Radial turbine cooling

Radial turbines have been used extensively in many applications including small ground based electrical power generators, automotive engine turbochargers and aircraft auxiliary power units. In all of these applications the turbine inlet temperature is limited to a value commensurate with the material strength limitations and life requirements of uncooled metal rotors. To take advantage of all the benefits that higher temperatures offer, such as increased turbine specific power output or higher cycle thermal efficiency, requires improved high temperature materials and/or blade cooling. Extensive research is on-going to advance the material properties of high temperature superalloys as well as composite materials including ceramics. The use of ceramics with their high temperature potential and low cost is particularly appealing for radial turbines. However until these programs reach fruition the only way to make significant step increases beyond the present material temperature barriers is to cool the radial blading

On Turbulence and its Effects on Aerodynamics of Flow through Turbine Stages

In reality, the flows encountered in turbines are highly three‐dimensional, viscous, turbulent, and often transonic. These complex flows will not yield to understanding or prediction of their behavior without the application of contemporary and strong modeling techniques, together with an adequate turbulence model, to reveal effects of turbulence phenomenon and its impact on flow past turbine blades. The discussion primarily targets the turbulence features and their impact on fluid dynamics; streaming of blades, and efficiency performance. Turbulence as a phenomenon, turbulence effects and the transition onset in turbine stages are discussed. Flow parameters distribution past turbine stages, approaches to turbulence modeling, and how turbulent effects change efficiency and require an innovative design, among others are presented. Furthermore, a comparison study regarding the application and availability of various turbulence models is fulfilled, showing that every aerodynamic effect, encountered of flow pass turbine blades can be predicted via different model. This work could be very helpful for researchers and engineers working on prediction of transition onset, turbulence effects, and their impact on the overall turbine performance

The Aero-Thermal Performance of Purge Flow and Discrete Holes Film Cooling of Rotor Blade Platform in Modern High Pressure Gas Turbines: A Review

Design of cooling systems for rotor platforms is critical due to the complex flow field and heat transfer phenomena related to the secondary flow structures originating at the blade leading edge. Horseshoe vortex and passage vortex are the fluid-dynamic features that largely influence the aerodynamic behaviour and the thermal protection level of the platform. The driving parameter is the coolant to mainstream momentum flux ratio, but several issues have to be considered in the design process of cooling technologies. As well acknowledged, an in-depth understanding of losses and heat transfer phenomena are deemed necessary to design effective cooling systems. In the present review, measurements and predictions on the behaviour of the HPT rotor cooled platform, obtained during the last two decades by several research groups, are gathered, described and analysed in terms of aerodynamic losses and heat transfer performance, and are compared with one another with respect to the effectiveness level that is ensured

The Design, Fabrication, and Validation of a Film Cooled Rotating Turbine Cascade with an Actively Cooled Shroud in a Closed Loop Wind Tunnel

To test shroud and blade cooling effectiveness, a closed loop, heated wind tunnel housing a film cooled rotating turbine cascade with prescribed blade and vane geometry surrounded by a fully cooled shroud with a leading edge gap were designed and assembled on Louisiana State University’s campus. Heat transfer coefficients and film cooling effectiveness results were computed using a 1-D semi-infinite solid conduction analysis of material temperatures obtained with liquid crystal thermography. Proper analysis required a step change in air temperature; so a bypass loop provided mainstream air heating while maintaining the shroud and blades at ambient temperature. Also, analysis required hollow tip film cooled turbine blades constructed of low thermal conductivity material, resulting in fabrication by 3-D plastic printing. An analytical stress model and finite element analysis validated plastic blade structural base design. Static-structural and dynamic fatigue loading analyses determined rotor shaft size. Heat transfer and pressure loss calculations verified the system’s required blade and shroud cooling characteristics. Finally, velocity vector measurements at the nozzle guide vane leading edge and recorded pressures in a nozzle guide vane passageway upstream of the turbine cascade location validated incoming freestream flow properties for the design condition at which heat transfer measurements were recorded. A total pressure loss analysis for varying rotor speed and blowing ratio was conducted with the development of total pressure contours downstream of the exit guide vanes to understand losses in the nozzle-rotor passage. Loss structures were found at the tip and root of the exit guide vane, attributed to the shed vortex and tip leakage vortex developed from the rotor blade. Total pressure loss decreased as blowing ratio increased due to energy added to the mainstream flow through the film cooling air. Rotor speed was varied from 55 to 655 RPM. Total pressure losses were lowest at 55 RPM, increased with increasing rotor speed past the 355 RPM design speed, and decreased as rotor speed approached 655 RPM. These results could be attributed to the introduction or extraction of shaft work in the rotating system along with aerodynamic losses associated with changes from the incidence angle design

Development of an Ultra-High Efficiency Gas Turbine Engine (UHEGT) with Stator Internal Combustion: Design, Off-Design, and Nonlinear Dynamic Operation

