Trailing Edge Thickness Research Articles

Numerical study of improving aerodynamic performance of low solidity LPT cascade through increasing trailing edge thickness

This paper presents a new idea to reduce the solidity of low-pressure turbine (LPT) blade cascades, while remain the structural integrity of LPT blade. Aerodynamic performance of a low solidity LPT cascade was improved by increasing blade trailing edge thickness (TET). The solidity of the LPT cascade blade can be reduced by about 12.5% through increasing the TET of the blade without a significant drop in energy efficiency. For the low solidity LPT cascade, increasing the TET can decrease energy loss by 23.30% and increase the flow turning angle by 1.86% for Reynolds number (Re) of 25,000 and freestream turbulence intensities (FSTI) of 2.35%. The flow control mechanism governing behavior around the trailing edge of an LPT cascade is also presented. The results show that appropriate TET is important for the optimal design of high-lift load LPT blade cascades.

Experimental Study of Drag-Reduction Devices on a Flatback Airfoil

Various trailing-edge drag-reduction devices, including a new flap device, were examined experimentally on a flatback airfoil in a wind tunnel. The tests concerned a 30% thick airfoil with 10.6% thick trailing edge. Pressure, hot wire, and stereo particle image velocimetry measurements were performed at a chord Reynolds number of . Results show that the best-performing devices decrease drag, increase the vortex shedding frequency, and reduce flow variation downstream of the wing trailing edge. The best-performing device was a combination of the flap with an offset cavity plate. Further investigation is required for the optimization of the new device to examine its effects on noise reduction, load mitigation, and control.

Investigation of wave phenomena on a blunt airfoil with straight and serrated trailing edges

An investigation of pressure waves in compressible subsonic and transonic flow around a generic airfoil is performed in a modified shock tube. New comprehensive results are presented on pressure waves in compressible flow. For the first time, the influence of trailing edge serration will be examined in terms of the reduction in pressure wave amplitude. A generic airfoil is tested in two main configurations, one with blunt trailing edges and the other one with serrated trailing edges in a Mach number range from 0.6 to 0.8 and at chord Reynolds numbers of 1 × 106 < Re c < 5 ×106. The flow of the blunt trailing edge is characterized by a regular vortex street in the wake creating a regular pattern of upstream-moving pressure waves along the airfoil. The observed pressure waves lead to strong pressure fluctuations within the local flow field. A reduction in the trailing edge thickness leads to a proportional increase in the frequency of the vortex street in the wake as well as the frequency of the waves deduced from constant Strouhal number. By serrating the trailing edge, the formation of vortices in the wake is disturbed. Therefore, also the upstream-moving waves are influenced and reduced in their strength resulting in a steadier flow. An increasing length of the saw tooth enhances the three dimensionality of the structures in the wake and causes a strong decrease in the wave amplitude.

Numerical simulation of the effect of relative thickness on aerodynamic performance improvement of asymmetrical blunt trailing-edge modification

In this paper, the aerodynamic performance of wind turbine airfoils with different relative thicknesses and their modifications has been numerically investigated to facilitate a greater understanding of the effects of maximum relative thickness and its position on the aerodynamic performance improvement of asymmetrical blunt trailing-edge modification. The lift and drag coefficients of airfoil NACA4415 are calculated with the k-ω SST turbulence model, and are compared with experimental data to validate the simulation accuracy of the Computational Fluid Dynamics (CFD) approach. The airfoils with different relative thicknesses are modified to be asymmetrical blunt trailing-edge airfoils by means of the software Xfoil. The best trailing-edge thickness distribution ratio is obtained by comparing the aerodynamic performance of the modifications with different distribution ratios. The aerodynamic performance of original airfoils and their asymmetrical modifications with the best thickness distribution ratio being 1:3 is investigated to analyze the increments of lift and drag coefficients and lift-drag ratio. Results indicate that with the increasing of relative thickness, the lift coefficient increment of NACA4418 airfoil is the smallest for the angle of attack more than 9°, and the drag coefficient increment as a whole decreases first and then increases, but the average lift-drag ratio increment of NACA4412 airfoil is the largest, closely followed by NACA4415 airfoil. It is also showed that with the relative thickness position close to the leading-edge, the increments of lift and drag coefficients decrease and increase for the angle of attack more than a certain value, respectively, and the average lift-drag ratio increment of NACA4415 airfoil is positive and larger than those of NACA4415-mod25 and NACA4415-mod20 airfoils. Therefore, the medium thickness airfoil whose relative thickness position is away from the leading-edge is more suited to the asymmetrical blunt trailing-edge modification.

