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  • Aircraft Propulsion and Power
  • 2005-2009  (37)
  • 1950-1954  (3)
  • 1930-1934  (1)
  • 1925-1929
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  • 1
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    Brush Seal Arrangement for the RS-68 Turbopump Set (2006)
    In:  CASI
    Details
    Publication Date: 2018-06-06
    Description: The nature of the RS-68 turbopumps requires that the hydrogen seals separating the pump from the turbine must have extremely low levels of leakage and be contained in small packages. Conventional seal technologies are not able to reasonably satisfy such design requirements. A review of experimental measurements and analysis publications suggests that brush seals are well suited for the design requirements. Brush seals are shown to have less leakage than conventional labyrinth and damper seals and have no adverse effects on the rotordynamics of the machine. The bulk-flow analysis presented by Hendricks et al. is used as a guideline to create a spreadsheet that provides mass flow through the seal and heat generated by the rubbing contact of the bristles on the shaft. The analysis is anchored to published data for LN2 and LH2 leakage tests. Finally, the analysis is used to design seals for both applications. It is observed that the most important analysis parameter is the thickness of the bristle pack and its relationship to seal clearance, lay angle and pressure drop.
    Keywords: Aircraft Propulsion and Power
    Type: Seals/Secondary Fluid Flows Workshop 1997; Volume I; 165-196; NASA/CP-2006-214329/VOL1
    Format: application/pdf
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  • 2
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    A Spreadsheet for the Mixing of a Row of Jets with a Confined Crossflow (2005)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An interactive computer code, written with a readily available software program, Microsoft Excel (Microsoft Corporation, Redmond, WA) is presented which displays 3 D oblique plots of a conserved scalar distribution downstream of jets mixing with a confined crossflow, for a single row, double rows, or opposed rows of jets with or without flow area convergence and/or a non-uniform crossflow scalar distribution. This project used a previously developed empirical model of jets mixing in a confined crossflow to create an Microsoft Excel spreadsheet that can output the profiles of a conserved scalar for jets injected into a confined crossflow given several input variables. The program uses multiple spreadsheets in a single Microsoft Excel notebook to carry out the modeling. The first sheet contains the main program, controls for the type of problem to be solved, and convergence criteria. The first sheet also provides for input of the specific geometry and flow conditions. The second sheet presents the results calculated with this routine to show the effects on the mixing of varying flow and geometric parameters. Comparisons are also made between results from the version of the empirical correlations implemented in the spreadsheet and the versions originally written in Applesoft BASIC (Apple Computer, Cupertino, CA) in the 1980's.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-2005-213137 , E-14651
    Format: application/pdf
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  • 3
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    Acoustic Characteristics of Various Treatment Panel Designs Specific to HSCT Mixer-Ejector Application (2006)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: The development process of liner design methodology is described in several reports. The results of the initial effort of concept development, screening, laboratory testing of various liner concepts, and preliminary correlation (generic data) are presented in a report Acoustic Characteristics of Various Treatment Panel Designs for HSCT Ejector Liner Acoustic Technology Development Program. The second phase of laboratory test results of more practical concepts and their data correlations are presented in this report (product specific). In particular, this report contains normal incidence impedance measurements of several liner types in both a static rig and in a high temperature flow duct rig. The flow duct rig allows for temperatures up to 400 F with a grazing flow up to Mach 0.8. Measurements of impedance, DC flow resistance, and in the flow rig cases, impact of the liner on boundary layer profiles are documented. In addition to liner rig tests, a limited number of tests were made on liners installed in a mixer-Ejector nozzle to confirm the performance of the liner prediction in an installed configuration.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/CR-2006-214401 , E-15680
    Format: application/pdf
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  • 4
