Optimization Study of Guide Vanes for the Intake Fan-Duct Connection Using CFD

The connection between an intake fan and a ventilation shaft must be designed in such a way that it minimizes the energy waste due to singularity losses. As a result, the questions of which radius of curvature to use and if guide vanes have to be included need to be answered. In that case, the variables such as the number, upstream and downstream penetration length, radius of curvature, and width of the vanes, need to be defined. Although this work is oriented to mine ventilation, these questions are usually valid in other engineering applications as well. The objective of this study is to define the previously mentioned variables to determine the optimal design combination for the radius/diameter relationship (r/D). Computational fluid dynamics was used to determine the shock loss factor of seven elbow curvature ratios for a 3 m diameter duct and fan, with and without guide vanes to estimate the best performing configuration and, therefore, to maximize the fan airflow volume. The methodology used consisted of initially developing models in 2D geometries, to optimize the meshing and the CPU use, and studying separately the number of vanes, upstream and downstream penetration, radius of curvature, and width of the vanes for each curvature ratio (r/D). Then, the best-performing variable combinations for each curvature ratio were selected to be simulated and studied with the 3D geometries. The application of the guide vane designs for three-dimensional simulated geometries is presented, first without and then with guide vanes, including the shock loss factors obtained. The methodology and obtained results allowed quantifying the energy savings and to reduce the CFD simulations steps required to optimize the design of the elbow and guide vanes. The results obtained cannot be used with elbows in exhaust fans, because fluid dynamics phenomena are different.

A Comprehensive Analytical Shock Loss Model for Axial Compressor Cascades

The approach applied in various research papers that model compressor shock losses is valid only for certain types of airfoil cascades operating in a narrow range of working conditions. Lately, more general shock loss models have been established that cover a wider variety of airfoils and operating regimes. However, owing to the complexity of the studied matter, the majority of such models are, to a certain extent, presented only in a descriptive manner. The lack of specific details can affect the end results when such a model is utilized since improvisation cannot be avoided. Some models also apply complex numerical procedures that can slow the calculations and be a source of computational instability. In this research, an attempt has been made to produce an analytical shock loss model that is simple enough to be described in detail while being universal and robust enough to find wide application in the fields of design and performance analysis of transonic compressors and fans. The flexible description of airfoil geometry encompasses a variety of blade shapes. Both unchoked and choked operating regimes are covered, including a precise prediction of choke occurrence. The model was validated using a number of numerical test cases.

Investigation of the flowfield on a vaneless counter-rotating turbine with seal cavity flow ejection

The sealing of the rotor-rotor gap and rotor disk cooling are vital to the safe operation of the vaneless counter-rotating turbine(VCRT). In order to quantifies the influence of the wheel-space cavity flow on the VCRT aerodynamic performance, and to improve turbine efficiency of the VCRT at certain rim seal ejection rates, numerical studies which considered the effects of rotor-rotor rim seal flow ejection are carried out in this paper. The three dimensional unsteady computational fluid dynamic analysis of a VCRT at the engine conditions are performed, and the seal flow ejected downstream of the high pressure rotor row at six sealing flow rate are modeled. The interaction among the high pressure rotor trailing shock wave, the downstream secondary flow and the seal flow has been studied and quantitatively characterized as a function of the purge ejection rate. Numerical results show that seal flow- mainstream flow interaction is entirely dominated by the high pressure rotor trailing edge shock at the hub, low pressure hub passage vortex and the mixing of the sealing flow from wheelspace and mainstream. When the mass flow rate of the coolant is smaller than some threshold value, the shock loss of the high pressure rotor and hub secondary flow loss of the low pressure rotor are decreased with the increasing of the coolant mass flow rate. It causes that the VCRT efficiency is gradually increased. On condition that the amount of the seal flows is beyond the threshold value, the key roles in modification of the VCRT performance are changed. The increment of the hub secondary flow loss and the mixing loss are gradually larger than the decrement of the shock loss. As a result, the turbine efficiency gradually decreases.