Distributed forcing flow control in the wake of a blunt trailing edge profiled body using plasma actuators

A modern flow control technique for reducing the drag associated with the periodic shedding of von Kármán vortices in the wake of a blunt trailing edge profiled body is presented. The technique involves distributed forcing of the wake flow using an array of dielectric barrier discharge plasma actuators, with a spanwise spacing matched to the spanwise wavelength of the dominant secondary wake instability. The experiments include measurement of the velocity field in multiple vertical and horizontal planes in the wake using particle image velocimetry, as well as base pressure, at Reynolds numbers of 2000, 3000, and 5000 based on trailing edge thickness. The flow control technique causes elongation of the vortex formation region across the span, and significant reduction of the fluctuating and total drag forces, up to a maximum of 94% and 18%, respectively. The effectiveness of the flow control technique is shown to be dependent on the induced momentum coefficient. Proper orthogonal decomposition analysis is used to investigate the mechanism of interaction of the flow control technique with the wake flow. Two distinct flow regimes are observed depending on the induced momentum coefficient. The effect of the control on the wake flow structure in the first regime is similar to those observed in previous studies involving mild spanwise-periodic geometric perturbations at the trailing edge, where control leads to streamwise displacement of the vortices and a shift in shedding frequency. However, an incremental increase in the momentum coefficient leads to a second flow regime similar to those previously observed in the case of large-amplitude geometric perturbations, with an almost complete attenuation of vortex shedding in the near-wake region.

익형 뒷전 형상이 날개 공력 특성에 미치는 영향

In the paper, a study on the analysis of the effects of trailing edge thickness on the aerodynamic characteristics of an airfoil is described. In this research, modification of the formula representing NACA symmetric airfoil is studied to change the airfoil shape with different trailing edge thickness of user’s choice. According to the result of aerodynamic characteristics, as the trailing edge thickness increases the maximum lift coefficient increases while the lift-to-drag ratio decreases. In this paper flow calculation results are demonstrated and the analysis on those results and findings on the effects of non-zero thickness of trailing edge are suggested.

Analysis of novel low specific speed pump designs

Centrifugal pumps with very low specific speed present significant design challenges. Narrow blade channels, large surface area of hub and shroud discs relative to the blade area, and the presence of significant of blade channel vortices are typical features linked with the difficulty to achieve head and efficiency requirements for such designs. This paper presents an investigation of two novel designs of very low specific speed impellers: impeller having blades with very thick trailing edges and impeller with thick trailing edges and recirculating channels, which are bored along the impeller circumference. Numerical simulations and experimental measurements were used to study the flow dynamics of those new designs. It was shown that thick trailing edges suppress local eddies in the blade channels and decrease energy dissipation due to excessive swirling. Furthermore the recirculating channels will increase the circumferential velocity component on impeller outlet thus increasing the specific energy, albeit adversely affecting the hydraulic efficiency. Analysis of the energy dissipation in the volute showed that the number of the recirculating channels, their geometry and location, all have significant impact on the magnitude of dissipated energy and its distribution which in turn influences the shape of the head curve and the stability of the pump operation. Energy dissipation within whole pump interior (blade channels, volute, rotor- stator gaps) was also studied.

