ALBERT

Description: University of Miami utilized NASA's spacecraft viewing technology to develop the optical profilometer provides more accurate measurements of cleft palate casts than has heretofore been possible, enabling better planning of corrective surgery. Lens like instrument electronically scans a palate cast precisely measuring its irregular contours by detecting minute differences in the intensity of a light beam reflected off the cast. Readings are computer processed and delivered to the surgeon by a teleprinter.

Description: Increased performance of rocket engine using fluorine-oxygen mixture with RP-1 fuel

Description: Supersonic flow of two turbine blade sections with turning angle of 160-degrees

Description: Cascade investigation of cooling characteristics of air-cooled turbine blade for use in turboprop engine

Description: Addition of liquid fluorine to liquid oxygen in Vanguard first stage oxygen-hydrocarbon rocket engine

Description: Thermal stability and decomposition products of propyl pentaborane heated at 147 to 190 deg C

Description: Combustion efficiency of hydrogen fuel for varying afterburner configurations

Description: Combustion efficiency of pentaborane fuel in 48- inch diameter ramjet engine

Description: Some results of recent experimental investigations at supersonic and transonic speeds are presented to show the present status in the estimation of load distributions on controls and adjacent wing surfaces resulting from the deflection of flap controls and spoiler controls. The results indicate that the development of methods for predicting loads associated with controls has not kept pace with the acquisition of experimental data. At low supersonic speeds sweeping the hinge line induces strong three-dimensional-flow characteristics which cannot be treated by the simplified methods previously developed for controls without sweep. At transonic speeds the estimation of loads associated with controls must usually be dependent upon experimental information inasmuch as the latest attempts to predict chordwise and spanwise loadings have met with only limited success.

Description: Boundary-layer-transition and heat-transfer measurements were obtained from flight tests of blunt and sharp cones having apex angles of 50 deg. The test Mach number range was from 1.7 to 4.7, corresponding to free-stream Reynolds numbers, based on cone base diameter, of 18. 3 x 10(exp 6) and 32.1 x 10(exp 6), respectively. Transition on both models occurred at a local Reynolds number of 1 x 10(exp 6) to 2 X 10(exp 6) based on distance from the stagnation point. Transition Reynolds numbers based on momentum thickness were between 320 and 380 for the blunt cone. The model surface roughness was 25 rms microinches or greater. Turbulent heat transfer to the conical surface of the blunt cone at a Mach number of 4 was 30 percent less than that to the surface of the sharp cone. Available theories predicted heat-transfer coefficients reasonably well for the fully laminar or turbulent flow conditions.

Description: Skin temperatures and surface pressures have been measured on a slightly blunted cone-cylinder-flare configuration to a maximum Mach number of 9.89 with a rocket-propelled model. The cone had a t o t a l angle of 25 deg and the flare had a 10 deg half-angle. Temperature data were obtained at eight cone locations, four cylinder locations, and seven flare locations; pressures were measured at one cone location, one cylinder location, and three flare locations. Four stages of propulsion were utilized and a reentry type of trajectory was employed in which the high-speed portion of flight was obtained by firing the last two stages during the descent of the model from a peak altitude of 99,400 feet. The Reynolds number at peak Mach number was 1.2 x 10(exp 6) per foot of model length. The model length was 6.68 feet. During the higher speed portions of flight, temperature measurements along one element of the nose cone indicated that the boundary layer was probably laminar, whereas on the opposite side of the nose the measurements indicated transitional or turbulent flow. Temperature distributions along one meridian of the model showed the flare to have the highest temperatures and the cylinder generally to have the lowest. A maximum temperature of 970 F was measured on the cone element showing the transitional or turbulent flow; along the opposite side of the model, the maximum temperatures of the cone, cylinder, and flare were 545 F, 340 F, and 680 F, respectively, at the corresponding time.

