Engineering topics

Kenji Miki

Publications and source records attributed to Kenji Miki.

At least 37 records · Page 2

NASA NTRSDate not supplied

Numerical Investigation of the Effect of Trailing Edge Thickness of Simulated CMC Blades on Loss Profiles

To understand such an aerodynamic characteristic of the CMC blades, we numerically simulated three different CMC blades with varying trailing edge thicknesses. We utilized 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. We considered two sets of experimental test points with different Reynolds numbers at a given high-free stream turbulence range (Tu=~10-13%). The LES was previously validated by comparing with experimental data. We validated the RANS transition model against the ERCOFTAC experimental zero-pressure gradient smooth flat plate boundary layer at low- and high freestream turbulence offered by the experimental data of the Stripf’s for a turbine blade. We confirmed that the predicted skin friction and the heat flux agree well with the data. Finally, we computed the predicted pressure loading profiles, total loss distributions, and integrated loss of the three different CMC blades and made comparisons against the experimental data acquired at NASA Glenn Transonic Turbine Blade Cascade Rig. Our paper will present the aforementioned results and discuss the ability of LES and RANS model of our choice to predict salient aerodynamic characteristics of the blades with the various trailing edge thickness effects.

LES↗

NASA NTRSDate not supplied

Numerical Study of Combustor-Turbine Interactions Considering a Two-Stage High-Pressure Turbine

In this study, preliminary results are presented from the Open National Combustion Code (OpenNCC) applied to the two-stage high-pressure turbine (HTP) from the energy efficient engine (E3) program. Although the combustor is designed for efficient fuel–air mixing and a relatively uniform exit temperature profile, temporal and spatial variation in combustor exit conditions inevitably lead to hot-streaks in the HPT. These hot streaks can cause local hot spots on the HPT blade surfaces that reduce blade life. To improve understanding of these effects and the overall design of combustor and HPT, we are developing a fully-coupled combustor-turbine computational fluid dynamics (CFD) capability. Previous work has simulated the E^3 combustor at the simulated Sea-Level Take-off (SLTO) condition (P(41) = 2.4 [atm]) with and without the first-stage HPT stator included. This work addresses HPT simulations using inflow conditions based on previous simulation results of the E^3 combustor, as well as investigating the effects of spatial and temporal temperature non-uniformity of the combustor-exit flow on the HPT performance for a more realistic SLTO (P(41) = 27.4 [atm])). First, we analyze existing results from the combustor and the HPT 1st-stage stator simulation, where the unsteady hot gas coming out of the combustor exit travels between the stationary stators and then directly exits the computational domain (i.e., no effect from the rotors). Second, we simulate the HPT including the 1st and 2nd stage stators and rotors with the time-averaged, spatially-nonuniform inflow to the HPT taken from the combustor (with HPT 1st-stage stators) solution. Differences in simulations results for the flow field in the 1st-stage stator region are presented. The results of these two simulations (combustor and HPT simulated in a sequential manner) will be compared to a fully-coupled EEE combustor/HPT simulation to understand the impacts of fully-coupled simulations on combustor and HPT performance for two different operating conditions.

LES↗

NASA NTRSDate not supplied

Numerical Study of Combustor-Turbine Interactions Considering a Two-Stage High-Pressure Turbine

In this study, preliminary results are presented from the Open National Combustion Code (OpenNCC) applied to the two-stage high-pressure turbine (HTP) from the energy efficient engine (E3) program. Although the combustor is designed for efficient fuel–air mixing and a relatively uniform exit temperature profile, temporal and spatial variation in combustor exit conditions inevitably lead to hot-streaks in the HPT. These hot streaks can cause local hot spots on the HPT blade surfaces that reduce blade life. To improve understanding of these effects and the overall design of combustor and HPT, we are developing a fully-coupled combustor-turbine computational fluid dynamics (CFD) capability. Previous work has simulated the E3 combustor at the simulated Sea-Level Take-off (SLTO) condition (P41 = 2.4 [atm]) with and without the first-stage HPT stator included. This work addresses HPT simulations using inflow conditions based on previous simulation results of the E3 combustor, as well as investigating the effects of spatial and temporal temperature non-uniformity of the combustor-exit flow on the HPT performance for a more realistic SLTO (P41 = 27.4 [atm])). First, we analyze existing results from the combustor and the HPT 1st-stage stator simulation, where the unsteady hot gas coming out of the combustor exit travels between the stationary stators and then directly exits the computational domain (i.e., no effect from the rotors). Second, we simulate the HPT including the 1st and 2nd stage stators and rotors with the time-averaged, spatially-nonuniform inflow to the HPT taken from the combustor (with HPT 1st-stage stators) solution. Differences in simulations results for the flow field in the 1st-stage stator region are presented. The results of these two simulations (combustor and HPT simulated in a sequential manner) will be compared to a fully-coupled EEE combustor/HPT simulation to understand the impacts of fully-coupled simulations on combustor and HPT performance for two different operating conditions.

