A careful introduction of coolant airflows, which create the film protecting the turbine blade, is critical for a more fuel-efficient and environmentally friendly jet engine. Accurate prediction of streams mixing with the cooling airflows is of great interest for achieving a better design of a jet engine. The main objective of this article is to numerically investigate the rate of heat transfer in a high-pressure turbine rotor passage with purge flow at the hub using a large eddy simulation (LES). An in-house computational fluid dynamics solver, Glenn-HT from the NASA Glenn Research Center, was utilized. The three-dimensional blade and the conditions are those of the Penn State University START rotating rig. A high-quality 125 million cell structured grid, which adequately resolves the high Reynolds number flow (Re similar to 350,000) and the complex secondary flow structures, was constructed. To evaluate the adiabatic wall temperature, two LES simulations with different isothermal wall temperatures were carried out. This method is more robust especially when cooling air injections at the blade surface need to be considered. Our numerical simulations were able to capture a very accurate representation of three-dimensional unsteady flow structures near the tip as well as the secondary flow originating from the purge. Several distinct high heat transfer areas were identified. In addition, temporal "energy separation" in the high vorticity level regions was observed, which has not been reported before.

A careful introduction of coolant airflows, which create the film protecting the turbine blade, is critical for a more fuel-efficient and environmentally friendly jet engine. Accurate prediction of streams mixing with the cooling airflows is of great interest for achieving a better design of a jet engine. The main objective of this paper is to numerically investigate the rate of heat transfer in a high-pressure turbine (HPT) rotor passage with purge flow at the hub using a large eddy simulation (LES.) An in-house CFD solver, Glenn-HT from the NASA Glenn Research Center was utilized. The three-dimensional blade and the conditions are those of the Penn State University START rotating rig. A high-quality 125 million cell structured grid, which adequately resolves the high-Reynolds number flow (Re similar to 350,000) and the complex secondary flow structures, was constructed. To evaluate the adiabatic wall temperature, two LES simulations with different isothermal wall temperatures were carried out. This method is more robust especially when cooling air injections at the blade surface need to be considered. Our numerical simulations were able to capture a very accurate representation of three-dimensional unsteady flow structures near the tip as well as the secondary flow originating from the purge. Several distinct high heat transfer areas were identified. In addition, temporal "energy separation" in the high vorticity level regions was observed, which has not been reported before.

There is great interest in the usage of ceramic matrix composites (CMC) as a turbine blade material. However, depending on the manufacturing process of the CMC, blades may have a thicker trailing edge. The design space therefore needs to be updated due to the resulting flow physics. Recently, experimental results acquired at the NASA Glenn Transonic Turbine Blade Cascade Rig showed that a loss measure generally increased with increasing trailing edge thickness. For some cases, however, the general downward trend of the loss with increasing Reynolds number (Re) was interrupted by a local loss peak around Re similar to 1.24 x 10(6), and then subsequently dropped to the original pre-peak trend. A possible cause of this intriguing phenomenon was speculated to be transonic vortex shedding, which is the mechanism of vortex shedding promoted by reflected shed pressure waves at the trailing edge at relatively high Reynolds numbers and transonic Mach numbers. A Reynolds-averaged Navier-Stokes analysis or an under-resolved large eddy simulation (LES) does not reproduce this apparent anomaly and thus a highly-resolved LES (the total mesh count of similar to 290 million cells) was performed to investigate the aerodynamics of the CMC blade. The numerical results at Re =1,246,350 show that the pressure waves generated by the vortex shedding in the wake travel upstream and significantly influence the transition and separation on the suction side thus enhancing the vortex shedding in the wake. This feedback does not hold under a low-Re condition (Re = 621,900). The Reynolds number dependence was also examined by numerical perturbation of the pressure waves in the wake and by examining how such perturbation attenuates or endures. It is confirmed that the perturbation of the pressure waves is quickly damped below a set Reynolds number.

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 large eddy simulation (LES) as well as the Reynolds Averaged Navier-Stokes (RANS) coupled with an intermittency function-based transition model to simulate aerodynamic performance of three different CMC blades and compared the predictions against the data acquired at NASA Glenn Transonic Turbine Blade Cascade Rig. Although these 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. The spanwise extent of the grid was increased by 50% at the same time to accommodate large vortices which might not have been captured by the original domain due to size limitations. 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. The increase in the span was found to be inconsequential.

