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A gas turbine engine includes a core flow passage, a bypass flow passage, and a propulsor arranged at an inlet of the bypass flow passage and the core flow passage. The propulsor includes a row of propulsor blades. The row includes no more than 20 of the propulsor blades. The propulsor has a pressure ratio between about 1.2 and about 1.7 across the propulsor blades.
A gas turbine engine includes a bypass flow passage that has an inlet and defines a bypass ratio in a range of approximately 8.5 to 13.5. A fan is arranged within the bypass flow passage. A first turbine is a 5-stage turbine and is coupled with a first shaft, which is coupled with the fan. A first compressor is coupled with the first shaft and is a 3-stage compressor. A second turbine is coupled with a second shaft and is a 2-stage turbine. The fan includes a row of fan blades that extend from a hub. The row includes a number (N) of the fan blades, a solidity value (R) at tips of the fab blades, and a ratio of N/R that is from 14 to 16.
A gas turbine engine includes a gear assembly and a bypass flow passage that includes an inlet and an outlet that define a design pressure ratio between 1.3 and 1.55. A fan is arranged at the inlet. A first turbine is coupled with a first shaft such that rotation of the first turbine will drive the fan, through the first shaft and the gear assembly, at a lower speed than the first shaft. The fan includes a row of fan blades. The row includes 12-16 (N) fan blades, a solidity value (R) that is from 1.0 to 1.3, and a ratio of N/R that is from 10.0 to 16.
Advances in the gas turbine technology level and the corresponding advances in measurement instruments and technique are reviewed for each of the past decades, starting with the forties. The review provides a picture of the gradual development from the earlier, relatively simple systems, to the present sophisticated machines. A look in the future indicates that continued advances in gas turbine technology will be needed, used, and supported and that substantial changes are underway as to how these advances will be achieved. Some projections of the type of advances and in technology and measurements to be expected in the decade of the eighties are presented.
A system which would allow a substantially increased output from a turboshaft engine for brief periods in emergency situations with little or no loss of turbine stress rupture life is proposed and studied analytically. The increased engine output is obtained by overtemperaturing the turbine; however, the temperature of the compressor bleed air used for hot section cooling is lowered by injecting and evaporating water. This decrease in cooling air temperature can offset the effect of increased gas temperature and increased shaft speed and thus keep turbine blade stress rupture life constant. The analysis utilized the NASA-Navy-Engine-Program or NNEP computer code to model the turboshaft engine in both design and off-design modes. This report is concerned with the effect of the proposed method of power augmentation on the engine cycle and turbine components. A simple cycle turboshaft engine with a 16:1 pressure ratio and a 1533 K (2760 R) turbine inlet temperature operating at sea level static conditions was studied to determine the possible power increase and the effect on turbine stress rupture life that could be expected using the proposed emergency cooling scheme. The analysis showed a 54 percent increse in output power can be achieved with no loss in gas generator turbine stress rupture life. A 231 K (415 F) rise in turbine inlet temperature is required for this level of augmentation. The required water flow rate was found to be .0109 kg water per kg of engine air flow.
Gas emissions were studied from a staged Rich Burn/Quick-Quench Mix/Lean Burn combustor were studied under test conditions encountered in High Speed Research engines. The combustor was modeled at conditions corresponding to different engine power settings, and the effect of primary dilution airflow split on emissions, flow field, flame size and shape, and combustion intensity, as well as mixing, was investigated. A mathematical model was developed from a two-equation model of turbulence, a quasi-global kinetics mechanism for the oxidation of propane, and the Zeldovich mechanism for nitric oxide formation. A mixing-controlled combustion model was used to account for turbulent mixing effects on the chemical reaction rate. This model assumes that the chemical reaction rate is much faster than the turbulent mixing rate.
Aircraft gas turbine engine developments
The Non-Metallic Gas Turbine Engine project, funded by NASA Aeronautics Research Institute (NARI), represents the first comprehensive evaluation of emerging materials and manufacturing technologies that will enable fully nonmetallic gas turbine engines. This will be achieved by assessing the feasibility of using additive manufacturing technologies for fabricating polymer matrix composite (PMC) and ceramic matrix composite (CMC) gas turbine engine components. The benefits of the proposed effort include: 50 weight reduction compared to metallic parts, reduced manufacturing costs due to less machining and no tooling requirements, reduced part count due to net shape single component fabrication, and rapid design change and production iterations. Two high payoff metallic components have been identified for replacement with PMCs and will be fabricated using fused deposition modeling (FDM) with high temperature capable polymer filaments. The first component is an acoustic panel treatment with a honeycomb structure with an integrated back sheet and perforated front sheet. The second component is a compressor inlet guide vane. The CMC effort, which is starting at a lower technology readiness level, will use a binder jet process to fabricate silicon carbide test coupons and demonstration articles. The polymer and ceramic additive manufacturing efforts will advance from monolithic materials toward silicon carbide and carbon fiber reinforced composites for improved properties. Microstructural analysis and mechanical testing will be conducted on the PMC and CMC materials. System studies will assess the benefits of fully nonmetallic gas turbine engine in terms of fuel burn, emissions, reduction of part count, and cost. The proposed effort will be focused on a small 7000 lbf gas turbine engine. However, the concepts are equally applicable to large gas turbine engines. The proposed effort includes a multidisciplinary, multiorganization NASA - industry team that includes experts in ceramic materials and CMCs, polymers and PMCs, structural engineering, additive manufacturing, engine design and analysis, and system analysis.
