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The preparation and physical properties of several aliphatic hydrocarbons and intermediates
No Description Available.
Recovery of Alumina From Kaolin by the Lime-Soda Sinter Process
Report issued by the U.S. Bureau of Mines on exploration of kaolin clay deposits for alumina. Properties of the kaolin clay and alumina recovered are listed. This report includes tables, graphs, and illustrations.
Rocket Power Plants Based on Nitric Acid and their Specific Propulsive Weights
Two fields are reserved for the application of rocket power plants. The first field is determined by the fact that the rocket power plant is the only type of power plant that can produce thrust without dependence upon environment. For this field,the rocket is therefore the only possible power plant and the limit of what may be done is determined by the status of the technical development of these power plants at the given moment. The second field is that in which the rocket power plant proves itself the most suitable as a high-power drive in free competition with other types of power plants. The exposition will be devoted to the demarcation of this field and its division among the various types of rocket power plants.
Some considerations of the lateral stability of high-speed aircraft
No Description Available.
Some investigations of the general instability of stiffened metal cylinders VIII : stiffened metal cylinders subjected to pure torsion
An experimental investigation of the general instability of reinforced thin-walled metal cylinders was carried out at the California Institute of Technology. The basic parameters involved were the spacing and sectional properties of the stiffening elements, the wall thickness, and the diameter of the cylinder. An analysis of the experimental data led to a suitable parameter for estimating the general instability stress of reinforced metal cylinders when subjected to pure torsion loading.
Stresses in and general instability of monocoque cylinders with cutouts III : calculation of the buckling load of cylinders with symmetric cutout subjected to pure bending
No Description Available.
Tests of the NACA 64(SUB 1)A212 airfoil section with a slat, a double slotted flap, and boundary-layer control by suction
No Description Available.
A transonic propeller of triangular plan form
No Description Available.
Two-dimensional wind-tunnel investigation of the NACA 64(sub 1)-012 airfoil equipped with two types of leading-edge flap
No Description Available.
Wind-tunnel investigation of the air load distribution on two combinations of lifting surface and fuselage
No Description Available.
Wind-tunnel investigation of the effects of surface-covering distortion on the characteristics of a flap having undistorted contour maintained for various distances ahead of the trailing edge
No Description Available.
Application of an ultraviolet spectrophotometric method to the estimation of alkylnaphthalenes in 10 experimental jet-propulsion fuels
No Description Available.
Fuel tests on an I-16 jet-propulsion engine at static sea-level conditions
No Description Available.
Investigation of carbon deposition in an I-16 jet-propulsion engine at static sea-level conditions
No Description Available.
A Preliminary Study of Ram-Actuated Cooling Systems for Supersonic Aircraft
An analysis has been made of the characteristics of several cooling cycles suitable for cockpit cooling of supersonic aircraft. All the cycles considered utilize the difference between dynamic and ambient static pressure to actuate the cooling system and require no additional power source. The results of the study indicate that as flight speeds become greater, increasingly complex systems are required to reduce the altitudes above approximately 35,000 feet, a system composed of an externally loaded expansion turbine in conjunction with a supersonic diffuser would maintain tolerable ventilating air temperature, at least up to a flight Mach number of 2. The most complex system considered,composed of compressor, intercooler, and expansion turbine with the intercooler cooling air decreased in temperature by expansion through an auxiliary turbine is capable of maintaining a ventilation air temperature less than ambient temperature up to a flight Mach number of 3.7.
An Investigation of Convergent-Divergent Diffusers at Mach Number 1.85
An investigation has been conducted in the Cleveland 18- by 18-inch supersonic tunnel at a Mach number of 1.85 and angles of attack from 0 deg to 5 deg to determine optimum design configurations for a convergent-divergent type of supersonic diffuser with a subsonic diffuser of 5 deg included divergence angle. Total pressure recoveries in excess of theoretical recovery across a normal shock at a free-stream Mach number of 1.85 wore obtained with several configurations. The highest recovery for configurations without a cylindrical throat section was obtained with an inlet having an included convergence angle of 20 deg. Insertion of a 2-inch throat section between a 10 deg included angle inlet and the subsonic diffuser stabilized the shock inside the diffuser and resulted in recoveries as high as 0.838 free-stream total pressure at an angle of attack of 0 deg, corresponding to recovery of 92.4 percent of the kinetic energy of the free air stream. Use of the throat section also lessened the reduction in recovery of all configurations due to angle of attack.