Description: The multistage turbine from the J73 turbojet engine has previously been investigated with standard and with reduced-chord rotor blading in order to determine the individual performance characteristics of each configuration over a range of over-all pressure ratio and speed. Because both turbine configurations exhibited peak efficiencies of over 90 percent, and because both units had relatively wide efficient operating ranges, it was considered of interest to determine the performance of the first stage of the turbine as a separate component. Accordingly, the standard-bladed multistage turbine was modified by removing the second-stage rotor disk and stator and altering the flow passage so that the first stage of the unit could be operated independently. The modified single-stage turbine was then operated over a range of stage pressure ratio and speed. The single-stage turbine operated at a peak brake internal efficiency of over 90 percent at an over-all stage pressure ratio of 1.4 and at 90 percent of design equivalent speed. Furthermore, the unit operated at high efficiencies over a relatively wide operating range. When the single-stage results were compared with the multistage results at the design operating point, it was found that the first stage produced approximately half the total multistage-turbine work output.

Description: The heat transfer and pressures on the surfaces of several flat-plate models with various external crosswise stiffener arrangements are presented. The tests were made in a free jet at Mach numbers of 0.77, 1.39, and 1.98 for Reynolds numbers of 3 x 10(exp 6), 7 x 10(exp 6), and 14 x 10(exp 6), respectively, based on a length of 1 foot. The addition of external crosswise stiffeners to the flat-plate models caused large pressure and heat-transfer variations on the surfaces of the models.

Description: A temperature sensor comprising an optical transducer member having an array of birefringent crystals. The length and, accordingly, the sensitivity to temperature change of successive birefringent crystals varies according to a particular relationship. The transducer is interconnected with a fiber optic transmission and detecting system. Respective optical output signals that are transmitted from the birefringent crystals via the fiber optic transmission system are detected and decoded so as to correspond to digits of a numbering system, whereby an accurate digital representation of temperature can ultimately be provided.

Description: An inexpensive, lightweight fiber optic micro-sensor that is suitable for applications which may require remote temperature sensing. The disclosed temperature sensor includes a phosphor material that, after receiving incident light stimulation, is adapted to emit phosphorescent radiation output signals, the amplitude decay rate and wavelength of which are functions of the sensed temperature.

Description: Rotating-stall limits in supersonic turbofan engine

Description: In the investigation of stability of a two-dimensional laminar flow with respect to small disturbances, a disturbance of the stream function moving downstream (in the direction of the x-axis) by the "partial wave formula" is described.

Description: The literature on turbulent heat transfer has in the course of years attained a considerable volume. Since this very complicated problem has not as yet found a complete solution, further studies in this field may be expected. The heat engineer must therefore accomodate himself to a constantly increasing number of theories and formulas. Since the theories generally start from hypothetical assumptions, and since they contain true and false assertions, verified knowledge and pure suppositions often being intermingled in a manner difficult to tell them apart, the specialist had difficulty in forming a correct evaluation of the individual studies. The need therefore arises for a presentation of the problem of turbulent heat transfer which is not initially bound by hypothetical assumptions and in which uninvestigated can be clearly distinguished form each other. Such a presentation will be given in the present treatment. Brief remarks with regard to the development of the theory of local heat transfer are included.

Description: The Mass Flowmeter is a device used to measure flow of oxygen in spacecraft's life support system. Tylan Corporation's Mass Flow Controller's major application is accurate control of reactive gases-- such as hydrogen, phosphine and silane as they are diffused at extremely high temperatures into silicon wafers. Wafers are ultimately cut up into integrated circuits or "chips" for electronic products. Precise process control afforded by the Mass Flow Controller makes it possible to produce circuit chips of greater performance at lower cost.

Description: Measurements of aerodynamic heat transfer have been made along the hemisphere and cylinder of a hemisphere-cylinder rocket-propelled model in free flight up to a Mach number of 3.88. The test Reynolds number based on free-stream condition and diameter of model covered a range from 2.69 x l0(exp 6) to 11.70 x 10(exp 6). Laminar, transitional, and turbulent heat-transfer coefficients were obtained. The laminar data along the body agreed with laminar theory for blunt bodies whereas the turbulent data along the cylinder were consistently lower than that predicted by the turbulent theory for a flat plate. Measurements of heat transfer at the stagnation point were, in general, lower than the theory for stagnation-point heat transfer. When the Reynolds number to the junction of the hemisphere-cylinder was greater than 6 x l0(exp 6), the transitional Reynolds number varied from 0.8 x l0(exp 6) to 3.0 x 10(exp 6); however, than 6 x l(exp 6) when the Reynolds number to the junction was less, than the transitional Reynolds number varied from 7.0 x l0(exp 6) to 24.7 x 10(exp 6).