Ceramic Matrix Composite (CMC) is an enabling material allowing higher turbine inlet temperatures and possibly resulting in better thermal efficiency of gas turbine engines. This is because CMC layers with environmental barrier coating (EBC) have significantly higher thermal limits (∼1755K) compared with the more conventional alloyed metallic blades. CMCs possess a complex fabrication process resulting in different geometrical characteristics than metallic blades (e.g., larger trailing edge thickness, large leading-edge curvature, etc.). It is therefore desirable to assess the aerodynamic performance of the CMC blades using experimental and numerical simulations. To investigate, three different CMC blades with varying trailing edge thicknesses were numerically simulated. Both Large Eddy simulation (LES) with the dynamic subgrid closure as well a recently implemented Reynolds-averaged Navier-Stokes (RANS) coupled with an intermittency function-based transition model were utilized. Two sets of experimental test points with different Reynolds numbers at a given high-free stream turbulence range (Tu = ∼10–13%) were considered. The LES was previously validated by comparing with experimental data. The RANS transition model was validated against the experimental data of the Stripf’s for a turbine blade. The predicted pressure loading profiles, total loss distributions, and integrated loss of the three different CMC blades were computed and comparisons against the experimental data are presented herein. The mixing and boundary layer loss was evaluated by both the RANS and LES and was found that the mixing loss forms a larger proportion of the overall loss as the trailing edge thickness increases.

Experimental and numerical investigations were performed to evaluate the effects of the exit nozzle configuration of a fluidic oscillator on the flowfield and impingement heat transfer performance downstream on a flat plate. A conventional wall-attachment-type fluidic oscillator with a fan-shaped exit configuration was studied. Eight different exit fan angles (0 deg <= theta <= 130 deg) were considered, and the effects of each nozzle configuration on the theta(jet) oscillation frequency, average fluid fan angle theta(jet), and discharge coefficient C-d were studied. Results show that the exit fan angle has a significant effect on the average fluid fan angle. Initially, theta(jet) increases along with the fan angle up to (theta <= 70 deg); beyond which, theta(jet )eventually plateaus with further increase in theta. The jet oscillation frequency appears to be independent of nozzle configurations. A case study was performed that includes sweeping jet impingement heat transfer in order to assess the practical implication of the exit geometry. Heat transfer experiments were performed for each nozzle configuration at a jet-to-wall spacing of HID = 5 and three coolant mass flow rates (<(m)over dot> = 0.97, 1.48, and 1.97 g/s). The average heat transfer tends to deteriorate with increasing fan angle. However, a pronounced improvement in cooling uniformity zeta was observed at large-fan-angle configurations (theta >= 70 deg). The time-resolved velocity field obtained from an unsteady Reynolds-averaged Navier-Stokes simulation revealed that separation bubble starts to appear at the diverging wall of the nozzle exit when theta > 70 deg that prevents the lateral motion of the jet, thus reduces the average fluid fan angle.

The structural and aerodynamic performance of a low aspect ratio SiC/SiC ceramic matrix composite (CMC) high pressure turbine (HPT) blade was determined. The application was a NASA notional single aisle aircraft engine to be available in the N+3, beyond 2030, time frame. The notional rpm was maintained, and to satisfy stress constraints, the annulus area was constrained. This led to a low span blade. For a given clearance, low span blade is likely to have improved efficiency when shrouded. The efficiency improvement due to shrouding was found to strongly depend on the axial gap between the shroud and casing. Axial gap, unlike clearance or reaction, is not a common parameter used to correlate the efficiency improvement due to shrouding. The zero clearance stage efficiency of the low aspect ratio turbine was 0.920. Structural analyses showed that the rotor blade could be shrouded without excessive stresses. The goal was to have blade stresses less than 100MPa (14.5 ksi) for the unshrouded blade. Under some not very restrictive circumstances, such as blade stacking, a one-dimensional radial stress equation accurately predicted area averaged Von Mises stress at the blade hub. With appropriate stacking, radial and Von Mises stresses were similar.

The Paper involves experimental and numerical heat transfer analyses of four pin-fin cooling configurations in a rectangular channel. Two design concepts were studied with two separate pin-fin (cylindrical and triangular) geometries. The first design includes a single-wall pin-fin configuration where the coolant flows through an array of full-length cylindrical (PF) and triangular pin fins (TF). The second configuration involves a double-wall pin-fin-jet configuration where the coolant jet impinges on the target wall, then flows through the array of partial-length cylindrical pin fins with jets (PFC) and partial-length triangular pin fins with jets (TFC). IR thermography was used to measure the steady state wall temperature of the target surface to estimate the overall cooling effectiveness phi. A transient heat transfer experiment was conducted to evaluate the internal heat transfer coefficient from the transient wall temperature. It is found that the double-wall pin-fin-jet (PFC and TFC) configurations show higher cooling effectiveness compared to the single-wall pin-fin (PF and TF) designs and the triangular pin fin (TF and TFC) showed a higher heat transfer augmentation compared to the cylindrical pin fin (PF and PFC) design. Numerical study later confirmed that the triangular pin fin creates a pair of streamwise vortices that augments the local heat transfer. The presence of the pin fin in the double-wall configuration prevents the development of a strong crossflow that negatively affects the heat transfer of the impinging jet. In addition, the triangular pin fin not only prevents the strong crossflow but redirects the coolant in the lateral direction that eventually contributes to the formation of a counterrotating vortex pair in the streamwise direction and thus augments local heat transfer and improves overall cooling effectiveness. Pressure drop measurement showed that the full-length triangular pin fin has a higher-pressure loss but the partial-length pin fin can lower the pressure loss considerably without compromising additional cooling performance.

