ALBERT

Description: In the design of a combustor, information is necessary for the mixing of the fuel and air in order to determine the optimum combustor length. In scramjet combustors the mixing often takes place in a shear layer that is formed between the fuel and air. This research was an experimental study of shear layers in supersonic flows aimed at determining what mechanisms affect the shear layer so that the mixing could be better predicted. A second goal was to provide sufficient instream information for use in checking existing Computational Fluid Dynamic (CFD) codes. The shear layer between a supersonic two-dimensional air stream (M = 2 or M = 3) was mixed with a near sonic two-dimensional air stream (M = 1.2). Instream measurements of pitot pressure and cone static pressure were used to determine mean velocity profiles at various axial locations. These velocity profiles were used to determine the shear layer spreading rate and are compared with various predictions. Wall measurements of static pressure, temperature and skin friction were also taken and are presented. The instream measurements were also used for comparison with an existing CFD code. The upstream velocity, pressure and temperature profiles were used as a starting profile and the code was used to calculate downstream profiles for comparison with the experimental results. Reasonable agreement between the measured and calculated results was obtained.

Description: This paper presents a set of second-order closure models for low-Reynolds-number turbulence near the wall. Existing closure models for the Reynolds-stress equations were modified to show proper near-wall behavior. A dissipation-rate equation for the turbulent kinetic energy is also reformulated. The proposed models satisfy realizability and will not produce unphysical behavior. Fully developed channel flows are used for model testing. The equations are solved for the mean velocity, the Reynolds stresses, and the dissipation rate of the turbulent kinetic energy. The calculations are compared with both direct numerical simulations and with measurements. It is shown that the present models perform well in predicting the behavior of the turbulence near a wall. Significant improvements over previous models in predicting the components of the Reynolds stress tensor are obtained in the present models.

Description: Hi-Tech Inc., a company which manufactures water jetting equipment, needed a high pressure rotating swivel, but found that available hardware for the system was unsatisfactory. They were assisted by Marshall, which had developed water jetting technology to clean the Space Shuttles. The result was a completely automatic water jetting system which cuts rock and granite and removes concrete. Labor costs have been reduced; dust is suppressed and production has been increased.

Description: Stars and Stripes racing yacht brought the American's Cup back to the United States. Originating from NASA's drag reduction technology, the boats "secret weapon" was that the hull's underside was coated with riblets. Riblets are small, barely visible grooves on the surface of an airplane intended to reduce skin friction by smoothing the turbulent airflow next to the skin. Grooves are V-shaped with the angle pointing in the direction of the airflow. No deeper than a scratch, they have a pronounced beneficial influence on air turbulence. *No longer commercially available.

Description: Macrodyne, Inc.'s laser velocimeter (LV) is a system used in wind tunnel testing of aircraft, missiles and spacecraft employing electro optical techniques to probe the flow field as the tunnel blows air over a model of flight vehicle and to determine velocity of air and its direction at many points around the model. However, current state-of-the-art minicomputers cannot handle the massive flow of real time data from several sources simultaneously. Langley developed instrument Laser Velocimeter Autocovariance Buffer Interface (LVABI). LVABI is interconnecting instrument between LV and computer. It acquires data from as many as six LV channels at high real time data rates, stores it in memory and sends it to computer on command. LVABI has application in variety of research, industrial and defense functions requiring precise flow measurement.

Description: An approximate method for development of flow and thermal boundary layers in laminar regime on cylinders with arbitrary cross section and transpiration-cooled walls is obtained by use of Karman's integrated momentum equation and an analogous heat-flow equation. Incompressible flow with constant property values throughout boundary layer is assumed. Shape parameters for approximated velocity and temperature profiles and functions necessary for solution of boundary-layer equations are presented as charts, reducing calculations to a minimum. The method is applied to determine local heat-transfer coefficients and surface temperature-cooled turbine blades for a given flow rate. Coolant flow distributions necessary for maintaining uniform blade temperatures are also determined.

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: Numerical solutions of the differential equation obtained from the momentum theorem for the development of a turbulent boundary layer along a thermally insulated surface in two-dimensional and in radial shock-free flow are presented in tabular form for a range of Mach numbers from 0.100 to 10. The solution can be used in a step-wise procedure with any given distribution of favorable pressure gradients and for zero pressure gradients. Solutions are also given for use with moderate adverse pressure gradients. The mean velocity in the boundary layer is approximated by a power-law profile. In view of the stepwise integration methods to be used, the exponent designated the profile shape can be varied along the surface between the integral fraction limits 1/5 and 1/11 through interpolation. Agreement obtained between theoretical and experimental boundary-layer development in a supersonic nozzle at a nominal Mach number of 2 indicates the general validity of the approximations used in the analysis - in particular, the method of extrapolating low-speed skin-friction relations to high Mach number flows. The extrapolation method used assumes that the skin-friction coefficient depend primarily on Reynolds number, provided that the density and the kinematic viscosity are evaluated at surface conditions.

