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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.

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At least 667 records · Page 37
All sourcesNASA NTRSDOE OSTI
NASA NTRS2002

Advanced Nacelle Acoustic Lining Concepts Development

The work reported in this document consisted of six distinct liner technology development subtasks: 1) Analysis of Model Scale ADP Fan Duct Lining Data (Boeing): An evaluation of an AST Milestone experiment to demonstrate 1995 liner technology superiority relative to that of 1992 was performed on 1:5.9 scale model fan rig (Advanced Ducted Propeller) test data acquired in the NASA Glenn 9 x 15 foot wind tunnel. The goal of 50% improvement was deemed satisfied. 2) Bias Flow Liner Investigation (Boeing, VCES): The ability to control liner impedance by low velocity bias flow through liner was demonstrated. An impedance prediction model to include bias flow was developed. 3) Grazing Flow Impedance Testing (Boeing): Grazing flow impedance tests were conducted for comparison with results achieved at four different laboratories. 4) Micro-Perforate Acoustic Liner Technology (BFG, HAE, NG): Proof of concept testing of a "linear liner." 5) Extended Reaction Liners (Boeing, NG): Bandwidth improvements for non-locally reacting liner were investigated with porous honeycomb core test liners. 6) Development of a Hybrid Active/Passive Lining Concept (HAE): Synergism between active and passive attenuation of noise radiated by a model inlet was demonstrated.

Bielak, G.↗
NASA NTRS1978

High angle canard missile test in the Ames 11-foot transonic wind tunnel

Four blunted ogive-cylinder missile models with a length-to-diameter ratio of 10.4 were tested at transonic speeds and large angles of attack. The configurations are: body, body with tail panels, body with canards, and body with canards and tails. Forces and moments from the entire model and each of the eight fins were measured over the pitch range of 20 deg to 50 deg and 0 deg to 45 deg roll. Canard deflection angles between 0 deg and 15 deg were tested. Exploratory vapor screen flow visualization testing was also performed. Sample force and moment data are reported along with observations from the vapor screen tests.

Schwind, R. G.↗
NASA NTRS1993

Actuated forebody strake controls for the F-18 high alpha research vehicle

A series of ground-based studies have been conducted to develop actuated forebody strake controls for flight test evaluations using the NASA F-18 High-Alpha Research Vehicle. The actuated forebody strake concept has been designed to provide increased levels of yaw control at high angles of attack where conventional rudders become ineffective. Results are presented from tests conducted with the flight-test strake design, including static and dynamic wind-tunnel tests, transonic wind-tunnel tests, full-scale wind-tunnel tests, pressure surveys, and flow visualization tests. Results from these studies show that a pair of conformal actuated forebody strakes applied to the F-18 HARV can provide a powerful and precise yaw control device at high angles of attack. The preparations for flight testing are described, including the fabrication of flight hardware and the development of aircraft flight control laws. The primary objectives of the flight tests are to provide flight validation of the groundbased studies and to evaluate the use of this type of control to enhance fighter aircraft maneuverability.

Murri, Daniel G.↗
NASA NTRS1995

Upwind-Differencing Computations Of Incompressible Flows

Report discusses application of upwind-differencing numerical-integration scheme in conjunction with method of pseudocompressibility to computations of flows of incompressible fluids, with emphasis on steady-state flows. Three test cases presented: driven flow in square cavity, flow over backward-facing step, and flow around circular cylinder.

Rogers, Stuart↗
NASA NTRS1998

Modular Rocket Engine Control Software (MRECS)

The Modular Rocket Engine Control Software (MRECS) Program is a technology demonstration effort designed to advance the state-of-the-art in launch vehicle propulsion systems. Its emphasis is on developing and demonstrating a modular software architecture for advanced engine control systems that will result in lower software maintenance (operations) costs. It effectively accommodates software requirement changes that occur due to hardware technology upgrades and engine development testing. Ground rules directed by MSFC were to optimize modularity and implement the software in the Ada programming language. MRECS system software and the software development environment utilize Commercial-Off-the-Shelf (COTS) products. This paper presents the objectives, benefits, and status of the program. The software architecture, design, and development environment are described. MRECS tasks are defined and timing relationships given. Major accomplishments are listed. MRECS offers benefits to a wide variety of advanced technology programs in the areas of modular software architecture, reuse software, and reduced software reverification time related to software changes. MRECS was recently modified to support a Space Shuttle Main Engine (SSME) hot-fire test. Cold Flow and Flight Readiness Testing were completed before the test was cancelled. Currently, the program is focused on supporting NASA MSFC in accomplishing development testing of the Fastrac Engine, part of NASA's Low Cost Technologies (LCT) Program. MRECS will be used for all engine development testing.