There is a need to improve predictions of losses resulting from large eddy simulations (LES) of low-pressure turbines (LPT) in gas turbines. This may be done by assessing the accuracy of predictions against validation data and understanding the source of any inaccuracies. LES is a promising approach for capturing the laminar/turbulent transition process in a LPT. In previous studies, the authors utilized LES to model the flow field over a Variable Speed Power Turbine (VSPT) blade and successfully captured characteristic features of separation/reattachment and transition on the suction side at both the cruise (positive incidence) and take-off conditions (negative incidence) and as well, simulated the effect of free-stream turbulence (FST) on those phenomena. The predicted pressure loading profiles agreed well with the experimental data for both a high and a low FST case at a Reynolds number of Reex = 220,000. In this paper, we present wake profiles resulting from computations for a range of FST values. Although the predicted wake profiles for the lowest FST case (Tu = 0.5%) matched the experimental data, at higher FST (Tu = 10–15%,) the wake was wider than the experimentally measured wake and for both cases were displaced laterally when compared to the experimental measurements. In our investigation of the causes of the said discrepancies we have identified important effects which could strongly influence the predicted wake profile. Predicted losses were improved by assuring the validity of the flow solution. This was done by utilizing spectral analysis to scrutinize the dynamic behavior of the wake and determine solution accuracy resulting from low mesh density and low accuracy of convective modeling.

The objective of this research effort was to explore innovative cooling architectures enabled by additive manufacturing techniques for improved turbine cooling performance. The ability to create complex internal geometries was leveraged to better distribute coolant as well as to integrate inherently unstable flow devices to enhance internal and external heat transfer. This was accomplished with a multi-faceted approach including analytical, experimental, and computational components. This final report documents progress during the total 4.25 year effort that was extended (at no additional cost) from the original 3-year cooperative agreement. Prior to starting the project, significant effort was invested in investigating innovative cooling designs from the open literature. When the project started, we quickly down-selected to 4 configurations: impingement jet fluidic oscillators (for internal leading edge cooling), reverse film cooling (for pressure surface), sweeping film cooling jets (for suction surface), and trailing edge slot cooling with microchannels. During the 2nd year, the reverse film cooling design was shown to be sub-optimal and work on that topic was halted. The three remaining technologies were integrated into a large scale nozzle guide vane and installed into a low-speed linear cascade facility for further interrogation. Initial results showed the sweeping film cooling jets (on the suction surface) to be the most promising technology compared to a baseline diffusion shaped film hole (777 design). The benefit was particularly evident at high blowing ratios (>1.5) when the 777 coolant film separated from the downstream surface. Benefits were also evident at elevated freestream turbulence levels. The leading edge cooling with unsteady jets was less effective in terms of peak or average cooling – however it was superior to round hole impingement cooling in terms of spatial uniformity. Finally, the trailing edge cooling design with micro-channels was scrapped in favor of a pinned arrangement with centerbody. During the 3rd year, the same three technologies were integrated into a transonic linear cascade for an assessment of compressibility effects. Again, the sweeping film cooling jets proved superior at high blowing ratios while the unsteady leading edge impingement and trailing edge pin-fins with centerbody designs yielded mixed results. Finally, during the 4th and final year, these 3 technologies were integrated into a direct metal laser sintered (DMLS) nozzle guide vane for testing in a high temperature, transonic annular vane cascade. The facility matches the flow temperature, Mach number, and coolant pressure ratios of an actual gas turbine. Before doing this, an additional series of tests were completed to validate a redesigned trailing edge cooling architecture. This final design included a centerbody with triangular pins between the centerbody and the vane external skin. The majority of these triangular pins are fabricated with a 30% gap to the centerbody – reducing their pressure drop considerably while still providing excellent heat transfer augmentation. This additional test campaign required an additional 3-month extension request to complete testing in the high temperature NGV test facility (Turbine Reacting Flow Rig – TuRFR). Testing of the DMLS vane in TuRFR pitted the 3 innovative cooling technologies against more traditional technologies (shaped 777 film hole, round impingement jet, and full pin arrangement in the trailing edge. The sweeping film cooling jets (on the suction surface) showed approximately 15% improvement in overall effectiveness compared to the 777 film hole while the trailing edge showed up to 20% improvement. The leading edge impingement was disappointing with a substantially lower effectiveness than traditional direct impingement. In summary, the study was successful in demonstrating that DMLS-enabled cooling technologies can yield significant gains in cooling performance.

