Experimental investigation of multiple-jet liquid injection into hypersonic flow
Experimental investigation of multiple-jet liquid injection into hypersonic flow
Bushnell, D. M.↗
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An experimental investigation of the internal methane pressure in hydrogen attack
An experimental investigation of the internal methane pressure that is the driving force for bubble growth in hydrogen attack (HA) was done on pure iron (204 ppm C) and on two low carbon steels of slightly different compositions. The methane content N (c.c gas/g. material) in attacked specimens was measured by a vacuum extraction technique. The total void volume V (c.c) was determined from density measurements before and after HA exposure. The two values, N and V, were then used in an equation of state to calculate an average methane pressure P for the attack stages beyond a density loss (d.l.) greater than 0.05 pct. It was determined that N and P depend on hydrogen exposure conditions and the presence of traces of strong carbide forming alloying elements (in steel). They are independent of specimen size and grain size over a limited range. P varies as the bubble volume increases, showing a generally decreasing trend which brings it to values lower than calculated equilibrium pressures, although well within the same order of magnitude. Possible reasons for this behavior are discussed.
Natan, M.↗
Experimental investigation of freezing blowby in a copper/water heat pipe
An experimental investigation designed to evaluate and better define the overall characteristics of freezing blowby in a copper/water heat pipe was conducted. The results from various rates of restart heat addition and channel blockage, indicate that upon breakthrough the depressurization of the evaporator may result in an effective heat transport capacity far in excess of the steady-state transport limit. The resulting transient conditions imposed on the heat pipe by the effective increased heat transport capacity can cause a loss of liquid in the evaporator and potential dryout. Evidence is presented which indicates that in order to prevent either temporary or permanent dryout, sufficient liquid inventory must be present in the evaporator wicking structure to accommodate the increased transient thermal load and allow sufficient time for the capillary wicking structure to reprime.
Ochterbeck, J. M.↗
Experimental investigation of crossflow jet mixing in a rectangular duct
An experimental investigation of the mixing of nonreacting opposed rows of jets injected normal to a confined rectangular crossflow has been conducted. Planar Mie-scattering was used to measure the time-average concentration distribution of the jet fluid in planes perpendicular to the duct axis. The mixing effectiveness of round orifice injectors was measured as a function of orifice spacing and orifice diameter. Mixing effectiveness was determined using a spatial unmixedness parameter based on the variance of mean jet concentration distributions. Optimum mixing was obtained when the spacing-to-duct height ratio was inversely proportional to the square root of the jet-to-mainstream momentum-flux ratio. For opposed rows of round holes with centerlines inline, mixing was similar for blockages up to 75 percent. Lower levels of unmixedness were obtained as a function of downstream location when axial injection length was minimized. Mixing may be enhanced if orifice centerlines of opposed rows are staggered, but note that blockage must be less than 50 percent for this configuration.
Liscinsky, D. S.↗
Experimental investigation of crossflow jet mixing in a rectangular duct
An experimental investigation of the mixing of nonreacting opposed rows of jets injected normal to a confined rectangular crossflow has been conducted. Planar Mie-scattering was used to measure the time-average concentration distribution of the jet fluid in planes perpendicular to the duct axis. The mixing effectiveness of round orifice injectors was measured as a function of orifice spacing and orifice diameter. Mixing effectiveness was determined using a spatial unmixedness parameter based on the variance of mean jet concentration distributions. Optimum mixing was obtained when the spacing-to-duct height ratio was inversely proportional to the square root of the jet-to-mainstream momentum-flux ratio. For opposed rows of round holes with centerlines inline, mixing was similar for blockages up to 75 percent. Lower levels of unmixedness were obtained as a function of downstream location when axial injection length was minimized. Mixing may be enhanced if orifice centerlines of opposed rows are staggered, but note that blockage must be less than 50 percent for this configuration.