Description: The condensation pressure of air was determined over the range of temperature from 60 to 85 K. The experimental results were slightly higher than the calculated values based on the ideal solution law. Heat of vaporization of oxygen was determined at four temperatures ranging from about 68 to 91 K and of nitrogen similarly at four temperatures ranging from 62 to 78 K.

Description: The heat requirements for the icing protection of two radome configurations have been studied over a range of design icing conditions. Both the protection limits of a typical thermal protection system and the relative effects of the various icing variables have been determined. For full evaporation of all impinging water, an effective heat density of 14 watts per square inch was required. When a combination of the evaporation and running wet surface systems was employed, a heat requirement of 5 watts per square inch provided protection at severe icing and operating conditions.

Description: The Navier-Stokes equations of motion and the equation of continuity are transformed so as to apply to an orthogonal curvilinear coordinate system rotating with a uniform angular velocity about an arbitrary axis in space. A usual simplification of these equations as consistent with the accepted boundary-layer theory and an integration of these equations through the boundary layer result in boundary-layer momentum-integral equations for three-dimensional flows that are applicable to either rotating or nonrotating fluid boundaries. These equations are simplified and an approximate solution in closed integral form is obtained for a generalized boundary-layer momentum-loss thickness and flow deflection at the wall in the turbulent case. A numerical evaluation of this solution carried out for data obtained in a curving nonrotating duct shows a fair quantitative agreement with the measures values. The form in which the equations are presented is readily adaptable to cases of steady, three-dimensional, incompressible boundary-layer flow like that over curved ducts or yawed wings; and it also may be used to describe the boundary-layer flow over various rotating surfaces, thus applying to turbomachinery, propellers, and helicopter blades.

Description: The trajectories of droplets in the air flowing past NACA 65(1)-208 airfoil and an NACA 65(1)-212 airfoil, both at an angle of attack of 4 degrees, were determined. The amount of water in droplet form impinging on the airfoils, the area of droplet impingement, and the rate of droplet impingement per unit area on the airfoil surface affected were calculated from the trajectories and are presented. The amount, extent, and rate of impingement of the NACA 65(1)-208 airfoil are compared with the results for the NACA 65(1)1-212 airfoil. Under similar conditions of operation, the NACA 65(1)-208 airfoil collects less water than the NACA 65(1)-212 airfoil. The extent of impingement on the upper surface of the NACA 65(1)-208 airfoil is much less than on the upper surface of the NACA 65(1)-212 airfoil, but on the lower surface the extents of impingement are about the same.

Description: A conservative zoning technique, wherein the flow field for a finite-difference calculation is divided into several regions to simplify grid generation, is discussed and is applied in the solution of a two-dimensional problem of complex topology. Calculations are performed on two zonal, or patched, grid systems for the supersonic flow over a double-airfoil configuration. The solution is smooth and continuous across the zonal interfaces, and shock waves pass through the boundaries without distortion. In addition, the time accuracy of the zonal boundary method is verified by a two-zone cylinder calculation with a stationary inner and a rotating outer mesh. The feasibility of the zonal approach for use in the solution of geometrically complex and unsteady problems is thus demonstrated.

