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The Theory of the Pitot and Venturi Tubes, Part 2
Report discussing several aspects of pitot and venturi tubes, including the energy equation for steady adiabatic flow, introduction of mean speed into the energy equation, isentropic flow of an ideal gas, the theory of the pitot tube, and the theory of the venturi meter.
Elements of the Wing Section Theory and of the Wing Theory
Results are presented of the theory of wings and of wing sections which are of immediate practical value. They are proven and demonstrated by the use of the simple conceptions of kinetic energy and momentum only.
Flow and Drag Formulas for Simple Quadrics
The pressure distribution and resistance found by theory and experiment for simple quadrics fixed in an infinite uniform stream of practically incompressible fluid are calculated. The experimental values pertain to air and some liquids, especially water; the theoretical refer sometimes to perfect, again to viscid fluids. Formulas for the velocity at all points of the flow field are given. Pressure and pressure drag are discussed for a sphere, a round cylinder, the elliptic cylinder, the prolate and oblate spheroid, and the circular disk. The velocity and pressure in an oblique flow are examined.
Flow and Force Equations for a Body Revolving in a Fluid
A general method for finding the steady flow velocity relative to a body in plane curvilinear motion, whence the pressure is found by Bernoulli's energy principle is described. Integration of the pressure supplies basic formulas for the zonal forces and moments on the revolving body. The application of the steady flow method for calculating the velocity and pressure at all points of the flow inside and outside an ellipsoid and some of its limiting forms is presented and graphs those quantities for the latter forms.
General Theory of Aerodynamic Instability and the Mechanism of Flutter
The aerodynamic forces on an oscillating airfoil or airfoil-aileron combination of three independent degrees of freedom were determined. The problem resolves itself into the solution of certain definite integrals, which were identified as Bessel functions of the first and second kind, and of zero and first order. The theory, based on potential flow and the Kutta condition, is fundamentally equivalent to the conventional wing section theory relating to the steady case. The air forces being known, the mechanism of aerodynamic instability was analyzed. An exact solution, involving potential flow and the adoption of the Kutta condition, was derived. The solution is of a simple form and is expressed by means of an auxiliary parameter k. The flutter velocity, treated as the unknown quantity, was determined as a function of a certain ratio of the frequencies in the separate degrees of freedom for any magnitudes and combinations of the airfoil-aileron parameters.
The Inertia Coefficients of an Airship in a Frictionless Fluid
The apparent inertia of an airship hull is examined. The exact solution of the aerodynamical problem is studied for hulls of various shapes with special attention given to the case of an ellipsoidal hull. So that the results for the ellipsoidal hull may be readily adapted to other cases, they are expressed in terms of the area and perimeter of the largest cross section perpendicular to the direction of motion by means of a formula involving a coefficient kappa which varies only slowly when the shape of the hull is changed, being 0.637 for a circular or elliptic disk, 0.5 for a sphere, and about 0.25 for a spheroid of fineness ratio. The case of rotation of an airship hull is investigated and a coefficient is defined with the same advantages as the corresponding coefficient for rectilinear motion.
The Minimum Induced Drag of Aerofoils
Equations are derived to demonstrate which distribution of lifting elements result in a minimum amount of aerodynamic drag. The lifting elements were arranged (1) in one line, (2) parallel lying in a transverse plane, and (3) in any direction in a transverse plane. It was shown that the distribution of lift which causes the least drag is reduced to the solution of the problem for systems of airfoils which are situated in a plane perpendicular to the direction of flight.
The Aerodynamic Forces on Airship Hulls
The new method for making computations in connection with the study of rigid airships, which was used in the investigation of Navy's ZR-1 by the special subcommittee of the National Advisory Committee for Aeronautics appointed for this purpose is presented. The general theory of the air forces on airship hulls of the type mentioned is described and an attempt was made to develop the results from the very fundamentals of mechanics.