Liscinsky, D. S.↗
Experimental Investigation of a Preloaded Spring-tab Flutter Model
An experimental investigation was made of a preloaded spring-tab flutter model to determine the effects on flutter speed of aspect ratio, tab frequency, and preloaded spring constant. The rudder was mass-balanced, and the flutter mode studied was essentially one of three degrees of freedom (fin bending coupled with rudder and tab oscillations). Inasmuch as the spring was preloaded, the tab-spring system was a nonlinear one. Frequency of the tab was the most significant parameter in this study, and an increase in flutter speed with increasing frequency is indicated. At a given frequency, the tab of high aspect ratio is shown to have a slightly lower flutter speed than the one of low aspect ratio. Because the frequency of the preloaded spring tab was found to vary radically with amplitude, the flutter speed decreased with increase in initial displacement of the tab.
VIBRATION AND FLUTTER - TABS↗
An Experimental Investigation of Hollow Turbine Blades for Expandable Jet Engines
An experimental investigation was made to determine the feasibility of using hollow turbine blades in engines designed for short service life. Airfoils were fabricated from sheet material and techniques of welding and brazing attachment were investigated. The airfoils were not intended to be cooled. A principal objective was the reduction of strategic material requirements primarily as a function of direct weight reduction. The materials considered were in the density range of 0.28 to 0.31 pound per cubic inch. Three materials were used for airfoils: N-155, Inconel-X and L-605. These were attached to J47 turbine blade bases and operated to destruction at maximum service conditions of turbine speed and temperature. It was found that L-605 airfoils brazed to suitable bases satisfied the requirements for expendable engines. Service life varied from 11 to 40 hours.
Morgan, W C↗
An Experimental Investigation of the Flow Structure of Supersonic Impinging Jets
An experimental investigation into the jet structure associated with sound production by a supersonic impinging jet is presented. Large plate impinging tones are investigated for a nozzle pressure ratio (NPR) of 4 and nozzle-to-plate spacings between 1 and 5 nozzle exit diameters, where NPR is equal to the ratio of the stagnation pressure to the pressure at the nozzle lip. Results from phase-locked shadowgraph and phase-averaged digital particle image velocimetry (DPIV) studies indicate that, during the oscillation cycle, the Mach disk oscillates axially, a well defined recirculation zone is created in the subsonic impingement region and moves toward the plate, and the compression and expansion regions in the outer supersonic flow move downstream, Sound appears to be generated in the wall jet at approximately 2.6R from the jet axis, where R is the nozzle exit radius. The oscillatory motion in the wall jet is the result of the periodic fluid motion in the near wall region.
Henderson, Brenda↗
Experimental Investigation of Convoluted Contouring for Aircraft Afterbody Drag Reduction
An experimental investigation was performed in the NASA Langley 16-Foot Transonic Tunnel to determine the aerodynamic effects of external convolutions, placed on the boattail of a nonaxisymmetric nozzle for drag reduction. Boattail angles of 15 and 22 were tested with convolutions placed at a forward location upstream of the boattail curvature, at a mid location along the curvature and at a full location that spanned the entire boattail flap. Each of the baseline nozzle afterbodies (no convolutions) had a parabolic, converging contour with a parabolically decreasing corner radius. Data were obtained at several Mach numbers from static conditions to 1.2 for a range of nozzle pressure ratios and angles of attack. An oil paint flow visualization technique was used to qualitatively assess the effect of the convolutions. Results indicate that afterbody drag reduction by convoluted contouring is convolution location, Mach number, boattail angle, and NPR dependent. The forward convolution location was the most effective contouring geometry for drag reduction on the 22 afterbody, but was only effective for M < 0.95. At M = 0.8, drag was reduced 20 and 36 percent at NPRs of 5.4 and 7, respectively, but drag was increased 10 percent for M = 0.95 at NPR = 7. Convoluted contouring along the 15 boattail angle afterbody was not effective at reducing drag because the flow was minimally separated from the baseline afterbody, unlike the massive separation along the 22 boattail angle baseline afterbody.
