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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 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

Procedures for shape optimization of gas turbine disks

Two procedures, the feasible direction method and sequential linear programming, for shape optimization of gas turbine disks are presented. The objective of these procedures is to obtain optimal designs of turbine disks with geometric and stress constraints. The coordinates of the selected points on the disk contours are used as the design variables. Structural weight, stress and their derivatives with respect to the design variables are calculated by an efficient finite element method for design senitivity analysis. Numerical examples of the optimal designs of a disk subjected to thermo-mechanical loadings are presented to illustrate and compare the effectiveness of these two procedures.

Cheu, Tsu-Chien↗
NASA NTRS1990

Supercomputer applications in gas turbine flowfield simulation

The numerical simulation of the unsteady three-dimensional viscous flow in a gas turbine stage is considered. Results from a three-dimensional time-accurate Navier-Stokes simulation of rotor-stator interaction in an axial turbine stage are presented. The present study uses a fine grid in the spanwise direction to better resolve the complex three-dimensional flowfield, and complements earlier reported coarse-grid calculations. Several different features of the flowfield are analyzed and compared to earlier calculations and to experimental data whenever possible. Computer animation techniques are used to visualize various unsteady three-dimensional features of the flow. The results demonstrate the capabilities of current computing hardware in obtaining accurate simulations of unsteady flows in turbomachines.

Madavan, Nateri K.↗
NASA NTRS1990

Development of impact design methods for ceramic gas turbine components

Impact damage prediction methods are being developed to aid in the design of ceramic gas turbine engine components with improved impact resistance. Two impact damage modes were characterized: local, near the impact site, and structural, usually fast fracture away from the impact site. Local damage to Si3N4 impacted by Si3N4 spherical projectiles consists of ring and/or radial cracks around the impact point. In a mechanistic model being developed, impact damage is characterized as microcrack nucleation and propagation. The extent of damage is measured as volume fraction of microcracks. Model capability is demonstrated by simulating late impact tests. Structural failure is caused by tensile stress during impact exceeding material strength. The EPIC3 code was successfully used to predict blade structural failures in different size particle impacts on radial and axial blades.

Song, J.↗
NASA NTRS1993

An efficient liner cooling scheme for advanced small gas turbine combustors

A joint Army/NASA program was conducted to design, fabricate, and test an advanced, small gas turbine, reverse-flow combustor utilizing a compliant metal/ceramic (CMC) wall cooling concept. The objectives of this effort were to develop a design method (basic design data base and analysis) for the CMC cooling technique and then demonstrate its application to an advanced cycle, small, reverse-flow combustor with 3000 F burner outlet temperature. The CMC concept offers significant improvements in wall cooling effectiveness resulting in a large reduction in cooling air requirements. Therefore, more air is available for control of burner outlet temperature pattern in addition to the benefits of improved efficiency, reduced emissions, and lower smoke levels. The program was divided into four tasks. Task 1 defined component materials and localized design of the composite wall structure in conjunction with development of basic design models for the analysis of flow and heat transfer through the wall. Task 2 included implementation of the selected materials and validated design models during combustor preliminary design. Detail design of the selected combustor concept and its refinement with 3D aerothermal analysis were completed in Task 3. Task 4 covered detail drawings, process development and fabrication, and a series of burner rig tests. The purpose of this paper is to provide details of the investigation into the fundamental flow and heat transfer characteristics of the CMC wall structure as well as implementation of the fundamental analysis method for full-scale combustor design.

Paskin, Marc D.↗
NASA NTRS1956

Methods for Measuring Temperatures of Thin-walled Gas-turbine Blades

The accuracy and durability of two means for measuring the metal temperature of gas-turbine blades were investigated experimentally in a turbojet engine. Cemented thermocouples that could be placed in walls as thin 0.010 inch and commercial temperature-indicating paints were evaluated at turbine-inlet temperatures up to about 1650 degrees F and for centrifugal forces equivalent to about 42,000 times gravity. The cemented thermocouples operated satisfactorily. The temperature-indicating paints were neither as accurate nor as durable as the cemented thermocouples.

Stepka, Francis S↗
NASA NTRS1995

Thermal barrier coating life modeling in aircraft gas turbine engines

Analytical models useful for predicting ceramic thermal barrier coating (TBC) spalling life in aircraft gas turbine engines are presented. Electron beam-physical vapor deposited (EB-PVD) and plasma sprayed TBC systems are discussed. TBC spalling was attributed to a combination of mechanisms such as metal oxidation at the ceramic-metal interface, ceramic-metal interface stress concentrations at free surfaces due to dissimilar materials, ceramic-metal interface stresses caused by local radius of curvature and interface roughness, material properties and mechanical behavior, transient temperature gradients across the ceramic layer and component design features. TBC spalling life analytical models were developed based on observations of TBC failure modes and plausible failure theories. TBC failure was assumed to occur when the imposed stresses exceeded the material strength (at or near the ceramic-metal interface). TBC failure knowledge gaps caused by lack of experimental evidence and analytical understanding are noted. The analytical models are considered initial engineering approaches that capture observed TBC failure trends.

Nissley, D. M.↗
NASA NTRS1999

Coupled, Transient Simulations of the Interaction Between Power and Secondary Flowpaths in Gas Turbines

This paper presents a coupled analysis of the interaction between mainpath and secondary flowpaths in gas turbines using transient simulations. Some of the topics include: 1) Need for Coupled Analysis; 2) Primary-Secondary Coupling Schematic; 3) Secondary Flow Requirement; 4) Objectives of Present Methodology; 5) Current Methodologies Recap; 6) Proposed Coupled Code Methodology; 7) Description of SCISEAL Code; 8) Description of Turbo Code; 9) Code Coupling/Interface Issues; and 10) Current Interface Strategy. This paper is presented in viewgraph form.