Abstract The structural and aerodynamic performance of a a low aspect ratio SiC/SiC CMC High Pressure Turbine blade was determined. The application was a NASA notional single aisle aircraft engine to be available in the N+3, beyond 2030, time frame. The notional rpm was maintained, and to satisfy stress constraints the annulus area was constrained. This led to a low span blade. For a given clearance low span blade are likely to have improved efficiency when shrouded. The efficiency improvement due to shrouding was found to strongly depend on the axial gap between the shroud and casing. Axial gap, unlike clearance or reaction, is not a common parameter used to correlate the efficiency improvement due to shrouding. The zero clearance stage efficiency of the low aspect ratio turbine was 0.920. Structural analyses showed that the rotor blade could be shrouded without excessive stresses. The goal was to have blade stresses less than 100 MPa (14.5 ksi) for the unshrouded blade. Under some not very restrictive circumstances, such as blade stacking, a one-dimensional radial stress equation accurately predicted area averaged Von Mises stress at the blade hub. With appropriate stacking radial and Von Mises stresses were similar.

This study includes the design, validation, and fabrication via direct metal laser sintering (DMLS) of a gas turbine nozzle guide vanes (NGV) that incorporates three innovative cooling schemes specifically enabled by additive manufacturing. The novel NGV design is the culmination of an extensive research and development effort over a period of 4 years that included low- and high-speed cascade testing coupled with unsteady computational fluid dynamics for numerous candidate innovative cooling architectures. The final vane design (SJ-vane) consists of sweeping jet (SJ) film cooling holes on the suction surface (SS), sweeping jet impingement holes at the leading edge and double-wall partial length triangular pin-fin with impinging jet at the trailing edge. For comparison purposes, a second DMLS enabled vane (777-vane) was designed and fabricated with prototypical cooling circuits to serve as a baseline. This vane consists of a shaped film cooling holes on the suction surface, circular impingement holes at the leading edge, and full-length cylindrical pin-fins at the trailing edge. Experiments with the two DMLS enabled vanes were performed at the Ohio State University Turbine Reacting Flow Rig (TuRFR) at engine-relevant temperature (1375 K) and Mach number conditions. Infrared (IR) thermography was utilized to measure the wall temperature of the pressure and suction surface at several coolant mass flowrates to estimate the overall cooling effectiveness (ϕ). Results showed improved cooling performance for the advanced cooling schemes (sweeping jet film cooling, impingement cooling, and triangular pin-fin cooling) compared with the baseline cooling schemes.

Experimental and numerical investigations were conducted 7 study the effects of high blowing ratios and high freestream turbulence on sweeping jet film cooling. Experiments were conducted on a nozzle guide vane suction surface in a low-speed near cascade. Experiments were performed at blowing ratios of 5-3.5 and free stream turbulence of 0.6% and 14.3%. Infrared thermography was used to estimate the adiabatic cooling effectiveness. Thermal field and boundary layer measurement ere conducted at a cross-plane (x/D = 12) downstream of the hole exit. Results were compared with a baseline 777-shaped hole and showed that sweeping jet hole has a better cooling performance at high blowing ratios. The Thermal field data revealed that the coolant separates from the surface at high blowing ratios for the 777-shaped hole while the coolant remains attached for the sweeping jet hole. Boundary layer measurement further confirmed that due to the sweeping action of the jet, the jet omentum of the sweeping jet hole is much lower than that of a 777-shaped hole. Thus the coolant remains closer to the wall even at high blowing ratios. Large Eddy Simulations (LES) were performed for both sweeping jet and the 777-shaped hole to evaluate the interaction between the coolant and the freestream at the near hole regions. Results showed that 777-shaped hole has strong jetting action at high blowing ratio that originates inside The hole breakout edges thus causing the jet to blow off from the surface. In contrast, the sweeping jet hole does not show this behavior due to its internal geometry and the sweeping action of the jet.

