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Sulfur removed from nickel-base superalloy used to make gas-turbine blades by heating alloy and simultaneously subjecting it to sputtering by directed Ar(Sup+) ions from ion gun or from glow discharge. Reduction of sulfur content of superalloy by factor of 10 increases lifetime of turbine blade made of alloy by similar factor, because stability of protective surface oxide formed during operation of turbine increased.
Foreign object damage (FOD) behavior of two commercial gas-turbine grade silicon nitrides, AS800 and SN282, was determined at ambient temperature through strength testing of flexure test specimens impacted by steel ball projectiles with a diameter of 1.59 mm in a velocity range from 220 to 440 m/s. AS800 silicon nitride exhibited a greater FOD resistance than SN282, primarily due to its greater value of fracture toughness (K(sub lc)). The use of an additional equiaxed, fine-grained silicon nitride (NC132) showed that fracture toughness was a key material parameter affecting FOD resistance. The observed damages generated by projectile impact were typically in the forms of well- or ill-developed ring and cone cracks with little presence of radial cracks.
A method and apparatus for suppressing jet noise in a gas turbine engine powerplant 10 is disclosed. Various construction details are developed for providing sound suppression at sea level take-off operative conditions and not providing sound suppression at cruise operative conditions. In one embodiment, the powerplant 10 has a lobed mixer 152 between a primary flowpath 44 and a second flowpath 46, a diffusion region downstream of the lobed mixer region (first mixing region 76), and a deployable ejector/mixer 176 in the diffusion region which forms a second mixing region 78 having a diffusion flowpath 72 downstream of the ejector/mixer and sound absorbing structure 18 bounding the flowpath throughout the diffusion region. The method includes deploying the ejector/mixer 176 at take-off and stowing the ejector/mixer at cruise.
Lean direct injection (LDI) is a combustion concept to reduce oxides of nitrogen (NOx) for next generation aircraft gas turbine engines. These newer engines have cycles that increase fuel efficiency through increased operating pressures, which increase combustor inlet temperatures. NOx formation rates increase with higher temperatures, and the LDI strategy is to avoid high temperature by staying fuel lean and away from stoichiometric burning. Thus, LDI relies on rapid and uniform fuel/air mixing. Additionally, in combustors for smaller core engines, the area and volume bring about mixing and time scale challenges not found in larger engines. As part of our parametric study in which we vary swirler angle and orientation and look at their effect on fluid mixing and combustion, we examine one configuration of a 7-point lean direct injector by looking at the non-combusting 2-D velocity field using PIV, and combusting system for chemical species using chemiluminescent imaging and flame spectroscopy. The circular 7-point array consists of axial swirlers, with the center 60° swirler surrounded by six 52° swirlers. The velocity results for this configuration show that the outer swirlers serve to isolate the center flow field near the injector exit. A recirculation zone forms downstream of the center swirler, but not behind the outer swirlers. The combusting results also show an isolated zone directly downstream of the center injector. The flame spectra show variation in speciation of combustion species such as OH*, C3*, CH*, and C2* as a function of position within the combustor.
Experiments are conducted for the noise power and spectra emitted from a gas turbine combustor can exhausting to the atmosphere. The theory of combustion noise is applied to the results to determine the noise generating capability of the flame in the absence of reflecting can surfaces. The results show that for a fixed fuel (JP-4) the noise output is independent of fuel/air ratio for well stabilized can-type flames and heavily dependent on airflow while the spectra are dominated by the can acoustics, primarily through sound absorption by the liner. In an installed configuration the noise output depends heavily on the enclosure acoustics. Scaling rules are presented for installed configurations.
Statistics on gas turbine rotor failures that have occurred in U.S. commercial aviation during 1975 are presented. The compiled data were analyzed to establish: (1) The incidence of rotor failures and the number of contained and uncontained rotor bursts; (2) The distribution of rotor bursts with respect to engine rotor component; i.e., fan, compressor or turbine; (3) The type of rotor fragment (disk, rim or blade) typically generated at burst; (4) The cause of failure; (5) The type of engines involved; and (6) The flight condition at the time of failure.
