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Concept increases productivity and wafer quality. Cutting wafers from silicon ingots produces smooth wafers at high rates with reduced blade wear. Involves straight reciprocating saw blade and slight rotation of ingot between cutting strokes. Many parallel blades combined to cut many wafers simultaneously from ingot.
The technique and uses of the multiple blade slurry (MBS) saw are considered. Multiple bands of steel are arranged in a frame and the frame is reciprocated with the steel bands to a workpiece, while simultaneously applying abrasive at the point of contact. The blades wear slots in the workpiece and progress through the piece resulting in several parts of wafers. The transition to MBA from diamond slicing is justified by savings resulting from minimized kerf losses, minimized subsurface damage, and improved surface quality off the saw. This allows wafering much closer to finished thickness specifications. The current state of the art MBS technology must be significantly improved if the low cost solar array (LSA) goals are to be attained. It is concluded that although MBS will never be the answer to every wafering requirement, the economical production of wafers to LSA project specifications will be achieved.
Microwave sensor technology is being investigated by the NASA Glenn Research Center as a means of making non-contact structural health measurements in the hot sections of gas turbine engines. This type of sensor technology is beneficial in that it is accurate, it has the ability to operate at extremely high temperatures, and is unaffected by contaminants that are present in turbine engines. It is specifically being targeted for use in the High Pressure Turbine (HPT) and High Pressure Compressor (HPC) sections to monitor the structural health of the rotating components. It is intended to use blade tip clearance to monitor blade growth and wear and blade tip timing to monitor blade vibration and deflection. The use of microwave sensors for this application is an emerging concept. Techniques on their use and calibration needed to be developed. As a means of better understanding the issues associated with the microwave sensors, a series of experiments have been conducted to evaluate their performance for aero engine applications. This paper presents the results of these experiments.
Size reduction systems used in biomass processing break biomass into smaller pieces by utilizing the kinetic energy from the sharp rotating blades. Abrasive and/or erosive wear caused by biomass comminution results in blade wear of the sharp edged cutters, deteriorating the process efficiency. Here, this study aims to optimize the blade design and improve the system efficiency by attempting to understand the interactions between the blades and biomass particles. Since real-time monitoring of these interactions is impractical during operation, mechanical simulations offer a viable alternative for investigating the shredding process. Yet, the irregular geometry and complex mechanical properties of biomass—such as the anisotropic nature of woodchips and their nonlinear fracture behavior—pose significant challenges for accurately simulating contact pressure. In this work an anisotropic yield material model, along with a damage initiation and evolution function, is applied to the woodchip particle to study the contact pressure on shredder blade, offering a scientific basis for improved blade design and process efficiency. This approach can be extended to other biomass processing systems with similar anisotropic feedstocks, making it a valuable tool for advancing sustainable biomass utilization.
Mechanical preprocessing of biomass, including size reduction, is a crucial step in converting biomass into biofuel. However, feedstock inevitably contains abrasive intrinsic and extrinsic inorganics that may cause excessive tool wear in preprocessing. Here, this work demonstrates that performance of a knife mill can be significantly improved by applying a more wear-resistant blade material. A series of full-scale knife mill tests were performed for size reduction of forest residue using blades of tungsten carbide (WC–Co), iron-borided tool steel, and diamond-like carbon (DLC) coated tool steel. Blade material loss was quantified in correlation to the amount of feedstock processed and wear mechanisms were investigated via worn surface characterization. While the thin DLC coating was removed quickly, the WC-Co and iron-borided blades improved the tool life by 8X and 3X compared with the M2 tool steel blades (baseline), respectively. The in-situ throughput and power consumption measurements provided additional insights. The WC-Co and iron-borided blades had ~3X higher throughput than the baseline blades by the end of the test with lower normalized power consumption. The experimental results were then used as input for a techno-economic analysis, which suggested that the more wear resistant blades could cut the knife milling cost by $2–3 per ton of biomass processed with downtime reduced by 65–85%.
Experimental results are presented for the case of titanium blade tip specimens of various geometrical configurations rubbing at 100 m/s against specimens of nickel-chromium sintered powder metal seal material, the latter being fed toward the rotating blades at an incursion rate of 0.0254 mm/s. Blade tips in the form of orthogonal cutting tools with about 85 deg negative rake angles exhibited desirable abrading capabilities, as measured by the tear-free appearance of the grooves they generated in the seal material, little wear of blade tips, low forces of interaction and low seal densification. Similar results have been obtained for blade specimens with tips of small radius of curvature, as well as for square-ended and slanted blade tips that are plasma-sprayed with abrasive particles. The relationship between the size of these particles and their abrading effectiveness is considered.
