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At least 469 records · Page 26

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NASA NTRS1974

The Design of Pumpjets for Hydrodynamic Propulsion

A procedure for use in the design of a wake adapted pumpjet mounted on the aft end of a body of revolution is presented. To this end, a pumpjet is designed for the Akron airship. The propulsor mass flow is selected to minimize kinetic energy losses through the duct and in the discharge jet. The shaft speed and disk size are selected to satisfy specified limits of cavitation performance and to provide acceptable blade loading. The streamtubes which pass through a propulsor mounted on a tapered afterbody follow essentially conical surfaces. A method is provided for defining these surfaces as a function of shroud geometry, rotor head distribution, and the energy distribution of the ingested mass flow. The three-dimensional effects to which the conical flow subjects the cylindrical blade design sections are described and a technique is presented which permits incorporation of these effects in the blade design procedure.

Bruce, E. P.↗

DOE OSTI2021

Ducted Assembly Steady-State Heat Transfer Software (DASSH)

The Ducted Assembly Steady-State Heat transfer software (DASSH) is being developed at Argonne National Laboratory to perform a steady-state thermal fluids calculation to determine the coolant flow and temperature distribution for a hexagonal reactor core configuration with ducted assemblies. DASSH is intended to be used in the early stages of the reactor design process when assembly components can be undefined or are undergoing considerable design changes. DASSH provides a rapid assessment of the temperature and flow distribution to allow quick characterization of the system and identification of problem areas. This document is a guide for DASSH users. It summarizes the core capabilities of DASSH and highlights differences between DASSH and SE2-ANL. Instructions for obtaining and installing DASSH are provided along with some guidelines and recommendations for newer Python users. It covers the structure of the input file, provides directions for running DASSH, describes the various output files produced by the code, and provides instructions for visualizing data. Examples demonstrating various DASSH options are included. DASSH is under active development, so it is anticipated that this guide will grow and evolve as the code is updated.

22 GENERAL STUDIES OF NUCLEAR REACTORS↗

NASA NTRS1973

Analytical method for predicting the pressure distribution about a nacelle at transonic speeds

The formulation and development of a computer analysis for the calculation of streamlines and pressure distributions around two-dimensional (planar and axisymmetric) isolated nacelles at transonic speeds are described. The computerized flow field analysis is designed to predict the transonic flow around long and short high-bypass-ratio fan duct nacelles with inlet flows and with exhaust flows having appropriate aerothermodynamic properties. The flow field boundaries are located as far upstream and downstream as necessary to obtain minimum disturbances at the boundary. The far-field lateral flow field boundary is analytically defined to exactly represent free-flight conditions or solid wind tunnel wall effects. The inviscid solution technique is based on a Streamtube Curvature Analysis. The computer program utilizes an automatic grid refinement procedure and solves the flow field equations with a matrix relaxation technique. The boundary layer displacement effects and the onset of turbulent separation are included, based on the compressible turbulent boundary layer solution method of Stratford and Beavers and on the turbulent separation prediction method of Stratford.

Keith, J. S.↗

NASA NTRS1975

High subsonic flow tests of a parallel pipe followed by a large area ratio diffuser

Experiments were performed on a pilot model duct system in order to explore its aerodynamic characteristics. The model was scaled from a design projected for the high speed operation mode of the Aircraft Noise Reduction Laboratory. The test results show that the model performed satisfactorily and therefore the projected design will most likely meet the specifications.

Barna, P. S.↗

NASA NTRS1979

Computerized systems analysis and optimization of aircraft engine performance, weight, and life cycle costs

The computational techniques utilized to determine the optimum propulsion systems for future aircraft applications and to identify system tradeoffs and technology requirements are described. The characteristics and use of the following computer codes are discussed: (1) NNEP - a very general cycle analysis code that can assemble an arbitrary matrix fans, turbines, ducts, shafts, etc., into a complete gas turbine engine and compute on- and off-design thermodynamic performance; (2) WATE - a preliminary design procedure for calculating engine weight using the component characteristics determined by NNEP; (3) POD DRG - a table look-up program to calculate wave and friction drag of nacelles; (4) LIFCYC - a computer code developed to calculate life cycle costs of engines based on the output from WATE; and (5) INSTAL - a computer code developed to calculate installation effects, inlet performance and inlet weight. Examples are given to illustrate how these computer techniques can be applied to analyze and optimize propulsion system fuel consumption, weight, and cost for representative types of aircraft and missions.

