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Brief description of some of the recent trends in turbine aerodynamics research. Areas covered include cooled turbine aerodynamics, high work-factor turbines, pneumatic variable geometry, and computer analysis. Most of this research is shown to have been directed primarily toward allowing the turbine to accept efficiently higher inlet temperatures and blade loading.
The boundary-layer losses associated with the flow process in the blade channel are discussed. To determine the overall design point efficiency of a turbine, other losses must also be considered; these include tip clearance loss and disk friction loss. The sum of these losses normally comprises all the losses that are considered in the design of a full admission axial flow turbine. If, however, a partial admission turbine is being considered, there are additional losses that must be included. The partial admission losses usually considered are the pumping loss in the inactive blade channels and the filling-and-emptying loss in the blade passages as they pass through the admission arc. Finally, a loss that occurs at off-design operation of any turbine is the incidence loss.
Geothermal power conversion cycles are compared with respect to recovery of the available wellhead power. The cycles compared are flash steam, in which steam turbines are driven by steam separated from one or more flash stages; binary, in which heat is transferred from the brine to an organic turbine cycle; flash binary, in which heat is transferred from flashed steam to an organic turbine cycle; and dual steam, in which two-phase expanders are driven by the flashing steam-brine mixture and steam turbines by the separated steam. Expander efficiencies assumed are 0.7 for steam turbines, 0.8 for organic turbines, and 0.6 for two-phase expanders. The fraction of available wellhead power delivered by each cycle is found to be about the same at all brine temperatures: 0.65 with one stage and 0.7 with four stages for dual stream; 0.4 with one stage and 0.6 with four stages for flash steam; 0.5 for binary; and 0.3 with one stage and 0.5 with four stages for flash binary.
Aircraft gas turbine performance and efficiency are related to airfoil tip clearance. The possibility has been considered to obtain optimum performance and efficiency by reducing clearance to a safe minimum with the aid of a closed-loop tip clearance control system, which utilizes a tip clearance sensor. The use of optical sensing methods appears to represent a potential solution to the tip clearance measurement problem. Principles of sensor operation are discussed along with dimensional considerations and diffraction limitations. A description is presented of the study of the feasibility of a certain sensor, taking into account the test rig system, the optical components, and the mounting tube. Attention is also given to the operation of the feasibility-study sensor, performance estimations, the optical fiber bundle, light beam refraction, and aspects of aircraft engine implementation.
The performance of a variable-area stator, axial flow power turbine was determined in a cold-air component research rig for two inlet duct configurations. The two ducts were an interstage diffuser duct and an accelerated-flow inlet duct which produced stator inlet boundary layer flow blockages of 11 percent and 3 percent, respectively. Turbine blade total efficiency at design point was measured to be 5.3 percent greater with the accelerated-flow inlet duct installed due to the reduction in inlet blockage. Blade component measurements show that of this performance improvement, 35 percent occurred in the stator and 65 percent occurred in the rotor. Analysis of inlet duct internal flow using an Axisymmetric Diffuser Duct Code (ADD Code) were in substantial agreement with the test data.
The performance of a variable-area stator, axial flow power turbine was determined in a cold-air component research rig for two inlet duct configurations. The two ducts were an interstage diffuser duct and an accelerated-flow inlet duct which produced stator inlet boundary layer flow blockages of 11 percent and 3 percent, respectively. Turbine blade total efficiency at design point was measured to be 5.3 percent greater with the accelerated-flow inlet duct installed due to the reduction in inlet blockage. Blade component measurements show that of this performance improvement, 35 percent occurred in the stator and 65 percent occurred in the rotor. Analysis of inlet duct internal flow using an Axisymmetric Diffuser Duct Code (ADD Code) were in substantial agreement with the test data.