Wind-tunnel development of optimum double-slotted-flap configurations for seven thin NACA airfoil sections
No Description Available.
Effects of compressibility on the characteristics of five airfoils
Report presenting the results of pressure-distribution tests to determine the effects of compressibility on the characteristics of the NACA 66,2-215, 66,2-015, 65(216)-418, 16-212, and 23015 airfoil sections. Schileren photographs of the air flow and data on the wake characteristics was also obtained.
Flight observations of aileron flutter at high Mach numbers as affected by several modifications
No Description Available.
Air-flow behavior over the wing of an XP-51 airplane as indicated by wing-surface tufts at subcritical and supercritical speeds
No Description Available.
Measurements of Static and Total Pressure Throughout the Transonic Speed Range as Obtained from an Airspeed Head Mounted on a Freely Falling Body
No abstract available.
Partial Measurements in Flight of the Flying Qualities of a Grumman XF7F-1 Airplane with a Modified Vertical Tail
Partial measurements in flight of the handling qualities of a Grumman XF7F-1 airplane were investigated. Results are for low altitude and normal center of gravity conditions only.
A study of several parameters controlling the trajectories of a supersonic antiaircraft missile powered with solid- or liquid-fuel rockets
No Description Available.
The calculation of drag for airfoil sections and bodies of revolution at subcritical speeds
No Description Available.
Flight-test measurements of aileron control surface behavior at super critical Mach numbers
No Description Available.
Tests of Submerged Duct Installation on the Ryan FR-1 Airplane in the Ames 40- by 80-Foot Wind Tunnel
An investigation of an NACA submerged intake installation on the Ryan FR-1 was conducted to determine the full-scale aerodynamic characteristics of this installation. In addition, tests were conducted on the submerged inlet with revised entrance lips and deflectors to determine the configuration which would result in the best dynamic pressure recovery measured at the inlet for this installation without a major rework of the entrance. Stalling of the air flow over the inner lip surface created excessive dynamic pressure losses with the original entrance. The revised entrance produced a 12-percent increase in dynamic pressure recovery at the design high-speed inlet-velocity ratio and resulted in an improvement of thte critical-speed characteristics of the entrance lip. A complete redesign of the entrance including a decrease in ramp angle and adjustment of lip camber is necessary to secure optimum results from this submerged duct installation.
Correlation of the Trim Limits of Stability Obtained for a PB2Y-3 Flying Boat and a 1/8-Size Powered Dynamic Model
Tests of a PB2Y-3 flying boat were made at the U.S> Naval Air Station, Patuxent River, Md., to determine its hydrodynamic trim limits of stability. Corresponding tests were also made of a 1/8-size powered dynamic model of the same flying boat in Langley tank no. 1. During the tank tests, the full-size testing procedure was reproduced as closely as possible in order to obtain data for a direct correlation of the results. As a nominal gross load of 66,000 pounds, the lower trim limits of the full-size and model were in good agreement above a speed of 80 feet per second. As the speed decreased below 80 feet per second, the difference between the model trim limits and full-scale trim limits gradually became larger. The upper trim limit of the model with flaps deflected 0 deg was higher than that of the full-size, but the difference was small over the speed range compared. At flap deflections greater than 0 deg, it was not possible to trim either the model of the airplane to the upper limit with the center of gravity at 28 percent of the mean aerodynamic chord. The decrease in the lower trim limits with increase in flap deflection showed good agreement for the airplane and model. The lower trim limits obtained at different gross loads for the full-size airplane were reduced to approximately a single curve by plotting trim against the square root of C(sub delta (sub o)) divided by C(sub V).
Drag characteristics of rectangular and swept-back NACA 65-009 airfoils having various aspect ratios as determined by flight tests at supersonic speeds
No Description Available.
An investigation of the downwash at the probable tail location behind a high-aspect-ratio wing in the Langley 8-foot high-speed tunnel
No Description Available.
Cooling of Gas Turbines, IV - Calculated Temperature Distribution in the Trailing Part of a Turbine Blade Using Direct Liquid Cooling
A theoretical analysis of the temperature distribution through the trailing portion of a blade near the coolant passages of liquid cooled gas turbines was made. The analysis was applied to obtain the hot spot temperatures at the trailing edge and influence of design variables. The effective gas temperature was varied from 2000 degrees to 5000 degrees F in each investigation.