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    Acoustic Characteristics of Various Treatment Panel Designs for HSCT Ejector Liner Acoustic Technology Development Program (2006)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: The objectives of the initial effort (Phase I) of HSR Liner Technology Program, the selection of promising liner concepts, design and fabrication of these concepts for laboratory tests, testing these liners in the laboratory by using impedance tube and flow ducts, and developing empirical impedance/suppression correlation, are successfully completed. Acoustic and aerodynamic criteria for the liner design are established. Based on these criteria several liners are designed. The liner concepts designed and fabricated include Single-Degree-of-Freedom (SDOF), Two-Degree-of-Freedom (2DOF), and Bulk Absorber. Two types of SDOF treatment are fabricated, one with a perforated type face plate and the other with a wiremesh (woven) type faceplate. In addition, special configurations of these concepts are also included in the design. Several treatment panels are designed for parametric study. In these panels the facesheets of different porosity, hole diameter, and sheet thickness are utilized. Several deep panels (i.e., 1 in. deep) are designed and instrumented to measure DC flow resistance and insitu impedance in the presence of grazing flow. Basic components of these panels (i.e., facesheets, bulk materials, etc.) are also procured and tested. The results include DC flow resistance, normal impedance, and insertion loss.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/CR-2006-214399 , E-15678
    Format: application/pdf
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  • 5
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    Explicit Finite Element Modeling of Multilayer Composite Fabric for Gas Turbine Engine Containment Systems, Phase II (2009)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: A team consisting of Arizona State University, Honeywell Engines, Systems & Services, the National Aeronautics and Space Administration Glenn Research Center, and SRI International collaborated to develop computational models and verification testing for designing and evaluating turbine engine fan blade fabric containment structures. This research was conducted under the Federal Aviation Administration Airworthiness Assurance Center of Excellence and was sponsored by the Aircraft Catastrophic Failure Prevention Program. The research was directed toward improving the modeling of a turbine engine fabric containment structure for an engine blade-out containment demonstration test required for certification of aircraft engines. The research conducted in Phase II began a new level of capability to design and develop fan blade containment systems for turbine engines. Significant progress was made in three areas: (1) further development of the ballistic fabric model to increase confidence and robustness in the material models for the Kevlar(TradeName) and Zylon(TradeName) material models developed in Phase I, (2) the capability was improved for finite element modeling of multiple layers of fabric using multiple layers of shell elements, and (3) large-scale simulations were performed. This report concentrates on the material model development and simulations of the impact tests.
    Keywords: Aircraft Propulsion and Power
    Type: PB2010-103422 , DOT/FAA/AR-08/37,P3
    Format: text
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    NASA TECHNICAL REPORTS
  • 6
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    Adaptive Technologies Developed and Demonstrated for Extending the Stable Operating Range of Compressors (2005)
    In:  CASI
    Details
    Publication Date: 2018-06-02
    Description: The operational envelope of gas turbine engines is constrained by the stability limit of the compression system. The dangers of exceeding this limit are severe, with the potential for engine failure and loss of the aircraft. To avoid such failures, compressor designers provide an adequate stability (stall) margin in the compressor design to account for inlet distortions, degradation due to wear, throttle transients, and other factors that reduce compressor stability from the original design intent. In some cases, the required stall margin results in the compressor operating line being below the maximum efficiency potential of the compression system. Current approaches to increasing stability tend to decrease the efficiency of the compressor. The focus of this work is to increase the stall margin of compressors without decreasing their efficiency.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 2004; NASA/TM-2005-213419
    Format: application/pdf
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  • 7