An Ultra-High Efficiency Gas Turbine (UHEGT) technology is developed in this study. In UHEGT, the combustion process is no longer contained in isolation between the compressor and turbine, rather distributed in multiple stages and integrated within the High-Pressure (HP)-turbine stator rows. Fundamental issues of aero-thermodynamic design, combustion, and heat transfer are addressed in this study. The aero-thermodynamic study shows that the UHEGT-concept improves the thermal efficiency of gas turbines by 5-7% above the current most advanced gas turbine engines, such as Alstom GT24. The designed thermodynamic cycle has a 45% thermal efficiency and includes a six-stage turbine with three stages of stator internal combustion. Meanline approach is used to preliminary design the entire flow path in the turbine. Multiple configurations are designed and simulated via Computational Fluid Dynamics (CFD) to achieve the optimum combustion system for UHEGT. Flow patterns, temperature distributions, secondary losses, etc. are among the parameters studied in the results. The final configuration for the combustion system includes two rows of injectors placed before the stator rows in the first three turbine stages. The current injector configuration provides a highly uniform temperature distribution at the rotor inlet, low pressure loss, and low emissions compared to the other cases. Different approaches are numerically studied to lower the stator blade surface temperature distribution in UHEGT from which indexing (clocking) is shown to be very effective. In the final part of this study, a dynamic simulation is performed on the entire engine using the nonlinear generic code GETRAN developed by Schobeiri. The simulations are in 2D (space-time) and include the complete gas turbine engine. The system performance is studied under variable design and off-design conditions. The results show that most of the system parameters fluctuate with similar patterns to the fuel schedule. However, the amplitudes of the fluctuations are different and there is a time lag in the response profiles relative to the fuel schedules. It is shown that thermal efficiency variations are smaller compared to the other parameters which means the system performs in efficiencies close to the design point throughout the entire cycle

Aeronautical engineering: A special bibliography with indexes, supplement 82, April 1977

This bibliography lists 311 reports, articles, and other documents introduced into the NASA scientific and technical information system in March 1977

Numerical Investigations of Flow and Film Cooling with Endwall Contouring and Blade Tip Ejection under Rotating Turbine Conditions

An effort is made to numerically study the impact of rotating turbine conditions on the aerodynamic performance, film cooling effectiveness and heat transfer with the application of the endwall contouring and blade tip ejection. For this purpose, the three-stage HP turbine research facility at the Turbomachinery Performance and Flow Research Laboratory (TPFL), Texas A&M University, was newly designed and equipped. Using the geometry of this three-stage research turbine rig, comprehensive numerical simulations are performed to systematically study the impact of the rotation from the perspectives of both aerodynamics and heat transfer. Introducing endwall contouring has become a promising means to reduce the secondary flow losses. Thus TPFL developed a physics-based method which enables researchers and engineers to design endwall contours for any arbitrary blade types regardless of the blade loading, degree of reaction, stage load and flow coefficients. Using this approach, TPFL designed the new endwall contouring which was implemented on the platform of both the first-stage and second-stage rotors. The rotation impacts on the aerodynamics performance due to the endwall contouring were numerically studied using four different rotational speeds namely, 2000 rpm, 2400 rpm, 2600 rpm and 3000 rpm. Meanwhile, the influence on film cooling effectiveness and heat transfer caused by the endwall contouring was investigated for the first-stage rotor. Different purge-to-mainstream mass flow ratios of MFR = 0.5%, 1.0% and 1.5% were taken into account at the design rotational speed of 3000rpm. The effect of rotational speed (2400rpm, 2550rpm and 3000rpm) was investigated at typical MFR=1.0%. To investigate the flow characteristics and film cooling at high pressure turbine blade tips, four different rotor-blade tip configurations are designed and studied at TPFL: the plane and squealer tips with tip hole cooling and the plane and squealer tips with pressure-side-edge compound angle hole cooling. Seven perpendicular holes that are evenly distributed along the camber line are used for the tip hole cooling, whilst eight compound-angle holes for the pressure-side-edge cooling. The coolant was ejected through the cooling holes with low, medium and high global blowing ratios at 3000 rpm to study the impact of the blowing ratio on both the cooling effectiveness and heat transfer. Effects of rotation on the cooling effectiveness and heat transfer were calculated at the rotational speeds of 2000rpm, 2550 rpm, and 3000 rpm

Effect of Velocity and Temperature Distribution at the Hole Exit on Film Cooling of Turbine Blades

An existing three-dimensional Navier-Stokes code (Arnone et al, 1991), modified Turbine Branch, to include film cooling considerations (Garg and Gaugler, 1994), has been used to study the effect of coolant velocity and temperature distribution at the hole exit on the heat transfer coefficient on three film-cooled turbine blades, namely, the C3X vane, the VKI rotor, and the ACE rotor. Results are also compared with the experimental data for all the blades. Moreover, Mayle's transition criterion (1991), Forest's model for augmentation of leading edge heat transfer due to free-stream turbulence (1977), and Crawford's model for augmentation of eddy viscosity due to film cooling (Crawford et al, 1980) are used. Use of Mayle's and Forest's models is relevant only for the ACE rotor due to the absence of showerhead cooling on this rotor. It is found that, in some cases, the effect of distribution of coolant velocity and temperature at the hole exit can be as much as 60 percent on the heat transfer coefficient at the blade suction surface, and 50 percent at the pressure surface. Also, different effects are observed on the pressure and suction surface depending upon the blade as well as upon the hole shape, conical or cylindrical

Effect of velocity and temperature distribution at the hole exit on film cooling of turbine blades

An existing three-dimensional Navier-Stokes code, modified to include film cooling considerations, has been used to study the effect of coolant velocity and temperature distribution at the hole exit on the heat transfer coefficient on three-film-cooled turbine blades, namely, the C3X vane, the VKI rotor, and the ACE rotor. Results are also compared with the experimental data for all the blades. Moreover, Mayle's transition criterion, Forest's model for augmentation of leading edge heat transfer due to freestream turbulence, and Crawford's model for augmentation of eddy viscosity due to film cooling are used. Use of Mayle's and Forest's models is relevant only for the ACE rotor due to the absence of showerhead cooling on this rotor. It is found that, in some cases, the effect of distribution of coolant velocity and temperature at the hole exit can be as much as 60% on the heat transfer coefficient at the blade suction surface, and 50% at the pressure surface. Also, different effects are observed on the pressure and suction surface depending upon the blade as well as upon the hole shape, conical or cylindrical