Deere, Karen A.↗
Experimental Investigation of Ice Accretion Effects on a Swept Wing
An experimental investigation was conducted to study the effects of 2-, 5-, 10-, and 22.5-min ice accretions on the aerodynamic performance of a swept finite wing. The ice shapes tested included castings of ice accretions obtained from icing tests at the NASA Glenn Icing Research Tunnel (IRT) and simulated ice shapes obtained with the LEWICE 2.0 ice accretion code. The conditions used for the icing tests were selected to provide five glaze ice shapes with complete and incomplete scallop features and a small rime ice shape. The LEWICE ice shapes were defined for the same conditions as those used in the icing tests. All aerodynamic performance tests were conducted in the 7- x 10-ft Low-Speed Wind Tunnel Facility at Wichita State University. Six component force and moment measurements, aileron hinge moments, and surface pressures were obtained for a Reynolds number of 1.8 million based on mean aerodynamic chord and aileron deflections in the range of -15o to 20o. Tests were performed with the clean wing, six IRT ice shape castings, seven smooth LEWICE ice shapes, and seven rough LEWICE ice shapes. Roughness for the LEWICE ice shapes was simulated with 36-size grit. The experiments conducted showed that the glaze ice castings reduced the maximum lift coefficient of the clean wing by 11.5% to 93.6%, while the 5-min rime ice casting increased maximum lift by 3.4%. Minimum iced wing drag was 133% to 3533% greater with respect to the clean case. The drag of the iced wing near the clean wing stall angle of attack was 17% to 104% higher than that of the clean case. In general, the aileron remained effective in changing the lift of the clean and iced wings for all angles of attack and aileron deflections tested. Aileron hinge moments for the iced wing cases remained within the maximum and minimum limits defined by the clean wing hinge moments. Tests conducted with the LEWICE ice shapes showed that in general the trends in aerodynamic performance degradation of the wing with the simulated ice shapes were similar to those obtained with the IRT ice shape castings. However, in most cases, the ice castings resulted in greater aerodynamic performance losses than those obtained with the LEWICE ice shapes. For the majority of the LEWICE ice shapes, the addition of 36-size grit roughness to the smooth ice shapes increased aerodynamic performance losses.
Ice shape↗
An experimental investigation of supersonic flow past a wedge-cylinder configuration
An experimental investigation of supersonic flow past double-wedge configurations was conducted. Over the range of geometries tested, it was found that, while theoretical solutions both for a Type V pattern and for a Type VI pattern could be generated for a particular flow condition (as defined by the geometry and the free-stream conditions), the weaker, Type VI pattern was observed experimentally. More rigorous flow-field solutions were developed for the flow along the wing leading-edge. Solutions were developed for the three-dimensional flow in the plane of symmetry of a swept cylinder (which represented the wing leading-edge) which was mounted on a wedge (which generated the "bow" shock wave). A numerical code was developed using integral techniques to calculate the flow in the shock layer upstream of the interaction region (i.e., near the wing root). Heat transfer rates were calculated for various free stream conditions. The present investigation was undertaken to examine the effects of crossflow on the resultant flow-field and to verify the flow model used in theoretical calculations.
Barnette, D. W.↗
An experimental investigation of the radial displacements of a thin-walled cylinder
Experimental investigation of thin walled cylinder radial displacements
THIN WALL↗
Experimental investigation of model variable-geometry and ogee tip rotors.
An experimental investigation was conducted to systematically explore the effects of inter-blade spatial relationships and pitch variations on rotor performance and wake geometry. Variable-geometry rotors consisting of various combinations of blade length, axial spacing, azimuth spacing, and collective pitch were tested at model scale in hover and forward flight. In addition, a hover test of a rotor with an ogee blade tip design was conducted to determine its performance and wake characteristics. The results of this investigation indicate that several variable-geometry rotor configurations can offer substantial improvements in hover performance without adversely affecting forward flight performance.
Landgrebe, A. J.↗
Experimental investigation of model variable-geometry and ogee tip rotors
An experimental investigation was conducted to systematically explore the effects of inter-blade spatial relationships and pitch variations on rotor performance and wake geometry. Variable-geometry rotors consisting of various combinations of blade length, axial spacing, azimuth spacing, and collective pitch were tested at model scale in hover and forward flight. In addition, a hover test of a model rotor with an ogee blade tip design was conducted to determine its performance and wake characteristics. The results of this investigation indicate that properly selected variable geometry rotor configurations can offer substantial improvements in hover performance without adversely affecting forward flight performance. Axial spacing of alternate blades was found to provide the greatest performance benefit, and further improvements were achieved by combining azimuth spacing with axial spacing. The performance benefit appears to be related to the relief of local adverse aerodynamic phenomena produced by vortex interference. The ogee tip design was found to substantially reduce the concentrated core intensity of the tip vortex, and could thus prove beneficial for the relief of blade-vortex interaction problems. However, the ogee tip was found to reduce hover performance at model scale.