LES↗

NASA NTRSDate not supplied

The Effect of Trailing Edge Grid Resolution in Large-Eddy Simulations

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.

CMC↗

NASA NTRSDate not supplied

Computational Study of Modeling Fully-Coupled Combustor-Turbine Interactions by the Open National Combustion Code (OpenNCC)

LES↗

NASA NTRSDate not supplied

Combustor-Turbine Interactions By Using the Open National Combustion Code (OpenNCC) and the Glenn-HT Code

We develop the methodology of coupling two different computational fluid dynamic (CFD) codes, OpenNCC and Glenn-HT, in order to investigate the unsteady flow fields inside the combustor and around the first-stage stator of a high-pressure turbine (HPT) from the energy efficient engine (E 3 ) program. In our coupling strategy, OpenNCC, a time accurate unstructured mesh multi-phase combustion CFD code, is applied to the combustor region and Glenn-HT, a structured mesh CFD code designed for turbine applications, is utilized inside the HPT region. As a proof of concept, we first simulate the three-dimensional airflow over the backward facing step and compared the predicted pressure coefficient against experimental data. We then considered a Jet-A/Air combustor with the NASA single learn direct injector (LDI) as a reacting flow test case. Here, Glenn-HT is used in the downstream of the combustor and assumes that the working fluid is a single gas (i.e., air) where the properties of the gas are taken from time- and area-averaged values at the combustor exit. A fully coupled combustor-turbine simulation using both OpenNCC and Glenn-HT was conducted on the E 3 combustor and a detailed analysis done pertaining to how a realistic combustor outflow affects migration of hot-streaks and its impact on the aerodynamic behavior inside the HPT. It is found that the combustion dynamics (i.e., the switching main and pilot flame strength) significantly alters the aerodynamic behavior inside the HPT including hot-streak migration. Variations of the temperature and the velocity magnitude in the passage could be up to ±75 and ±15. These large variations are an important part of the combustor-turbine interactions, which are overlooked by a single component simulation of the HPT while imposing the steady inflow boundary condition at the HPT inlet. Finally, it is observed the strong density gradient associated with the hot-streaks and the pressure gradient at the passage significantly enhances the baroclinic torque, affecting the vorticity field.

CFD hot-streaks↗

NASA NTRSDate not supplied

The Effect of Trailing Edge Grid Resolution in Large-Eddy Simulations

LES↗

NASA NTRSDate not supplied

LES Analysis of Loss Anomaly in a Turbine Blade with Large Trailing Edge Radius

There is great interest in the usage of Ceramic Matrix Composites (CMC) as a turbine blade material. CMCs can tolerate higher temperatures than alloyed metallic blades, resulting in better thermal efficiency of gas turbine engines. However, depending on the manufacturing process of the CMC, blades may have a larger trailing edge thickness. The design space therefore needs to be updated due to the resulting flow physics. Recently, experimental results acquired at NASA Glenn Transonic Turbine Blade Cascade Rig showed that a loss measure generally increased with increasing trailing edge thickness. For one case however, the losses initially decreased with increasing Reynolds number but rose and peaked at Re ~1.24x10 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 higher Mach numbers. A Reynolds-averaged Navier-Stokes (RANS) analysis or a low-resolution large eddy simulation (LES) does not reproduce this apparent anomaly and thus it is worth performing a wall-resolved LES (the total mesh count of ~290 million cells) to investigate the aerodynamics of the CMC blade in the context of transonic vortex shedding. Our 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 said vortex shedding in the wake. This feedback seems not to 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.