The continuing improvement of aircraft gas turbine engine operating efficiencies involves increases in overall engine pressure ratio increases that will result in combustor inlet pressure and temperature increases, greater combustion temperature rises, and higher combustor exit temperatures. These conditions entail the development of fuel injectors generating uniform circumferential and radial temperature patterns, as well as combustor liner configurations and materials capable of withstanding increased thermal radiation even as the amount of cooling air is reduced. Low NO(x)-emitting combustor concepts are required which will employ staged combustion. The development status of component technologies answering these requirements are presently evaluated.
The continuing improvement of aircraft gas turbine engine operating efficiencies involves increases in overall engine pressure ratio increases that will result in combustor inlet pressure and temperature increases, greater combustion temperature rises, and higher combustor exit temperatures. These conditions entail the development of fuel injectors generating uniform circumferential and radial temperature patterns, as well as combustor liner configurations and materials capable of withstanding increased thermal radiation even as the amount of cooling air is reduced. Low NO(x)-emitting combustor concepts are required which will employ staged combustion. The development status of component technologies answering these requirements are presently evaluated.
Thermally sprayed coatings have been extensively used to enhance materials properties and provide surface protection against their working environments in a number of industrial applications. Thermal barrier coatings (TBC) are used to reduce the thermal conductivity of aerospace turbine blades and improve the turbine overall thermal efficiency. TBC allows higher gas operating temperatures and lower blade material temperatures due to the thermal insulation provided by these ceramic coatings. In the automotive industry, coatings are currently applied to a number of moving parts that are subjected to friction and wear inside the engine such as pistons, cylinder liners, valves and crankshafts to enhance their wear resistance and prolong their useful operation and lifetime.
High-temperature shape-memory alloys (HTSMAs) based on nickel-titanium (NiTi) with significant ternary additions of palladium (Pd), platinum (Pt), gold (Au), or hafnium (Hf) have been identified as potential high-temperature actuator materials for use up to 500 C. These materials provide an enabling technology for the development of "smart structures" used to control the noise, emissions, or efficiency of gas turbine engines. The demand for these high-temperature versions of conventional shape-memory alloys also has been growing in the automotive, process control, and energy industries. However these materials, including the NiPtTi alloys being developed at the NASA Glenn Research Center, will never find widespread acceptance unless they can be readily processed into useable forms.
Cooling Flow Reduction: a) Focus on improving turbine hot gas path part cooling efficiency. b) Applicable to current metallic turbine components and synergistic with advanced materials. c) Address challenges of IGCC/hydrogen fuel environment (for example, possible cooling hole plugging). Leakage Flow Reduction: a) Focus on decreasing turbine parasitic leakages, i.e. between static-to-static, static-to-rotating turbine parts. b) Develop improved seal designs in a variety of important areas. Purge Flow Reduction: a) Focus on decreasing required flows to keep rotor disk cavities within temperature limits. b) Develop improved sealing at the cavity rims and modified flow geometries to minimize hot gas ingestion and aerodynamic impact.
To enhance efficiency of gas turbines, new thermal barrier coatings (TBCs) must be designed which improve upon the thermal stability limit of 7 wt% yttria stabilized zirconia (7YSZ), approximately 1200 C. This tenant has led to the development of new TBC materials and microstructures capable of improved high temperature performance. This study focused on increasing the erosion durability of cubic zirconia based TBCs, traditionally less durable than the metastable t' zirconia based TBCs. Composite TBC microstructures composed of a low thermal conductivity/high temperature stable cubic Low-k matrix phase and a durable t' Low-k secondary phase were deposited via APS. Monolithic coatings composed of cubic Low-k and t' Low-k were also deposited, in addition to a 7YSZ benchmark. The thermal conductivity and erosion durability were then measured and it was found that both of the Low-k materials have significantly reduced thermal conductivities, with monolithic t' Low-k and cubic Low-k improving upon 7YSZ by approximately 13 and approximately 25%, respectively. The 40 wt% t' Low-k composite (40 wt% t' Low-k - 60 wt% cubic Low-k) showed a approximately 22% reduction in thermal conductivity over 7YSZ, indicating even at high levels, the t' Low-k secondary phase had a minimal impact on thermal in the composite coating. It was observed that a mere 20 wt% t' Low-k phase addition can reduce the erosion of a cubic Low-k matrix phase composite coating by over 37%. Various mixing rules were then investigated to assess this non-linear composite behavior and suggestions were made to further improve erosion durability.