Landgrebe, A. J.↗
An experimental investigation of reducing advanced turboprop cabin noise by wing shielding
An experimental investigation was undertaken to determine if wing shielding could reduce the noise impacting the fuselage of an advanced turboprop airplane. Four wings were tested behind two eight-bladed propeller models. Significant shielding of the propeller noise was observed and a particular wing-propeller geometry was identified to provide the most shielding. Specifically, an up-inboard rotation would be needed for a low-wing airplane and a down-inboard rotation for a high-wing airplane. As the axial Mach number was increased, the position where the shielding starts moved farther downstream. This shift in the start of shielding was roughly a straight line with respect to Mach number between M = 0.7 and M = 0.8. At M = 0.85 the start of shielding does not shift any farther downstream. A simple barrier noise-reduction model gave the same trends with transducer positions as did the data, and, if corrected for Mach number shift, the model might be used to provide estimates of the wing shielding. Besides providing a barrier to the noise reaching the shielded area, the wing also reflects some of the noise back onto the unshielded area. This can make the noise difference between the unshielded and shielded areas of the fuselage larger than would be expected by simple wind shielding.
Dittmar, J. H.↗
An experimental investigation of reducing advanced turboprop cabin noise by wing shielding
An experimental investigation was undertaken to determine if wing shielding could reduce the noise impacting the fuselage of an advanced turboprop airplane. Four wings were tested behind two eight-bladed propeller models. Significant shielding of the propeller noise was observed and a particular wing-propeller geometry was identified to provide the most shielding. Specifically, an up-inboard rotation would be needed for a low-wing airplane and a down-inboard rotation for a high-wing airplane. As the axial Mach number was increased, the position where the shielding starts moved farther downstream. This shift in the start of shielding was roughly a straight line with respect to Mach number between M = 0.7 and M = 0.8. At M = 0.85 the start of shielding does not shift any farther downstream. A simple barrier noise-reduction model gave the same trends with transducer positions as did the data, and, if corrected for Mach number shift, the model might be used to provide estimates of the wing shielding. Besides providing a barrier to the noise reaching the shielded area, the wing also reflects some of the noise back onto the unshielded area. This can make the noise difference between the unshielded and shielded areas of the fuselage larger than would be expected by simple wind shielding.
Dittmar, J. H.↗
Experimental Investigation of Stage Performance of J71 Three-stage Turbine
An experimental investigation of the stage performance of the J71 three-stage turbine at design speed and pressure ratio was conducted as part of a program involving research on the design and operational characteristics of high-work-output multistage axial-flow turbines. The investigation indicated that the over-all work output of the turbine based on temperature measurements at design speed and pressure ratio was 95 percent of the design value, with the greatest loss in work output and efficiency occurring in the first stage. Choking in the second-stage stator over most of the pressure ratios investigated prevented the attainment of design first-stage work output and also prevented any increase in work output of the first stage with increasing over-all pressure ratio at design speed. High losses in the first stage were a contributing factor to this choking condition.
Forrette, Obert E↗
An Experimental Investigation of Hypergolic Ignition Delay of Hydrogen Peroxide with Fuel Mixtures
An experimental investigation of hypergolicity and ignition delay of fuel mixtures with hydrogen peroxide is presented. Example results of high speed photography and schleiren from drop tests are shown. Also, a discussion of the sensitivity to experimental parameters such as drop size and subsequent uncertainty considerations of ignition delay results is presented. It is shown that using the described setup on the mixtures presented, the precision uncertainty is on the order of 6% of average ignition delay and 5% of average decomposition delay. This represents sufficient repeatability for first order discrimination of ignition delay for propellant development and screening. Two mixtures, each using commonly available amines and transition metal compounds, are presented as examples that result in ignition delays on the order of 10 milliseconds.
Blevins, John A.↗