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    Some Characteristics of Sprays Obtained from Pintle-type Injection Nozzles (1933)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This report presents the results of tests made with the pintle-type injection nozzles, one having a pintle angle of 8 degrees, the other a pintle angle of 30 degrees. The fuel was injected into a glass-windowed pressure chamber and the spray photographed by means of the N.A.C.A. spray photography apparatus. Curves are presented that give the penetration of the spray tips when fuel oil is injected by pressures of 1,500 to 4,000 pounds per square inch into air at room temperature and densities of 11 to 18 atmospheres. High-speed spark photographs show the appearance of the sprays in air at a density of 18 atmospheres. The results indicate that the pintle angles have little effect on the size of the spray cone angle, which is about the same as that of sprays from plain round hole orifices. The penetration of the spray from the nozzle with an 8 degree pintle is slightly higher than that of the spray from the nozzle with a 30 degree pintle. The penetration of the sprays from the pintle nozzles, for comparable conditions of injection pressure and air density, is about the same as that of sprays from round-hole orifices. Increase in air density decreases the penetration in about the same ratio with all the injection pressures.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-TN-465
    Format: application/pdf
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  • 8
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    Damage-Tolerant, Affordable Composite Engine Cases Designed and Fabricated (2005)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: An integrated team of NASA personnel, Government contractors, industry partners, and university staff have developed an innovative new technology for commercial fan cases that will substantially influence the safety and efficiency of future turbine engines. This effective team, under the direction of the NASA Glenn Research Center and with the support of the Federal Aviation Administration, has matured a new class of carbon/polymer composites and demonstrated a 30- to 50-percent improvement in specific containment capacity (blade fragment kinetic energy/containment system weight). As the heaviest engine component, the engine case/containment system greatly affects both the safety and efficiency of aircraft engines. The ballistic impact research team has developed unique test facilities and methods for screening numerous candidate material systems to replace the traditional heavy, metallic engine cases. This research has culminated in the selection of a polymer matrix composite reinforced with triaxially braided carbon fibers and technology demonstration through the fabrication of prototype engine cases for three major commercial engine manufacturing companies.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 2004; NASA/TM-2005-213419
    Format: application/pdf
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  • 9
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    Optimization of Turbine Rim Seals (2006)
    In:  CASI
    Details
    Publication Date: 2018-06-06
    Description: Experiments are being conducted to gain an understanding of the physics of rim scale cavity ingestion in a turbine stage with the high-work, single-stage characteristics envisioned for Advanced Subsonic Transport (AST) aircraft gas turbine engines fo the early 21st century. Initial experimental measurements to be presented include time-averaged turbine rim cavity and main gas path static pressure measurements for rim seal coolant to main gas path mass flow ratios between 0 and 0.02. The ultimate objective of this work is develop improved rim seal design concepts for use in modern high-work, single sage turbines n order to minimize the use of secondary coolant flow. Toward this objective the time averaged and unsteady data to be obtained in these experiments will be used to 1) Quantify the impact of the rim cavity cooling air on the ingestion process. 2) Quantify the film cooling benefits of the rim cavity purge flow in the main gas path. 3) Quantify the impact of the cooling air on turbine efficiency. 4) Develop/evaluate both 3D CFD and analytical models of the ingestion/cooling process.
    Keywords: Aircraft Propulsion and Power
    Type: Seals/Secondary Fluid Flows Workshop 1997; Volume I; 253-268; NASA/CP-2006-214329/VOL1
    Format: application/pdf
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  • 10
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    Effects of 'Cooled' Cooling Air on Pre-Swirl Nozzle Design (2006)
    In:  CASI
    Details
    Publication Date: 2018-06-06
    Description: It is common practice to use Pre-Swirl Nozzles to facilitate getting the turbine blade cooling air onboard the rotating disk with minimum pressure loss and reduced temperature. Higher engine OPR's and expanded aircraft operating envelopes have pushed cooling air temperatures to the limits of current disk materials and are stressing the capability to cool the blade with practical levels of cooling air flow. Providing 'Cooled' Cooling Air is one approach being considered to overcome these limitations. This presentation looks at how the introduction of 'Cooled' Cooling Air impacts the design of the Pre-Swirl Nozzles, specifically in relation to the radial location of the nozzles.
    Keywords: Aircraft Propulsion and Power
    Type: Seals/Secondary Fluid Flows Workshop 1997; Volume I; 83-92; NASA/CP-2006-214329/VOL1
    Format: application/pdf
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