Fishbach, L. H.↗

NASA NTRS1984

Development of laminar flow control wing surface porous structure

It was concluded that the chordwise air collection method, which actually combines chordwise and spanwise air collection, is the best of the designs conceived up to this time for full chord laminar flow control (LFC). Its shallower ducting improved structural efficiency of the main wing box resulting in a reduction in wing weight, and it provided continuous support of the chordwise panel joints, better matching of suction and clearing airflow requirements, and simplified duct to suction source minifolding. Laminar flow control on both the upper and lower surfaces was previously reduced to LFC suction on the upper surface only, back to 85 percent chord. The study concludes that, in addition to reduced wing area and other practical advantages, this system would be lighter because of the increase in effective structural wing thickness.

Klotzsche, M.↗

NASA NTRS1951

Wind-Tunnel Investigation of Air Inlet and Outlet Openings on a Streamline Body

In connection with the general problem of providing air flow to an aircraft power plant located within a fuselage, an investigation was conducted in the Langley 8-foot high-speed tunnel to determine the effect on external drag and pressure distribution of air inlet openings located at the nose of a streamline body. Air outlet openings located at the tail and at the 21-percent and 63-percent stations of the body were also investigated. Boundary layer transition measurements were made and correlated with the force and the pressure data. Individual openings were investigated with the aid of a blower and then practicable combinations of inlet and outlet openings were tested. Various modifications to the internal duct shape near the inlet opening and the aerodynamic effects of a simulated gun in the duct were also studied. The results of the tests suggested that outlet openings should be designed so that the static pressure of the internal flow at the outlet would be the same as the static pressure of the external flow in the vicinity of the opening.

Becker, John V↗

NASA NTRS2005

Benchmark Data for Evaluation of Aeroacoustic Propagation Codes With Grazing Flow

Increased understanding of the effects of acoustic treatment on the propagation of sound through commercial aircraft engine nacelles is a requirement for more efficient liner design. To this end, one of NASA s goals is to further the development of duct propagation and impedance reduction codes. A number of these codes have been developed over the last three decades. These codes are typically divided into two categories: (1) codes that use the measured complex acoustic pressure field to reduce the acoustic impedance of treatment that is positioned along the wall of the duct, and (2) codes that use the acoustic impedance of the treatment as input and compute the sound field throughout the duct. Clearly, the value of these codes is dependent upon the quality of the data used for their validation. Over the past two decades, data acquired in the NASA Langley Research Center Grazing Incidence Tube have been used by a number of researchers for comparison with their propagation codes. Many of these comparisons have been based upon Grazing Incidence Tube tests that were conducted to study specific liner technology components, and were incomplete for general propagation code validation. Thus, the objective of the current investigation is to provide a quality data set that can be used as a benchmark for evaluation of duct propagation and impedance reduction codes. In order to achieve this objective, two parallel efforts have been undertaken. The first of these is the development of an enhanced impedance eduction code that uses data acquired in the Grazing Incidence Tube. This enhancement is intended to place the benchmark data on as firm a foundation as possible. The second key effort is the acquisition of a comprehensive set of data selected to allow propagation code evaluations over a range of test conditions.

Jones, Michael G.↗

NASA NTRS2009

Device for Measuring Low Flow Speed in a Duct

A multiple-throat venturi system has been invented for measuring laminar flow of air or other gas at low speed (1 to 30 cm/s) in a duct while preserving the laminar nature of the flow and keeping the velocity profile across the duct as nearly flat as possible. While means for measuring flows at higher speeds are well established, heretofore, there have been no reliable means for making consistent, accurate measurements in this speed range. In the original application for which this system was invented, the duct leads into the test section of a low-speed wind tunnel wherein uniform, low-speed, laminar flow is required for scientific experiments. The system could also be used to monitor a slow flow of gas in an industrial process like chemical vapor deposition. In the original application, the multiple- throat venturi system is mounted at the inlet end of the duct having a rectangular cross section of 19 by 14 cm, just upstream of an assembly of inlet screens and flow straighteners that help to suppress undesired flow fluctuations (see Figure 1). The basic venturi measurement principle is well established: One measures the difference in pressure between (1) a point just outside the inlet, where the pressure is highest and the kinetic energy lowest; and (2) the narrowest part (the throat) of the venturi passage, where the kinetic energy is highest and the pressure is lowest. Then by use of Bernoulli s equation for the relationship between pressure and kinetic energy, the volumetric flow speed in the duct can be calculated from the pressure difference and the inlet and throat widths. The design of this system represents a compromise among length, pressure recovery, uniformity of flow, and complexity of assembly. Traditionally, venturis are used to measure faster flows in narrower cross sections, with longer upstream and downstream passages to maintain accuracy. The dimensions of the passages of the present venturi system are sized to provide a readily measurable pressure drop. Multiple throats are used to minimize the length needed to recover internal energy and enable the velocity profile to recover to near flatness.

