Search NASA⌕ Search

SEARCH · Search NASA

Results for “plenum”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 415 records · Page 23
All sourcesNASA NTRSDOE OSTI
NASA NTRS1977

Tubular sublimatory evaporator heat sink

An evaporative refrigerator or cooler comprising a bundle of spaced, porous walled tubes closed at one of their ends and vented to a vacuum at the other end is disclosed. The tube bundle is surrounded by a water jacket having a hot water inlet distribution manifold and a cooled water outlet through a plenum chamber. Hot water is pumped into the jacket to circulate around the tubes, and when this water meets the vacuum existing inside the tubes, it evaporates thereby cooling the water in the jacket. If cooling proceeds to the point where water penetrating or surrounding all or part of the tubes freezes, operation continues with local sublimation of the ice on the tubes while the circulating water attempts to melt the ice. Both sublimation and evaporation may take place simultaneously in different regions of the device.

Webbon, B. W.↗
NASA NTRS1978

Design and Performance Evaluation of Slotted Walls for Two-Dimensional Wind Tunnels

A procedure for designing slotted walls for two dimensional wind tunnels is is presented. The design objective is the minimization of blockage of streamline curvature or the reduction of both. The slotted wall boundary condition is derived both for flow from the tunnel into the plenum and vice versa, and the procedure for evaluating wall interference is described. A correlation of experimental data for the slotted wall boundary condition is given. Results are given for several designs and evaluations of slotted wind tunnel walls.

Barnwell, R. W.↗
NASA NTRS1978

A procedure for computation of fully stalled flows in two-dimensional passages

A procedure is described for computation of incompressible, steady, two-dimensional flows in fully stalled diffusers with plenum exit. The procedure is successful in predicting pressure distributions and patterns to the accuracy of the data. The procedure employs a zonal model; this maintains close connections between the modeling and the physics thereby providing insight into critical aspects of modeling separated flows. The procedure presented is also convenient for computing unstalled flows in passages with turbulent boundary layers for either direct or indirect design problems. Computing times are well within engineering feasibility. The concepts developed can be extended to other classes of separated flows; some of these extensions have already been completed and are referenced.

Woolley, R. L.↗
NASA NTRS1979

Self stabilizing sonic inlet

An inlet suitable for a turbine engine in a STOL, VTOL or CTOL aircraft is described. A circumferentially extended slot is provided in the inner surface of the air inlet at the windward side and downstream of the throat region. The slot communicates with a circumferential plenum chamber formed in the front of the air inlet just behind the lip. Circumferentially extending rows of apertures are provided on the lip establishing two sets of apertures spaced circumferentially away from the slot in opposite directions. The slot removes the boundary layer from the critical portion of the diffuser to minimize or eliminate flow pressure loss or separations resulting from diffusion or tuning. The apertures are in a region of low static pressure on the lip of the inlet and serve as a source of suction to cause air flow into the slot.

Miller, B. A.↗
NASA NTRS1980

Spectral properties of gaseous uranium hexafluoride at high temperature

A study to determine relative spectral emission and spectral absorption data for UF6-argon mixtures at elevated temperatures is discussed. These spectral data are required to assist in the theoretical analysis of radiation transport in the nuclear fuel-buffer gas region of a plasma core reactor. Relative emission measurements were made for UF6-argon mixtures over a range of temperatures from 650 to 1900 K and in the wavelength range from 600 to 5000 nanometers. All emission results were determined for a total pressure of 1.0 atm. Uranium hexafluoride partial pressures varied from about 3.5 to 12.7 mm Hg. Absorption measurements were attempted at 600, 625, 650 and 675 nanometers for a temperature of 1000 K. The uranium partial pressure for these determinations was 25 mm Hg. The results exhibit appreciable emission for hot UF6-argon mixtures at wavelengths between 600 and 1800 nanometers and no measurable absorption. The equipment used to evaluate the spectral properties of the UF6-argon mixtures included a plasma torch-optical plenum assembly, the monochromator, and the UF6 transfer system. Each is described.