The Effects of Trailing Edge Thickness on the Losses of Ultrahigh Lift Low Pressure Turbine Blades

Experimental, numerical and analytical methods were used to investigate the effects of the blade trailing edge thickness on the profile loss of ultrahigh-lift low-pressure turbine blades. Two cascades, the T106C and the T2, were studied. The loss obtained based on the data at the blade trailing edge plane and the plane 0.3 Chord downstream of the trailing edge agree with each other for T106C blade with and without upstream wakes at different Reynolds numbers. The blade profile losses were broken down as the suction surface boundary loss, the pressure side boundary loss and the mixing loss downstream of the trailing edge for six Reynolds numbers. Trailing edge thicknesses varying from 1.4% to 4.7% pitch were investigated at a Reynolds number of 210,000. It was found that the flow distributions across the passage at the trailing edge planes were highly nonuniform. In particular, and as a result, the trailing edge base pressure was higher than the mixed-out static pressure, so the contribution of the base pressure to the mixing loss downstream of the trailing edge plane was to reduce the loss. When the trailing edge thickness increases, there are three main effects: (1) the area with high base pressure region increases, which tends to reduce the downstream mixing loss; (2) the base pressure reduces, which tends to increase the loss; and (3) the flow diffusion downstream of the trailing edge, which tends to increase the loss. The overall result is the combined effect. For the T106C cascade, increasing the trailing edge thickness from 1.9% pitch to 2.8% pitch has a small effect on the loss. Further increasing the trailing edge thickness increases the loss. The T2 blade has a higher lift than the T106C blade, so the effects of the base pressure in reducing the mixing loss downstream of the trailing edge is more evident. The experimental results show that the profile loss first decreases and then increases as the trailing edge thickness increases. CFD, using the transition k-ω SST model and the k-ω SST model, provides good predictions of the aerodynamic performance. It was used to study the cases with trailing edge thicknesses of 1.4% pitch and 2.9% pitch. The profile loss is almost the same for these two trailing edge thickness. The results show that it is possible to use thicker blade trailing edges in low pressure turbines without aerodynamic penalty. This can lead to benefits in terms of mechanical integrity and manufacturing cost reductions.

火星探査飛行機用プロペラ開発のための翼型に関する研究

Aerodynamic performances of four kinds of airfoils are examined in flows with low-Reynolds number. The aim of the present study is to find the airfoil shape, which provides a large value of a maximum lift-drag ratio in order to develop a high efficiency propeller for the Mars exploration airplane. A triangle airfoil is a base contour of the test airfoils and effects of blunt trailing edges on airfoil characteristics are examined using numerical simulations. Furthermore, a variation of aerodynamic characteristics of one airfoil with a blunt trailing edge due to decreased thickness and increased camber is also examined. The numerical results are verified by experiments. As the trailing edge thickness is increased, the lift-drag ratio becomes larger and values of the lift and the drag forces become smaller at the angle of attack where the lift-drag ratio becomes maximum. By increasing the trailing edge thickness, the boundary layer separation is suppressed and the aerodynamic forces are reduced. The increased camber and decreased thickness yield a significant lift increase of the airfoil with the blunt trailing edge and a small amount of drag increase. As a result, the airfoil with the blunt trailing edge yields large value of the lift-drag ratio in the flows with low-Reynolds numbers.

Hydraulic losses in the spiral case of low specific speed pumps

This contribution is focused on analysis of pressure losses in spiral case of centrifugal pump with thick trailing edges and with recirculation channels. Recirculation channels have different geometrical configuration and influence the size of available specific energy as well as hydraulic efficiency. Subsequently, the contribution analyses the flow in spiral case itself with respect to its function and its filling with liquid. Studied phenomena affect the research of pumps with low specific speed, the stability of specific energy characteristic curves and also the configuration of recirculation channels.

Application of Circulation Controlled Blades for Vertical Axis Wind Turbines

The blades of a vertical axis wind turbine (VAWT) rotor see an inconsistent angle of attack through its rotation. Consequently, VAWT blades generally use symmetrical aerofoils with a lower lift-to-drag ratio than cambered aerofoils tailored to maximise horizontal axis wind turbine rotor performance. This paper considers the feasibility of circulation controlled (CC) VAWT blades, using a tangential air jet to provide lift and therefore power augmentation. However CC blade sections require a higher trailing-edge thickness than conventional sections giving rise to additional base drag. The choice of design parameters is a compromise between lift augmentation, additional base drag as well as the power required to pump the air jet. Although CC technology has been investigated for many years, particularly for aerospace applications, few researchers have considered VAWT applications. This paper considers the feasibility of the technology, using Computational Fluid Dynamics to evaluate a baseline CC aerofoil with different trailing-edge ellipse shapes. Lift and drag increments due to CC are considered within a momentum based turbine model to determine net power production. The study found that for modest momentum coefficients significant net power augmentation can be achieved with a relatively simple aerofoil geometry if blowing is controlled through the blades rotation.