Gas turbine engines for military/commercial fixed-wing and rotary wing aircraft use thermal barrier coatings in the high-temperature sections of the engine for improved efficiency and power. The desire to further make improvements in gas turbine engine efficiency and high power-density is driving the research and development of thermal barrier coatings and the effort of improving their tolerance to fine foreign particulates that may be contained in the intake air. Both commercial and military aircraft engines often are required to operate over sandy regions such as in the Middle-East nations, as well as over volcanic zones. For rotorcraft gas turbine engines, the sand ingestion is adverse during take-off, hovering near ground, and landing conditions. Although, most of the rotorcraft gas turbine engines are fitted with inlet particle separators, they are not 100 percent efficient in filtering fine sand particles of size 75 microns or below. The presence of these fine solid particles in the working fluid medium has an adverse effect on the durability of turbine blade thermal barrier coatings and overall performance of the engine. Typical turbine blade damages include blade coating wear, sand glazing, Calcia-Magnesia-Alumina-Silicate (CMAS) attack, oxidation, plugged cooling holes, all of which can cause rapid performance deterioration including loss of aircraft. The objective of this research is to understand the fine particle interactions with typical ceramic coatings of turbine blades at the microstructure level. A finite-element based microstructure modeling and analysis has been performed to investigate particle-surface interactions, and restitution characteristics. Experimentally, a set of tailored thermal barrier coatings and surface treatments were down-selected through hot burner rig tests and then applied to first stage nozzle vanes of the Gas Generator Turbine of a typical rotorcraft gas turbine engine. Laser Doppler velocity measurements were performed during hot burner rig testing to determine sand particle incoming velocities and their rebound characteristics upon impact on coated material targets. Further, engine sand ingestion tests were carried out to test the CMAS tolerance of the coated nozzle vanes. The findings from this on-going collaborative research to develop the next-gen sand tolerant coatings for turbine blades are presented in this paper.
Friction and fretting wear tests on materials used as friction dampers to reduce vibrational stresses in compressor blades for turbine engines
The rub mechanics of aircraft gas turbine engine compressor abradable blade tip seals was studied at simulated engine conditions. In 12 statistically planned, instrumented rub tests using single titanium blades and fiber-metal rubstrips the rub velocity, incursion rate, incursion depth, blade thickness, and abradable strength were varied to determine the effects on rub energy, heat split between the blade, rubstrip surface and rub debris, and blade and seal wear. The rub energies were found to be most significantly affected by the incursion rate while rub velocity and blade thickness were of secondary importance. In five additional rub tests using single nickel alloy blades and multiple titanium alloy blades, rub energy and wear effects were found to be similar for titanium and nickel alloy blades while rub energies increased for multiple blades relative to single blade test results.
A plasma sprayed graded layered ceramic/metallic (ZrO2/CoCrAlY) seal was evaluated for JT9D turbine application by rig and engine tests. Four cyclic thermal shock rig tests were conducted during the program. Three completed 1000 simulated engine thermal cycle tests and the fourth completed 500 cycles without severe cracking or spalling. Three ceramic seals were installed in a JT9D experimental engine to evaluate the effect of the engine thermal environment on the seals. All three seals completed the test successfully without severe cracking or spalling. The three seals did have slight laminar cracks at the 85/15-40/60 ZrO2/CoCrAlY interface. The second engine test evaluated the rub capabilities of the seal. Six ceramic seals were installed in the engine with fourteen abrasive tip blades. Three of the six seals rubbed to a depth of 24 mils. Eight of the fourteen abrasive tip blades showed evidence of wear. Three of the eight blades wore a maximum of five mils. Engine rub test results demonstrated the potential of reducing turbine clearances and thereby improving engine performance by use of sprayed ceramic seals.
Significant improvements in engine readiness with reductions in maintenance costs and turn-around times can be achieved with an engine condition monitoring systems (CMS). The CMS provides health status of critical engine components, without disassembly, through monitoring with advanced sensors. Engine failure reports over 35 years were categorized into 20 different modes of failure. Rotor bearings and turbine blades were determined to be the most critical in limiting turbopump life. Measurement technologies were matched to each of the failure modes identified. Three were selected to monitor the rotor bearings and turbine blades: the isotope wear detector and fiberoptic deflectometer (bearings), and the fiberoptic pyrometer (blades). Signal processing algorithms were evaluated for their ability to provide useful health data to maintenance personnel. Design modifications to the Space Shuttle Main Engine (SSME) high pressure turbopumps were developed to incorporate the sensors. Laboratory test fixtures have been designed for monitoring the rotor bearings and turbine blades in simulated turbopump operating conditions.
Significant improvements in engine readiness with attendant reductions in maintenance costs and turnaround times can be achieved with an engine condition monitoring system (CMS). The CMS provides real time health status of critical engine components, without disassembly, through component monitoring with advanced sensor technologies. Three technologies were selected to monitor the rotor bearings and turbine blades: the isotope wear detector and fiber optic deflectometer (bearings), and the fiber optic pyrometer (blades). Signal processing algorithms were evaluated and ranked for their utility in providing useful component health data to unskilled maintenance personnel. Design modifications to current configuration Space Shuttle Main Engine (SSME) high pressure turbopumps and the MK48-F turbopump were developed to incorporate the sensors.