Krascella, N. L.↗
NASA NTRS1980

Reduction of nitric oxide emissions from a combustor

A turbojet combustor and method for controlling nitric oxide emissions by employing successive combustion zones is described. After combustion of an initial portion of the fuel in a primary combustion zone, the combustion products of the primary zone are combined with the remaining portion of fuel and additional plenum air and burned in a secondary combustion zone under conditions that result in low nitric oxide emissions. Low nitric oxide emissions are achieved by a novel turbojet combustor arrangement which provides flame stability by allowing stable combustion to be accompanied by low nitric oxide emissions resulting from controlled fuel-lean combustion (ignited by the emission products from the primary zone) in a secondary combustion zone at a lower combustion temperature resulting in low emission of nitric oxide.

Craig, R. A.↗
NASA NTRS1980

Supersonic flow development in slotted wind tunnels

The development of test section slot shapes for achieving smooth supersonic Mach number distribution without overexpansion or waviness has, in the past, been largely an experimentally iterative or 'cutand try' procedure for each wind tunnel. To overcome the obvious disadvantages of time and expense involved in such an experimental approach, a simple analytical method has been developed to predict the supersonic flow development in a two-dimensional slotted tunnel given only the variation of open area ratio with downstream distance and the Mach number corresponding to the plenum static pressure. The well known method of characteristics is used with the constraint that it be compatible with the quadratic cross-flow pressure drop boundary condition at the slotted wall. The predicted results from this method agree remarkably well with the experimental calibration data available for some of the existing facilities. The flow mechanism responsible for causing overexpansion in the centerline Mach number distribution with some slot shapes has been brought to light.

Ramaswamy, M. A.↗
NASA NTRS1981

Effects of a modified leading edge on noise and boundary-layer transition in a rod-wall sound shield at Mach 5

A version of a rod wall sound shield was tested in the Mach 5 pilot quiet tunnel over a range of unit Reynolds numbers from 10 to 35 million per meter. The model was modified by inclining the leading edge plates to produce an initial 2 deg expansion to ascertain the sensitivity of boundary layer transition to leading edge disturbances. Rod surface pitot pressures, mean free stream pitot pressures, and static pressures on the rods and plenum walls were measured. Hot-wire measurements were also made in the model and nozzle free stream at a unit Reynolds number of 15 million per meter. The surface pitot pressures indicated that transition was much farther forward than for the previous tests due to the leading edge modification and minor fabrication flaws in the model. Early boundary layer transition on the rods was confirmed by hot-wire measurements which showed much higher noise levels in the free stream shield flow when compared with results from previous tests. Mean pitot pressure surveys within the shielded region inside the model indicated that there was an overexpansion and recompression that would limit the streamwise length of undisturbed flow to about 13 cm along the centerline.

Creel, T. R., Jr.↗
NASA NTRS1981

Description of recent changes in the Langley 6- by 28-inch transonic tunnel

Calibrations were obtained in the Langley 6 by 28-inch transonic tunnel with newly installed controllable reentry flaps and test section floor and ceiling. Using available theory, the top and bottom slotted walls were redesigned for minimum wind tunnel interference errors of blockage and stream-line curvature. To minimize Mach number gradients along the tunnel axis downstream of the model, controllable flaps were installed to regulate the flow reentering the test section through the slotted walls. The flap setting is independent of stagnation pressure and varies only with Mach number. The freestream Mach number is determined from the pressrue measured at a station 66.04 cm upstream of the model station. The model has no significant influence on the vertical Mach number distribution at this station. This method of Mach number determination appears to be more accurate than one using the plenum pressure.