Applications of Modern Hydrodynamics to Aeronautics. Part 1: Fundamental Concepts and the Most Important Theorems. Part 2: Applications
A discussion of the principles of hydrodynamics of nonviscous fluids in the case of motion of solid bodies in a fluid is presented. Formulae are derived to demonstrate the transition from the fluid surface to a corresponding 'control surface'. The external forces are compounded of the fluid pressures on the control surface and the forces which are exercised on the fluid by any solid bodies which may be inside of the control surfaces. Illustrations of these formulae as applied to the acquisition of transformations from a known simple flow to new types of flow for other boundaries are given. Theoretical and experimental investigations of models of airship bodies are presented.
One-dimensional flows of an imperfect diatomic gas
With the assumptions that Berthelot's equation of state accounts for molecular size and intermolecular force effects, and that changes in the vibrational heat capacities are given by a Planck term, expressions are developed for analyzing one-dimensional flows of a diatomic gas. The special cases of flow through normal and oblique shocks in free air at sea level are investigated. It is found that up to a Mach number 10 pressure ratio across a normal shock differs by less than 6 percent from its ideal gas value; whereas at Mach numbers above 4 the temperature rise is considerable below and hence the density rise is well above that predicted assuming ideal gas behavior. It is further shown that only the caloric imperfection in air has an appreciable effect on the pressures developed in the shock process considered. The effects of gaseous imperfections on oblique shock-flows are studied from the standpoint of their influence on the life and pressure drag of a flat plate operating at Mach numbers of 10 and 20. The influence is found to be small. (author).
Boron and zirconium from crucible refractories in a complex heat-resistant alloy
In a laboratory study of the factors involved in the influence of induction vacuum melting on 55ni-20cr-15co-4mo-3ti-3al heat resistant alloy, it was found that the major factor was the type of ceramic used as the crucible. The study concluded that trace amounts of boron or zirconium derived from reaction of the melt with the crucible refactories improved creep-rupture properties at 1,600 degrees F. Boron was most effective and, in addition, markedly improved hot-workability.
Blowing-type boundary-layer control as applied to the trailing-edge flaps of a 35 degree swept-wing airplane
A wind-tunnel investigation was made to determine the effects on the aerodynamic characteristics of a 35 degree swept-wing airplane of applying blowing-type boundary-layer control to the trailing-edge flaps. Flight tests of a similar airplane were then conducted to determine the effects of boundary-layer control on the handling qualities and operation of the airplane, particularly during landing and take-off. The wind-tunnel and flight tests indicated that blowing over the flaps produced large increases in flap lift increment, and significant increases in maximum lift. The use of blowing permitted reductions in the landing approach speeds of as much as 12 knots.
Systematic two-dimensional cascade tests of NACA 65-series compressor blades at low speeds
The performance of NACA 65-series compressor blade section in cascade has been investigated systematically in a low-speed cascade tunnel. Porous test-section side walls and for high-pressure-rise conditions, porous flexible end walls were employed to establish conditions closely simulating two-dimensional flow. Blade sections of design lift coefficients from 0 to 2.7 were tested over the usable angle-of-attack range for various combinations of inlet-flow angle. A sufficient number of combinations were tested to permit interpolation and extrapolation of the data to all conditions within the usual range of application. The results of this investigation indicate a continuous variation of blade-section performance as the major cascade parameters, blade camber, inlet angle, and solidity were varied over the test range. Summary curves of the results have been prepared to enable compressor designers to select the proper blade camber and angle of attack when the compressor velocity diagram and desired solidity have been determined.
An analysis of pressure studies and experimental and theoretical downwash and sidewash behind five pointed-tip wings at supersonic speeds
Flow-angle and pressure surveys behind five, thin, pointed-tip wings of varying plan form have been made at Mach numbers 1.62 and 2.41. Schlieren studies at a Mach number 1.93 for the same five plan-form wings were made to illustrate the behavior of the vortex sheet. The surveys were conducted at 1.5, 3, and 4 root chords behind three triangular wings of 50 degree, 63 degree, and 72 degree leading-edge sweep angle, and behind the 50 degree triangular wing reversed. The flow behind a pointed-tip wing having a sweptback leading edge and a sweptforward trailing edge (both 50 degrees) was also surveyed. Experiment and one of the theoretical methods are compared for the reversed triangular wing and the pointed-tip wing with the 50 degree sweptback leading edge and sweptforward trailing edge.
