Trailing Edge Separation Research Articles

Receptivity-orientated drag reduction of twin cylinders by steady leading-edge suction control based on adjoint method

This study investigates the drag reduction of two tandem square cylinders under steady suction control at Reynolds numbers 50–200. The position where the suction flow should be placed is determined by using a receptivity analysis based on the adjoint method, and we investigate how control affects the fluid force and flow structures. High-order dynamic mode decomposition (HODMD) is applied to analyze the dynamic coherence modes and uncover the underlying control mechanism. The adjoint modes show that the regions of maximum receptivity to momentum forcing are localized on each side of the up-cylinder (UC) near the leading edge (LE). Thus, the suction flow is placed on the LE. The drag can be significantly reduced at wide gap distances, especially for the co-shedding regime. Under suction flow control, the separation is suppressed near the LE, and the gap vortices are no longer fed by the vorticity generated by the separated shear layer; they only result from the trailing-edge separation, which weakens and shrinks. Subsequently, the interaction between the gap flow and the down-cylinder (DC) is weakened, which reduces the drag and lift forces. The decrease in drag exceeds 66.4% for the UC and reaches 81.6% for the DC. The fluctuating reduction in the lift for the UC (DC) exceeds 59.0% (75.7%). HODMD results show that, as the suction flow velocity increases, the LE suction flow modifies the local time-averaged modes rather than the global mode energy. Conversely, the dynamic mode energy decreases significantly, whereas the mode shape remains unchanged except for a phase shift.

Influences of trailing-edge synthetic jets on longitudinal aerodynamic characteristics of a flying wing aircraft

For realizing the nice aerodynamic maneuverability of flying wing aircrafts (FWAs), a longitudinal aerodynamic control technology based on circulation control (CC) using trailing-edge synthetic jet actuators was proposed. Influences on longitudinal characteristics of a FWA were investigated. Results show that synthetic jets could improve the lift, drag, and nose-down moment, having potential of flight control at entire area of attack of angles (AOAs). Cl increment dips and then rises with the growth of AOAs, reaching the minimum at AOA of 12°. The maximum percentage of Cl enhancement and ΔCl/Cμ is separately 64.5% and 50.74%, respectively. Before 12°, synthetic jets could narrow the area of “dead zone,” improve flow velocities along the upper surface, and then move the trailing-edge separation point and the leading-edge (LE) stagnation point downward, enhancing the circulation. Moreover, synthetic jets grow rapidly through entrainment of the local flow, leading to the improvement of valid camber and, hence, the increase in Cl. CC efficiency decreases with the augmentation of AOAs, and leading-edge vortex (LEV) is weakened, causing the drop of Cl increment. After 12°, synthetic jets could enhance the longitudinal velocity of LEV and reduce the swirling number, improving the strength and stability of LEV, which results in larger suction near the leading edge. Moreover, the strengthened LEV could promote flow mixing and then weaken reverse pressure gradients along the wing section, thus improving flow velocities and CC efficiency at the wing section. It is above two factors that make Cl increment rise after 12°.

Large eddy simulations of a utility‐scale horizontal axis wind turbine including unsteady aerodynamics and fluid‐structure interaction modelling

AbstractGrowing horizontal axis wind turbines are increasingly exposed to significant sources of unsteadiness, such as tower shadowing, yawed or waked conditions and environmental effects. Due to increased dimensions, the use of steady tabulated airfoil coefficients to determine the airloads along long blades can be questioned in those numerical fluid models that do not have the sufficient resolution to solve explicitly and dynamically the flow close to the blade. Various models exist to describe unsteady aerodynamics (UA). However, they have been mainly implemented in engineering models, which lack the complete capability of describing the unsteady and multiscale nature of wind energy. To improve the description of the blades' aerodynamic response, a 2D unsteady aerodynamics model is used in this work to estimate the airloads of the actuator line model in our fluid–structure interaction (FSI) solver, based on 3D large eddy simulation. At each section along the actuator lines, a semi‐empirical Beddoes‐Leishman model includes the effects of noncirculatory terms, unsteady trailing edge separation, and dynamic stall in the dynamic evaluation of the airfoils' aerodynamic coefficients. The aeroelastic response of a utility‐scale wind turbine under uniform, laminar and turbulent, sheared inflows is examined with one‐ and two‐way FSI coupling between the blades' structural dynamics and local airloads, with and without the enhanced aerodynamics' description. The results show that the external half of the blade is dominated by aeroelastic effects, whereas the internal one is dominated by significant UA phenomena, which was possible to represent only thanks to the additional model implemented.