Tarrant, C.↗
NASA NTRS1987

Performance of laminar-flow leading-edge test articles in cloud encounters

An extensive data bank of concurrent measurements of laminar flow (LF), particle concentration, and aircraft charging state was gathered for the first time. From this data bank, 13 flights in the simulated airline service (SAS) portion were analyzed to date. A total of 6.86 hours of data at one-second resolution were analyzed. An extensive statistical analysis, for both leading-edge test articles, shows that there is a significant effect of cloud and haze particles on the extent of laminar flow obtained. Approximately 93 percent of data points simulating LFC flight were obtained in clear air conditions; approximately 7 percent were obtained in cloud and haze. These percentages are consistent with earlier USAF and NASA estimates and results. The Hall laminar flow loss criteria was verified qualitatively. Larger particles and higher particle concentrations have a more marked effect on LF than do small particles. A particle spectrometer of a charging patch are both acceptable as diagnostic indicators of the presence of particles detrimental to laminar flow.

Davis, Richard E.↗
NASA NTRSDate not supplied

Simulations of the Modular Axisymmetric Scramjet Test Rig Under Reacting Flow Conditions

Simulations of the Air Force Research Laboratory (AFRL) Modular Axisymmetric Scramjet Test Rig (MASTeR) are presented. MASTeR is a parametric test article capable of investigating various scramjet cavity flameholder designs and fueling strategies with the goal to characterize and optimize flameholding capability. In the current work, three cavity aspect ratios, and two fueling strategies (upstream and in-cavity) with ethylene at a nominal facility pressure and temperature condition are evaluated. The reacting simulations are performed, and the resulting flow characteristics are discussed. For each configuration and fueling strategy, cavity residence time, entrainment rate, and fuel-air equivalence ratio are computed. The MASTeR geometry is defined in Engineering Sketch Pad (ESP) and the simulations use a sketch-to-solution (S2S) automated unstructured grid adaptation tool in VULCAN-CFD. This tool automatically generates a simulation grid from the ESP geometry and systematically adapts it to the numerical solution based on the Hessian error estimate of a specified flow field parameter. Reynolds averaged simulations (RAS) are used with a two-equation linear eddy viscosity and diffusivity model. The resulting database can be compared with the experimental data as those become available, and explored to develop models for cavity performance for scramjet propulsion design applications.

hypersonics↗
NASA NTRSDate not supplied

Simulations of the Modular Axisymmetric Scramjet Test Rig Under Reacting Flow Conditions

Simulations of the Air Force Research Laboratory (AFRL) Modular Axisymmetric Scramjet Test Rig (MASTeR) are presented. MASTeR is a parametric test article capable of investigating various scramjet cavity flameholder designs and fueling strategies with the goal to characterize and optimize flameholding capability. In the current work, three cavity aspect ratios, three depths, and two fueling strategies (upstream and in-cavity) with ethylene at a nominal facility pressure and temperature conditions are evaluated. The simulations are performed for mixing-only and reacting flows, and the resulting flow characteristics are compared. For each configuration and fueling strategy, cavity residence time, entrainment rate, and fuel-air equivalence ratio are computed. The MASTeR geometry is defined in the Engineering Sketch Pad (ESP) and the simulations use a sketch-to-solution (S2S) automated unstructured grid adaptation tool in VULCAN-CFD. This tool automatically generates a simulation grid from the ESP geometry and systematically adapts it to the numerical solution based on the Hessian error estimate of a specified flow field parameter. Reynolds averaged simulations (RAS) are used with typical two-equation linear eddy viscosity and diffusivity model. The resulting database can be compared with the experimental data as those becomes available and explored to develop models for cavity performance for scramjet propulsion design applications.

hypersonics↗
NASA NTRS1985

Nacelle/pylon/wing integration on a transport model with a natural laminar flow nacelle

Tests were conducted in the Langley 16-Foot Transonic Tunnel at free-stream Mach numbers from 0.70 to 0.82 and angles of attack from -2.5 deg to 4.0 deg to determine if nacelle/pylon/wing integration affects the achievement of natural laminar flow on a long-duct flow-through nacelle for a high-wing transonic transport configuration. In order to fully assess the integration effect on a nacelle designed to achieve laminar flow, the effects of fixed and free nacelle transitions as well as nacelle longitudinal position and pylon contouring were obtained. The results indicate that the ability to achieve laminar flow on the nacelle is not significantly altered by nacelle/pylon/wing integration. The increment in installed drag between free and fixed transition for the nacelles on symmetrical pylons is essentially the calculated differences between turbulent and laminar flow on the nacelles. The installed drag of the contoured pylon is less than that of the symmetrical pylon. The installed drag for the nacelles in a rearward position is greater than that for the nacelles in a forward position.