Synthetic jet actuation strategies for momentumless trailing edge wake

Periodic addition of momentum employing a Trailing Edge Synthetic Jet Actuation (TESJA) to control the wake of a blunt trailing edge plate was investigated experimentally at Reynolds number Reh=U∞h/ν=7200, based on the trailing edge thickness, h. A pair of synthetic jets were symmetrically placed at the model base. Two-component PIV measurements were reported for natural wake and actuated wakes under different actuation strategies namely asymmetric actuation, symmetric synchronous dual-actuation, and out-of-phase dual-actuation. A pair of synthetic jets were symmetrically placed at the model base. Two positions of the TESJA on the model base (peripheral or central position) were also studied and discussed. The forcing was applied at a frequency of 0.75 the natural vortex shedding frequency, while a high momentum addition coefficient of up to 24% was applied to balance the momentum deficit and to accomplish a near momentumless state. The measurements revealed that the application of TESJA reduced the extent of the separation bubble and redistributed energy among the Reynolds stresses. A perfect lock-in of vortex shedding in the wake was achieved. The actuation efficiency depended on both the applied control strategy and the actuators' positions. Forcing was found to be most effective in achieving a near momentumless state when applied at base model extremities (peripheral configuration) to permit the interaction of synthetic jet with the neighbouring shear layers and a large entrainment. The above strategies and positions led generally to three actuated wake states: (1) vortex pair shedding vectoring for asymmetric forcing, (2) symmetric vortex shedding under in-phase synchronous forcing, and (3) alternate vortex shedding with increased vorticity and turbulence, combined with a large momentum addition when out-of-phase forcing was applied.

Performance prediction of NREL (National Renewable Energy Laboratory) Phase VI blade adopting blunt trailing edge airfoil

This study uses Computational Fluid Dynamics (CFD) to evaluate the performance of a blade with blunt airfoil which is adapted at the root. FLUENT software is used of numerical analysis, and ICEM-CFD is used for generating mesh representations. The experimental conditions of the blade used for this study were based on experimental data presented by the National Renewable Energy Laboratory. The models for numerical analysis are blades in which the blunt trailing-edge thickness was 1%, 5% and 10% of the chord. These blades are modifications of the NREL (National Renewable Energy Laboratory) Phase VI blade. For detailed data, the numerical analysis has been performed for wind speeds of 7, 10, 15, 20, and 25 m/s. Also, the comparison analysis of results of a baseline and modified blades has been performed.

Investigation of Trailing Edge Cooling Concepts in a High Pressure Turbine Cascade—Aerodynamic Experiments and Loss Analysis

As part of a European research project, the aerodynamic and thermodynamic performance of a high pressure turbine cascade with different trailing edge cooling configurations was investigated in the wind tunnel for linear cascades at DLR in Go¨ttingen. A transonic rotor profile with a relative thick trailing edge was chosen for the experiments. Three trailing edge cooling configurations were applied, first central trailing edge ejection, second a trailing edge shape with a pressure side cut-back and slot equipped with a diffuser rib array, and third pressure side film cooling through a row of cylindrical holes. For comparison, aerodynamic investigations on a reference cascade with solid blades (no cooling holes or slots) were performed. The experiments covered the subsonic, transonic and supersonic exit Mach number range of the cascade while varying cooling mass flow ratios up to 2 %. This paper analyzes the effect of coolant ejection on the airfoil losses. Emphasis was given on separating the different loss contributions due to shocks, pressure, and suction side boundary layer, trailing edge, and mixing of the coolant flow. Employed measurement techniques are schlieren visualization, blade surface pressure measurements, and traverses by pneumatic probes in the cascade exit flow field and around the trailing edge. The results show that central trailing edge ejection significantly reduces the mixing losses and therefore decreases the overall loss. Higher loss levels are obtained when applying the configurations with pressure side blowing. In particular, the cut-back geometry reveals strong mixing losses due to the low momentum coolant fluid, which is decelerated by the diffuser rib array inside the slot. The influence of coolant flow rate on the trailing edge loss is tremendous, too. Shock and boundary layer losses are major contributions to the overall loss but are less affected by the coolant. Finally a parameter variation changing the temperature ratio of coolant to main flow was performed, resulting in increasing losses with decreasing coolant temperature.