Low-Order Modeling of Dynamic Stall on Airfoils in Incompressible Flow

Airfoil dynamic stall in incompressible flow is characterized by two interacting viscous flow phenomena: time-varying trailing-edge separation and the shedding of intermittent leading-edge-vortex structures. In the current work, a physics based low-order method capable of modeling the interactions between the two flow phenomena is developed with the aim of predicting dynamic stall with only a few empirical tuning parameters. Large computational datasets are used to understand the flow physics of unsteady airfoils so as to augment an inviscid, unsteady airfoil theory to model the time-dependent viscous effects. The resulting model requires only three empirical coefficients for a given airfoil and Reynolds number, which could be obtained from a single moderate-pitch-rate unsteady motion for that airfoil/Reynolds number combination. Results from the low-order model are shown to compare excellently with computational and experimental solutions, in terms of both aerodynamic loads and flow-pattern predictions. In addition to formulating a method with limited empirical dependencies, the current research provides valuable insights into the flow physics of unsteady airfoils and their connection to rapidly predictable theoretical parameters.

Steady and unsteady aerodynamic loading of a NACA 16-616 aerofoil in a uniform flow

AbstractThis paper investigates the hydrodynamic near-field of a NACA 16-616 aerofoil over a range of angles-of-attack, encompassing the pre-stall, stall and post-stall flow regimes. In both the static pressure and the pressure fluctuation results, it is shown that each flow regime is easily distinguished, and it is further shown that each regime has different spectral behaviour and boundary layer characteristics. It is found that the NACA 16-616 aerofoil stalls by an abrupt leading-edge mechanism, characterised by a sudden change in the static pressure and unsteady surface pressure spectra between $16^\circ $ and $17^\circ $ angles-of-attack, but of more interest is that there is a secondary yet significant trailing-edge flow separation mechanism occurring upstream of the trailing-edge and moving further upstream as the angle-of-attack increases in the pre-stall regime. A comparison is made between the spectra and coherence of the unsteady surface pressure of the NACA 16-616 aerofoil and those of the classic NACA 0012 aerofoil and shows that such a secondary mechanism has a significant impact for large pre-stall angles-of-attack on the unsteady surface pressure. This will have a significant impact on the radiated far-field sound, distinguishing the NACA 16-616 aerofoil from aerofoils that do not have this secondary mechanism. The existence and extent of this secondary trailing-edge separation mechanism is further shown by the hot-wire anemometry boundary layer velocity results that indicate separation within the pre-stall regime.

Energy extraction in the dynamic modes of flow for airfoil's laminar separation flutter

This paper aims to gain new insight into the physical mechanism of laminar separation flutter (LSF) from the perspective of energy transfer and dynamic mode decomposition (DMD) modes of flow. An online DMD method accounting for the airfoil's pitch motion is developed, and the relationship between the topology of energy map and DMD modes is established. Simulation results indicate that there are two limit cycle branches in energy map, but only one branch is stable. The LSF time response state can be predicted accurately by the stable limit cycle branch. The topology of an energy map is dominated by the DMD mode corresponding to the airfoil's pitch frequency. The developed DMD method can extract the variation of flow structures effectively. The pressure distribution of DMD mode corresponding to the pitch frequency is dominated by the leading-edge suction and bubble's suction. The bubble's suction is induced by the trailing-edge laminar separation bubble or laminar separation bubble (LSB). When the pitch amplitude is larger than 4°, the trailing-edge laminar separation bubble transforms to LSB. The inherent mechanism is that increasing the trailing-edge separation bubble's intensity promotes the energy extraction while the occurrence of LSB mitigates it.