LES↗

NASA NTRSDate not supplied

Application of Machine Learning Techniques in Calibration and Data Reduction of Multi-Hole Probes

This work presents procedures for implementing machine learning methods into existing algorithms for multi-hole probe calibration and data reduction. It demonstrates that using artificial neural networks (ANNs) can decrease the amount of calibration data needed to achieve a specific calibration uncertainty by over 50%, while also significantly reducing data reduction times. Instead of surface fitting methods, ANNs are employed. Initially, directional calibration coefficients related to flow angles are computed based on pressure measurements, and then these flow angles serve as input parameters for subsequent ANNs to iteratively define Mach number, static pressure, and total pressure. In an alternative approach, new calibration coefficients directly relate pressure measurements from the five-hole probe to the quantities of interest, thereby eliminating the need for iterative algorithms used in conventional surface fitting methods. This method offers several advantages: an average increase of less than 1%in calibration uncertainty for flow angles and a significant reduction in data reduction times to a few seconds on average. Additionally, the methodology is confirmed to avoid both over- and under-fitting.

Machine Learning↗

NASA NTRSDate not supplied

Application of Machine Learning Techniques in Calibration and Data Reduction of Multi-Hole Probes

This work presents procedures to implement machine learning methods in the existing algorithms for multi-hole probe calibrations and data reduction. It is shown here, that utilizing artificial neural networks (ANNs) can reduce the amount of calibration data that needs to be acquired in order to obtain a specific calibration uncertainty, by more than 50% while simultaneously reducing data reduction times significantly. ANNs were used instead of the surface fitting methods, where first, the directional calibration coefficients related to the flow angles are calculated based on the pressure measurements, and then the flow angles are used as a set of the input parameters for the following ANNs to define Mach number and static and total pressure iteratively. In a second approach, novel calibration coefficients were used to directly relate the pressure measurements from five-hole probe to the quantities of interest thus, eliminating the need for iterative algorithms used in the conventional surface fitting methods. The advantageous features of this method are an average increase of less than 1% in the calibration uncertainty for flow angles and significant reduction of the data reduction times (few seconds). In addition, we confirmed the methodology to avoid over-fitting and under-fitting.

Machine Learning↗

NASA NTRSDate not supplied

Cryogenic Extension of NASA Species Polynomials Using Hydrogen and Oxygen at Stoichiometry

The Helmholtz free energy for a mixture has been used in many chemistry and combustion studies for mixtures. An approach is described where polynomial fits intended for a thermally perfect ideal gases can be modified and implemented into a Helmholtz free energy model, such as those used by NIST in their REFPROP program. However, the limitation of the polynomial fits must be respected and should only used in a valid temperature interval where the flow field can be treated as a thermally perfect ideal gas when these species are present. This approach can allow a CFD simulation to have cryogenic regions of a flow field where high fidelity cryogenic species models are needed and allows for the use of species in other areas at higher temperatures where combustion species do not have such a model available or is not needed. Presented in the paper are CFD 1D detonation wave and species mass fraction simulation results which have cryogenic upstream conditions of Hydrogen and Oxygen at stoichiometry.

Combustion↗

NASA NTRSDate not supplied

Numerical Investigation of the Effect of Cooling Airflow inside High-Pressure Turbine using the Source Term Approach

In this study, the Open National Combustion code (OpenNCC) is applied to simulate the airflow inside the high-pressure turbine (HPT) of the energy efficient engine (EEE). The main objective of this study is to compare the characteristics of the HPT performance obtained by using three different approaches (Case 1: the realistic combustion products are considered at the HPT inlet, but the cooling airflows are not included, Case 2: the working fluid is assumed to be a single gas (i.e., air), and the cooling airflows are considered, Case 3: the cooling airflow and the combustion products at the HPT inlet are included). To mimic the cooling airflows at the HPT surface, we use the surface source team approach, in which we impose the source term at a specific area of a cooling airflow hole at the solid surfaces by specifying an injection angle, temperature, turbulent intensity, and mass flowrate of each cooling airflow. For the validation of the model, the data for the EEE model was obtained from the General Electric (GE) test campaign (RDG10), and the full scale warm-air rig test condition representing a test point close to the integrated core/low spool design condition was considered. This validation test was done using the stationary uniform inflow condition (i.e., the reported total temperature and total pressure). Secondly, we considered the Sea-Level Take-off (SLTO) engine condition and separately performed the combustor simulation to obtain a more realistic time-averaged, spatially-nonuniform HPT inflow condition. This work provides a better understanding the effect of the cooling airflow model with different assumptions on the HPT performance. It is found that using different working gas have a minor effect on the performance metric, however the presence of the cooling airflows significantly affects the HPT performance such as the turbine efficiency.