Area-suction boundary-layer control as applied to the trailing-edge flaps of a 35 degree swept-wing airplane
A wind-tunnel investigation was made to determine the effects on the aerodynamic characteristics of a 35 degree swept-wing airplane of applying area-suction boundary-layer control to the trailing-edge flaps. Flight tests of a similar airplane were then conducted to determine the effect of boundary-layer control in the handling qualities and operation of the airplane, particularly during landing. The wind-tunnel and flight tests indicated that area suction applied to the trailing-edge flaps produced significant increases in flap lift increment. Although the flap boundary-layer control reduced the stall speed only slightly, a reduction in minimum comfortable approach speed of about 12 knots was obtained.
Calculated and measured stresses in simple panels subject to intense random acoustic loading including the near noise field of a turbojet engine
Flat 2024-t3 aluminum panels measuring 11 inches by 13 inches were tested in the near noise fields of a 4-inch air jet and turbojet engine. The stresses which were developed in the panels are compared with those calculated by generalized harmonic analysis. The calculated and measured stresses were found to be in good agreement. In order to make the stress calculations, supplementary data relating to the transfer characteristics, damping, and static response of flat and curved panels under periodic loading are necessary and were determined experimentally. In addition, an appendix containing detailed data on the near pressure field of the turbojet engine is included.
Characteristics of the Langley 8-foot transonic tunnel with slotted test section
A large wind tunnel, approximately 8 feet in diameter, has been converted to transonic operation by means of slots in the boundary extending in the direction of flow. The usefulness of such a slotted wind tunnel, already known with respect to the reduction of the subsonic blockage interference and the production of continuously variable supersonic flows, has been augmented by devising a slot shape with which a supersonic test region with excellent flow quality could be produced. Experimental locations of detached shock waves ahead of axially symmetric bodies at low supersonic speeds in the slotted test section agreed satisfactorily with predictions obtained by use of existing approximate methods.
Compressible laminar boundary layer over a yawed infinite cylinder with heat transfer and arbitrary Prandtl number
The equations are presented for the development of the compressible laminar boundary layer over a yawed infinite cylinder. For compressible flow with a pressure gradient the chordwise and spanwise flows are not independent. Using the Stewartson transformation and a linear viscosity-temperature relation yields a set of three simultaneous ordinary differential equations in a form yielding similar solutions. These equations are solved for stagnation-line flow for surface temperatures from zero to twice the free-stream stagnation temperature and for a wide range of yaw angle and free-stream Mach number. The results indicate that the effect of yaw on the heat-transfer coefficient at the stagnation line depends markedly on the free-stream Mach number. An unusual result of the solutions is that for large yaw angles and stream Mach numbers the chordwise velocity within the boundary layer exceeds the local external chordwise velocity, even for a highly cooled wall.
A correlation of results of flight investigation with results of an analytical study of effects of wing flexibility on wing strains due to gusts
An analytical study of the effects of wing flexibility on wing strains due to gusts has been made for four spanwise stations of a four-engine bomber airplane, and the results have been correlated with results of a previous flight investigation.
Drag minimization for wings and bodies in supersonic flow
The minimization of inviscid fluid drag is studied for aerodynamic shapes satisfying the conditions of linearized theory, and subject to imposed constraints on lift, pitching moment, base area, or volume. The problem is transformed to one of determining two-dimensional potential flows satisfying either Laplace's or Poisson's equations with boundary values fixed by the imposed conditions. A general method for determining integral relations between perturbation velocity components is developed. This analysis is not restricted in application to optimum cases; it may be used for any supersonic wing problem.