Lamb, M.↗
NASA NTRS1973

A mercury flow meter for ion thruster testing

The theory of operation of the thermal flow meter is presented, and a theoretical model is used to determine design parameters for a device capable of measuring mercury flows in the range of 0 to 5 gm/hr. Flow meter construction is described. Tests performed using a positive displacement mercury pump as well as those performed with the device in the feed line of an operating thruster are discussed. A flow meter response time of about a minute and a sensitivity of about 10 mv/gm/hr are demonstrated. Additional work to relieve a sensitivity of the device to variations in ambient temperature is indicated to improve its quantitative performance.

Wilbur, P. J.↗
DOE OSTI2022

Optimized Test Section for Helium Flow Loop Experiment

Phase 3 of this project uses the knowledge and techniques gained from Phase 1 and Phase 2. The first outcome of Phase 3 is a new test section which can be manufactured and tested in the Helium Flow Loop Experiment to verify the results of the optimization study. The second outcome of Phase 3 is analyzing the features of this design and determining its performance under simulated TBM conditions: heat flux, pressure, mass flow rate, one-sided and volumetric heating. This will be accomplished by first using the results from the Phase 1 analyses and then applying the optimization strategy from Phase This phase uses the STAR-CCM+ Design Manager tool to apply the optimization strategy to the Helium Flow Loop Experiment (HFLE). The optimization algorithm used by the Design Manager tool is Simultaneous Hybrid Exploration that is Robust, Progressive, and Adaptive (SHERPA search algorithm). In addition to demonstrating the optimization features, a simple parameter sweep is also included in this phase. This report details the applied single objective optimization.

42 ENGINEERING↗
NASA NTRS2003

High-Flow Jet Exit Rig Designed and Fabricated

The High-Flow Jet Exit Rig at the NASA Glenn Research Center is designed to test single flow jet nozzles and to measure the appropriate thrust and noise levels. The rig has been designed for the maximum hot condition of 16 lbm/sec of combustion air at 1960 R (maximum) and to produce a maximum thrust of 2000 lb. It was designed for cold flow of 29.1 lbm/sec of air at 530 R. In addition, it can test dual-flow nozzles (nozzles with bypass flow in addition to core flow) with independent control of each flow. The High- Flow Jet Exit Rig was successfully fabricated in late 2001 and is being readied for checkout tests. The rig will be installed in Glenn's Aeroacoustic Propulsion Laboratory. The High-Flow Jet Exit Rig consists of the following major components: a single component force balance, the natural-gas-fueled J-79 combustor assembly, the plenum and manifold assembly, an acoustic/instrumentation/seeding (A/I/S) section, a table, and the research nozzles. The rig will be unique in that it is designed to operate uncooled. The structure survives the 1960 R test condition because it uses carefully selected high temperature alloy materials such as Hastelloy-X. The lower plenum assembly was designed to operate at pressures to 450 psig at 1960 R, in accordance with the ASME B31.3 piping code. The natural gas-fueled combustor fires directly into the lower manifold. The hot air is directed through eight 1-1/2-in. supply pipes that supply the upper plenum. The flow is conditioned in the upper plenum prior to flowing to the research nozzle. The 1-1/2-in. supply lines are arranged in a U-shaped design to provide for a flexible piping system. The combustor assembly checkout was successfully conducted in Glenn's Engine Component Research Laboratory in the spring of 2001. The combustor is a low-smoke version of the J79 combustor used to power the F4 Phantom military aircraft. The natural gas-fueled combustor demonstrated high-efficiency combustion over a wide range of operating conditions. This wide operating envelope is required to support the testing of both single- and dual-flow nozzles. Key research goals include providing simultaneous, highly accurate acoustic, flow, and thrust measurements on jet nozzle models in realistic flight conditions, as well as providing scaleable acoustic results. The High-Flow Jet Exit Rig is a second-generation high-flow test rig. Improvements include cleaner flow with reduced levels of particulate, soot, and odor. Choked-flow metering is required with plus or minus 0.25-percent accuracy. Thrust measurements from 0 to 2000 lbf are required with plus or minus 0.25-percent accuracy. Improved acoustics will be achieved by minimizing noise through large pipe bend radii, lower internal flow velocities, and microdrilled choke plates with thousands of 0.040-in.- diameter holes.