CFD Hot-streaks↗

NASA NTRSDate not supplied

Numerical Investigation of Combustor-Turbine Interactions with a Two-Stage High-Pressure Turbine Including Cooling Airflows

Combustor–turbine interactions (CTIs) are investigated by performing three-dimensional unsteady simulations using a realistic combustor and high-pressure turbine (HPT) configurations from the Energy Efficient Engine (E3) program. To understand CTIs, we compare the numerically predicted flow fields from single-component simulations (Step 1: the combustor + the first-stage stator of turbine; Step 2: the two-stage HPT imposing the time-averaged flow solution from Step 1 as the inflow boundary condition) and a fully coupled combustor–turbine simulation (Step 3) at the sea-level takeoff (SLTO) condition. In addition to three previous simulations where the cooling airflows inside the HPT had been neglected for all Step 1, Step 2, and Step 3, two new simulations of Step 2 and Step 3 take into account the cooling airflows using the source team approach. In this approach, to mimic the cooling airflows, we impose the source term at a specific area of a cooling airflow hole at the solid surfaces without making a mesh of each hole. The objective of this study is twofold. One is to investigate the effect of the presence of the cooling airflows on the aerodynamics of the combustor and HPT as well as the HPT performance. The second is to perform a detailed comparison among the calculated flowfields by two different numerical schemes, the central-difference with the standard Jameson–Schmidt–Turkel (CD-JST) scheme and the AUSM scheme. There is a noticeable difference in the hot-streak distributions at the first-stage stator. In addition, depending on the choice of the numerical scheme and the presence of the cooling airflows, an occurrence of shock waves appearing at the first-stage stator is greatly influenced. Thus, this has a noticeable impact on the HPT performance. It is shown that the estimated turbine efficiencies from Step 3 are about 7 % less than the ones from Step 2.

Combustor-turbine interaction↗

NASA NTRSDate not supplied

Validated Reduced Computational Methods for Realistic RDC Injection Modeling

This paper outlines a simplified computational method without chemistry for single and multi-injector performance for rotating detonation combustion (RDC) environment as well as relevant performance metrics. For non-reactive cases, an unsteady pressure profile from a Method of Characteristics (MOC) simulation was used to simulate an RDC combustion wave travelling around the annulus. In reactive cases, a pre-mixed stoichiometric mixture of hydrogen and air is injected into a 2D unwrapped combustion chamber with a specified injector area. The performance is evaluated based on several performance metrics including unsteady discharge coefficient, average mass flux coming into the combustor, and percentage time obstructed. Finally, the computational method was validated through an experimental RDC test setup with a single element injector in a hydrogen-air detonation driver environment.

Propulsion↗

NASA NTRSDate not supplied

Numerical Investigation of Combustor-Turbine Interactions with Cooling Air Flows Included in High-Pressure Turbine

In this paper, we would like to include the cooling airflows inside the turbine and perform the fully-coupled combustor–turbine simulation at a realistic sea-level takeoff (SLTO) condition. The new results are compared with the previous study,in which we performed the fully-coupled combustor–turbine simulation without the cooling airflows and the purge flows inside the HPT. This paper also serves to provide a better understanding about the effect of the unsteadiness in the turbine inflow condition on the turbine performance in the presence of the cooling airflows.

LES↗

NASA NTRSDate not supplied

Dynamic Analysis of Reynolds Number Effects on Trailing Edge Transonic Vortex Shedding and Its Impact on Turbine Blade Aerodynamic Performance

Time-resolved, high-speed self-aligned focusing schlieren images were acquired in the NASA Glenn Research Center Transonic Turbine Blade Cascade facility to help understand the aerodynamic behavior of high-pressure, thick trailing edge turbine blades. The trailing edge thickness of 9% of axial chord tested represents simulated ceramic matrix composite fabrication constraints, which was verified previously to possess a high-loss flow regime at high inlet turbulence conditions over a narrow range of Reynolds numbers and at a fixed design exit Mach number of 0.74. Our high-speed images, which were acquired at 10 distinct Reynolds numbers, show a significant increase in energy from flow oscillations due to transonic vortex shedding at Reynolds numbers corresponding to the high loss conditions. For those conditions, strong acoustic waves turn into shock waves. Spectral Proper Orthogonal Decomposition of the high-speed images shows acoustic waves from trailing edge vortex shedding at all conditions, with increased spectral energy at the high-loss conditions and slightly increasing frequency (about 6%) as a function of Reynolds number. Analysis of potential feedback timing is performed using velocity fields from a previous LES simulation, considering different feedback mechanisms. Most noteworthy is the acoustic/shock-boundary layer interaction mechanism on the suction surface at the blade geometric throat, which likely plays an important role in realistic curved blade passages.