Elliptic cones alone and with wings at supersonic speed
To help fill the gap in the knowledge of aerodynamics of shapes intermediate between bodies of revolution and flat triangular wings, force and moment characteristics for elliptic cones have been experimentally determined for Mach numbers of 1.97 and 2.94. Elliptic cones having cross-sectional axis ratios from 1 through 6 and with lengths and base areas equal to circular cones of fineness ratios 3.67 and 5 have been studied for angles of bank of 0 degree and 90 degrees. Elliptic and circular cones in combination with triangular wings of aspect ratios 1 and 1.5 also have been considered. The angle-of-attack range was from 0 degree to about 16 degrees, and the Reynolds number was 8 x 10(6), based on model length. In addition to the forces and moments at angle of attack, pressure distributions for elliptic cones at zero angle of attack have been determined. The results of this investigation indicate that there are distinct aerodynamic advantages to the use of elliptic cones.
Experimental determination of effects of frequency and amplitude on the lateral stability derivatives for a delta, a swept, and unswept wing oscillating in yaw
Three wing models were oscillated in yaw about their vertical axes to determine the effects of systematic variations of frequency and amplitude of oscillation on the in-phase and out-of-phase combination lateral stability derivatives resulting from this motion. The tests were made at low speeds for a 60 degree delta wing, a 45 degree swept wing, and an unswept wing; the swept and unswept wings had aspect ratios of 4. The results indicate that large changes in the magnitude of the stability derivatives due to the variation of frequency occur at high angles of attack, particularly for the delta wing. The greatest variations of the derivatives with frequency take place for the lowest frequencies of oscillation; at the higher frequencies, the effects of frequency are smaller and the derivatives become more linear with angle of attack. Effects of amplitude of oscillation on the stability derivatives for delta wings were evident for certain high angles of attack and for the lowest frequencies of oscillation. As the frequency became high, the amplitude effects tended to disappear.
An experimental investigation of sting-support effects on drag and a comparison with jet effects at transonic speeds
Various dummy stings were tested on the rear of a related series of afterbody shapes for Mach numbers from 0.80 to 1.10 and Reynolds numbers based on body length from 15.0 x 16 to the 6th power to 17.4 x 10 to the 6th power. A method is presented whereby approximate sting interference corrections can be made to models having afterbody shapes and sting supports similar to those of these tests if the Reynolds numbers are of the same order of magnitude and a turbulent boundary layer exists at the model base. Also presented is an analysis of jet duplication by use of a sting.
Full-scale investigation of several jet-engine noise-reduction nozzles
A number of noise-suppression nozzles were tested on full-scale engines. In general, these nozzles achieved noise reduction by the mixing interference of adjacent jets, that is, by using multiple-slot-nozzles. Several of the nozzles achieved reductions in sound power of approximately 5 decibels (nearly 70 percent) with small thrust losses (approx. 1 percent). The maximum sound-pressure level was reduced by as much as 18 decibels in particular frequency bands. Some of the nozzles showed considerable spatial asymmetry; that is, the sound field was not rotationally symmetrical. A method of calculating the limiting frequency effected by such nozzles is presented. Furthermore data are shown that appear to indicate that further reductions in sound power will not be easily achieved from nozzles using mixing interference as a means of noise suppression.
Ground simulator studies of the effects of valve friction, stick friction, flexibility, and backwash on power control system quality
Report presents results of tests made on a power control system by means of a ground simulator to determine the effects of various combinations of valve friction and stick friction on the ability of the pilot to control the system. Various friction conditions were simulated with a rigid control system, a flexible system, and a rigid system having some backlash. For the tests, the period and damping of the simulated airplane were held constant.
Growth of disturbances in a flame-generated shear region
Results are presented of an experimental and theoretical investigation of the growth of transverse velocity disturbances in the shear region caused by a flame in a duct. In the theoretical stability analysis, a flow field arising from a flame in a duct was analyzed. The flow was neutrally stable to symmetric disturbances and unstable to antisymmetric ones. In the experimental part of the program disturbances of various frequencies were imposed on a flame stabilized in a duct, and the effects were measured by shadow photography and photomultiplier-probe surveys.
Incompressible flutter characteristics of representative aircraft wings
This report gives the results of a detailed study of the flutter characteristics of four representative aircraft wings. This study was made using the electric-analog computer at the California Institute of Technology. During the course of this investigation eight important parameters of each wing were varied and, in addition, the effects of mass, inertia, pitching spring, and location of a concentrated mass were investigated for all four wings and at several sweepback angles.