Influence of open trailing edge on laminar aerofoils at low Reynolds number

This article deals with the effect of open trailing edge on the aerodynamic characteristics of laminar aerofoils at low Reynolds numbers, the attention being focussed on the influence of such a trailing-edge imperfection on the aerodynamic efficiency. Wind tunnel tests have been performed at different Reynolds numbers and angles of attack, and global aerodynamic as well as pressure distributions were measured (in these tests two types of open trailing edges, either sharp or rounded were considered). From experimental results, a quantitative analysis of the influence of the trailing-edge thickness on the degradation of aerofoil aerodynamic performances has been obtained, which allows the establishment of a criterion for an acceptance limit for this kind of imperfection.

The Effects of a Tripped Turbulent Boundary Layer on Vortex Shedding from a Blunt Trailing Edge Hydrofoil

Experiments on vortex shedding from a blunt trailing edge symmetric hydrofoil operating at zero angle of attack in a uniform high speed flow, Reh=16.1·103-96.6·103, where the reference length h is the trailing edge thickness, are reported. The effects of a tripped turbulent boundary layer on the wake characteristics are analyzed and compared with the condition of a natural turbulent transition. The foil surface is hydraulically smooth and a fully effective boundary layer tripping at the leading edge is achieved with the help of a distributed roughness. The vortex shedding process is found to be strongly influenced by the boundary layer development: the tripped turbulent transition promotes the re-establishment of organized vortex shedding. In the context of the tripped transition and in comparison with the natural one, significant increases in the vortex span-wise organization, the vortex-induced hydrofoil vibration, the wake velocity fluctuations, and the vortex strength are revealed. Although the vortex shedding frequency is decreased, a modified Strouhal number based on the wake width at the end of the vortex formation region is constant and evidences the similarity of the wakes in terms of spatial distribution for the two considered boundary layer transition processes.

Aerofoil tones at moderate Reynolds number

AbstractIt is known experimentally that an aerofoil immersed in a uniform stream at a moderate Reynolds number emits tones. However, there have been major differences in the experimental observations in the past. Some experiments reported the observation of multiple tones, with strong evidence that these tones are most probably generated by a feedback loop. There is also an experiment reporting the observation of a single tone with no tonal jump or other features associated with feedback. In spite of the obvious differences in the experimental observations published in the literature, it is noted that all the dominant tone frequencies measured in all the investigations are in agreement with an empirically derived Paterson formula. The objective of the present study is to perform a direct numerical simulation (DNS) of the flow and acoustic phenomenon to investigate the tone generation mechanism. When comparing with experimental studies, numerical simulations appear to have two important advantages. The first is that there is no background wind tunnel noise in numerical simulation. This avoids the signal-to-noise ratio problem inherent in wind tunnel experiments. In other words, it is possible to study tones emitted by a truly isolated aerofoil computationally. The second advantage is that DNS produces a full set of space–time data, which can be very useful in determining the tone generation processes. The present effort concentrates on the tones emitted by three NACA0012 aerofoils with a slightly rounded trailing edge but with different trailing edge thickness at zero degree angle of attack. At zero degree angle of attack, in the Reynolds number range of$2\ensuremath{\times} 1{0}^{5} $to$5\ensuremath{\times} 1{0}^{5} $, the boundary layer flow is attached nearly all the way to the trailing edge of the aerofoil. Unlike an aerofoil at an angle of attack, there is no separation bubble, no open flow separation. All the flow separation features tend to increase the complexity of the tone generation processes. The present goal is limited to finding the basic tone generation mechanism in the simplest flow configuration. Our DNS results show that, for the flow configuration under study, the aerofoil emits only a single tone. This is true for all three aerofoils over the entire Reynolds number range of the present study. In the literature, it is known that Kelvin–Helmholtz instabilities of free shear layers generally have a much higher spatial growth rate than that of the Tollmien–Schlichting boundary layer instabilities. A near-wake non-parallel flow instability analysis is performed. It is found that the tone frequencies are the same as the most amplified Kelvin–Helmholtz instability at the location where the wake has a minimum half-width. This suggests that near-wake instability is the energy source of aerofoil tones. However, flow instabilities at low subsonic Mach numbers generally do not cause strong tones. An investigation of how near-wake instability generates tones is carried out using the space–time data provided by numerical simulations. Our observations indicate that the dominant tone generation process is the interaction of the oscillatory motion of the near wake, driven by flow instability, with the trailing edge of the aerofoil. Secondary mechanisms involving unsteady near-wake motion and the formation of discrete vortices in regions further downstream are also observed.