Sewall, W. G.↗
NASA NTRS1980

Laser angle sensor development

Electrical and optical parameters were developed for a two axis (pitch/roll) laser angle sensor. The laser source and detector were mounted in the plenum above the model. Two axis optical distortion measurements of flow characteristics in a 0.3 transonic cryogenic tunnel were made with a shearing interferometer. The measurement results provide a basis for estimating the optical parameters of the laser angle sensor. Experimental and analytical information was generated on model windows to cover the reflector. A two axis breadboard was assembled to evaluate different measurement concepts. The measurement results were used to develop a preliminary design of a laser angle sensor. Schematics and expected performance specifications are included.

Pond, C. R.↗
NASA NTRS1981

Solar Heating System installed at Belz Investment Company, Memphis, Tennessee

A hot air solar system which utilizes flat plate air collectors is discussed. Collector areas for each of four buildings cover 780 sq ft, with storage capacity of 390 cu ft per building. The air system has a special air handling unit to move air through the collectors and into and out of the rock storage, with connection to the air duct distribution system. The heat of the motor is added to the heat delivered to the system. The solar system also includes four motorized special low leakage dampers and two gravity fabric dampers. The system is automatically controlled by a solid state controller with three thermistors: one located in the collectors, one in the rock box to plenum, one in the return air duct from the heated space. A three stage heating thermostat, located in the conditioned space, controls the operation.

Source record↗
NASA NTRS1980

A system for model access in tunnels with an unbreathable test medium

In many specialty wind tunnels, test gases other than ambient air are used to meet special testing requirements. A typical example is the use of freon as the test gas to achieve a realistic density ratio between gas and model for exploring flutter stability boundaries. Another example is the use of pressurized air to elevate the stream density and enhance Reynolds number or dynamic pressure simulation. Such specialty tunnels require a system of access o the model which will allow services and changes to the model without exposing personnel to the unnatural and perhaps hostile environment or requiring the venting and purging of the entire tunnel circuit. This paper will describe the plenum and model access systems for the forthcoming U.S. National Transonic facility where gaseous nitrogen (N2) at temperatures between 338 and 78 K and at pressures to 9 bars is used as the test medium. The operation at cold temperatures imposes some additional requirements which make the access systems more difficult to design and time consuming to operate than for conventional wind tunnels.

Howell, R. R.↗
NASA NTRS1981

Advanced Low-Emissions Catalytic-Combustor Program, phase 1

Six catalytic combustor concepts were defined, analyzed, and evaluated. Major design considerations included low emissions, performance, safety, durability, installations, operations and development. On the basis of these considerations the two most promising concepts were selected. Refined analysis and preliminary design work was conducted on these two concepts. The selected concepts were required to fit within the combustor chamber dimensions of the reference engine. This is achieved by using a dump diffuser discharging into a plenum chamber between the compressor discharge and the turbine inlet, with the combustors overlaying the prediffuser and the rear of the compressor. To enhance maintainability, the outer combustor case for each concept is designed to translate forward for accessibility to the catalytic reactor, liners and high pressure turbine area. The catalytic reactor is self-contained with air-cooled canning on a resilient mounting. Both selected concepts employed integrated engine-starting approaches to raise the catalytic reactor up to operating conditions. Advanced liner schemes are used to minimize required cooling air. The two selected concepts respectively employ fuel-rich initial thermal reaction followed by rapid quench and subsequent fuel-lean catalytic reaction of carbon monoxide, and, fuel-lean thermal reaction of some fuel in a continuously operating pilot combustor with fuel-lean catalytic reaction of remaining fuel in a radially-staged main combustor.