Ceramic Matrix Composite (CMC) with environmental barrier coating can tolerate significantly higher temperatures compared with the conventional metal alloy blades and thus can help achieve higher thermal efficiency in gas turbine engines. Due to the complex fabrication process of CMC blades, larger blade trailing edge thicknesses and larger leading-edge radii of curvature are expected. These features significantly alter aerodynamic performance especially near the trailing edge and thus the loss profile. In previous studies, the authors used the large eddy simulation (LES) as well as the Reynolds Averaged Navier-Stokes (RANS) coupled with an intermittency function-based transition model to simulate an aerodynamics performance of three different CMC blades and compared the predictions against the data acquired at NASA Glenn Transonic Turbine Blade Cascade Rig. Although we were able to successfully predict different characteristics of aerodynamic performance among three different blades (e.g., pressure loading and integrated losses), LES consistently predicted a small recirculation bubble at the end of the suction side. This was not inconsistent with the physics of the flow, but the pressure loading, locally, near the trailing edge, and as a result, the measured wake profiles did not match the experimental data. To delve further into these issues, a new LES grid with a much finer mesh near the trailing edge was generated where the spanwise extent of the grid was increased by 50%. An LES simulation for the CMC blade with the largest trailing edge thicknesses at the high-freestream turbulence (Tu=13%) condition showed that using such a mesh is necessary for LES to accurately capture the transition phenomena on the suction side. The resulting total pressure coefficient profile and the shift of the wake profile were greatly improved.
Ceramic Matrix Composite (CMC) with environmental barrier coating can tolerate significantly higher temperatures compared with the conventional metal alloy blades and thus can help achieve higher thermal efficiency in gas turbine engines. Due to the complex fabrication process of CMC blades, larger blade trailing edge thicknesses and larger leading-edge radii of curvature are expected. These features significantly alter aerodynamic performance especially near the trailing edge and thus the loss profile. In previous studies, the authors used the large eddy simulation (LES) as well as the Reynolds Averaged Navier-Stokes (RANS) coupled with an intermittency function-based transition model to simulate an aerodynamics performance of three different CMC blades and compared the predictions against the data acquired at NASA Glenn Transonic Turbine Blade Cascade Rig. Although we were able to successfully predict different characteristics of aerodynamic performance among three different blades (e.g., pressure loading and integrated losses), LES consistently predicted a small recirculation bubble at the end of the suction side. This was not inconsistent with the physics of the flow, but the pressure loading, locally, near the trailing edge, and as a result, the measured wake profiles did not match the experimental data. To delve further into these issues, a new LES grid with a much finer mesh near the trailing edge was generated where the spanwise extent of the grid was increased by 50%. An LES simulation for the CMC blade with the largest trailing edge thicknesses at the high-freestream turbulence (Tu=13%) condition showed that using such a mesh is necessary for LES to accurately capture the transition phenomena on the suction side. The resulting total pressure coefficient profile and the shift of the wake profile were greatly improved.
Thermal efficiency of gas turbine engine increases as the temperature and pressure at the combustor increases. Consequently, the materials used inside a combustor must survive an increasingly challenging environment. For this reason, accurate assessment of heat transfer is crucial for combustor design. While all three modes of heat transfer are present inside a combustor, the focus of this paper is the thermal radiation. Radiative heat transfer in a gas turbine combustors are particularly interesting from three reasons. Firstly, the radiative heat loss from the combustion region may affect the emission performance. Secondly, the cooling air will protect the liner from convection but not necessary from radiation. Finally, it is less frequently incorporated in CFD analysis than other forms of heat transfer. In this work, radiative heat transfer using discrete ordinate method has been incorporated in OpenNCC (a publicly releasable version of the National Combustion Code) developed at NASA Glenn Research Center. Aside from massively parallel computation capability using MPI and the ability to utilize unstructured mesh, the current implementation includes two types of spectral models, namely, the weighted some of gray gas model and the full spectrum correlated k-distribution model. After presenting the theory and the strategy of implementation, results of validation cases for gray gas and spectral models will be presented. While the implementation of the radiation solver is intended for gas turbine application, the radiation solver can run independently from the convection/combustion solver and the same theory can be applied to other application.