The development of ceramic components for automobile gas turbine engines is described with attention given to the steady and unsteady thermal conditions the ceramics will experience, and their anti-corrosion and strain-resistant properties. The ceramics considered for use in the automobile turbines include hot-pressed Si3N4, reaction-sintered, isostatically pressed Si3N4, hot-pressed SiC, reaction-bonded SiC, and glass ceramics. Attention is given to the stress analysis of ceramic structures and the state of the art of ceramic structural technology is reviewed, emphasizing the use of ceramics for combustion chambers and ceramic shrouded turbomachinery (a fully ceramic impeller).
Fuel economy potentials were calculated and compared among ten turbomachinery configurations. All gas turbine engines were evaluated with a continuously variable transmission in a 1978 compact car. A reference fuel economy was calculated for the car with its conventional spark ignition piston engine and three speed automatic transmission. Two promising engine/transmission combinations, using gasoline, had 55 to 60 percent gains over the reference fuel economy. Fuel economy sensitivities to engine design parameter changes were also calculated for these two combinations.
An innovative annular combustor configuration is being developed for aircraft and other gas turbine engines. This design has the potential of permitting higher turbine inlet temperatures by reducing the pattern factor and providing a major reduction in NO(x) emission. The design concept is based on a Variable Residence Time (VRT) technique which allows large fuel particles adequate time to completely burn in the circumferentially mixed primary zone. High durability of the combustor is achieved by dual function use of the incoming air. The feasibility of the concept was demonstrated by water analogue tests and 3-D computer modeling. The computer model predicted a 50 percent reduction in pattern factor when compared to a state of the art conventional combustor. The VRT combustor uses only half the number of fuel nozzles of the conventional configuration. The results of the chemical kinetics model require further investigation, as the NO(x) predictions did not correlate with the available experimental and analytical data base.
Fluid mechanical investigations simulating the flow in the primary zone of a gas turbine combustor are presented using three generic test rigs: (1) rotating pipe yielding a swirling jet of air; (2) primary zone model with a single swirler and various primary jet configurations, operated with air; and (3) two rectangular models of a (stretched-out) annular combustor with five swirlers in the backwall and with various primary jet configurations, one operated with air and the other with water. Concentration measurements are obtained using laser sheet imaging techniques and velocity measurements using a laser Doppler velocimeter. The results show recirculation zones, intense mixing, instabilities of the interacting jets and the presence of large random vortical motions. The flowfields are shown to exhibit bimodal behavior, have asymmetries despite symmetrical geometry and inlet conditions and display strong jet/swirler and swirler/swirler interactions.
An innovative annular combustor configuration is being developed for aircraft and other gas turbine engines. This design has the potential of permitting higher turbine inlet temperatures by reducing the pattern factor and providing a major reduction in NO(x) emission. The design concept is based on a Variable Residence Time (VRT) technique which allows large fuel particles adequate time to completely burn in the circumferentially mixed primary zone. High durability of the combustor is achieved by dual function use of the incoming air. The feasibility of the concept was demonstrated by water analogue tests and 3-D computer modeling. The computer model predicted a 50 percent reduction in pattern factor when compared to a state of the art conventional combustor. The VRT combustor uses only half the number of fuel nozzles of the conventional configuration. The results of the chemical kinetics model require further investigation, as the NO(x) predictions did not correlate with the available experimental and analytical data base.
Ceramic thermal barrier coatings (TBCS) have been developed for advanced gas turbine and diesel engine applications to improve engine reliability and fuel efficiency. However, durability issues of these thermal barrier coatings under high temperature cyclic conditions are still of major concern. The coating failure depends not only on the coating, but also on the ceramic sintering/creep and bond coat oxidation under the operating conditions. Novel test approaches have been established to obtain critical thermomechanical and thermophysical properties of the coating systems under near-realistic transient and steady state temperature and stress gradients encountered in advanced engine systems. This paper presents detailed experimental and modeling results describing processes occurring in the ZrO2-Y2O3 thermal barrier coating systems, thus providing a framework for developing strategies to manage ceramic coating architecture, microstructure and properties.