Sturgess, G. J.↗
NASA NTRS1982

Adaptive-wall wind-tunnel research at NASA-Ames Research Center

Adaptive wall wind tunnel research is summarized. Small scale two and three dimensional wind tunnel experiments and numerical experiments with a three dimensional adaptive wall simulator are included. A NACA 0012 airfoil was tested in a 25 by 13 cm slotted wall test section. Airflow through the test section walls was controlled by adjusting the pressures in segmented plenums. Interference free conditions were successfully attained in subsonic and transonic flows. For the three dimensional experiment, the 25 by 13 cm wind tunnel was modified to permit cross stream wall adjustments. The test model was a semispan wing mounted to one sidewall. Wall interference was substantially reduced at several angles of attack at Mach 0.60. A wing on wall configuration was also modeled in the numerical experiments. These flow simulations showed that free air conditions can be approximated by adjusting boundary conditions at only the floor and ceiling of the test section. No sidewall control was necessary. Typical results from these experiments are discussed.

Schairer, E. T.↗
NASA NTRS1982

Seeding considerations for an LV system in a large transonic wind tunnel

When it was decided to use a laser velocimeter to measure the properties of propellers, seeding was a great concern since large particles fail to flow and small ones are too small to be seen. Many methods were tried and weeded out by using a Malvern particle sizer. The most promising ones were tested in the tunnel and the laser velocimeter (LV) measurements compared to theoretical values of veocity as the particle approached a blunt nose body along a stagnation streamline. Data obtained from the LV system were compared with the one dimensional particle lag calculation. This figure showed the theoretical velocity over the blunt nose and a velocity profile for 5 um particles. This indicated the particles were approximately 5 um. The seeding method is shown. The seed was atomized by 2 seeders run with all 12 available atomizer jets on. The atomizer seed traveled from these two seeders through four 1 inch tubes 20 feet long to the plenum chamber where this cluster of tubes injected the seed into the air stream. The tubes were located 60 feet from the model and could be moved only by shutting the tunnel down. Future seeding plans are shown.

Freedman, R. J.↗
NASA NTRS1981

V/STOL status from the engine technology viewpoint

Directions for designs and powerplant configurations for V/STOL aircraft are explored. The Harrier is the only V/STOL aircraft in service, having performed for over 12 yr. Rolls-Royce has studied engines with powered lift separation from forward propulsion, and engines with composite lift/propulsion modes, resulting in concentration on vectored thrust using rotating or deflected nozzles. The necessity for higher combat performance creates a need for higher specific thrust for supersonic flight (Mach 2), which is achievable with plenum chamber burning. The core flows are exhausted through separate nozzle systems for independent control. Intake in a V/STOL follows a short path and results in nonuniform flow, and testing is described to account for the effects of nonuniform static pressures and temperatures.

Lewis, G. M.↗
NASA NTRS1983

Air-Lubricated Lead Screw

Air lubricated lead screw and nut carefully machined to have closely matched closely fitting threads. Compressed air injected into two plenums encircle nut and flow through orifices to lubricate mating threads. Originally developed to position precisely interferometer retroreflector for airborne measurement of solar infrared radiation, device now has positioning accuracy of 0.25 micron.

Perkins, G. S.↗
NASA NTRS1983

Comparison of predicted and experimental external heat transfer around a film cooled cylinder in crossflow

Calculations were made of the film cooling provided by rows of holes around the circumference of a cylinder in crossflow and the results were compared to experimental data. The calculations and experimental data were for conditions that simulate most of those that are typical of air cooled turbine vane leading edges. Injection was from a single and multiple rows of holes located at different angular locations from the stagnation line. The holes in the rows were angled normal to the flow direction and at a 25 degree angle to the cylinder wall. The calculations and experimental data were for several constant values of blowing ratios for all rows and for different blowing ratios for each row, representing a simulation of a common coolant plenum supply to multiple rows of holes. The calculations were made using a finite difference boundary layer code, STAN5. Contrary to initial expectations that injection would trip the boundary layer flow into the turbulent regime, the results indicated that the high free stream acceleration apparently kept the flow laminar for holes in the first 45 degrees past stagnation. The trend in Stanton number reduction due to coolant injection was predicted with generally good agreement at the lower blowing rates, but for multile rows of holes, agreement was poor beyond the first row.

Stepka, F. S.↗