Description: Internal performance of an XJ79-GE-1 variable ejector was experimentally determined with the primary nozzle in two representative after-burning positions. Jet-thrust and air-handling data were obtained in quiescent air for 4 selected ejector configurations over a wide range of secondary to primary airflow ratios and primary-nozzle pressure ratios. The experimental ejector data are presented in both graphical and tabulated form.

Description: The unsteady aerodynamic forces, [based on two-dimensional incompressible flow considerations], are determined for an unstaggered cascade, the blades of which are vibrating in phase in an approach flow parallel to the blades.

Description: Nine divergent-shroud ejector configurations were investigated to determine the effect of shroud divergence angle on ejector internal performance. Unheated dry air was used for both the primary and secondary flows. The decrease in the design-point thrust coefficient with increasing flow divergence angle (angle measured from primary exit to shroud exit) followed very closely a simple relation involving the cosine of the angle. This indicates that design-point thrust performance for divergent-shroud ejectors can be predicted with reasonable accuracy within the range investigated. The decrease in design-point thrust coefficient due to increasing the flow divergence engle from 120deg to 30deg (half-singles) was approximately 6 percent. Ejector air-handling characteristics and the primary-nozzle flow coefficient were not significantly affected by change in shroud divergence angle.

Description: The net heat of combustion of the product formed by the reaction of diborane with a mixture of acetylene and ethylene was found to be 20,440 +/- 150 Btu per pound for the reaction of liquid fuel to gaseous carbon dioxide, gaseous water, and solid boric oxide. The measurements were made in a Parr oxygen-bomb calorimeter, and the combustion was believed to be 98 percent complete. The estimated net-heat of combustion for complete combustion would therefore be 20,850 +/- 150 Btu per pound.

Description: The heat of combustion of the product formed by the reaction acetylene and diborane was found to be 20,100 +/- 100 Btu per pound for the reaction of liquid fuel to gaseous carbon dioxide, gaseous water, and solid boric oxide. The measurements were made in a Parr oxygen-bomb calorimeter, and chemical analyses both of the sample and of the combustion products indicated combustion in the bomb calorimeter to have been 97 percent complete. The estimated net heat of combustion for complete combustion would therefore be 20,700 +/- 100 Btu per pound.

Description: A two-spool turbojet engine was operated in the Lewis altitude wind tunnel to study the inception of compressor surge. In addition to the usual steady-state pressure and temperature measurements, the compressors were extensively instrumented with fast-response interstage pressure transducers. Thus it was possible to obtain maps for both compressors, pressure oscillations during rotating stall, effects of stall on efficiency, and stage-loading curves. In addition, with the transient measurements, it was possible to record interstage pressures and then compute stage performance during accelerations to the stall limit. Rotating stall was found to exist at low speeds in the outer spool. Although the stall arose from poor flow conditions at the inlet-stage blade tips, the low-energy air moved through the machine from the tip at the inlet to the outer spool to the hub at the inlet to the inner spool. This tip stall ultimately resulted in compressor surge in the mid-speed region, and necessitated inter-compressor air bleed. Interstage pressure measurements during acceleration to the compressor stall limit indicated that rotating stall was not a necessary condition for compressor surge and that, at the critical stall point, the circumferential interstage pressure distribution was uniform. The exit-stage group of the inner spool was first t o stall; then, the stages upstream stalled in succession until the inlet stage of the outer spool was stalled. With a sufficiently high fuel rate, the process repeated with a cycle time of about 0.1 second. It was possible to construct reproducible stage stall lines as a function of compressor speed from the stage stall points of several such compressor surges. This transient stall line was checked by computing the stall line from a steady-state stage-loading curve. Good agreement between the stage stall lines was obtained by these two methods.