A low-speed experimental investigation of the effect of a sandpaper type of roughness on boundary-layer transition
Report describing the effects of an area of roughness on the velocity and turbulence measurements of an airfoil. It details the effects of the size, location, and height of the roughness on the Reynolds number. From Summary: "An investigation was made in the Langley low-turbulence pressure tunnel to determine the effect of size and location of a sandpaper type of roughness on the Reynolds number for transition."
Measurement of static pressure on aircraft
Existing data on the errors involved in the measurement of static pressure by means of static-pressure tubes and fuselage vents are presented. The errors associated with the various design features of static-pressure tubes are discussed for the condition of zero angle of attack and for the case where the tube is inclined to flow. Errors which result from variations in the configuration of static-pressure vents are also presented. Errors due to the position of a static-pressure tube in the flow field of the airplane are given for locations ahead of the fuselage nose, ahead of the wing tip, and ahead of the vertical tail fin. The errors of static-pressure vents on the fuselage of an airplane are also presented. Various methods of calibrating static-pressure installations in flight are briefly discussed.
Measurements of free-space oscillating pressures near propellers at flight Mach numbers to 0.72
In the course of a short flight program initiated to check the theory of Garrick and Watkins (NACA rep. 1198), a series of measurements at three stations were made of the oscillating pressures near a tapered-blade plan-form propeller and rectangular-blade plan form propeller at flight Mach numbers up to 0.72. In contradiction to the results for the propeller studied in NACA rep. 1198, the oscillating pressures in the plane ahead of the propeller were found to be higher than those immediately behind the propeller. Factors such as variation in torque and thrust distribution, since the blades of the present investigation were operating above their design forward speed, may account for this contradiction. The effect of blade plan form shows that a tapered-blade plan-form propeller will produce lower sound-pressure levels than a rectangular-blade plan-form propeller for the low blade-passage harmonics (the frequencies where structural considerations are important) and produce higher sound-pressure levels for the higher blade-passage harmonics (frequencies where passenger comfort is important).
The mechanism of thermal-gradient mass transfer in the sodium hydroxide-nickel system
"Thermal-gradient mass transfer" was investigated in the molten sodium hydroxide-nickel system. Possible mechanisms (physical, electrochemical, and chemical) are discussed in terms of experimental and theoretical evidence. Experimental details are included in appendixes.
NACA research on slurry fuels
An extensive program was conducted to investigate the use of concentrated slurries of boron and magnesium in liquid hydrocarbon as fuels for afterburners and ramjet engines. Analytical calculations indicated that magnesium fuel would give greater thrust and that boron fuel would give greater range than are obtainable from jet hydrocarbon fuel alone. It was hoped that the use of these solid elements in slurry form would permit the improvement to be obtained without requiring unconventional fuel systems or combustors. Small ramjet vehicles fueled with magnesium slurry were flown successfully, but the test flights indicated that further improvement of combustors and fuel systems was needed.
Origin and Prevention of Crash Fires in Turbojet Aircraft
The tendency for the jet engine rotor to continue to rotate after crash presents the probability that crash-spilled combustibles suspended in the air or puddled on the ground at the engine inlet may be sucked into the engine. Studies with jet engines operating on a test stand and full-scale crashes of turbojet-powered airplanes showed that combustibles drawn into the engine in this way ignite explosively within the engine. Experiment showed that the gas flow through the engine is too rapid to permit the ignition of ingested combustibles on the hot metal in contact with the main gas stream. Ignition will occur on those hot surfaces not in the main gas stream. The portion of the engine airflow is diverted for cooling and ventilation to these zones where the gas moves slowly enough for ignition to occur.
The similarity rules for second-order subsonic and supersonic flow
The similarity rules for linearized compressible flow theory (Gothert's rule and its supersonic counterpart) are extended to second order. It is shown that any second-order subsonic flow can be related to "nearly incompressible" flow past the same body, which can be calculated by the Janzen-Rayleigh method.