A Viscous/Inviscid Interactive Approach for the Prediction of Performance of Hydrofoils and Propellers with Nonzero Trailing Edge Thickness

A viscous/inviscid interactive (VII) approach is applied to predict the performance of hydrofoils and propellers with nonzero trailing edge thickness. The emphasis has been put on developing VII models for flow separation. The investigation starts from a two-dimensional (2D) hydrofoil. The current method uses an iterative scheme to find a nonlifting closing extension behind the trailing edge. Two kinds of schemes are applied for the iteration process:a non-lifting extension with 1 or 2 degrees of freedom, in fully wetted condition andan extension which is treated like a cavity surface, but with a nonconstant cavity pressure distribution. The results from these schemes are compared with those from a commercial RANS Solver (Fluent). Next, the current schemes using flap extensions are extended to three-dimensional (3D) propeller flows. The 3D models are developed so that all the span-wise strips of the propeller satisfy similar conditions to those used in 2D. A propeller with significant nonzero trailing edge thickness is analyzed, using several 3D models, and the results are compared with existing experimental data.

CFD Analysis of Blunt Trailing Edge Airfoils Obtained with Several Modification Methods

generar bordes de salida con espesor ha probado que es posible obtener buen rendimiento aerodinamico sin comprometer los requerimientos estructurales. Dos diferentes metodos son analizados: el metodo de recorte y el metodo de adicion de espesor. Simulaciones bidimensionales fueron realizadas usando Ansys CFX® para valores del numero de Reynolds de 3.2 millones. Se presenta un estudio basado en los metodos de Taguchi para distintos espesores de borde de salida es presentado. Los resultados obtenidos muestran que el mientras que se probo que en el metodo de recorte + de sustentacion debida a las alteraciones geometricas causadas. En general, el metodo de adicion de critico de ataque

Effects of Splitter Plates on Turbulent Quantities on a Boundary Layer Developing on a Flat Plate near the Trailing Edge

An experimental investigation was made on a turbulent boundary layer near the trailing edge on a long flat plate with a blunt trailing edge. The flow was controlled by an additional splitter plate attached to the trailing edge along the wake center line. The length of the splitter plate, l, was varied from 0.3 to 3 times the trailing edge thickness, h. Measurement of fluctuating velocity was made under the freestream zero pressure gradient. Almost independent of the installation of the splitter plate, the turbulent intensity and the Reynolds shear stress in the inner layer decreased when approaching the trailing edge of the flat plate (x/h=0), compared with that in the fully-developed turbulent boundary layer under the zero pressure gradient. The only exception was the distribution at x/h=0 near the wall in the case without the splitter plate (l/h=0), where the degree of the decrease in urms/Ue was relatively small, and vrms/Ue and -uv / Ue2 were larger than those of the zero pressure gradient. The effect of l/h on turbulent quantities was shown just upstream of the trailing edge upon close comparison, and was classified into two groups: l/hl/h≧1. The tendency of the turbulent energy production terms approaching x/h=0 coincided with the decrease in the streamwise turbulent intensity and in the Reynolds shear stress.