Open-cycle recuperated gas turbine plant with inlet temperatures of 1255 to 1644 K (1800 to 2500 F) and recuperators with effectiveness values of 0, 70, 80 and 90% are considered. A 1644 K (2500 F) gas turbine would have a 33.5% plant efficiency in a simple cycle, 37.6% in a recuperated cycle and 47.6% when combined with a sulfur dioxide bottomer. The distillate burning recuperated plant was calculated to produce electricity at a cost of 8.19 mills/MJ (29.5 mills/kWh). Due to their low capital cost $170 to 200 $/kW, the open cycle gas turbine plant should see duty for peaking and intermediate load duty.
Particle interactions in gas turbine engines can be multicomponent, complex phenomena leading to the degradation of thermal (TBCs) and environmental barrier coatings (EBCs) meant to protect engine components. Ingestion of particles into the engine can lead to recession of coatings due to particle erosion. Similarly, these same particles can become molten, adhere to coatings and result in thermochemical corrosion of coating materials. Particle erosion testing is often carried out where the particles are injected into a gas stream, accelerated within a nozzle, and impinge on sample. Conversely, most molten particle corrosion testing is often done in static laboratory furnaces, which does not capture the dynamic nature of deposition in application. Nevertheless, these damage mechanisms are often tested separately and no single standard exists to test both erosive and corrosive particle interactions with coating materials under relevant operating conditions for gas turbine engines. Understanding the synergies of particle interactions in engines is crucial in determining operating lifetimes of potential coating materials. Such considerations emphasize the need for realistic approaches in standardizing particle interaction testing in combustion environments. The current study outlines initial efforts at NASA Glenn’s Erosion Burner Rig Facility in improving dynamic erosion/corrosion testing methods by assessing the durability of state-of-the-art (SOA) TBC material 7 wt.% yttria stabilized zirconia (7YSZ) as a function of particle deposition rate, burner temperature, and particle size. Calibration data to determine particle deposition rate will be presented, and mass and optical profilometry measurements were utilized to estimate mass/volume loss versus deposition per increment of particulate used over time. Electron microscopy analyses were then carried out to assess coating damage after testing.
Particle interactions in gas turbine engines can be multicomponent, complex phenomena leading to the degradation of thermal (TBCs) and environmental barrier coatings (EBCs) meant to protect engine components. Ingestion of particles into the engine can lead to recession of coatings due to particle erosion. Similarly, these same particles can become molten, adhere to coatings and result in thermochemical corrosion of coating materials. Particle erosion testing is often carried out where the particles are injected into a gas stream, accelerated within a nozzle, and impinge on sample. Conversely, most molten particle corrosion testing is often done in static laboratory furnaces, which does not capture the dynamic nature of deposition in application. Nevertheless, these damage mechanisms are often tested separately and no single standard exists to test both erosive and corrosive particle interactions with coating materials under relevant operating conditions for gas turbine engines. Understanding the synergies of particle interactions in engines is crucial in determining operating lifetimes of potential coating materials. Such considerations emphasize the need for realistic approaches in standardizing particle interaction testing in combustion environments. The current study outlines initial efforts at NASA Glenn’s Erosion Burner Rig Facility in improving dynamic erosion/corrosion testing methods by assessing the durability of state-of-the-art (SOA) TBC material 7 wt.% yttria stabilized zirconia (7YSZ) as a function of particle deposition rate, burner temperature, and particle size. Calibration data to determine particle deposition rate will be presented, and mass and optical profilometry measurements were utilized to estimate mass/volume loss versus deposition per increment of particulate used over time. Electron microscopy analyses were then carried out to assess coating damage after testing.
Particle interactions in gas turbine engines can be multicomponent, complex phenomena leading to the degradation of thermal (TBCs) and environmental barrier coatings (EBCs) meant to protect engine components. Ingestion of particles into the engine can lead to recession of coatings due to particle erosion. Similarly, these same particles can become molten, adhere to coatings and result in thermochemical corrosion of coating materials. Particle erosion testing is often carried out where the particles are injected into a gas stream, accelerated within a nozzle, and impinge on sample. Conversely, most molten particle corrosion testing is often done in static laboratory furnaces, which does not capture the dynamic nature of deposition in application. Nevertheless, these damage mechanisms are often tested separately and no single standard exists to test both erosive and corrosive particle interactions with coating materials under relevant operating conditions for gas turbine engines. Understanding the synergies of particle interactions in engines is crucial in determining operating lifetimes of potential coating materials. Such considerations emphasize the need for realistic approaches in standardizing particle interaction testing in combustion environments. The current study outlines initial efforts at NASA Glenn’s Erosion Burner Rig Facility in improving dynamic erosion/corrosion testing methods by assessing the durability of state-of-the-art (SOA) TBC material 7 wt.% yttria stabilized zirconia (7YSZ) as a function of particle deposition rate, burner temperature, and particle size. Calibration data to determine particle deposition rate will be presented, and mass and optical profilometry measurements were utilized to estimate mass/volume loss versus deposition per increment of particulate used over time. Electron microscopy analyses were then carried out to assess coating damage after testing.