A simplified method for approximating the transient motion in angles of attack and sideslip during a constant rolling maneuver
The transient motion in angles of attack and sideslip during a constant rolling maneuver has been analyzed. Simplified expressions are presented for the determination of the pertinent modes of motion as well as the modal coefficient corresponding to each mode. Calculations made with and without the derivatives for side force due to sideslip and lift-curve slope indicate that although these derivatives increase the total damping of the system they do not markedly affect the transient motion.
A summary of preliminary investigations into the characteristics of combustion screech in ducted burners
Increasing demands for higher afterburner performance have required operation at progressively higher fuel-air ratios, which has increased the occurrence and intensity of screeching combustion. The onset of screech may be followed by rapid destruction of the combustor shell and other combustor parts. Because of its destructive characteristics, considerable effort has been expended to understand and eliminate screech. NACA work on the screeching combustion problem prior to 1954 is summarized herein. These studies showed that resonant acoustic oscillations are a primary component of the screech mechanism in the burners thus far investigated.
Theoretical analysis of total-pressure loss and airflow distribution for tubular turbojet combustors with constant annulus and liner cross-sectional areas
Compressible and incompressible flow calculations were made of the combustor total-pressure-loss coefficient and liner airflow distribution for tubular turbojet combustors having constant annulus and liner cross-sectional areas along the combustor axis. Information on static and total pressure distribution and liner air-jet entrance angles along the length of the combustor was obtained as an intermediate step in the calculations. The calculations include the effects of heat release, annulus wall friction, and variation in discharge coefficients of the liner wall openings along the combustor. The combustor total-pressure-loss coefficient and liner air-flow distribution are presented graphically in terms of the following dimensionless parameters: (1) combustor reference Mach number, (2) ratio of combustor-exit to inlet total temperature, (3) fraction of total airflow passing through the liner dome, (4) ratio of total open hole area in liner wall to total combustor cross-sectional area, and (5) ratio of liner cross-sectional area to total combustor cross-sectional area.
Theory of self-excited mechanical oscillations of helicopter rotors with hinged blades
Vibrations of rotary-wing aircraft may derive their energy from the rotation of the rotor rather than from the air forces. A theoretical analysis of these vibrations is described and methods for its application are explained in Chapter one. Chapter two reports the results of an investigation of the mechanical stability of a rotor having two vertically hinged blades mounted upon symmetrical supports, that is, of equal stiffness and mass in all horizontal directions. Chapter three presents the theory of ground vibrations of a two-blade helicopter rotor on anisotropic flexible supports.
Thin airfoil theory based on approximate solution of the transonic flow equation
A method is presented for the approximate solution of the nonlinear equations of transonic flow theory. Solutions are found for two-dimensional flows at a Mach number of 1 and for purely subsonic and purely supersonic flows. Results are obtained in closed analytic form for a large and significant class of nonlifting airfoils. At a Mach number of 1 general expressions are given for the pressure distribution on an airfoil of specified geometry and for the shape of an airfoil having a prescribed pressure distribution. Extensive comparisons are made with available data, particularly for a Mach number of 1, and with existing solutions.
Effect of fuel variables on carbon formation in turbojet-engine combustors
Report presents the results of an investigation of the effects of fuel properties and of a number of fuel additives on combustion-chamber carbon deposition and exhaust-gas smoke formation in a single tubular turbojet-engine combustor. Limited tests were conducted with a number of the fuels in several full-scale turbojet engines to verify single-combustor data.
Investigation of a nonlinear control system
A discontinuous variation of coefficients of the differential equation describing the linear control system before nonlinear elements are added is studied in detail. The nonlinear feedback is applied to a second-order system. Simulation techniques are used to study performance of the nonlinear control system and to compare it with the linear system for a wide variety of inputs. A detailed quantitative study of the influence of relay delays and of a transport delay is presented.
Relation of Turbojet and Ramjet Combustion Efficiency to Second-Order Reaction Kinetics and Fundamental Flame Speed
Theoretical studies of the turbojet and ramjet combustion process are summarized and the resulting equations are applied to experimental data obtained from various combustor tests. The theoretical treatment assumes that one step in the over-all chain of processes which constitute jet-engine combustion is sufficiently slow to be the rate-controlling step that determines combustion efficiency.