Aerodynamic study of low Reynolds number airfoil and mini-unmanned aerial vehicle in simulated rain environment

PurposeRainfall is one of the main atmospheric conditions that significantly affect the aerodynamic performance of the low Reynolds number flights. In this paper, the adverse effects of rain on the aerodynamic performance of a two-dimensional (2D) airfoil with a chord-based low Reynolds number of 2 × 105 and the mini-unmanned aerial vehicle (UAV) for various flight conditions, i.e. 0°–40° at Mach number 0.04 were studied numerically. The purpose of this study is to explore the aerodynamic penalties that affect the liquid water content (LWC = 5.33) of the airfoil and UAV performance in rain under different flying conditions.Design/methodology/approachThe Eulerian–Lagrangian two-phase flow method is adopted to simulate the rain environment over an airfoil and mini-UAV aerodynamic performances. The Reynolds Averaged Navier–Stokes equations are considered to solve the time-averaged equations of motion for fluid flow.FindingsThe effect of rainfall on the airfoil and mini-UAV is studied numerically and validated experimentally. For 2D airfoil, the lift and drag coefficients for both numerical and experimental results show a very good correlation at Reynolds number 2 × 105. For three-dimensional (3D) mini-UAV, the lift and drag coefficients for both numerical and experimental results show a very good correlation at Mach number 0.04. The raindrops distribution around the airfoil, premature trailing edge separation, boundary-layer velocity profiles at five different chord positions (i.e. LE, 0.25c, 0.5c, 0.75c and 0.98c) on the upper surface of the airfoil, water film height and the location of rivulet formation on the upper surface of the airfoil are also presented.Originality/valueFor 2D airfoil, the recorded maximum variation of the coefficient of lift and lift-to-drag (L/D) ratio is observed to be 5.33% at an 8° and 10.53% at a 4° angle of attack (AOA) between numerical and experimental results under the influence of rainfall effect for LWC = 5.33. The L/D ratio percentage degradation is seen to be 61.9% at an AOA of 0°–2° for the rain environment. For 3D mini-UAV, the recorded maximum variation of the coefficient of lift and L/D ratio are observed to be 2.84% and 4.60% at a 30° stall AOA under the influence of rainfall effect for LWC = 5.33. The numerical results are impressively in agreement with the experimental results. UAV designers will benefit from the findings presented in this paper. This will be also helpful for training the pilots to control the airplanes in a rain environment.

Low Reynolds Number Effects on the Separation and Wake of a Compressor Blade

Abstract This paper investigates the surface boundary layer and wake development of a compressor blade at a range of low Reynolds number from 45,000 to 120,000. Experiments in a miniature linear compressor cascade facility have been performed with detailed surface pressure measurements and flow visualization to track variations in the separation bubble size. These have been combined with high-resolution pneumatic pressure and hot-wire probe traverses in the downstream wake. High-fidelity direct numerical simulations have been completed on the same compressor blade section across the same range of operating conditions. The results show that large laminar separation bubbles exist on both blade surfaces. As Reynolds number increases, these separation bubbles shorten in length and reduce in thickness. Correspondingly, the downstream wake narrows, although the peak wake loss coefficient remains approximately constant. As the Reynolds number is increased from 45,000 to 120,000, the bubble length on the suction side reduced from 48% to 28% chord and on the pressure side reduced from 35% to 20% chord, while the loss coefficient reduced from 9% to 5%. The flow features are examined further within the high-fidelity computations, which reveal the dependence of the wake turbulence on the laminar separation bubbles. The separation bubbles are found to generate turbulent kinetic energy, which convects downstream to form the outer part of wake. As Re increases, a shorter bubble produces less turbulence in the outer part of the boundary layer leading to a narrower wake. However, the trailing edge separation is largely independent of Reynolds number, leading to the constant peak loss coefficient observed. The overall loss is shown to vary linearly with the total turbulence production, and this depends on the size of the separation bubbles. Overall, this research provides new insight into the connection between the blade surface flow field and the wake characteristics at low Reynolds number. The findings suggest that changes that minimize the extent of the blade separation bubbles could provide significant improvements to both the steady and unsteady properties of the wake.