Description: An investigation of forced-convection heat transfer and associated pressure drops was conducted with air flowing through electrically heated Inconel tubes having various degrees of square-thread-type roughness, an inside diameter of 1/2 inch, and a length of 24 inches. were obtained for tubes having conventional roughness ratios (height of thread/radius of tube) of 0 (smooth tube), 0.016, 0.025, and 0.037 over ranges of bulk Reynolds numbers up to 350,000, average inside-tube-wall temperatures up to 1950deg R, and heat-flux densities up to 115,000 Btu per hour per square foot. Data The experimental data showed that both heat transfer and friction increased with increase in surface roughness, becoming more pronounced with increase in Reynolds number; for a given roughness, both heat transfer and friction were also influenced by the tube wall-to-bulk temperature ratio. Good correlation of the heat-transfer data for all the tubes investigated was obtained by use of a modification of the conventional Nusselt correlation parameters wherein the mass velocity in the Reynolds number was replaced by the product of air density evaluated at the average film temperature and the so-called friction velocity; in addition, the physical properties of air were evaluated at the average film temperature. The isothermal friction data for the rough tubes, when plotted in the conventional manner, resulted in curves similar to those obtained by other investigators; that is, the curve for a given roughness breaks away from the Blasius line (representing turbulent flow in smooth tubes) at some value of Reynolds number, which decreases with increase in surface roughness, and then becomes a horizontal line (friction coefficient independent of Reynolds number). A comparison of the friction data for the rough tubes used herein indicated that the conventional roughness ratio is not an adequate measure of relative roughness for tubes having a square-thread-type element. The present data, as well as those of other investigators, were used to isolate the influence of ratios of thread height to width, thread spacing to width, and the conventional roughness ratio on the friction coefficient. A fair correlation of the friction data was obtained for each tube with heat addition when the friction coefficient and Reynolds number were defined on the basis of film properties; however, the data for each tube retained the curve characteristic of that particular roughness. The friction data for all the rough tubes could be represented by a single line for the complete turbulence region by incorporating a roughness parameter in the film correlation. No correlation was obtained for the region of incomplete turbulence.

Description: Research was conducted to determine the effect of the electrode parameters of spacing, configuration, and material' on the energy required for ignition of a flowing propane-air mixture. In addition, the data were used to indicate the energy distribution along the spark length and to confirm previous observations concerning the effect of spark duration on ignition energy requirements. The data were obtained with a mixture at a fuel-air ratio of 0.0835 (by weight), a pressure of 3 inches of mercury absolute, a temperature of 80 F, and a mixture velocity of 5 feet per second. Results showed that the energy required for ignition decreased as the electrode spacing was increased; a minimum energy occurred at. a spacing of 0.65 inch for large electrodes. For small electrodes, the spacing for minimum energy was not sharply defined. Small-diameter electrodes required less energy than large-diameter electrodes if the spacing was less than the optimum distance of 0.65 inch; at a spacing equal to the optimum distance, no difference was noted. Significant effects of electrode material on ignition energy were ascribed to differences in the type of spark discharges produced; glow discharges required higher energy than the arc-glow discharges. With pure glow discharges, the ignition energy was substantially constant for lead, cadmium, brass, aluminum, and tungsten electrodes. A method is described for determining the energy distribution along a glow discharge. It was found that one-third to one-half of the energy in the spark was concentrated in a small region near the cathode electrode, and the remainder was uniformly distributed across the spark gap. It was impossible to ascertain the dependence of ignition on. this distribution. It was also observed that long-duration (600 microsec) sparks required much less energy for ignition than did short-duration (1 microsec) sparks.

Description: No abstract available

Description: No abstract available

Description: A vapor block resistant liquid artery structure for heat pipes. A solid tube artery with openings is encased in the sintered material of a heat pipe wick. The openings are limited to that side of the artery which is most remote from the heat source. The liquid in the artery can thus exit the artery through the openings and wet the sintered sheath, but vapor generated at the heat source is unlikely to move around the solid wall of the artery and reverse its direction in order to penetrate the artery through the openings. An alternate embodiment uses finer pore size wick material to resist vapor entry.

Description: A thermal microstructure measurement system (TMMS) operates autonomously h its own internal power supply and telemeters data to a platform. A thermal array is mounted on a cross-braced frame designed to orient itself normal to existing currents with fixed sensor positioning bars protruding from the cross bars. A plurality of matched thermistors, conductivity probes and inclinometers are mounted on the frame. A compass and pressure transducer are contained in an electronics package suspended below the array. The array is deployed on a taut mooring below a subsurface float. Data are digitized, transmitted via cable to a surface buoy and then telemetered to the platform where the data is processed via a computer, recorded and/or displayed. The platform computer also sends commands to the array via telemetry.

Description: Tests of a 1/5 scale model of a proposed 153-foot high-speed submarine have been conducted in the Langley full-scale tunnel at the request of the Bureau of Ships, Department of the Navy. The test program included: (1) force tests to determine the drag, control effectiveness, and static stability characteristics for a number of model configurations, both in pitch and in yaw, (2) pressure measurements to determine the boundary-layer conditions and flow characteristics in the region of the propeller, and (3) an investigation of the effects of propeller operation on the model aerodynamic characteristics. In response to oral requests from the Bureau of Ships representatives t hat the basic data obtained in these tests be made available to them as rapidly as possible, this data report has been prepared to present some of the more pertinent results. All test results given in the present paper are for the propeller-removed condition and were obtained at a Reynolds number of approximately 22,300,000 based on model length.