Trailing Edge↗

NASA NTRSDate not supplied

Dynamic Analysis of Reynolds Number Effects on Trailing Edge Transonic Vortex Shedding and Its Impact on Turbine Blade Aerodynamic Performance

In this work, we will discuss observations from images acquired from a time-resolved, high-speed self-aligned focusing Schlieren campaign that was performed at the CW-22 facility at NASA Glenn Research Center to understand the dynamic behavior of thick trailing-edge high-pressure turbine blades simulating a ceramic matrix composite (CMC) construction at high inlet turbulence conditions. For the CMC-9 blade, which has a trailing edge thickness of 9% of the axial chord, we identified a regime where an excessive total pressure loss (loss anomaly) is observed only for a narrow range of Reynolds numbers at a fixed exit Mach number of 0.74. The loss anomaly is qualitatively explained by our images, which were taken at 10 distinct blade Reynolds numbers spanning a factor of 6. The images show a significant increase in energy related to the oscillations due to transonic vortex shedding at the Reynolds numbers related to the high loss conditions. This increased energy leads to the formation of strong acoustic waves that turn into shock waves at the highest loss conditions. From our observations stemming from Spectral POD analysis of the high-speed images, we see the acoustic waves produced by the trailing edge vortex shedding exist in all conditions tested; but the shedding frequency has a very slight trend upwards as the Reynolds number is increased, varying about 6% in the range tested. Considering this variation of shedding frequency as a function of Reynolds number, which is well-established for other bluff-body flows, we stipulate there may be a potential feedback mechanism involving an acoustic information transfer path across neighboring blades that may explain why only a narrow range of Reynolds numbers displays strong, shock-forming vortex shedding. We consider a few feedback paths and examine the timing based on the mean flow field from a high-resolution LES simulation. It appears that all feedback mechanisms are viable, presenting an integer number of delay cycles with respect to disturbances generated at the trailing edge. Most noteworthy, however, is the acoustic/shock-boundary layer interaction mechanism at the blade geometric throat. Based on our analysis, this mechanism likely plays an important role in realistic, curved turbine blade passages.

Aerodynamics↗

NASA NTRSDate not supplied

LES-Based Analysis of Film Cooling in a High-Pressure Turbine Rotor Including Effects of Purge and Tip Flow

Accurate modeling of coolant airflows, which form protective films over turbine blades, is essential for designing fuel-efficient and environmentally sustainable gas turbine engines. Excessive coolant reduces thermal efficiency, while insufficient coolant leads to blade overheating and causes damage. Therefore, precise prediction of flow field interactions with cooling air is critical for optimizing turbine performance. This study numerically investigates the cooling effectiveness of purge and film cooling flows within a high-pressure turbine (HPT) rotor using Large Eddy Simulation (LES). The study utilizes NASA Glenn Research Center’s Glenn-HT solver. The simulation models ethe conditions of the Penn State University START rotating rig. A high-fidelity structured mesh comprising up to 800 million cells is employed to resolve high-Reynolds number flow (Re ≈ 350,000) and to capture intricate secondary flow structures, including tip leakage and purge-induced vortices. Film cooling effectiveness computations are highly sensitive to boundary conditions at the cooling holes and to grid resolution. Even with well-resolved grids and included plena, strong mixing challenges traditional eddy viscosity models. To address this, a simplified configuration is simulated: a truncated row of shaped holes on the suction side near the leading edge and a row on the pressure side, both fed from internal plena while the purge slot and tip clearance are also modeled. Two isothermal LES cases are conducted at two distinct wall temperatures, which yield the adiabatic wall temperature and the heat transfer coefficient. The definition and means of computation of the effectiveness is discussed in this paper. The simulations reveal detailed three-dimensional unsteady flow features, including coherent vortical structures and secondary flows originating from the purge cavity. Film cooling effectiveness and Nusselt number distributions are presented for both the blade surface and tip, highlighting regions of elevated heat transfer and complex thermal behavior. These findings underscore the importance of high-resolution LES and realistic boundary conditions in capturing the dynamics of purge and film cooling, offering valuable insights for improving turbine blade design and thermal management strategies.

Secondary Flows↗