An advanced blade/shroud system designed to maintain close clearance between blade tips and turbine shrouds and at the same time, be resistant to environmental effects including high temperature oxidation, hot corrosion, and thermal cycling is described. Increased efficiency and increased blade life are attained by using the advanced blade tip seal system. Features of the system include improved clearance control when blade tips preferentially wear the shrouds and a superior single crystal superalloy tip. The tip design, joint location, characterization of the single crystal tip alloy, the abrasive tip treatment, and the component and engine test are among the factors addressed. Results of wear testing, quality control plans, and the total manufacturing cycle required to fully process the blades are also discussed.
The level of performance deterioration that occurred during early service was studied and also the contribution of each engine module to the overall level of engine performance deterioration. To assist in the distribution of the performance losses to the module level, testing with expanded experimental instrumentation was accomplished to permit direct measurement of module performance. An analytical teardown of the engine was accomplished to permit assignment of module performance losses to individual mechanical damage mechanisms including blade and seal wear, roughness, and thermal distortion of turbine parts. The results of the testing show that the engine deteriorated 1.5 percent in thrust specific fuel consumption from its performance when new. Initial testing, which included water washing, showed that 0.2 percent in performance deterioration was due to surface contamination (dirt) and 0.1 percent was due to drift of the engine vane control system, leaving 1.2 percent in basic TSFC deterioration. This 1.2 percent TSFC loss was distributed among the engine modules with 0.6 percent caused by clearance changes, 0.4 percent loss due to thermal distortion, and 0.2 percent loss due to increased roughness of the fan and low-pressure compressor airfoils.
This presentation reviews the following: (i) Cause and effect of gas turbine blade tip seal wear (ii) Current clearance control practices (iii) Present approaches under investigation at GRC.
In 1984, the Vacuum Plasma Spray Lab was built at NASA/Marshall Space Flight Center for applying durable, protective coatings to turbine blades for the space shuttle main engine (SSME) high pressure fuel turbopump. Existing turbine blades were cracking and breaking off after five hot fire tests while VPS coated turbine blades showed no wear or cracking after 40 hot fire tests. Following that, a major manufacturing problem of copper coatings peeling off the SSME Titanium Main Fuel Valve Housing was corrected with a tenacious VPS copper coating. A patented VPS process utilizing Functional Gradient Material (FGM) application was developed to build ceramic lined metallic cartridges for space furnace experiments, safely containing gallium arsenide at 1260 degrees centigrade. The VPS/FGM process was then translated to build robust, long life, liquid rocket combustion chambers for the space shuttle main engine. A 5K (5,000 Lb. thrust) thruster with the VPS/FGM protective coating experienced 220 hot firing tests in pristine condition with no wear compared to the SSME which showed blanching (surface pulverization) and cooling channel cracks in less than 30 of the same hot firing tests. After 35 of the hot firing tests, the injector face plates disintegrated. The VPS/FGM process was then applied to spraying protective thermal barrier coatings on the face plates which showed 50% cooler operating temperature, with no wear after 50 hot fire tests. Cooling channels were closed out in two weeks, compared to one year for the SSME. Working up the TRL (Technology Readiness Level) to establish the VPS/FGM process as viable technology, a 40K thruster was built and is currently being tested. Proposed is to build a J-2X size liquid rocket engine as the final step in establishing the VPS/FGM process TRL for space flight.
The rub mechanics of compressor abradable blade tip seals at simulated engine conditions were investigated. Twelve statistically planned, instrumented rub tests were conducted with titanium blades and Feltmetal fibermetal rubstrips. The tests were conducted with single stationary blades rubbing against seal material bonded to rotating test disks. The instantaneous rub torque, speed, incursion rate and blade temperatures were continuously measured and recorded. Basic rub parameters (incursion rate, rub depth, abradable density, blade thickness and rub velocity) were varied to determine the effects on rub energy and heat split between the blade, rubstrip surface and rub debris. The test data was reduced, energies were determined and statistical analyses were completed to determine the primary and interactive effects. Wear surface morphology, profile measurements and metallographic analysis were used to determine wear, glazing, melting and material transfer. The rub energies for these tests were most significantly affected by the incursion rate while rub velocity and blade thickness were of secondary importance. The ratios of blade wear to seal wear were representative of those experienced in engine operation of these seal system materials.
Ceramic material systems are being considered for potential use as turbine blade tip gas path seals at temperatures up to 1370 1/4 C. Silicon carbide and silicon nitride structures were selected for study since an initial analysis of the problem gave these materials the greatest potential for development into a successful materials system. Segments of silicon nitride and silicon carbide materials over a range of densities, processed by various methods, a honeycomb structure of silicon nitride and ceramic blade tip inserts fabricated from both materials by hot pressing were tested singly and in combination. The evaluations included wear under simulated engine blade tip rub conditions, thermal stability, impact resistance, machinability, hot gas erosion and feasibility of fabrication into engine components. The silicon nitride honeycomb and low-density silicon carbide using a selected grain size distribution gave the most promising results as rub-tolerant shroud liners. Ceramic blade tip inserts made from hot-pressed silicon nitride gave excellent test results. Their behavior closely simulated metal tips. Wear was similar to that of metals but reduced by a factor of six.