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The design and testing of gas turbine engines employing ceramic components is discussed. Thermal shock and vibration test results as well as spin tests of various engine components are discussed.
Nitrogen oxide levels of low gas turbine engine emissions are correlated in this paper. The predictions of NO(x) emissions are based on a review of the literature of previous low NO(x) combustor programs and analytical chemical kinetic calculations. Concepts included in the literature review consisted of Lean-Premixed-Prevaporized (LPP), Rich Burn/Quick Quench/Lean/Burn (RQL), and Direct Injection. The NO(x) emissions were found to be an exponential function of adiabatic combustion temperature over a wide range of inlet temperatures, pressures and (lean) fuel-air ratios. It is stated that the low NO(X) potential of LPP is offset by the operational disadvantages of its narrow stability limits and its susceptibility to autoignition/ flashback. The RQL and the Direct Injection concepts have the advantage of wider stability limits comparable to conventional combustors. Results were obtained primarily with gaseous fuels and it is noted that the challenge will be to produce the same low levels of NO(x) with liquid fuels.
A mixer assembly for a gas turbine engine is provided, including a main mixer with fuel injection holes located between at least one radial swirler and at least one axial swirler, wherein the fuel injected into the main mixer is atomized and dispersed by the air flowing through the radial swirler and the axial swirler.
A mixer assembly for a gas turbine engine is provided, including a main mixer with fuel injection holes located between at least one radial swirler and at least one axial swirler, wherein the fuel injected into the main mixer is atomized and dispersed by the air flowing through the radial swirler and the axial swirler.
A mixer assembly for a gas turbine engine is provided, including a main mixer, and a pilot mixer having an annular housing. The pilot mixer can further include features to cool portions of the annular housing, such as a radial edge of the annular housing.
The current practice in aerospace turbine engine maintenance is to remove components defined as life-limited parts after a fixed time, on the basis of a predetermined number of flight cycles. Under this schedule-based maintenance practice, the worst-case usage scenario is used to determine the usable life of the component. As shown, this practice often requires removing a part before its useful life is fully consumed, thus leading to higher maintenance cost. To address this issue, the NASA Glenn Research Center, in a collaborative effort with Pratt & Whitney, has developed a generic modular toolset that uses data-mining technology to parameterize life usage models for maintenance purposes. The toolset enables a "condition-based" maintenance approach, where parts are removed on the basis of the cumulative history of the severity of operation they have experienced. The toolset uses data-mining technology to tune life-consumption models on the basis of operating and maintenance histories. The flight operating conditions, represented by measured variables within the engine, are correlated with repair records for the engines, generating a relationship between the operating condition of the part and its service life. As shown, with the condition-based maintenance approach, the lifelimited part is in service until its usable life is fully consumed. This approach will lower maintenance costs while maintaining the safety of the propulsion system. The toolset is a modular program that is easily customizable by users. First, appropriate parametric damage accumulation models, which will be functions of engine variables, must be defined. The tool then optimizes the models to match the historical data by computing an effective-cycle metric that reduces the unexplained variability in component life due to each damage mode by accounting for the variability in operational severity. The damage increment due to operating conditions experienced during each flight is used to compute the effective cycles and ultimately the replacement time. Utilities to handle data problems, such as gaps in the flight data records, are included in the toolset. The tool was demonstrated using the first stage, high-pressure turbine blade of the PW4077 engine (Pratt & Whitney, East Hartford, CT). The damage modes considered were thermomechanical fatigue and oxidation/erosion. Each PW4077 engine contains 82 first-stage, high-pressure turbine blades, and data from a fleet of engines were used to tune the life-consumption models. The models took into account not only measured variables within the engine, but also unmeasured variables such as engine health parameters that are affected by degradation of the engine due to aging. The tool proved effective at predicting the average number of blades scrapped over time due to each damage mode, per engine, given the operating history of the engine. The customizable tools are available to interested parties within the aerospace community.
The design practices for the gas turbine are traced throughout history with particular emphasis on the calculational or analytical methods. Three principal components of the gas turbine engine will be considered: namely, the compressor, the combustor and the turbine.
The design practices for the gas turbine are traced throughout history with particular emphasis on the calculational or analytical methods. Three principal components of the gas turbine engine will be considered: namely, the compressor, the combustor and the turbine.
The environmental stability and thermal gradient cyclic durability performance of SA Tyrannohex composites were investigated for turbine engine component applications. The work has been focused on investigating the combustion rig recession, cyclic thermal stress resistance and thermomechanical low cycle fatigue of uncoated and environmental barrier coated Tyrannohex SiC SA composites in simulated turbine engine combustion water vapor, thermal gradients, and mechanical loading conditions. Flexural strength degradations have been evaluated, and the upper limits of operating temperature conditions for the SA composite material systems are discussed based on the experimental results.