The impact of gas-turbine-engine-powered aircraft on worldwide pollution was defined within two major areas of contribution. First, the contribution of aircraft to the local air pollution of metropolitan areas and, second, the long-term effects on the chemical balance of the stratosphere of pollutants emitted from future generations of high-altitude, supersonic commercial and military aircraft. Preliminary findings indicate that stratospheric oxides of nitrogen (NOx) emissions may have to be limited to very low levels if, for example, ozone depletion with concomitant increases in sea-level radiation, are to be avoided. Theoretical considerations suggest that (NOx) levels as low as 1.0 gram per kilogram of fuel and less should be attainable from a idealized premixed type of combustor. Experimental rig studies were intended to explore new combustor concepts designed to minimize the formation of (NOx) in aircraft gas turbines and to define their major operational problems and limitations.
The requirements for hot-gas-path materials in gas turbine engines are demanding. These materials must maintain high strength and creep resistance in a particularly aggressive environment. A typical gas turbine environment involves high temperatures, rapid gas flow rates, high pressures, and a complex mixture of aggressive gases. Over the past forty years, a wealth of information on the behavior of ceramic materials in heat engine environments has been obtained. In the first part of the talk we summarize the behavior of monolithic SiC and Si3N4. These materials show excellent baseline behavior in clean, oxygen environments. However the aggressive components in a heat engine environment such as water vapor and salt deposits can be quite degrading. In the second part of the talk we discuss SiC-based composites. The critical issue with these materials is oxidation of the fiber coating. We conclude with a brief discussion of future directions in ceramic corrosion research.
Both recuperated and bottomed closed cycle gas turbine systems in electric power plants were studied. All systems used a pressurizing gas turbine coupled with a pressurized furnace to heat the helium for the closed cycle gas turbine. Steam and organic vapors are used as Rankine bottoming fluids. Although plant efficiencies of over 40% are calculated for some plants, the resultant cost of electricity was found to be 8.75 mills/MJ (31.5 mills/kWh). These plants do not appear practical for coal or oil fired plants.
The effect of direct water injection on the exhaust gas emissions of a turbojet combustor burning natural gas fuel was investigated. The results are compared with the results from similar tests using ASTM Jet-A fuel. Increasing water injection decreased the emissions of oxides of nitrogen (NOX) and increased the emissions of carbon monoxide and unburned hydrocarbons. The greatest percentage decrease in NOX with increasing water injection was at the lowest inlet-air temperature tested. The effect of increasing inlet-air temperature was to decrease the effect of the water injection. The reduction in NOX due to water injection was almost identical to the results obtained with Jet-A fuel. However, the emission indices of unburned hydrocarbons, carbon monoxide, and percentage nitric oxide in NOX were not.
Hydrogen, in the gas state, has been proposed to replace Jet-A (the fuel used for commercial jet engines) as a fuel for gas turbine combustion. For the combustion of hydrogen and oxygen only, water is the only product and the main greenhouse gas, carbon dioxide, is not produced. This is an obvious benefit of using hydrogen as a fuel. The situation is not as simple when air replaces oxygen in the combustion process. (Air is mainly a mixture of oxygen, nitrogen, and argon. Other components comprise a very small part of air and will not be mentioned.) At the high temperatures found in the combustion process, oxygen reacts with nitrogen, and this produces nitrogen oxide compounds, or NOx--the main component of atmospheric smog. The production of NOx depends mainly on two variables: the temperature at which combustion occurs, and the length of time that the products of combustion stay, or reside, in the combustor. Starting from a lean (excess air) air-to-fuel ratio, the goal of this research was to minimize hot zones caused by incomplete premixing and to keep the residence time short while producing a stable flame. The minimization of these two parameters will result in low- NOx hydrogen combustion.
This paper describes a gas turbine simulation system which utilizes the newly developed Java language environment software system. The system provides an interactive graphical environment which allows the quick and efficient construction and analysis of arbitrary gas turbine propulsion systems. The simulation system couples a graphical user interface, developed using the Java Abstract Window Toolkit, and a transient, space- averaged, aero-thermodynamic gas turbine analysis method, both entirely coded in the Java language. The combined package provides analytical, graphical and data management tools which allow the user to construct and control engine simulations by manipulating graphical objects on the computer display screen. Distributed simulations, including parallel processing and distributed database access across the Internet and World-Wide Web (WWW), are made possible through services provided by the Java environment.