Regulation of Vacuole Morphology by PIEZO Channels in Spreading Earth Moss

The perception of mechanical force is a fundamental property of most, if not all cells. PIEZO channels are plasma membrane-embedded mechanosensitive calcium channels that play diverse and essential roles in mechanobiological processes in animals1,2. PIEZO channel homologs are found in plants3,4, but their role(s) in the green lineage are almost completely unknown. Plants and animals diverged approximately 1.5 billion years ago, independently evolved multicellularity, and have vastly different cellular mechanics5. Here, we investigate PIEZO channel function in the moss Physcomitrium patens, a representative of one of the first land plant lineages. PpPIEZO1 and PpPIEZO2 were redundantly required for normal growth, size, and shape of tip-growing caulonema cells. Both were localized to vacuolar membranes and facilitated the release of calcium into the cytosol in response to hypoosmotic shock. Loss-of-function (ΔPppiezo1/2) and gain-of-function (PpPIEZO2-R2508K and -R2508H) mutants revealed a role for moss PIEZO homologs in regulating vacuole morphology. Our work here shows that plant and animal PIEZO homologs have diverged in both subcellular localization and in function, likely co-opted to serve different needs in each lineage. The plant homologs of PIEZO channels thus provide a compelling lens through which to study plant mechanobiology and the evolution of mechanoperceptive strategies in multicellular eukaryotes.

New Supersonic Loss Model for the Preliminary Design of Transonic Turbine Blades and the Influence of Pitch

In order to produce a more efficient design of a compact turbine driving a cryogenic engine turbo-pump for a satellite delivering rocket, a new supersonic loss model is proposed. The new model was constructed based on high-quality published data, composed of Schlieren photographs and experimental measurements, that combined provided a unique insight into the mechanisms driving supersonic losses. Using this as a cornerstone, model equations were formulated that predict the critical Mach number and shock loss and shock-induced mixing loss as functions of geometrical (i.e., blade outlet and uncovered turning angle and trailing edge thickness) and operational parameters (i.e., exit Mach number). A series of highly resolved CFD numerical simulations were conducted on an in-house designed state-of-the-art transonic turbine rotor row (around unity aspect ratio (AR)) to better understand changes in the shock system for varying parameters. The main outcome showed that pitch to chord ratio has a powerful impact on the shock system, and thus on the manner by which shock loss and shock-induced mixing loss is distributed to compose the overall supersonic losses. The numerical loss estimates for two pitch to chord ratios—t⁄c = 0.70 and t⁄c = 0.98—were compared with absolute loss data of a previously published similar blade with satisfactory agreement. Calibrated equations are provided to allow hands-on integration into existing overall turbine loss models, where supersonic losses play a key role, for further enhancement of preliminary turbine design.

An Analytical Shock-Loss Model Fully Integrated to a Streamline Curvature Method for Axial-Flow Compressors

The constant and increasing demand to obtain more accurate turbomachinery performance prediction in the design and analysis process has led to the development of higher fidelity flow field models. Despite extensive flow field information can be collected from 3-D RANS numerical simulations, the computational cost is expensive in terms of time and resources, especially if they are used as solvers within a design-optimisation framework. In contrast, 2-D throughflow methods, such as streamline curvature (SLC), provide an acceptable flow solution in minutes. The use of modern and advanced-design transonic axial-flow compressors and fans has been expanding due to their high shock-induced single-stage pressure ratios while being light, compact and robust. Transonic-flow analysis in blading is complex due to the shock structures involved and associated phenomena. Previous 2-D SLC tools have failed to replicate the real compressible-flow physics, assuming and oversimplifying the shock-system shape and location. The situation aggravates, when the assumed overall shock configuration applies only for design point at unstarted operations, requiring of empirical correlations to estimate the shock-loss coefficient for off-design operations. The overall compressor performance prediction is thence highly-dependent on the shock modelling quality. For this reason, a physics-based shock -structure and -loss model was developed and implemented into an existing in-house 2-D SLC compressor performance simulator to enhance the aerodynamic prediction in transonic axial-flow compressors. The novel shock-loss model is fully coupled to the 2-D SLC software, for which a blade-element-layout method was adapted to obtain the profile geometry definition. The analytical shock-loss model possesses the capability to operate at started and unstarted passages utilizing an iterative-solution method to position the choke-induced passage-shock. A significant contribution of the new shock-loss model is the solution of the relative total-pressure loss for the entire blade span, comprising the inlet relative subsonic supercritical and supersonic regions. In this manner, shock losses were determined throughout the blade span and for various off-design operating conditions, including those at choking. 2-D SLC simulations were conducted for the NASA Rotor 67 Fan to validate the models accordingly against test-rig data and verify against previous model estimations and 3-D CFD results. The analytical shock - structure and -loss model improved the shock-loss prediction between 40–50% with respect of the state-of-the-art models and showed satisfactory agreement against measured data within 0.6% at the blade tip and 0.3% at mid-span sections.

