency of this compressor. The low design-speed over-all pressure ratio at surge was caused by the first and the twelfth to fifteenth stages. The multiple over-all performance curves in the intermediate-speed range were at least partly the result of double-branched characteristic curves for the third and seventh stages.

Description: With some airplanes the approach to the stall is accompanied by changes in the behavior, such as tail buffeting or changes in the control characteristics of the airplane so that the pilot obtains a warning of the impending stall. Vith other airplanes it is possible to approach the stall without any perceptible warning other than the reading of the air-speed meter, in which case the danger of inadvertent stalling is considerably greater. Although it is not within the scope of this paper to discuss stalling characteristics, it is desired to point out that in general the danger of inadvertent stalling is greatest with those airplanes that behave worse when the stalling occurs; that is, with airplanes in which the stall starts at the wing tips. A warning of the impending stall is desirable in any case, but is particularly desirable with airplanes of the latter type.

Description: The transonic longitudinal aerodynamic characteristics of a 0.0858-scale model of the Lockheed XF-104 airplane have been obtained from tests at the Langley 16-foot transonic tunnel. The results of the investigation provide some general information applicable to the transonic properties of thin, low-aspect-ratio, unswept wing configurations utilizing a high horizontal tail . The model employs a horizontal tail mounted at the top of the vertical tail and a wing with an aspect ratio of 2.5, a taper ratio of 0.385, and 3.4-percent-thick airfoil sections. The lift, drag, and static longitudinal pitching moment were measured at Mach numbers from 0.80 t o 1.09 and angles of attack from -2.5 deg to 22.5 deg. Some of the dynamic longitudinal stability properties of the airplane have been predicted from the test results. In addition, some visual flow studies on the wing surfaces obtained at Mach numbers of 0.80 and 1.00 are included. Results of the investigation show that the transonic rise in drag coefficient at zero lift is about 0.030. At high angles of attack, the model becomes longitudinally unstable at Mach numbers from 0.80 t o 0.90, whereas a reduction in static stability is experienced when very high angles of attack are reached at Mach numbers above 0.90. Longitudinal dynamic stability calculations show that the longitudinal control is good at angles of attack below the unstable break in the static pitching-moment curves, but a typical corrective control applied after the occurrence of neutral stability has little effect in averting pitch-up.

Description: The effects of elevator deflections from 0deg to -20deg on the force and moment characteristics of a 1/20-scale model of the Convair F-102 airplane with chordwise fences have been determined a t Mach numbers from 0.6 to 1.1 for angles of attack up to 20deg in the Langley 8-foot transonic tunnel. The configuration exhibited static longitudinal stability throughout the range tested, although a mild pitch-up tendency was indicated a t Mach numbers from 0.85 to 0.95. Elevator pitch effectiveness decreased rapidly between the Mach numbers of 0.9 and 1.0, however, no complete loss or reversal was indicated for all conditions tested. Because of the type of longitudinal control used, trimming the configuration from the zero elevator condition resulted in substantial decreases in lift-curve slope and maximum lift-drag ratio and increases in drag due to lift. The drag at zero lift, drag due to lift, and trim drag were high for this configuration.

Description: A 1/10-scale rocket model of the Lockheed XF-104 with faired inlets has been flown over a Mach number range from 0.80 to 1.45 to determine low-lift drag and a limited amount of stability data. The center-of-gravity locations were 4.0 and 1.5 percent of the mean aerodynamic chord before and after sustainer firing, respectively. Oscillations induced by pulse rockets were used to determine stability data. The external transonic drag coefficient increased from a value of 0.0160 at Mach number 0.80 to a maximum of 0.0432 near Mach number 1-13, with a drag rise Mach number of about 0.93. At Mach numbers where it could be determined, the model exhibited stable dynamic and static stability characteristics at low lift.

Description: A flight test has been conducted to determine the longitudinal stability and control characteristics of a 0.133-scale model of the Consolidated Vultee XFY-1 airplane with windmilling propellers for the Mach number range between 0.70 and 1.13. The variation of lift-curve slope C(sub L(sub alpha) with Mach number was gradual with a maximum value of 0.074 occurring at a Mach number of 0.97. Propellers had little effect upon the values of lift-curve slope or the linearity of lift coefficient with angle of attack. At lift coefficients between approximately 0.25 and 0.45 with an elevon angle of approximately -l0 deg, there was a region of neutral longitudinal stability at Mach numbers below 0.93 introduced by the addition of windmilling propellers. Below a lift coefficient of 0.10 and above a lift coefficient of 0.45, the model was longitudinally stable throughout the Mach number range of the test. There was a forward shift in the aerodynamic center of about 3-percent mean aerodynamic chord introduced by the addition of propellers. The aerodynamic center as determined at low lift moved gradually from a value of 28.5-percent mean aerodynamic chord at a Mach number of 0.75 to a value of 47-percent mean aerodynamic chord at a Mach number of 1.10. There was an abrupt decrease in pitch damping between Mach numbers of 0.88 and 0.99 followed by a rapid increase in damping to a Mach number of 1.06. The propellers had little effect upon the pitch damping characteristics . The transonic trim change was a large pitching-down tendency with and without windmilling propellers. The elevons were effective pitch controls throughout the speed range; however, their effectiveness was reduced about 50 percent at supersonic speeds. The propellers had no appreciable effect upon the control effectiveness.

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.

Description: Profile-drag coefficients published from tests in the N.A.C.A. variable-density tunnel (Technical Reports Nos. 460, 537, 586, and 610, references 1 to 4) have tended to appear high as compared with results from the N.A.C.A. full-scale tunnel (Technical Report No. 530, reference 5) and from foreign sources (references 6 to 8). Such discrepancies were considered in Technical Report No. 586, and corrections for turbulence and tip effects were derived that tended to reduce the profile-drag coefficients, particularly for the thicker airfoils. The corrected profile-drag coefficients, designated by the lower-case symbol cdo as contrasted with the older CDO, have been employed in the airfoil reports published since Technical Report No. 460, but even these corrected results continued to appear high, particularly for the thicker sections. The important practical result is that a smaller increase of drag with airfoil thickness is indicated, which may be of primary importance to the airplane designer in choosing the optimum airfoil sections for actual wings. Further investigations of this subject were, of course, undertaken, one of the most important being an investigation of three symmetrical sections N.A.C A. 0009, 0012, and 0018 under conditions of low turbulence in the full-scale tunnel. Preliminary results from this investigation also indicate a smaller increase in drag with airfoil thickness than the results from the variable-density tunnel. Furthermore, comparative tests made in the two tunnels by applying strings to the surface of the N.A.C.A. 0012 airfoil to move the transition point to a predetermined position indicated that the effective reynolds Number concept would account approximately for the drag as affected by the position of transition from laminar to turbulent flow in the boundary layer.