A method of computing the transient temperature of thick walls from arbitrary variation of adiabatic-wall temperature and heat-transfer coefficient
A method of calculating the temperature of thick walls has been developed in which the time series and the response to a unit triangle variation of surface temperature concepts are used, together with essentially standard formulas for transient temperature and heat flow into thick walls. The method can be used without knowledge of the mathematical tools of its development. The method is particularly suitable for determining the wall temperature in one-dimensional thermal problems in aeronautics where there is a continuous variation of the heat-transfer coefficient and adiabatic-wall temperature. The method also offers a convenient means for solving the inverse problem of determining the heat-flow history when temperature history is known.
Approximate analysis of effects of large deflections and initial twist on torsional stiffness of a cantilever plate subjected to thermal stresses
From Summary: "An approximate analysis of the nonlinear effects of initial twist and large deflections on the torsional stiffness of a cantilever plate subjected to a nonuniform temperature distribution is presented. The Von Karman large-deflection equations are satisfied through the use of a variational principle. The results show that initial twist and applied moments can have significant effects on the changes in stiffness produced by nonuniform heating, particularly in the region of the buckling temperature difference. Results calculated by this approximate analysis are in satisfactory agreement with measured torsional deformations and changes in natural frequency."
Survey of hydrogen combustion properties
This literature digest of hydrogen-air combustion fundamentals presents data on flame temperature, burning velocity, quenching distance, flammability limits, ignition energy, flame stability, detonation, spontaneous ignition, and explosion limits. The data are assessed, recommended values are given, and relations among various combustion properties are discussed. New material presented includes: theoretical treatment of variation in spontaneous ignition lag with temperature, pressure, and composition, based on reaction kinetics of hydrogen-air composition range for 0.01 to 100 atmospheres and initial temperatures of 0 degrees to 1400 degrees k.
On flow of electrically conducting fluids over a flat plate in the presence of a transverse magnetic field
The use of a magnetic field to control the motion of electrically conducting fluids is studied. The incompressible boundary-layer solutions are found for flow over a flat plate when the magnetic field is fixed relative to the plate or to the fluid. The equations are integrated numerically for the effect of the transverse magnetic field on the velocity and temperature profiles, and hence, the skin friction and rate of heat transfer. It is concluded that the skin friction and the heat-transfer rate are reduced when the transverse magnetic field is fixed relative to the plate and increased when fixed relative to the fluid. The total drag is increased in all of the areas.
A variational theorem for creep with applications to plates and columns
A variational theorem is presented for a body undergoing creep. Solutions to problems of the creep behavior of plates, columns, beams, and shells can be obtained by means of the direct methods of the calculus of variations in conjunction with the stated theorem. The application of the theorem is illustrated for plates and columns by the solution of two sample problems.
A phenomenological relation between stress, strain rate, and temperature for metals at elevated temperatures
Report presenting a phenomenological relation between stress, strain rate, and temperature to account for the behavior of polycrystalline metals above the equicohesive temperature. This is because the behavior of metals at elevated temperatures is constantly becoming a more important problem. The validity of the relation suggested in the equation appears to be accurate for 7075-T6 aluminum alloy.
Differential equations of motion for combined flapwise bending, chordwise bending, and torsion of twisted nonuniform rotor blades
The differential equations of motion for the lateral and torsional deformations of twisted rotating beams are developed for application to helicopter rotor and propeller blades. No assumption is made regarding the coincidence of the neutral, elastic, and mass axes, and the generality is such that previous theories involving various simplifications are contained as subcases to the theory presented in this paper. Special attention is given the terms which are not included in previous theories. These terms are largely coupling-type terms associated with the centrifugal forces. Methods of solution of the equations of motion are indicated by selected examples.
An analysis of the effects of aeroelasticity on static longitudinal stability and control of a swept-wing airplane
From Introduction: "The results of the aforementioned study are presented in this report together with the method of analysis employed. The net stability change is shown together with the individual contributions due to flexibility of wing, tail, and fuselage, both including and neglecting the effect of inertial loads."