Thermal efficiency of gas turbine engine increases as the temperature and pressure at the combustor increases. Consequently, the materials used inside a combustor must survive an increasingly challenging environment. For this reason, accurate assessment of heat transfer is crucial for combustor design. While all three modes of heat transfer are present inside a combustor, the focus of this paper is the thermal radiation. Radiative heat transfer in a gas turbine combustors are particularly interesting from three reasons. Firstly, the radiative heat loss from the combustion region may affect the emission performance. Secondly, the cooling air will protect the liner from convection but not necessary from radiation. Finally, it is less frequently incorporated in CFD analysis than other forms of heat transfer. In this work, radiative heat transfer using discrete ordinate method has been incorporated in OpenNCC (a publicly releasable version of the National Combustion Code) developed at NASA Glenn Research Center. Aside from massively parallel computation capability using MPI and the ability to utilize unstructured mesh, the current implementation includes two types of spectral models, namely, the weighted some of gray gas model and the full spectrum correlated k-distribution model. After presenting the theory and the strategy of implementation, results of validation cases for gray gas and spectral models will be presented. While the implementation of the radiation solver is intended for gas turbine application, the radiation solver can run independently from the convection/combustion solver and the same theory can be applied to other application.
Efficiency investigations have been made on a single-stage modification of the turbine of a Mark 25 aerial torpedo to determine the performance of the unit with five different turbine nozzles. The output of the turbine blades was computed by analyzing the windage and mechanical-friction losses of the unit. The turbine was faund to be most efficient with a cast nozzle having sharp-edged inlets to the nine nozzle ports. An analysis af the effectiveness af the first and second stages of the standard Mark 25 torpedo turbine indicates that the first- stage turbine contributes nearly all the brake power produced at blade-jet speed ratios above 0.26.
The propulsion element of the NASA Advanced Subsonic Technology (AST) initiative is directed towards increasing the overall efficiency of current aircraft engines. This effort requires an increase in the efficiency of various components, such as fans, compressors, turbines etc. Improvement in engine efficiency can be accomplished through the use of lighter materials, larger diameter fans and/or higher-pressure ratio compressors. However, each of these has the potential to result in aeroelastic problems such as flutter or forced response. To address the aeroelastic problems, the Structural Dynamics Branch of NASA Glenn has been involved in the development of numerical capabilities for analyzing the aeroelastic stability characteristics and forced response of wide chord fans, multi-stage compressors and turbines. In order to design an engine to safely perform a set of desired tasks, accurate information of the stresses on the blade during the entire cycle of blade motion is required. This requirement in turn demands that accurate knowledge of steady and unsteady blade loading is available. To obtain the steady and unsteady aerodynamic forces for the complex flows around the engine components, for the flow regimes encountered by the rotor, an advanced compressible Navier-Stokes solver is required. A finite volume based Navier-Stokes solver has been developed at Mississippi State University (MSU) for solving the flow field around multistage rotors. The focus of the current research effort, under NASA Cooperative Agreement NCC3- 596 was on developing an aeroelastic analysis code (entitled TURBO-AE) based on the Navier-Stokes solver developed by MSU. The TURBO-AE code has been developed for flutter analysis of turbomachine components and delivered to NASA and its industry partners. The code has been verified. validated and is being applied by NASA Glenn and by aircraft engine manufacturers to analyze the aeroelastic stability characteristics of modem fans, compressors and turbines.