Athavale, M. M.↗
NASA NTRS2006

Development of a Thin Gauge Metallic Seal for Gas Turbine Engine Applications to 1700 F

The goal of doubling thrust-to-weight ratio for gas turbine engines has placed significant demands on engine component materials. Operating temperatures for static seals in the transition duct and turbine sections for instance, may well reach 2000 F within the next ten years. At these temperatures conventional age-hardenable superalloys lose their high strength via overaging and eventual dissolution of the gamma precipitate, and are well above their oxidation stability limit. Conventional solid-solution-strengthened alloys offer metallurgical stability, but suffer from rapid oxidation and little useful load bearing strength. Ceramic materials can theoretically be used at these temperatures, but manufacturing processes are in the developmental stages.

England, Raymond O.↗
NASA NTRS2005

Fuel premixing module for gas turbine engine combustor

A fuel-air premixing module is designed to reduce emissions from a gas turbine engine. In one form, the premixing module includes a central pilot premixer module with a main premixer module positioned thereround. Each of the portions of the fuel-air premixing module include an axial inflow swirler with a plurality of fixed swirler vanes. Fuel is injected into the main premixer module between the swirler vanes of the axial inflow swirler and at an acute angle relative to the centerline of the premixing module.

Chin, Jushan↗
NASA NTRS1976

Toward more environmentally resistant gas turbines - Progress in NASA-Lewis programs

The paper surveys programs for improving the environmental resistance of gas turbine and power system materials to oxidation and hot corrosion. A computer program Corest has been produced which provides an initial method for predicting long-time metal loss using short-time oxidation data. The role of the coefficient of thermal expansion mismatch on oxide spallation is being studied. A paralinear oxidation attack model in conjunction with regression analysis of weight change data has been applied to the study of oxidation and hot corrosion attack of some Ni-Cr-Al system alloys. Oxide-dispersion strengthened alloys based on the Ni-Cr-Al system are being developed. Sodium sulfate formation during turbine combustion is being investigated from a thermodynamic viewpoint.

Lowell, C. E.↗
NASA NTRS1979

A method to estimate weight and dimensions of large and small gas turbine engines

A computerized method was developed to estimate weight and envelope dimensions of large and small gas turbine engines within + or - 5% to 10%. The method is based on correlations of component weight and design features of 29 data base engines. Rotating components were estimated by a preliminary design procedure which is sensitive to blade geometry, operating conditions, material properties, shaft speed, hub tip ratio, etc. The development and justification of the method selected, and the various methods of analysis are discussed.

Onat, E.↗
NASA NTRS1979

The strainrange partitioning behavior of an advanced gas turbine disk alloy, AF2-1DA

The low-cycle, creep-fatigue characteristics of the advanced gas turbine disk alloy, AF2-1DA have been determined at 1400 F and are presented in terms of the method of strainrange partitioning (SRP). The mean stresses which develop in the PC and CP type SRP cycles at the lowest inelastic strainrange were observed to influence the cyclic lives to a greater extent than the creep effects and hence interfered with a conventional interpretation of the results by SRP. A procedure is proposed for dealing with the mean stress effects on life which is compatible with SRP.

Halford, G. R.↗
NASA NTRS1979

The strainrange partitioning behavior of an advanced gas turbine disk alloy, AF2-1DA

The low-cycle, creep-fatigue characteristics of the advanced gas turbine disk alloy, AF2-1DA have been determined at 1400 F and are presented in terms of the method of strainrange partitioning (SRP). The mean stresses which develop in the PC (tensile Plasticity reversed by compressive Creep) and CP (tensile Creep reversed by compressive Plasticity) type SRP cycles at the lowest inelastic strainrange were observed to influence the cyclic lives to a greater extent than the creep effects and hence interfered with a conventional interpretation of the results by SRP. A procedure is proposed for dealing with the mean stress effects on life which is compatible with SRP.

Halford, G. R.↗
NASA NTRS1983

The start-up of a gas turbine engine using compressed air tangentially fed onto the blades of the basic turbine

The use of compressed air was suggested to increase the reliability and motor lifetime of a gas turbine engine. Experiments were carried out and the results are shown in the form of the variation in circumferential force as a function of the entry angle of the working jet onto the turbine blade. The described start-up method is recommended for use with massive rotors.

Slobodyanyuk, L. K.↗
NASA NTRS1986

Optical and probe determination of soot concentrations in a model gas turbine combustor

An experimental program was conducted to track the variation in soot loading in a generic gas turbine combustor. The burner is a 12.7-cm dia cylindrical device consisting of six sheet-metal louvers. Determination of soot loading along the burner length is achieved by measurement at the exit of the combustor and then at upstream stations by sequential removal of liner louvers to shorten burner length. Alteration of the flow field approaching and within the shortened burners is minimized by bypassing flow in order to maintain a constant linear pressure drop. The burner exhaust flow is sampled at the burner centerline to determine soot mass concentration and smoke number. Characteristic particle size and number density, transmissivity of the exhaust flow, and local radiation from luminous soot particles in the exhaust are determined by optical techniques. Four test fuels are burned at three fuel-air ratios to determine fuel chemical property and flow temperature influences. Particulate concentration data indicate a strong oxidation mechanism in the combustor secondary zone, though the oxidation is significantly affected by flow temperature. Soot production is directly related to fuel smoke point.

Eckerle, W. A.↗