Numerical results are presented from the National Aeronautics and Space Administration (NASA) Glenn Research Center's in-house turbomachinery code Glenn-HT applied to the variable-speed power turbine (VSPT) experiment at the NASA Transonic Turbine Blade Cascade Facility. The main goal of this paper is to implement a digital filtering method to generate turbulence upstream and a subgrid model (localized dynamic k-equation model (LDKM)) in the framework of large-eddy simulation (LES) in order to investigate the effect of inflow turbulence on the transition seen in the VSPT experimental data at the cruise condition (incidence angle of 40 deg and Tu = 0.5%, 5%, 10%, and 15%). Although the boundary layer on the suction side and pressure side of the blades is initially laminar due to favorable pressure gradient, the laminar flow can transition to turbulent flow past a separation zone on the suction side or by natural or bypass transition. This process determines the total pressure losses in the wake. Therefore, it is desirable to develop a reliable prediction tool to accurately capture the transition mechanism in blade rows operated under the conditions of low Reynolds number and at a variety of freestream turbulence conditions. Our numerical studies reveal that the location of separation is rather insensitive to the level of Tu; however, the effect of increasing Tu seems to be in reducing the size and ultimately suppressing the separation bubble. In addition, we performed spectral analysis to identify the peak frequencies in the region where the separation bubble is formed, which provides valuable insights into the transition/separation mechanism.

A computational analysis is performed to determine if particulate impact events on the external surfaces of gas turbine engine rotor blades can be faithfully replicated in an experimental rotor cascade. The general electric (GE) energy efficient engine (E3) first-stage turbine flow-field at cruise conditions is first solved using a steady-state explicit mixing plane (MP) approach. To model flow in the cascade, a single E3 rotor periodic domain is then constructed with an inlet section matching the relative flow incidence angle from the mixing plane calculation. The mass-averaged relative flow conditions at the inlet and outlet of the mixing plane rotor section are imposed on the cascade boundaries and a steady solution is found. Particles with diameters ranging from 1 to 25 µm are tracked through each domain and the OSU deposition model is implemented to dictate the sticking and rebounding action of particles impacting solid surfaces. It is discovered that both the locations and parameters of the impacts in the cascade vary significantly from the engine environment. For smaller particles, this is credited to a stronger upstream influence of the blade on the cascade flow-field. As size increases, differences in deposition are instead driven by the interaction of the full-stage vane with the particles. The lack of a vane in the cascade causes drastically different particle inlet vectors over the rotor than are seen in the engine setting. The radial differences of particle impact locations are explored, and the role that pressure plays is considered.

A numerical study was conducted to evaluate the loss sensitivity of shrouded vs. unshrouded turbine rotor blades. Accuracy is demonstrated with a series of grid independence studies. Application of the methods is performed through various studies related to the effects of shrouding a High -Pressure Turbine (HPT) rotor blade for a NASA -specified N+3 timeframe single -aisle aircraft engine at takeoff conditions. Flat, Recessed, and Shrouded rotor configurations are evaluated at tip clearances from 0.25% to 4% of blade span. Mach # distributions, near-tip blade loading, and other flow characteristics are examined. Plots of stage efficiency vs. tip clearance are presented, with trends compared to available experimental data. It is shown that for the imposed boundary conditions, the addition of a shroud improves stage efficiency and significantly reduces sensitivity to tip clearance at higher clearance fractions. A casing recess is also shown to slightly increase sensitivity to tip clearance for tip clearances greater than 0.5%. Total pressure loss profiles vs. blade span are also compared, providing insight into the mechanisms behind the performance of the three configurations.

It is a challenge to simulate the flow in a Variable Speed Power Turbine (VSPT), or, for that matter, rear stages of low pressure turbines at low Reynolds numbers due to laminar flow separation or laminar/turbulent flow transition on the blades. At low Reynolds numbers, separation induced-transition is more prevalent which can result in efficiency lapse. LES has been used in recent years to simulate these types of flows with a good degree of success. In the present work, very low free stream turbulence flows at exit Reynolds number of 220k were simulated. The geometry was a cascade which was constructed with the midspan section of a VSPT design. Most LES simulations to date, have focused on the midspan region. As the endwall effect was significant in these simulations due to thick incoming boundary layer, full blade span computation was necessitated. Inlet flow angles representative of take-off and cruise conditions, dictated by the rotor speed in an actual design, were analyzed. This was done using a second order finite volume code and a high resolution grid. As is the case with Implicit-LES methods, no subgrid scale model was used. Blade static pressure data, at various span locations, and downstream probe survey measurements of total pressure loss coefficient were used to verify the results. The comparisons showed good agreement between the simulations and the experimental data.