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 improved efficiency arcjet thruster has a constrictor and electrically-conductive nozzle anode defining an arc chamber, and an electrically-conductive rod having a tip spaced upstream from the constrictor and defining a cathode spaced from the anode by a gap generally coextensive with the arc chamber. An electrical potential is applied to the anode and cathode to generate an electrical arc in the arc chamber from the cathode to anode. Catalytically decomposed hydrazine is supplied to the arc chamber with generation of the arc so as to produce thermal heating and expansion thereof through the nozzle. The constrictor can have a electrically insulative portion disposed between the cathode tip and the nozzle anode, and an electrically-conductive anode extension disposed along the insulative portion so as to define an auxiliary gap with the cathode tip substantially smaller than the gap defined between the cathode and nozzle anode for facilitating startup of arc generation. The constrictor can also include an electrically-conductive electrode with a variable electrical potential to vary the shape of the arc generated in the arc chamber. Also, the cathode is mounted for axial movement such that the gap between its tip and the nozzle anode can be varied to facilitate a generally nonerosive generation of the electrical arc at startup and reliable steady state operation. Further, the arc chamber can have a nonparallel subsonic-to-supersonic transition configuration, or alternatively solely a nonparallel supersonic configuration, for improved arc attachment.

Description: A vapor block resistant liquid artery structure for heat pipes. A solid tube artery with openings is encased in the sintered material of a heat pipe wick. The openings are limited to that side of the artery which is most remote from the heat source. The liquid in the artery can thus exit the artery through the openings and wet the sintered sheath, but vapor generated at the heat source is unlikely to move around the solid wall of the artery and reverse its direction in order to penetrate the artery through the openings. An alternate embodiment uses finer pore size wick material to resist vapor entry.

Description: Investigations were conducted of a 12 degree 21-inch conical diffuser of 2:l area ratio to determine the interrelation of boundary layer growth and performance characteristics. surveys were made of inlet and exit from, longitudinal static pressures were recorded, and velocity profiles were obtained through an inlet Reynolds number range, determined From mass flows and based on inlet diameter of 1.45 x 10(exp 6) to 7.45 x 10(exp 6) and a Mach number range of 0.11 to approximately choking. These investigations were made to two thicknesses of inlet boundary layer. The mean value, over the entire range of inlet velocities, of the displacement thickness of the thinner inlet boundary layer was approximately 0.035 inch and that of the thicker inlet boundary layer was approximately six times this value. The loss coefficient in the case of the thinner inlet boundary layer had a value between 2 to 3 percent of the inlet impact pressure over most of the air-flow range. The loss coefficient with the thicker inlet boundary layer was of the order of twice that of the thinner inlet boundary layer at low speeds and approximately three times at high speeds. In both cases the values were substantially less than those given in the literature for fully developed pipe flow. The static-pressure rise for the thinner inlet boundary layer was of the order of 95 percent of that theoretically possible over the entire speed range. For the thicker inlet boundary layer the static pressure rise, as a percentage of that theoretically possible, ranged from 82 percent at low speeds to 68 percent at high speeds.

Description: Performance and boundary-layer data were taken in a 12 degree 10-inch inlet-diameter conical diffuser of 2:1 exit- to inlet-area ratio. These data were taken for two inlet-boundary-layer conditions. The first condition was that of a thinner inlet boundary later (boundary-layer displacement thickness, delta* approximately equal to 0.034) produced by an inlet section approximately 1 inlet diameter in length between the entrance bell and the diffuser. The second condition was a thicker inlet boundary layer (delta* approximately equal to 0.120) produced by an additional inlet section length of approximately 6 diameters. Longitudinal static-pressure distributions were measured fro wall static orifices. Transverse total- and static-pressure surveys were made at the inlet and exit stations. Boundary-layer velocity distributions were measured at seven stations between the inlet and exit. These data were obtained for a Reynolds number (based on inlet diameter) range of 1 x 10(exp 6) to 3.9 x 10(exp 6). The corresponding Mach number range was from M = 0.2 to choking. At the maximum-power-available condition supersonic flow was obtained as far as 4.5 inches downstream from the diffuser inlet with a maximum Mach number of M approximately equal to 1.5. The total-pressure loss through the diffuser in percentage of inlet dynamic pressure was approximately 2.5 percent for the thinner inlet boundary later and 5.5 percent for the thicker inlet boundary later over the lower subsonic range. These valued increased with increasing flow rate- the values for the thicker inlet boundary later more than those for the thinner inlet boundary layer. The diffuser effectiveness, expressed as the ratio of the actual static-pressure rise to the ideal static-pressure rise, was about 85 percent for the thinner inlet boundary layer and about 67 percent for the thicker inlet boundary later in the lower subsonic range. These values decrease with increasing flow rate. Separated flow was observed for both inlet-boundary-layer conditions in the region of adverse pressure gradient just downstream of the transition curvature from inlet section to diffuser. The flow for the thinner-inlet-boundary-layer condition did not fully re-establish itself along the diffuser walls. The thicker inlet-boundary-layer flow, while not completely re-establishing the normal flow pattern downstream of the separated region, did re-establish more successfully than the thinner inlet boundary layer.