Description: A gas turbine engine combustor assembly of annular configuration has outer and inner walls made up of a plurality of axially extending multi-layered porous metal panels joined together at butt joints therebetween and each outer and inner wall including a transition panel of porous metal defining a combustor assembly outlet supported by a combustor mount assembly including a stiffener ring having a side undercut thereon fit over a transition panel end face; and wherein an annular weld joins the ring to the end face to transmit exhaust heat from the end face to the stiffener ring for dissipation from the combustor; a combustor pilot member is located in axially spaced, surrounding relationship to the end face and connector means support the stiffener ring in free floating relationship with the pilot member to compensate for both radial and axial thermal expansion of the transition panel; and said connector means includes a radial gap for maintaining a controlled flow of coolant from outside of the transition panel into cooling relationship with the stiffener ring and said weld to further cool the end face against excessive heat build-up therein during flow of hot gas exhaust through said outlet.

Description: To provide combustion stability in burners over a wide range of flow conditions, a coaxial injector is provided which includes an outer annular sleeve for introduction of one reactant and a hollow pintle with spaced slots for introduction of another reactant to impinge on the first reactant. The slots are staggered in location and size to permit initial interlocking of the reactants and, therefore, increased mixing and improved combustion performance.

Description: A gas turbine engine combustor assembly of unique configuration has an outer wall made up of a plurality of axially extending multi-layered porous metal panels joined together at butt joints therebetween by a reinforcing and heat dissipation ring and a unique weld configuration to prevent thermal erosion of the ends of the porous metal panels at the butt joints; the combustor further including a unique inner wall made up of a plurality of like axially extending multi-layered porous metal panels joined at butt joints by a reinforcing and heat dissipation ring on the inner surface of the inner wall panels and an improved butt weld joint that prevents thermal erosion of the ends of the porous metal inner wall panels.

Description: The performance characteristics of the 19B-8 and 19XB-1 turbojet engines and the windmilling-drag characteristics of the 19B-6 engine were determined in the Cleveland altitude wind tunnel. The investigations were conducted on the 19B-8 engine at simulated altitudes from 5000 to 25,000 feet with various free-stream ram-pressure ratios and on the 19XB--1 engine at simulated altitudes from 5000 to 30,000 feet with approximately static free-stream conditions. Data for these two engines are presented to show the effect of altitude, free-stream ram-pressure ratio, and tail-pipe-nozzle area on engine performance. A 21-percent reduction in tail-pipe-nozzle area of the 19B-8 engine increased the let thrust 43 percent the net thrust 72 percent, and the fuel consumption 64 percent. An increase in free-stream ram-pressure ratio raised the jet thrust and the air flow and lowered the net thrust throughout the entire range of engine speeds for the 19B-8 engine. At similar operating conditions, the corrected jet thrust and corrected air flow were approximately the same for both engines, and the corrected specific fuel consumption based on jet thrust was lower for the 19XB-1 engine than for the 19B-8 engine. The thrust and air-flow data obtained with both engines at various altitudes for a given free-stream rampressure ratio were generalized to standard sea-level atmospheric conditions. The performance parameters involving fuel consumption generalized only at high engine speeds at simulated altitudes as high as 15,000 feet. The windmilling drag of the 19B-8 engine increased rapidly as the airspeed was increased.

Description: In dusty air flows occurring in industrial practice in transport by air pressure of friable materials, in the drying, annealing, and so forth, of a pulverized solid mass in suspension, and in other processes, the concentration of solid particles usually has a magnitude of the order of 1 kg per 1 kg of air. At such a concentration, the ratio of the volume of the particles to the volume of the air is small (less than one-thousandth part). However, regardless of this, the presence of a solid admixture manifests itself in the rules for the velocity distribution of the air in a dusty air flow. As a result, the rules of velocity change are different for clean and for dusty air flows. The estimation of the influence of the admixture on the velocity of the motion of the flow presents a definitive interest. One of the attempts to estimate that influence on the axial velocity of a free axially symmetrical jet with admixtures was made by Abramovich. Abramovich assumed beforehand that the fine particles of the admixture in the jet are subject to the motion of the air (that is, that the velocity of the admixture is approximately equal to the local velocity of the air); he then took as the basis of his considerations, in solving the problem, the condition that the amount of motion of the two-phase jet must be constant.