Quinn, Frank↗

NASA NTRS2009

Measuring Thermal Conductivity of a Small Insulation Sample

Miller, Robert A.↗

NASA NTRSDate not supplied

Microstructure Controlled Multi-Layer Porous Material Liner Tested On The Advanced Noise Control Fan

A novel concept of multilayer micro structurally controlled noise mitigation porous material was developed and experimentally validated in the laboratory by normal incidence impedance test. Extensive parametric studies were made to understand various parameters that influence acoustic absorption and transmission loss based on the test specimen size, particle size, porosity, pore size, and its distribution through the thickness. The results were reported in many publications. This work was focused on applying the concept and assessing the validity of the technology to aircraft turbomachine engines. The NASA Glenn Research Center’s Advanced Noise Control Fan (ANCF) low-speed test rig was selected. The liner insert required a 4-ft internal diameter cylinder with a 2-inch thick and 9-inch axial length. The liner was fabricated in eight segments and assembled in an aluminum test rig. The liner was tested for both tonal and broadband noise at the Notre Dame Turbomachinery Lab White Field test facility, where the ANCF is located. The study concludes that: Microstructurally controlled porous material is very effective in dissipating the sound energy; the liner is effective for tonal noise (2 to 13 dB reduction in noise levels) that constitutes the major part of the overall fan noise, especially the blade passing frequency which is generally the dominating component; Also, the liner is effective for broadband noise with reductions in sound pressure level of about 3 dB for the shaft order greater than 20. The proposed two-layer liner model demonstrated the validity of the concept. More studies are required to optimize the design.

Liner↗

NASA NTRS1951

Preliminary Investigation of the Supersonic Flow Field Downstream of Wire-mesh Nozzles in a Constant-area Duct / Lawrence I. Gould

An investigation was conducted in a 3.4- by 3.4-inch duct to determine the characteristics of the supersonic flow downstream of four wire-mesh screen nozzles with nominal design Mach numbers in the range between 1.97 and 2.58. Two types of disturbances were observed in the flow field: a fine network of interacting expansion and compression waves which were formed immediately downstream of the screens and appeared to dissipate within 25 to 40 wave intersections; and relatively strong oblique shock waves that originated at the junctions of the screens and the walls and were reflected throughout the length of the duct. Regions of fairly uniform flow were found to exist. The total-pressure loss across the screens varied from 22 percent at Mach number 1.58 to 43 percent at Mach number 2.06.

WIND TUNNELS↗

NASA NTRS1993

Noise Trends of a 0.5 M (20 In.) Diameter Supersonic Throughflow Fan as Measured in an Unmodified Compressor Aerodynamic Test Facility

The tone noise levels of a supersonic throughflow fan were measured at subsonic and supersonic axial duct Mach numbers. The noise in the inlet plenum showed no blade passing and harmonic tones at subsonic or supersonic axial flow conditions. At subsonic axial flow conditions, the supersonic throughflow fan showed no inlet plenum tones at fan operating conditions where tone noise had been previously measured for a subsonic fan design. This lower inlet-quadrant noise level for the supersonic throughflow fan was the result of high subsonic inlet velocities acting to reduce the noise propagating out the inlet. The fan noise, which was prevented from propagating upstream by the high subsonic inlet velocities, appeared to increase the noise in the exhaust duct at subsonic throughflow conditions. The exhaust duct noise decreased at supersonic axial throughflow Mach numbers, with the lowest blade passing and harmonic tones levels being observed at the design axial Mach number of 2.0. Multiple pure tone noise was observed in the inlet duct at subsonic axial flow Mach numbers but was seen only in the exhaust duct at supersonic axial flow conditions.

Dittmar, James H.↗

DOE OSTI2024

Applying U.S. metal fuel experience to new fuel designs for fast reactors

With the increasing interest in small modular reactors or microreactors, developers are working to design and submit licensing approval requests of U–10Zr-fueled fast reactors. The developers and their proponents cite prior metal fuel experience (worldwide, but U.S. experience in particular for many developers) as the motivation and justification for their reactor concepts. The experience with metal fuel deployment in sodium-cooled fast reactors as well as the underlying irradiation testing database, provide a suitable basis for analytically justifying the use of metal fuel in new reactors. The evolution of metal fuel design and capability illustrates the importance of key fuel design parameters to consider in new applications of the prior experience: fuel smeared density, plenum-to-fuel volume ratio, the ratio of cladding radius to thickness, fuel composition, and cladding and duct materials. In-service operating and deployment conditions to be considered include fuel linear heat generation rate, fuel temperature, cladding temperature, peak burnup and peak fast fluence. Fuel designs and in-service conditions that are bounded by the database and experience are most easily addressed, but deviations from those previous parameters and conditions can be addressed by considering impacts on previously established behavior and applying other mitigating conservatisms, as appropriate. Here, the authors recommend any new deployment proceed with fuel surveillance and monitoring to mitigate risk, application of conservative measures to address uncertainties, and a fuel qualification program that addresses a range of in-service operating conditions with production fuel. The work reported should be of interest to students and regulators unfamiliar with metal fuel in fast reactors.