Description: Method and apparatus for obtaining dynamic calibrations of pressure transducers. A calibration head (15), a flexible tubing (23) and a bellows (20) enclose a volume of air at atmospheric pressure with a transducer (11) to be calibrated subject to the pressure inside the volume. All of the other apparatus in the drawing apply oscillations to bellows (20) causing the volume to change thereby applying oscillating pressures to transducer (11) whereby transducer (11) can be calibrated.

Description: The invention does away with the necessity of moving parts such as a check valve in a nuclear reactor cooling system. Instead, a jet pump, in combination with a TEMP, is employed to assure safe cooling of a nuclear reactor after shutdown. A main flow exists for a reactor coolant. A point of withdrawal is provided for a secondary flow. A TEMP, responsive to the heat from said coolant in the secondary flow path, automatically pumps said withdrawn coolant to a higher pressure and thus higher velocity compared to the main flow. The high velocity coolant is applied as a driver flow for the jet pump which has a main flow chamber located in the main flow circulation pump. Upon nuclear shutdown and loss of power for the main reactor pumping system, the TEMP/jet pump combination continues to boost the coolant flow in the direction it is already circulating. During the decay time for the nuclear reactor, the jet pump keeps running until the coolant temperature drops to a lower and safe temperature where the heat is no longer a problem. At this lower temperature, the TEMP/jet pump combination ceases its circulation boosting operation. When the nuclear reactor is restarted and the coolant again exceeds the lower temperature setting, the TEMP/jet pump automatically resumes operation. The TEMP/jet pump combination is thus automatic, self-regulating and provides an emergency pumping system free of moving parts.

Description: A space vehicle thermal heat rejection system 10 utilizing separate optimized heat pipe components for the functions of heat acquisition, heat transport, and heat rejection. A honeycomb panel heat pipe evaporator section 20 performs the function of heat acquisition, and forms a closed thermodynamic system with a dual channel heat pipe transport section 30, which performs the function of heat transport. A plurality of truss or channel core heat pipe rejection fins 41 form the condenser section 40, which performs the function of heat rejection. A common wall 32 separates the condenser section 40 from the transport section 30. Using the above heat pipe components and having efficient interfacing between them results in high performance factors for the overall system.

Description: A fluid leak indicator (30) for detecting and indicating leaks in visually inaccessible fluid tubing joints (20, 21), such as those obstructed by insulation (24), includes a bag system (25) and a wicking system (30) surrounding or wrapping the joints (20, 21) under the visual obstructing material (24). Leaking fluid is collected in the bag (25) or on the wicking material (34) where it is conducted along the wicking material (34) to a visibly accessible capturing transparent indicator bulb (35) for providing a visual indication of the leak without requiring a chemical change in the capturing indicator bulb (35).

Description: This document presents equations for the two-dimensional stationary problem of gas dynamics, and uses them to derive other equations, including equations for vorticity.

Description: The vortices forming in flowing water behind solid bodies are not represented correctly by the solution of the potential theory nor by Helmholtz's jets. Potential theory is unable to satisfy the condition that the water adheres at the wetted bodies, and its solutions of the fundamental hydrodynamic equations are at variance with the observation that the flow separates from the body at a certain point and sends forth a highly turbulent boundary layer into the free flow. Helmholtz's theory attempts to imitate the latter effect in such a way that it joins two potential flows, jet and still water, nonanalytical along a stream curve. The admissibility of this method is based on the fact that, at zero pressure, which is to prevail at the cited stream curve, the connection of the fluid, and with it the effect of adjacent parts on each other, is canceled. In reality, however, the pressure at these boundaries is definitely not zero, but can even be varied arbitrarily. Besides, Helmholtz's theory with its potential flows does not satisfy the condition of adherence nor explain the origin of the vortices, for in all of these problems, the friction must be taken into account on principle, according to the vortex theorem.