The Conference Proceedings contains 31 papers presented at the Propulsion and Energetics Panel Symposium on Erosion, Corrosion and Foreign Object Damage Effects in Gas Turbines which was held from 25-28th April 1994, in Rotterdam, The Netherlands. The Technical Evaluation Report and the Keynote Address are included at the beginning and discussions follow most papers. The Symposium was arranged in the following Sessions: Operational Experience and Requirements (7); Deposition and Erosion (7); Foreign Object Damage (5); Coatings, Repairs and Materials Aspects - 1 (4); Coatings, Repairs and Materials Aspects - 2 (7); and Testing and Certification Procedures (1).
Computations were made to determine the temperature distribution and cooling of solid gas-turbine blades.A range of temperatures was used from 1500 degrees to 2500 degrees F, blade-root temperatures from 100 degrees to 1000 degrees F, blade thermal conductivity from 8 to 220 BTU/(hr)(sq ft)(degrees F/ft), and net gas to metal heat transfer coefficients from 75 to 250 BTU/(hr)(sq ft)(degrees F).
This publication is the third part of a three part report of the project entitled "A Fully Nonmetallic Gas Turbine Engine Enabled by Additive Manufacturing" funded by NASA Aeronautics Research Institute (NARI). The objective of this project was to conduct additive manufacturing to produce ceramic matrix composite materials and aircraft engine components by the binder jet process. Different SiC powders with median sizes ranging from 9.3 to 53.0 microns were investigated solely and in powder blends in order to maximize powder packing. Various infiltration approaches were investigated to include polycarbosilane (SMP-10), phenolic, and liquid silicon. Single infiltrations of SMP-10 and phenolic only slightly filled in the interior. When the SMP-10 was loaded with sub-micron sized SiC powders, the infiltrant gave a much better result of filling in the interior. Silicon carbide fibers were added to the powder bed to make ceramic matrix composite materials. Microscopy showed that the fibers were well distributed with no preferred orientation on the horizontal plane and fibers in the vertical plane were at angles as much as 45deg. Secondary infiltration steps were necessary to further densify the material. Two to three extra infiltration steps of SMP-10 increased the density by 0.20 to 0.55 g/cc. However, the highest densities achieved were 2.10 to 2.15 g/cc. Mechanical tests consisting of 4 point bend tests were conducted. Samples from the two CMC panels had higher strengths and strains to failure than the samples from the two nonfiber reinforced panels. The highest strengths were from Set N with 65 vol% fiber loading which had an average strength of 66 MPa. Analysis of the fracture surfaces did not reveal pullout of the reinforcing fibers. Blunt fiber failure suggested that there was not composite behavior. The binder jet additive manufacturing method was used to also demonstrate the fabrication of turbine engine vane components of two different designs and sizes. The binder jet method has benefits over the conventional manufacturing of CMCs in that prototype and production parts can be fabricated quickly and economically with no tooling and extensive hand layup.
Computer analysis of total pressure loss and air flow distribution in annular flow gas turbine combustion chambers
Detroit Diesel Allison has been active in demonstrating the feasibility of stoichiometric gas turbine engines for many years. Results of this program are presented with emphasis placed on areas requiring additional development work. Cooling air usage, the effect of cooling air on turbine efficiency, fabrication (materials) methods, and control (instrumentation) techniques are the areas discussed and illustrated. Lamilloy, a pseudo-transpiring material developed by Allison, is discussed in detail.
Experiments were conducted to determine the noise power and spectra emitted from a gas turbine combustor can exhausting to the atmosphere. Limited hot wire measurements were made of the cold flow turbulence level and spectra within the can. The fuels used were JP-4, acetone and methyl alcohol burning with air at atmospheric pressure. The experimental results show that for a fixed fuel the noise output is dominated by the airflow rate and not the fuel/air ratio. The spectra are dominated by the spectra of the cold flow turbulence spectra which were invariant with airflow rate in the experiments. The effect of fuel type on the noise power output was primarily through the heat of combustion and not the reactivity. A theory of combustion noise based upon the flame radiating to open surroundings is able to reasonably explain the observed results. A thermoacoustic efficiency for noise radiation as high as .00003 was observed in this program for JP-4 fuel. Scaling rules are presented for installed configurations.