Effects of pulsed endwall air injection on corner separation and vortical flow of a compressor cascade

ABSTRACT Air injection is an effective method to eliminate flow separation and improve blade loading in compressors. In existing studies on unsteady air injection, achieving a better control effect than steady injection with lower injection consumption is a major difficulty and challenge. In this work, pulsed endwall air injection (PEAI) was carried out numerically in a highly loaded compressor cascade under different injection frequencies and injection amplitudes. The mechanism through which PEAI influences corner separation and vortical flow, especially the effect of initial injection phase on the flow field of compressor cascade, was revealed. The results showed that PEAI effectively suppressed the corner separation with lower injection mass flow rate compared with steady endwall air injection (SEAI). The time-averaged overall loss coefficient and endwall loss coefficient were reduced by 10.5% and 28.7%, respectively, under the optimal PEAI scheme. Higher injection amplitude exhibited a more effectively controlled range in the suction corner, which led to a decrease in corner separation and deterioration of the mid-span flow field. The interaction between trailing edge separation and low-momentum fluid at mid-span led to the formation of an injection vortex. When conducting PEAI, the effect of inertia was observed for flow in the injection hole, which was derived from the interactions between injection fluid at different velocities and will increase with the improvement of injection frequency. The effects of injection location and diameter of the injection hole were also investigated. The injection location of PEAI affected the flow field by changing the effect region of the injection and low-momentum fluid near the endwall. The injection hole diameter had a significant impact on the intensity of the passage vortex and mid-span flow field.

Insight into the effects of leading edge delamination on the aerodynamic performance of an airfoil and wind turbine

Severe erosion or delamination at the leading edge of blades has an adverse influence on the power loss of wind turbines. Obtaining relevant quantitative data and its mechanisms would help in the efficient management and maintenance of these turbines. In the present work, detailed analysis on the effects of various levels of leading edge delamination on the aerodynamics and flow characteristics of an S809 airfoil and, in turn, on the performance of the NREL phase VI rotor is conducted. The results indicated that an elongated leading edge separation bubble and two enclosed vortex systems appeared in the delaminated region of the airfoil when the delamination levels aggravate to a certain degree, while the trailing edge separation vortex can be observed for all delaminated airfoils. The aerodynamic performance of the wind turbine decreases with increasing delamination depth and increases with growing delamination depth. The most influential delamination length is 1%c, and the maximum power loss is 23.09% for the blade with the severe delamination case. A small area of deep leading edge delamination or defects may cause great power loss, which should not be ignored in practice engineering.

Performance of Wall-Modeled LES with Boundary-Layer-Conforming Grids for External Aerodynamics

We investigate the error scaling and computational cost of wall-modeled large-eddy simulation (WMLES) for external aerodynamic applications. The NASA Juncture Flow is used as representative of an aircraft with trailing-edge smooth-body separation. Two gridding strategies are examined: 1) constant-size grid, in which the near-wall grid size has a constant value and 2) boundary-layer-conforming grid (BL-conforming grid), in which the grid size varies to accommodate the growth of the boundary-layer thickness. Our results are accompanied by a theoretical analysis of the cost and expected error scaling for the mean pressure coefficient and mean velocity profiles. The prediction of is within less than 5% error for all the grids studied, even when the boundary layers are marginally resolved. The high accuracy in the prediction of is attributed to the outer-layer nature of the mean pressure in attached flows. The errors in the predicted mean velocity profiles exhibit a large variability depending on the location considered, namely, fuselage, wing-body juncture, or separated trailing edge. WMLES performs as expected in regions where the flow resembles a zero-pressure-gradient turbulent boundary layer such as the fuselage ( error). However, there is a decline in accuracy of WMLES predictions of mean velocities in the vicinity of wing-body junctions and, more acutely, in separated zones. The impact of the propagation of errors from the underresolved wing leading edge is also investigated. It is shown that BL-conforming grids enable a higher accuracy in wing-body junctions and separated regions due to the more effective distribution of grid points, which in turn diminishes the streamwise propagation of errors.