Description: The use of the linearized equations of Chaplygin to calculate the subsonic flow of a gas permits solving the problem of the flow about a wing profile for absence and presence of circulation. The solution is obtained in a practical convenient form that permits finding all the required magnitudes for the gas flow (lift, lift moment velocity distribution over the profile, and critical Mach number). This solution is not expressed in simple closed form; for a certain simplifying assumption, however, the equations of Chaplygin can be reduced to equations with constant coefficients, and solutions are obtained by using only the mathematical apparatus of the theory of functions of a complex variable. The method for simplifying the equations was pointed out by Chaplygin himself. These applied similar equations to the solution of the flow problem and obtained a solution for the case of the absence of circulation.

Description: In the flow about a body with large subsonic velocity if the velocity of the approaching flow is sufficiently large, regions of local supersonic velocities are formed about the body. It is known from experiment that these regions downstream of the flow are always bounded by shock waves; a continuous transition of the supersonic velocity to the subsonic under the conditions indicated has never been observed. A similar phenomenon occurs in pipes. If at two cross sections of the pipe the velocity is subsonic and between these sections regions of local supersonic velocity are formed without completely occupying a single cross section, these regions are always bounded by shock waves.

Description: A steady incompressible three-dimensional viscous flow analysis has been conducted for the Space Shuttle External Tank/Orbiter propellant feed line disconnect flapper valves with upstream elbows. The full Navier-Stokes code, INS3D, is modified to handle interior obstacles. Grids are generated by SVTGD3D code. Two dimensional initial grids in the flow cross section with and without the flappers are improved by elliptic smoothing to provide better orthogonality, clustering and smoothness to the three dimensional grid. The flow solver is tested for stability and convergence in the presence of interior flappers. An under-relaxation scheme has been incorporated to improve the solution stability. Important flow characteristics such as secondary flows, recirculation, vortex and wake regions, and separated flows are observed. Computed values for forces, moments, and pressure drop are in satisfactory agreement with water flow test data covering a maximum tube Reynolds number of 3.5 x 10(exp 6). The results will serve as a guide to improved design and enhanced testing of the disconnect.

Description: Combustion testing at total enthalpy conditions corresponding to flight Math numbers in excess of 12 requires the use of impulse facilities. The expansion tube is the only operational facility of its size which can provide these conditions without excessive oxygen dissociation or driver gas contamination. Expansio tube operation is described herein and the operational parameters having the largest impact on its performance are determined. These are: driver-to-intermediate chamber pressure ratio, driver gas molecular weight and specific heat ratio, and driver gas temperature. Increases in the lase named parameter will markedly affect the test section static pressure. Preliminary calibration tests are discussed and test gas conditions which have been achieved are presented. Calculated and experimental test times are compared and the parameters affecting test time are discussed. The direction of future work using this important experimental tool is indicated.

Description: The two-dimensional motion of an incompressible fluid about a closed contour with a definite velocity in magnitude and direction at infinity is considered. If, without changing the direction of the velocity at infinity, the magnitude is increased, the configuration of the streamlines remains unchanged and only the numbering of the stream function changes. There exists only one family of curves that can serve as streamlines in the incompressible flow about a given contour (at a given angle of attack); for example, the contour of an airplane wing. The case is quite different with a compressible fluid.

Description: The performance of a 16-stage axial-flow compressor, in which the mean-radius solidity was reduced from 1.28 to 1.02 in the fourteenth through sixteenth stage rotors was determined. The performance of this modification was compared with that of the compressor with original rotors. The reduced solidity resulted in slightly improved performance.

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: Inadvertent slot opening in a circulation controlled rotor due to plenum valve leakage is averted by venting duct pressure below a control minimum. A vent valve is disclosed.

Description: A freezing point method and device (10) are disclosed. The method and device pertain to an inflection point technique for determining the freezing points of mixtures. In both the method and device (10), the mixture is cooled to a point below its anticipated freezing point and then warmed at a substantially linear rate. During the warming process, the rate of increase of temperature of the mixture is monitored by, for example, thermocouple (28) with the thermocouple output signal being amplified and differentiated by a differentiator (42). The rate of increase of temperature data are analyzed and a peak rate of increase of temperature is identified. In the preferred device (10) a computer (22) is utilized to analyze the rate of increase of temperature data following the warming process. Once the maximum rate of increase of temperature is identified, the corresponding temperature of the mixture is located and earmarked as being substantially equal to the freezing point of the mixture. In a preferred device (10), the computer (22), in addition to collecting the temperature and rate of change of temperature data, controls a programmable power supply (14) to provide a predetermined amount of cooling and warming current to thermoelectric modules (56).