11 NUCLEAR FUEL CYCLE AND FUEL MATERIALS↗

DOE OSTI2024

Simulations of divertor designs that spatially separate power and particle exhaust using mid-leg divertor particle pumping

Predictive design modeling of a Dissipation-Focused Divertor for future operation in DIII-D reveals that increasing the poloidal distance of the pump duct entrance from the target surface along the low-field side divertor baffle increases neutral compression and modifies the spatial distribution of power dissipation. With a divertor pump located mid-leg between the target and the X-point, SOLPS-ITER boundary plasma simulations without drifts predict the formation of a dense neutral cloud near the target with > 30x higher neutral compression in detachment, a more stable detachment front located further from the target, and ~25% lower outer midplane separatrix density required for detachment onset, compared to a pump located in the scrape-off layer at the target surface. Up to 19 MW of power flowing into the divertors is modeled using the following two numerical implementations for particle pumping: a specified fraction of particles incident on variable wall sections of the plasma grid is removed from the computational domain (so-called albedo pumping), and a pump duct is modeled which includes dynamics of kinetic neutrals in the duct. The simulations show that the detachment front is located between the divertor target and the X-point and is relatively stable near the pump entrance, without a strong dependence on gas puff rate or injected power. The mid-leg pump design spatially separates the two primary functions of a divertor (power handling and particle exhaust), with the majority of power dissipation occurring near the target plate and particle exhaust taking place further upstream. The benefit of enhanced dissipation using mid-leg pumping comes at the cost of a higher outer midplane separatrix density for a given amount of particle injection.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗

NASA NTRS2005

Closed-Cycle Engine Program Used to Study Brayton Power Conversion

One form of power conversion under consideration in NASA Glenn Research Center's Thermal Energy Conversion Branch is the closed-Brayton-cycle engine. In the tens-of-kilowatts to multimegawatt class, the Brayton engine lends itself to potential space nuclear power applications such as electric propulsion or surface power. The Thermal Energy Conversion Branch has most recently concentrated its Brayton studies on electric propulsion for Prometheus. One piece of software used for evaluating such designs over a limited tradeoff space has been the Closed Cycle Engine Program (CCEP). The CCEP originated in the mid-1980s from a Fortran aircraft engine code known as the Navy/NASA Engine Program (NNEP). Components such as a solar collector, heat exchangers, ducting, a pumped-loop radiator, a nuclear heat source, and radial turbomachinery were added to NNEP, transforming it into a high-fidelity design and performance tool for closed-Brayton-cycle power conversion and heat rejection. CCEP was used in the 1990s in conjunction with the Solar Dynamic Ground Test Demonstration conducted at Glenn. Over the past year, updates were made to CCEP to adapt it for an electric propulsion application. The pumped-loop radiator coolant can now be n-heptane, water, or sodium-potassium (NaK); liquid-metal pump design tables were added to accommodate the NaK fluid. For the reactor and shield, a user can now elect to calculate a higher fidelity mass estimate. In addition, helium-xenon working-fluid properties were recalculated and updated.

Johnson, Paul K.↗

NASA NTRS2019

Over-the-Rotor Liner Investigation via the NASA Langley Normal Incidence Tube

NASA Langley and Glenn Research Centers have collaborated on the usage of acoustic liners mounted very near or directly over the rotor of turbofan aircraft engines. This collaboration began over a decade ago with the investigation of a metallic foam liner. Similar to conventional acoustic liner applications, this liner was designed to absorb sound generated by the rotor-alone and rotor-stator interaction sources within the fan duct. Given its proximity to the rotor tips, the expectation was that the liner would also serve as a pressure release and thereby inhibit the amount of noise generated. Initial acoustic results were promising, but there was concern regarding potential aerodynamic penalties. Nevertheless, there were sufficient positive results to warrant further investigation. To that end, the current report presents results obtained in the NASA Langley Normal Incidence Tube for 20 acoustic liner candidates for the OTR application. The majority contain grooves at their surface, designed to minimize aerodynamic penalties caused by placing the liner in close proximity to the fan rotor tips. The intent is to assess the acoustic properties of each liner configuration, and in particular to assess the effects of including the grooves on the overall acoustic performance. An additional intent of this paper is to provide documentation regarding recent enhancements to the NASA Langley Normal Incidence Tube.

Jones, Michael G↗

NASA NTRS1943

Improved Baffle Designs for Air-cooled Engine Cylinders

Explore the source record for details and available documents.

AIR FLOW-DUCTS↗