A leading-edge vortex initiation criteria for large amplitude foil oscillations using a discrete vortex model

A leading-edge vortex initiation criterion of an oscillating airfoil is shown to provide a means to predict the onset of leading-edge separation. This result is found to collapse the occurrence of separation during the oscillation cycle when scaled using the leading-edge shear velocity determined from the foil oscillating kinematic motion. This is of importance in developing low order models for predicting energy harvesting performance, as is shown in this study using an inviscid discrete vortex model (DVM). Results are obtained for a thin flat airfoil undergoing sinusoidal heaving and pitching motions with reduced frequencies of k=fc/U∞ in the range of 0.06–0.16, where f is the heaving frequency of the foil, c is the chord length, and U∞ is the freestream velocity. The airfoil pitches about the mid-chord, and the heaving and pitching amplitudes of the airfoil are ho=0.5c and θ0=70°, respectively, illustrating results for conditions near peak efficiency for energy harvesting. The DVM uses a panel method, applicable to a wide range of foil geometries. An empirical trailing-edge separation correction is also applied to the transient force results. The vortex shedding criterion is based on the transient local wall stress distribution determined using a computational fluid dynamics (CFD) simulation, indicating the time and location of zero stress at the foil surface. In addition, the local pressure gradient minimum is also used as a local indicator. The effects of a wide range of Reynolds numbers on separation are shown for the given range of reduced frequencies. The use of the effective angle of attack, when modified to include the pitching component, is also shown to correlate the leading-edge vortex initiation time. The advantage of the proposed separation criteria is that it can be fully determined from the motion kinematics and then applied to a wide range of low order models. Model results are given for the transient lift force and compare well with the CFD simulations. It is noted that at higher reduced frequencies the DVM overpredicts portions of the transient loads possibly caused by the reversed viscous flow from the trailing edge.

Combined flow control with full-span slot and end-wall boundary layer suction in a large-camber compressor cascade

To investigate combined flow control strategies with full-span slot and end-wall (EW) boundary layer suction, a large-camber compressor cascade was treated with a full-span slot configuration firstly. Then, two combined flow control schemes with different EW boundary layer suction slots were set up based on the full-span slot configuration. Finally, the performance of the baseline cascade and the slotted cascades was evaluated and compared numerically. The results show that the full-span slotted scheme can eliminate the trailing edge (TE) separation near the blade midspan and weakens the corner separation. However, the EW secondary flow still develops on the blade suction surface (SS) before the full-span slot outlet and separates to form the wall vortex (WV). In contrast, the two combined flow control schemes can effectively inhibit the development of the EW secondary flow on the blade SS, contributing to the smaller total pressure loss and stronger static pressure recovery capacity. Between them, the combined flow control scheme with full-chord-length EW suction slot behaves better. It can effectively eliminate the concentrated shedding vortex, WV and corner vortex in the cascade passage, and the passage vortex is also weakened to some extent. Therefore, the TE separation and the corner separation are basically eliminated. The total pressure loss, flow turning angle and static pressure coefficient can be increased by an average of −38.4%, 3.1° and 16.2% in the combined flow control cascade with full-chord-length EW suction slot respectively, which is beneficial for improving the stability margin and static pressure recovery capacity of compressor blades.