Buehrle, Robert J.↗
NASA NTRS1986

A new design concept for indraft wind-tunnel inlets with application to the national full-scale aerodynamic complex

The present inlet design concept for an indraft wind tunnel, which is especially suited to applications for which a specific test section flow quality is required with minimum inlet size, employs a cascade or vaneset to control flow at the inlet plane, so that test section total pressure variation is minimized. Potential flow panel methods, together with empirical pressure loss predictions, are used to predict inlet cascade performance. This concept has been used to develop an alternative inlet design for the 80 x 120-ft wind tunnel at NASA Ames Research Center. Experimental results show that a short length/diameter ratio wind tunnel inlet furnishing atmospheric wind isolation and uniform test section flow can be designed.

Ross, James C.↗
NASA NTRS1976

Studies with the USF/NASA toxicity screening test method - Effect of air flow and effect of fabric dye

One sample each of commercial polyurethane and polychloroprene flexible foams were evaluated using the USF/NASA toxicity screening test method. Air flow rates of 0, 0.16, 16, and 48 ml/sec were used to determine the effect of air flow on relative toxicity. Time to first sign of incapacitation and time to death were substantially reduced with both polyurethane and polychloroprene flexible foams by the introduction of 16 to 48 ml/sec air flow. The relative toxicity rankings of these materials were not altered by changes in air flow. Under these test conditions, the polyurethane foam consistently appeared more toxic than the polychloroprene foam. Samples of six different colors from the same fabric were evaluated separately, using the USF/NASA toxicity screening test method, to determine the effect of fabric dye, if any. The material was an upholstery fabric, consisting of 46 percent cotton, 33 percent wool, and 21 percent nylon. There appeared to be no significant effect of fabric dye on relative toxicity, for this material under these test conditions.

Hilado, C. J.↗
NASA NTRS2016

Space Launch System Base Heating Test: Environments and Base Flow Physics

The NASA Space Launch System (SLS) vehicle is composed of four RS-25 liquid oxygen-hydrogen rocket engines in the core-stage and two 5-segment solid rocket boosters and as a result six hot supersonic plumes interact within the aft section of the vehicle during flight. Due to the complex nature of rocket plume-induced flows within the launch vehicle base during ascent and a new vehicle configuration, sub-scale wind tunnel testing is required to reduce SLS base convective environment uncertainty and design risk levels. This hot-fire test program was conducted at the CUBRC Large Energy National Shock (LENS) II short-duration test facility to simulate flight from altitudes of 50 kft to 210 kft. The test program is a challenging and innovative effort that has not been attempted in 40+ years for a NASA vehicle. This paper discusses the various trends of base convective heat flux and pressure as a function of altitude at various locations within the core-stage and booster base regions of the two-percent SLS wind tunnel model. In-depth understanding of the base flow physics is presented using the test data, infrared high-speed imaging and theory. The normalized test design environments are compared to various NASA semi-empirical numerical models to determine exceedance and conservatism of the flight scaled test-derived base design environments. Brief discussion of thermal impact to the launch vehicle base components is also presented.

Mehta, Manish↗
NASA NTRS2001

Investigation of the Flow-Induced Vibration in the E2 Test Facility

An investigation of flow induced vibration due to coupling between the fluid flow and the propellants lines (LOX and RP-1) was performed. Various flow rate conditions were studied to check whether flow induced vibration was possible due to vortex shedding in both valves and pipe lines. Resonance test was conducted for all segments of the LOX-feedline for the preburner under test. In addition, critical values of frequency and velocity are calculated using a mass damping model. A simple chart characterizing the relation between frequency and velocity is developed for each component; i.e. propellant lines, valves and flow meters. It was found that flow induced vibration occurs for various segments with flow rates of 113 lb/s, 275 lb/s and 40 lb/s. Even more interesting using critical conditions for buckling, it was found that the valve or pipe may collapse for a flow rate of 275 lb/s and valve height of 10% of pipe diameter. Furthermore, two models for the acoustic pressure acting on the segments particularly for the valve are proposed.

Castillo, Luciano↗
NASA NTRS2001

Investigation of the Flow-Induced Vibration in the E2 Test Facility

An investigation of flow induced vibration due to coupling between the fluid flow and the propellants lines (LOX and RP-1) was performed. Various flow rate conditions were studied to check whether flow induced vibration was possible due to vortex shedding in both valves and pipe lines. Resonances test was conducted for all segments of the LOX-feedline for the preburner under test. In addition, critical values of frequency and velocity are calculated using a mass damping model. A simple chart characterizing the relation between frequency and velocity is developed for each component; i.e. propellant lines, valves and flow meters. It was found that flow induced vibration occurs for various segments with flow rates of 113 1b/s, 275 lb/s and 40 lb/s. Even more interesting using critical conditions for buckling, it was found that the valve or pipe may collapse for a flow rate of 275 lb/s and valve height of 10% of pipe diameter. Furthermore, two models for the acoustic pressure acting on the segments particularly for the valve are proposed.

Castillo, Luciano↗