Description: Low efficiency heat transfer in evaporators subject to unusual gravitational conditions is avoided through the use of a spiral evaporator conduit 12 receiving at an inlet 14 a vaporizable coolant at least partly in the liquid phase. Flow of the coolant through the conduit 12 demists the coolant by centrifuging the liquid phase against a pressurre wall 44 of the conduit 12. Vapor flow 40 induces counterrotating vortices 46, 48 which circulate the liquid phase coolant around the interior of the conduit 12 to wet all surfaces thereof.

Description: Surge characteristics of the XJ34-WE-32 turbojet engine were determined over a range of altitudes. Several typical oscillograph traces during which surge occurred are presented. The effect of altitude on the surge line and it's relation to the steady-state operating region are shown.

Description: An investigation of the nature of the flow field behind a rectangular circular-arc wing has been conducted in the Langley 9-inch supersonic tunnel. Pitot- and static-pressure surveys covering a region of flow behind the wing have been made together with detailed pitot surveys throughout the region of the wake. In addition, the flow direction has been measured using a weathercocking vane measurements. Theoretical calculations of the variation of both downwash and sidewash with angle of attack using Lagerstrom's superposition method have been made. In addition the effect of the wing thickness on the sidewash with the wing at 0 angle of attack has been evaluated. Near an angle of attack of 0, agreement between theory and experiment is good, particularly for the downwash results, except in the plane of the wing, inboard of the tip. In this region the proximity of the shed vortex sheet and the departure of the spanwise distribution of vorticity from theory would account for the disagreement. At higher angles of attack prediction of downwash depends on a knowledge of the location of the trailing vortex sheet, in order that the downwash may be corrected for its displacement and distortion. The theoretical location of the trailing vortex sheet, based on the theoretical downwash values integrated downstream from the wing trailing edge, is shown to differ widely from the experimental case. The rolling-up of the trailing vortex sheet behind the wing tip is evidenced by both the wake surveys and the flow-angle measurements.

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: A cooling assembly for fuel cells having a simplified construction whereby coolant is efficiently circulated through a conduit arranged in serpentine fashion in a channel within a member of such assembly. The channel is adapted to cradle a flexible, chemically inert, conformable conduit capable of manipulation into a variety of cooling patterns without crimping or otherwise restricting of coolant flow. The conduit, when assembled with the member, conforms into intimate contact with the member for good thermal conductivity. The conduit is non-corrodible and can be constructed as a single, manifold-free, continuous coolant passage means having only one inlet and one outlet. The conduit has an internal coil means which enables it to be bent in small radii without crimping.

Description: Means for sealing the tip 18 of a rotor turbine blade 10 against tip leakage flow comprising a multiplicity of recesses 30 formed in the surface of the tip 18. The recesses 30 are preferably formed in a labyrinthine or slaggered pattern which interposes at least one recess 30 in every leakage flow path across the tip 18 from the pressure side 26 to the suction side 28 of the blade 10.

Description: Among the concepts being considered for future Earth-to-orbit transport vehicles are fully reusable single-stage systems which take off vertically and land horizontally. Because these vehicles carry their own propellant internally, they are much larger than the present Space Shuttle Orbiter. One such single-stage vehicle under study is the circular body configuration which has the advantages of simple structural design and large volume-to-weight ratio. As part of an overall evaluation of this configuration, a series of heat transfer and surface flow tests were conducted. The phase-change paint and oil-flow tests were performed in the Langley 31-Inch Mach-10 Tunnel at angles of attack from 20 through 40 degrees in 5-degree increments. Heat-transfer coefficient data are presented for all angles of attack and detailed oil-flow photographs are shown for windward and leeward surfaces at 25 and 40 degrees angle of attack. In many ways, heating was similar to that previously determined for the Shuttle Orbiter so that, in a cursory sense, existing thermal protection systems would appear to be adequate for the proposed circular-body configurations.

Description: The present status of available information relative to the prediction of shock-induced boundary-layer separation is discussed. Experimental results showing the effects of Reynolds number and Mach number on the separation of both laminar and turbulent boundary layer are given and compared with available methods for predicting separation. The flow phenomena associated with separation caused by forward-facing steps, wedges, and incident shock waves are discussed. Applications of the flat-plate data to problems of separation on spoilers, diffusers, and scoop inlets are indicated for turbulent boundary layers.