Langley full-scale-tunnel investigation of the factors affecting the static lateral-stability characteristics of a typical fighter-type airplane
The factors that affect the rate of change of rolling moment with yaw of a typical fighter-type airplane were investigated in the Langley full-scale tunnel on a typical fighter-type airplane.Eight representative flight conditions were investigated in detail. The separate effects of propeller operation, of the wing-fuselage combination, and of the vertical tail to the effective dihedral of the airplane in each condition were determined. The results of the tests showed that for the airplane with the propeller removed, the wing-fuselage combination had positive dihedral effect which increased considerably with increasing angle of attack for all conditions. Flap deflection decreased the dihedral effect of the wing-fuselage combination slightly as compared with that with the flaps retracted. Flap deflection resulted in negative dihedral effect due to the vertical tail. Propeller operation decreased the lateral stability parameter of the airplane for all the conditions investigated with larger decreases being measured for the flaps deflected conditions.
Canopy loads investigation for the F6F-3 airplane
No Description Available.
A low-speed investigation of an annular transonic air inlet
Report presenting an investigation of three transonic fuselage-inlet installations designed to maintain substream velocities on the body ahead of air inlets. Surface pressures and inlet total pressures were measured at the tops of the test configurations for wide ranges of inlet-velocity ratio and angles of attack. Results indicated that substream velocities were maintained on all three noses over the angle of attack ranges and inlet-velocity ratio useful for high-speed flight.
Observations on an aileron-flutter instability encountered on a 45 degree swept-back wing in transonic and supersonic flight
No Description Available.
Tests of a Horizontal-Tail Model through the Transonic Speed Range by the NACA Wing-Flow Method
A 1/12-scale model of a horizontal tail of a fighter airplane was tested through the transonic speeds in the high-speed flow over an airplane wing, the surface of which served as a reflection plane for the model. Measurements of lift, elevator-hinge moment, angle of attack, and elevator angle were made in the Mach number range from 0.75 to 1.04 for elevator deflections ranging from 10 degrees to minus 10 degrees, and for angles of attack of minus 1.2 degrees, 0.4 degrees, and 3.4 degrees. The equipment used to measure the hinge moments of the model proved to be unsatisfactory, and for this reason the hinge-moment data are considered to be only qualitative.
The interaction of boundary layer and compression shock and its effect upon airfoil pressure distributions
No Description Available.
A Low-Speed Investigation of a Fuselage-Side Air Inlet for use at Transonic Flight Speeds
A low-speed investigation in the Langley propeller-research tunnel of annular air inlets designed to avoid compression shocks and attendant boundary-layer separation on the fuselage ahead of the inlets at transonic flight speeds by maintaining substream flow velocities on the fuselage nose was reported in NACA RM No. L6J04. In the present investigation, one of the original annular inlets was converted by the installation of a canopy and a nose-wheel fairing into a twin side inlet in order to study problems involved in applying such an inlet to a fighter-type airplane. Extensive measurements of pressures on the surface of the model and surveys of the internal flow were conducted at angles of attack of 0 degrees, 3 degrees, and 6 degrees over a wide range of inlet-velocity ratio.
Vibration Survey of Blades in 19XB Axial-Flow Compressor, 2, Dynamic Investigation
Strain-gage measurements were taken under operating conditions from blades of various stages of the 19XB axial-flow compressor in an effort to determine the reason for failures in the seventh and tenth stages. First bending-mode vibrations were detected in the first five stages of the compressor caused by each integral multiple of rotor speed from three through ten. Lead-wire failures in the last five stages resulted in incomplete data. The dynamic-vibration frequencies at various rotor speeds were compared with statically measured frequencies analytically corrected for the influence of centrifugal force. Large increases in vibration ani~litude with increased pressure ratio were observed. During surging operation, blade vibrations were not present. The effects of pressure ratio and surge indicate the existence of aerodynamic excitation as the cause of the blade vibrations.
Wind-tunnel investigation of air loads over a double slotted flap on the NACA 65(216)-215, a = 0.8 airfoil section
Report presenting an investigation at low speed and high Reynolds number to determine the air loads over a double slotted flap on the NACA 65(216)-215, a = 0.8 airfoil section. Results indicated that the loads on the flap change slowly with variation of angle of attack but quickly with the flap deflected. The main effect of the double-slotted flap is that it causes the airfoil to carry a greater load without stalling.