Passive Flow Control around NACA 0018 Airfoil Using Riblet at Low Reynolds Number

In this study, aerodynamic capabilities of NACA 0018 airfoil is numerically investigated by installing riblet on the suction side of airfoil. Numerical results were obtained by ANSYS Fluent using k-kl-kw transition model at Reynolds number of Re=100 000. Three different riblet airfoil configuration was performed at six different angles of attack (α=8°, 10°, 13°, 15°, 17° and 19°) and these results compared with the clean model. For M1 model the riblet was located at chord wise section of x/c=0.3 while it installed at x/c=0.7 for M2 model. For M3 model two riblets were used and they were located at both x/c=0.3 and x/c=0.7. Obtained numerical result show that the use of riblet remarkably affects the flow characteristics of airfoil. At α=8° the CL/CD value of M1 model is increased by 4.5% when compared to clean model. It is indicated that angle of attack at α=10o, lift coefficient is increased for all models with compared to clean model. Stall angle is delayed from α=13° to α=15° at M1 and M3 with compared to clean model and lift coefficient is increased about 37% because of the restriction of the laminar separation bubble and trailing edge separation.

Corner Separation Control Using a New Combined Slotted Configuration in a High-Turning Compressor Cascade under Different Solidities

In order to comprehensively control the corner separation and the blade trailing edge (TE) separation in a high-turning compressor stator cascade, this research proposes a new combined slotted configuration consisting of one full-span slot and two blade-end slots. Taking into account the effect of the blade solidity, the performance of the original cascade and the combined slotted cascade was calculated and evaluated in a wide incidence angle range at two blade solidities. The results indicated that the blade loading and the corner separation range of the original cascade becomes larger as the blade solidity decreases from 1.66 to 1.36, which leads to higher total pressure loss and lower pressure diffusing capacity under positive incidence angles. The low-momentum fluid in the boundary layer can be significantly re-energized by the high-momentum blade-end and full-span slots jets, hence the combined slotted configuration can eliminate the blade TE separation and reduce the corner separation remarkably in the full incidence angle range at the two blade solidities. By adopting the combined slotted configuration, the total pressure loss, turning angle and static pressure coefficient of the original cascade can be increased by −23.2%, 2.7° and 4.7% on average, respectively, when the blade solidity is 1.66, while they can be increased by −27.7%, 3.3° and 7.6% on average, respectively, when the blade solidity is 1.36. The combined slotted configuration has a significant adaptability to the low blade solidity (or high loading) condition and it shows a certain potential in increasing the aeroengine thrust-to-weight ratio by decreasing the compressor single-stage blade number.

Experimental study on the effect of bionic flap parameters on airfoil aerodynamic performance

The effects of bionic flap on airfoil performance were experimentally studied to provide theoretical support for the application of the bionic flap in aeronautical engineering. Seven kinds of bionic flaps were used to study the effects of the key flap parameters, including the flap angle, length, shape, and position, at a Reynolds number of Re = 0.8 × 106. At small angle of attacks (AoAs), the drag and pitching moment increased and the lift reduced when using the bionic flap. While at high AoAs, the lift increased and the drag reduced, which improved the airfoil stall characteristics. The configuration of deflection bionic flap had the smallest initial AoA for improving the airfoil stall characteristics in the seven kinds of bionic flaps. More than eight degrees of the effective AoA range for improving lift characteristics could be achieved. The maximum lift coefficient could be increased by 3.9%. Additionally, the control mechanisms of the flap under different flow conditions (attached flow and separated flow) were deeply studied. In the attached flow, the effective camber and thickness of the basic airfoil could be changed by the flap, resulting that the flow around the airfoil was affected, which in turn affected the Cl and the slope of the lift line. In the separated flow, the flap affected the flow around the airfoil by controlling the development of the trailing edge separation vortex. These research results confirmed the aerodynamic mechanisms for the formation of double layered feathers when birds land, and provided insight into application of bionic flaps in aeronautical engineering.

Full-Span Topology of Trailing-Edge Separation at Different Angles of Attack

The velocity field over a two-dimensional wing at Reynolds number of 672,000 was investigated to characterize the three-dimensional topology and evolution of the separate flow with variation of the angle of attack. Planar particle image velocimetry over the full span of the wing demonstrated that with increasing angle of attack, isolated pockets of backflow, which appeared near the trailing edge, merged and formed an asymmetric stall cell at angle of attack of 9.7°. The asymmetry was mainly associated with the dissimilar boundary layers developed on the wind tunnel walls at the spanwise ends of the wing. Secondary structures were also observed between the stall cell and the spanwise end. The stall cell topology was characterized using large-scale three-dimensional particle tracking velocimetry measurements using helium-filled soap bubbles. The results showed that the separation bubble had a small wall-normal height with two wall-normal counter-rotating vortices extended up to the edge of the separation bubble. In addition, the investigations demonstrated that vortex generators can induce a symmetric stall cell by removing the secondary structures and isolating the stall cell from the flows at the spanwise ends.