Loss Generation in Transonic Turbine Blading

The measured loss characteristic in a high-speed cascade tunnel of two turbine blades of different designs showed distinctly different trends with exit Mach number ranging from 0.8 to 1.4. Assessments using steady Reynolds-averaged Navier--Stokes equations (RANS) computation of the flow in the two turbine blades, complemented with control volume analyses and loss modeling, elucidate why the measured loss characteristic looks the way it is. The loss model categorizes the total loss in terms of boundary layer loss, trailing edge (TE) loss, and shock loss; it yields results in good agreement with the experimental data as well as steady RANS computed results. Thus, RANS is an adequate tool for determining the loss variations with exit isentropic Mach number and the loss model serves as an effective tool to interpret both the computational and the experimental data. The measured loss plateau in blade 1 for exit Mach number of 1–1.4 is due to a balance between a decrease of blade surface boundary layer loss and an increase in the attendant shock loss with Mach number; this plateau is absent in blade 2 due to a greater rate in shock loss increase than the corresponding decrease in boundary layer loss. For exit Mach number from 0.85 to 1, the higher loss associated with shock system in blade 1 is due to the larger divergent angle downstream of the throat than that in blade 2. However, when exit Mach number is between 1.00 and 1.30, blade 2 has higher shock loss. For exit Mach number above an approximate value of 1.4, the shock loss for the two blades is similar as the flow downstream of the throat is completely supersonic. In the transonic to supersonic flow regime, the turbine design can be tailored to yield a shock pattern the loss of which can be mitigated in near equal amount of that from the boundary layer with increasing exit Mach number, hence yielding a loss plateau in transonic-supersonic regime.

Loss Generation in Transonic Turbine Blading

The measured loss characteristic in a high-speed cascade tunnel of two turbine blades of different designs showed distinctly different trend with exit Mach number ranging from 0.8 to 1.4. Assessments using steady RANS computation of the flow in the two turbine blades, complemented with control volume analyses and loss modelling, elucidate why the measured loss characteristic looks the way it is. The loss model categorizes the total loss in terms of boundary layer loss, trailing edge loss and shock loss; it yields results in good agreement with the experimental data as well as steady RANS computed results. Thus RANS is an adequate tool for determining the loss variations with exit isentropic Mach number and the loss model serves as an effective tool to interpret both the computational and experimental data. The measured loss plateau in Blade 1 for exit Mach number of 1 to 1.4 is due to a balance between a decrease of blade surface boundary layer loss and an increase in the attendant shock loss with Mach number; this plateau is absent in Blade 2 due to a greater rate in shock loss increase than the corresponding decrease in boundary layer loss. For exit Mach number from 0.85 to 1, the higher loss associated with shock system in Blade 1 is due to the larger divergent angle downstream of the throat than that in Blade 2. However when exit Mach number is between 1.00 and 1.30, Blade 2 has higher shock loss. For exit Mach number above around 1.4, the shock loss for the two blades is similar as the flow downstream of the throat is completely supersonic. In the transonic to supersonic flow regime, the turbine design can be tailored to yield a shock pattern the loss of which can be mitigated in near equal amount of that from the boundary layer with increasing exit Mach number, hence yielding a loss plateau in transonic-supersonic regime.