Air-Stream Surveys in the Vicinity of the Tail of a 1/8.33-Scale Powered Model of the Republic XF-12 Airplane
The XF-12 airplane was designed by Republic Aviation Corporation to provide the Army Air Forces with a high performance, photo reconnaissance aircraft. A series of air-stream surveys were made n the vicinity of the empennage of a 1/8.33-scale powered model of the XF-12 airplane in the Langley 19-foot pressure tunnel. Surveys of the vortical-tail region were made through a range of yaw angles of plus or minus 20 degrees at a high and low angle of attack. The horizontal-tail surveys were made over a fairly wide range of angles of attack at zero degrees yaw. Several power and flap conditions were investigated. The results are presented in the form a dynamic pressure ratios, sidewash angles, and downwash angles plotted against vertical distance from the fuselage center line. The results of the investigation indicate that a vertical tail located in a conventional position would be in a field of flow where the dynamic pressure ratios at the horizontal tail to be increased; for equal lift coefficients, the effect of power or flap deflection on the direction of flow at any particular point in the region of the horizontal tail is small.
Comparative drag measurements at transonic speeds of 6-percent-thick airfoils of symmetrical double-wedge and circular-arc sections from tests by the NACA wing-flow method
No Description Available.
Drag measurements of a swept-back wing having inverse taper as determined by flight tests at supersonic speeds
Report discussing the results of flight tests to determine the drag at zero lift of a swept-back wing of inverse taper using an NACA 65-009 airfoil. The data was compared to untapered wings with a similar degree of sweepback. The tapered wing was found to have a lower drag coefficient than the 34-degree swept-back untapered wing but a higher drag coefficient than the 45-degree swept-back untapered wing.
Effect of number of fins on the drag of a pointed body of revolution at low supersonic velocities
No Description Available.
Altitude-Wind-Tunnel Investigation of the 19B-2, 19B-8 and 19XB-1 Jet- Propulsion Engines, 4, Analysis of Compressor Performance
Investigations were conducted in the Cleveland altitude wind tunnel to determine the performance and operational characteristics of the 19B-2, 19B-8, and 19XS-1 turbojet engines. One objective was to determine the effect of altitude, flight Mach number, and tail-pipe-nozzle area on the performance characteristics of the six-stage and ten-stage axial-flow compressors of the 19B-8 and 19XB-1 engines, respectively, The data were obtained over a range of simulated altitudes and flight Mach numbers. At each simulated flight condition the engine was run over its full operable range of speeds. Performance characteristics of the 19B-8 and 19XB-1 compressors for the range of operation obtainable in the turboJet-engine installation are presented. Compressor characteristics are presented as functions of air flow corrected to sea-level conditions, compressor Mach number, and compressor load coefficient. For the range of compressor operation investigated, changes in Reynolds number had no measurable effect on the relations among compressor Mach number, corrected air flow, compressor load coefficient, compressor pressure ratio, and compressor efficiency. The operating lines for the 19B-8 compressor lay on the low-air-flow side of the region of maximum compressor efficiency; the 19B-8 compressor operated at higher average pressure coefficients per stage and produced a lower over-all pressure ratio than did the 19XB-1 compressor.
Performance of a Mixed-Flow Impeller in Combination with Semivaneless Diffuser
The performance of a mixed-flow impeller in combination with a semivaneless diffuser were experimentally investigated. The diameter of the impeller was 11.0 inches and a maximum tip diameter of 14.74 inches. The semivaneless diffuser had an overall diameter of 28.00 inches. The performance properties of the mixed-flow impeller were also investigated with a 34.00 inch vane loss diffuser having a transition section of the same geometry as the semivaneless diffuser.
Experimental investigation of the effects of viscosity on the drag of bodies of revolution at a Mach number 1.5
No Description Available.
Increase in stable-air-flow operating range of a mixed-flow compressor by means of a surge inhibitor
No Description Available.
Free-fall measurements at transonic velocities of the drag of a wing-body configuration consisting of a 45 degree swept-back wing mounted forward of the maximum diameter on a body of fineness ratio 12
No Description Available.
An analysis of the compressive strength of honeycomb cores for sandwich construction
No Description Available.
An analysis of the factors that affect the exhaust process of a four-stroke-cycle reciprocating engine
From Introduction: "An investigation was made to determine the relative significance of the factors that affect the exhaust process; the effects that an exhaust-process change has on cylinder charging were given special attention. Differential equations of this nature have been developed by Kemble (reference 1) but the forms of these equations are such that general conclusions regarding the various factors affecting the exhaust process cannot be determined."