Unsteady Response of Natural Laminar Flow Airfoil Undergoing Small-Amplitude Pitch Oscillations

Large-eddy simulations are performed to investigate the dynamic response of a natural laminar flow airfoil undergoing harmonic small-amplitude pitch oscillations at a chord based Reynolds number of . Large changes in the transition location as well as trailing-edge separation are observed throughout the pitch cycles, which leads to a nonlinear response of the aerodynamic forces. Despite the highly nonlinear nature of the flow, the evolution of the boundary layer over the airfoil can be modeled by using a simple phase-lag concept, which suggests a quasi-steady evolution of the boundary layer. A simple empirical model is developed based on this phase-lag assumption, which fits very well with the measured experimental data and identifies the primary source of non-linearities in the unsteady aerodynamic forces.

Microjet Configuration Sensitivities for Active Flow Control on Multi-Element High-Lift Systems

Various characteristics of small surface-normal jets (microjets) located on a flap’s pressure-side near the trailing edge are investigated as an active aerodynamic load control technology for multi-element high-lift systems. Two-dimensional computational studies are carried out to investigate the sensitivity of microjet aerodynamic effectiveness to configuration. Initially, the effects of microjet location and width are explored. The study shows that lift enhancement decreases as the microjet is moved forward from the trailing edge, and that wider microjets are more effective in lift enhancement and drag reduction, whereas thinner microjets are more efficient. Microjet exit velocity profile effect on the integrated forces is also studied, and the effects are shown to be insignificant. To gain preliminary insight of the microjet system requirements, several plenum configurations are studied, and a preliminary analysis of the power requirements to supply the air mass flow is presented. Next, the study is extended to a geometry with moderate trailing edge separation to demonstrate microjet’s separation mitigation capability. To gain better understanding of the system requirements, pulsed versus continuous microjets are examined, and it is found that pulsed blowing causes a breakdown of the favorable effect of the microjet on the trailing edge flow and therefore reduces the circulation control effectiveness of the microjet.

Effect of optimal slot location on the performance of cascades of blades

Overloaded blades and lower solidities are two main characteristics of modern gas engines to reach lower specific fuel consumption and weight to thrust ratio. Hence, flow optimization methods are employed to maximize the performance. Slot treatment on blades is considered as a potential method to stand against the boundary layer separation for this issue. In the present study, the different locations of passive flow control with the slot along the compressor blade chord are simulated at four angles of attack (α), ranging from 10.93° to 28.93°. The purpose of this study is to discover the underlying reason in obtaining the best slot location, regarding the blade turning angle and total loss coefficient, in every tested inflow angle. Based on the results, the slot jet shows its utmost performance in eliminating the trailing edge separation when it is located upstream of the separation lines in the mid-span. The best slot location is able to improve the turning angle and the total loss coefficient up to 16.41% and 16.47% at 26.93°, respectively. The synthetic slot jet confines the growth of corner vortices and increases the turning angle near end walls up to 4°. In addition, the slot treatment pushes the backflow region over the suction surface up to 10% of the span toward the side walls. The importance of choosing the slot location decreases with an increase in inflow angle. However, the effect of slot locations on the performance in low angles is still considerable. With the tested blade geometry, the slotted profile in locations between 0.5 and 0.8 of axial chord length indicates a promising blade performance either in low or high angles. The slot treatment weakens the flow upstream of its location which is in contradiction with previous investigations. Two separation bubbles near the leading edge in the baseline profile have a greater size of about 2% of the blade chord for the slotted case.