Description: A plasma electrothermal thruster includes a capillary passage in which a plasma discharge is formed and directed out of an open end of the passage into a supersonic nozzle. Liquid supplied to the capillary passage becomes partially atomized to cool a confining surface of the passage. The plasma discharge is formed as the atomized liquid flows out of the open end into a supersonic equilibrium nozzle. The discharge can have a duration greater than the two way travel time of acoustic energy in the capillary to cause the plasma to flow continuously through the nozzle during the time of the discharge pulse.

Description: An improved heat pipe with an external artery. The longitudinal slot in the heat pipe wall which interconnects the heat pipe vapor space with the external artery is completely filled with sintered wick material and the wall of the external artery is also covered with sintered wick material. This added wick structure assures that the external artery will continue to feed liquid to the heat pipe evaporator even if a vapor bubble forms within and would otherwise block the liquid transport function of the external artery.

Description: The theory of Taylor and Maccoll (Ref,1) gives the surface pressure on an infinite cone in supersonic flow as a function of the cone vertex angle and the free stream Mach number and static pressure for a gas of vanishing viscosity. When a slender conical probe is used together with an impact pressure probe to determine the static pressure and Mach number in a low density gas stream, it is desirable to have some theoretical estimate of the effect of viscous boundary layer on the probe readings. Theoretical and experimental results with respect to impact probes have been presented in Refs. 5 and 6. A simple approximation for a conical probe based on linearized supersonic flow and compressible boundary layer theory is presented here.

Description: A rotor stage disposed within an annular fluid flowpath in a turbomachine is disclosed. The rotor stage includes outer and inner rings and a plurality of first and second airfoils. The outer and inner rings define flowpath surfaces for said flowpath. The first and second airfoils extend between the rings for transferring energy between the fluid and the rotor stage. At least one of the first airfoils surrounds means for transmitting energy across the flowpath.

Description: Fluidics flow control of a multiphase supply using a cylindrical chamber is achieved by introducing the supply flow radially into the chamber. The supply flow exits through a port in the center at the chamber. A control fluid is then introduced tangentially about 90.degree. upstream from the supply port. A second control fluid port may be added about 90.degree. upstream from the first control fluid port, but preferably two sets of supply and control ports are added with like ports diametrically opposite each other. The control fluid flows against the circular wall of the control chamber, which introduces a vortex in the flow of the supply flow that decays into a spiral path to the exit port in the center of the chamber. The control flow rate may thus be used to control the spiral path, and therefore the supply flow rate through the exit port.

Description: Systems are described for using multiple closely-packed spheres. In one system for passing fluid, a multiplicity of spheres lie within a container, with all of the spheres having the same outside diameter and with the spheres being closely nested in one another to create multiple interstitial passages of a known size and configuration and smooth walls. The container has an inlet and outlet for passing fluid through the interstitial passages formed between the nested spheres. The small interstitial passages can be used to filter out material, especially biological material such as cells in a fluid, where the cells can be easily destroyed if passed across sharp edges. The outer surface of the spheres can contain a material that absorbs a constitutent in the flowing fluid, such as a particular contamination gas, or can contain a catalyst to chemically react the fluid passing therethrough, the use of multiple small spheres assuring a large area of contact of these surfaces of the spheres with the fluid. In a system for storing and releasing a fluid such as hydrogen as a fuel, the spheres can include a hollow shell containing the fluid to be stored, and located within a compressable container that can be compressed to break the shells and release the stored fluid.

Description: A high efficiency forced air heat sink assembly employs a split feed transverse flow configuration to minimize the length of the air flow path through at least two separated fin structures. Different embodiments use different fin structure material configurations including honeycomb, corrugated and serpentine. Each such embodiment uses a thermally conductive plate having opposed exterior surfaces; one for receiving a component to be cooled and one for receiving the fin structures. The serpentine structured fin embodiment employs a plurality of fin supports extending from the plate and forming a plurality of channels for receiving the fin structures. A high thermal conductivity bondant, such as metal-filled epoxy, may be used to bond the fin structures to either the plate or the fin supports. Dip brazing and soldering may also be employed depending upon the materials selected.

Description: The Direct Simulation Monte Carlo (DSMC) method is applied in this paper to the study of rarefied, hypersonic, reentry flows. The assumptions and simplifications involved with the treatment of ionization, free electrons and the electric field are investigated. A new method is presented for the calculation of the electric field and handling of charged particles with DSMC. In addition, a two-step model for electron impact ionization is implemented. The flow field representing a 10 km/sec shock at an altitude of 65 km is calculated. The effects of the new modeling techniques on the calculation results are presented and discussed.