COMBUSTOR and CNOISE are computer codes that predict far-field noise that originates in the combustors of modern aircraft turbine engines -- especially modern, low-gaseous-emission engines, the combustors of which sometimes generate several decibels more noise than do the combustors of older turbine engines. COMBUSTOR implements an empirical model of combustor noise derived from correlations between engine-noise data and operational and geometric parameters, and was developed from databases of measurements of acoustic emissions of engines. CNOISE implements an analytical and computational model of the propagation of combustor temperature fluctuations (hot spots) through downstream turbine stages. Such hot spots are known to give rise to far-field noise. CNOISE is expected to be helpful in determining why low-emission combustors are sometimes noisier than older ones, to provide guidance for refining the empirical correlation model embodied in the COMBUSTOR code, and to provide insight on how to vary downstream turbinestage geometry to reduce the contribution of hot spots to far-field noise.
The Space Shuttle Main Engine (SSME) turbine environment stresses engine components to their design limits and beyond. The extremely high temperatures and rapid temperature cycling can easily cause parts to fail if they are not properly designed. Thin film heat flux sensors can provide heat loading information with almost no disturbance of gas flows or of the blade. These sensors can provide steady state and transient heat flux information. A thin film heat flux sensor is described which makes it easier to measure small temperature differences across very thin insulating layers.
An adaptive controls method for instability suppression in gas turbine engine combustors has been developed and successfully tested with a realistic aircraft engine combustor rig. This testing was part of a program that demonstrated, for the first time, successful active combustor instability control in an aircraft gas turbine engine-like environment. The controls method is called Adaptive Sliding Phasor Averaged Control. Testing of the control method has been conducted in an experimental rig with different configurations designed to simulate combustors with instabilities of about 530 and 315 Hz. Results demonstrate the effectiveness of this method in suppressing combustor instabilities. In addition, a dramatic improvement in suppression of the instability was achieved by focusing control on the second harmonic of the instability. This is believed to be due to a phenomena discovered and reported earlier, the so called Intra-Harmonic Coupling. These results may have implications for future research in combustor instability control.
The theory of the boundary element method is briefly reviewed with particular reference to the feasibility of elastic and inelastic three-dimensional stress analysis of complex structures characteristic of gas turbine engine components. Particular requirements of gas turbine analysis are defined, and examples of the use of a boundary element code designed for the three-dimensional stress analysis of turbine components are presented. It is shown that the general-purpose boundary element code can accurately and efficiently analyze many of the gas turbine engine structures.
Programs initiated by NASA to develop and demonstrate low emission advanced technology combustors for reducing aircraft gas turbine engine pollution are reviewed. Program goals are consistent with urban emission level requirements as specified by the U. S. Environmental Protection Agency and with upper atmosphere cruise emission levels as recommended by the U. S. Climatic Impact Assessment Program and National Research Council. Preliminary tests of advanced technology combustors indicate that significant reductions in all major pollutant emissions should be attainable in present generation aircraft gas turbine engines without adverse effects on fuel consumption. Preliminary test results from fundamental studies indicate that extremely low emission combustion systems may be possible for future generation jet aircraft. The emission reduction techniques currently being evaluated in these programs are described along with the results and a qualitative assessment of development difficulty.
Magnetic bearings are the subject of a new NASA Lewis Research Center and U.S. Army thrust with significant industry participation, and cooperation with other Government agencies. The NASA/Army emphasis is on high-temperature applications for future gas turbine engines. Magnetic bearings could increase the reliability and reduce the weight of these engines by eliminating the lubrication system. They could also increase the DN (diameter of bearing times the rpm) limit on engine speed and allow active vibration cancellation systems to be used, resulting in a more efficient, "more electric" engine. Finally, the Integrated High Performance Turbine Engine Technology (IHPTET) program, a joint Department of Defense/industry program, identified a need for a high-temperature (1200 F) magnetic bearing that could be demonstrated in their Phase III engine. This magnetic bearing is similar to an electric motor. It has a laminated rotor and stator made of cobalt steel. Wound around the stator's circumference are a series of electrical wire coils which form a series of electric magnets that exert a force on the rotor. A probe senses the position of the rotor, and a feedback controller keeps it centered in the cavity. The engine rotor, bearings, and casing form a flexible structure with many modes. The bearing feedback controller, which could cause some of these modes to become unstable, could be adapted to varying flight conditions to minimize seal clearances and monitor the health of the system.
The paper describes and discusses the results from some of the research and development programs for reducing aircraft gas turbine engine emissions. Although the paper concentrates on NASA programs only, work supported by other U.S. government agencies and industry has provided considerable data on low emission advanced technology for aircraft gas turbine engine combustors. The results from the two major NASA technology development programs, the ECCP (Experimental Clean Combustor Program) and the PRTP (Pollution Reduction Technology Program), are presented and compared with the requirements of the 1979 U.S. EPA standards. Emission reduction techniques currently being evaluated in these programs are described along with the results and a qualitative assessment of development difficulty.
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.
A gas turbine engine includes a core engine that includes a core flow path where air is compressed in a compressor section, communicated to a combustor section, mixed with an ammonia based fuel and ignited to generate a high energy combusted gas flow that is expanded through a turbine section. The turbine section is mechanically coupled to drive the compressor section. An ammonia flow path communicates an ammonia flow to the combustor section. A cracking device is disposed in the ammonia flow path. The cracking device is configured to decompose the ammonia flow into a fuel flow containing hydrogen (H2). At least one heat exchanger is upstream of the cracking device that provides thermal communication between the ammonia flow and a working fluid flow such that the ammonia fluid flow accepts thermal energy from the working fluid flow.