Lift Increment Research Articles

양력선 방법을 이용한 지면효과가 날개의 공력성능에 미치는 영향 분석

Weissinger의 방법을 기반으로 양력선 방법을 지면효과를 고려한 해석이 가능하도록 확장하였다. 이 방법을 이용하여 가로세로비 10의 타원형 날개와 인간동력항공기 날개에 대해 지면효과에 의한 공력성능 변화를 예측하였다. 날개가 지면에 가까워질수록 날개의 와류강도는 약간 증가하지만 내리흐름은 크게 감소한다. 인간동력항공기 날개의 경우에 날개가 지면에 가까울수록 양력은 증가하여 높이가 2m에서 양력계수는 지면효과를 받지 않을 때보다 5% 증가한다. 유도항력계수는 날개가 지면에서 스팬길이 높이에 있을 때에 지면효과가 없을 때보다 10% 감소하고 높이가 2m이면 55% 감소한다. The lifting-line method based on Weissinger's method is extended to be able to analyze the ground effect. The method is applied to predict the variation of aerodynamic performance due to ground effect for the elliptic wing with aspect ratio of 10 and the wing of human powered aircraft. While the vortex strength of the wing increases slightly, the downwash decreases significantly as the wing approaches to the ground. For the wing of human powered aircraft, the increment of lift at the height of 2m is 5% than the lift outside the influence of ground effect. The decrease of induced drag at the height of wing span is 10% and at the height of 2m is 55% than that out of ground effect.

Numerical Simulation of Laminar Flow over Slotted Airfoil

In this paper mainly focused effect of fixed slot on performance of NACA0012 airfoil. Using slot lower surface high pressurized air passes through slot to energized upper surface. Analysis has been done on NACA0012 of 1m cord length at 25C of air with 5m/s air velocity. First plain airfoil at different angle of attack has been analyzed to find out stall condition and Cl, Cd . After that same parameter and physics, only change in geometry of airfoil with 15%C leading edge slot. Flow separation is adverse effect for performance of airfoil. Due to flow separation adverse pressure gradient effect reverse flow is there, we cause reduction of lift coefficient and increment of drag coefficient so flow separation necessary to reduce. There are different techniques to reduce flow separation but effect of slot on airfoil has been studied. Keyword: Angle of attack (AOA), plain airfoil, slotted airfoil, SST model,

Aerodynamic Installation Effects of an Over-the-Wing Propeller on a High-Lift Configuration

Preliminary design studies indicate that a cruise-efficient short takeoff and landing aircraft has enhanced takeoff performance at competitive direct operating costs when using high-speed propellers in combination with internally blown flaps. The original tractor configuration is compared to an over-the-wing propeller, which allows for noise shielding. An additional geometry with partially embedded rotor similar to a channel wing is considered to increase the beneficial interaction. This paper shows the aerodynamic integration effects with a focus on climb performance and provides an assessment of the three aforementioned configurations for a simplified wing segment at takeoff conditions. Steady Reynolds-averaged Navier–Stokes simulations have been conducted using an actuator disk model and were evaluated based on the overall design. Interacting with the blown flap, the conventional tractor propeller induces large lift and drag increments due to the vectored sliptream. Although this effect is much smaller for an over-the-wing configuration, by halving the lift augmentation, the lift-to-drag ratio and the propulsive efficiency are considerably improved. Besides a moderate lift gain, the main advantage of a channel wing design is the location of the thrust vector close to the center of gravity resulting in a smaller nosedown pitching moment due to thrust. A disadvantage of over-the-wing propellers is the inhomogeneous inflow at higher velocity, which leads to oscillating blade loads and reduced efficiency.

The influence of simulated missile warhead fragment damage on the aerodynamic characteristics of two-dimensional wings

AbstractThe paper describes a method of representing damage on a wing due to multiple warhead fragments, and investigates two of the key variables: fragment impact density and hole diameter. The aerodynamic effects of the damage were quantified by wind-tunnel tests on a two-dimensional wing at a Reynolds number of 5 × 105. The wing was of hollow construction with leading and trailing-edge spars. In all of the cases tested, simulated fragment damage resulted in significant lift losses, drag increases and pitching moment changes. Increasing fragment density or hole size resulted in greater effects. To a first order approximation, both lift and drag increments at a given incidence were related to the percentage wing area removed. Surface flow visualisation showed that low fragment densities and small damage sizes resulted in a complex flow structure on the surface of the wing. This was made up of boundary-layer growth between the damage holes, attached wakes from the forward damage holes and separated surface flow over the rear of the wing. For these cases, individual hole patterns showed similar flow mechanisms to those seen for larger scale gunfire damage cases. Increased fragment density and hole size resulted in upper surface flow separation at the first row of holes. Behind this separation, the flow was attached and consisted of the combined wakes from the forward damage holes. Investigations into the influence of internal model structure indicated that trends in coefficient changes were similar for both hollow and solid wings. However, the magnitudes of the effects were found to be smaller for hollow wings than for solid wings.

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.

Design and Computational Studies on Plain Flaps

In this paper, Computational fluid Dynamics (CFD) is used to predict accurately the incremental lift due to use of passive lift -augmentation device 'Flap'. In a 2-D wing, lift at a given angle of attack can be increased by increasing camber. In aircraft design, values of Maximum Lift Coefficient (CL max ) are used to determine the take-off and landing distances, on the assumption of usage of some type of lift- enhancing device. Passive lift enhancement device, such as a trailing edge plain flap, is relevant to most of the aircraft designs except Short Take Off and Landing (STOL) However, accurate analytical solutions to complex non linear equations of aerodynamics are a far cry and one has to depend largely on support from experimental results. It is here that CFD analysis makes the task simpler. Computational study of a typical plain flap with NACA (National Advisory Committee for Aeronautics) 0012 airfoil is taken up to predict its influence on incrementing the lift during critical phases of flight at low speed. The flow analysis is done using CFD tool FLUENT 6.3 along with GAMBIT 2.2. Effect of Plain flaps on basic aerodynamic parameters such as CP, CL , CD are studied by analyzing the flow over an airfoil with plain flap. The results obtained from the analysis are compared with analytical solutions to establish the influence and relevance of flap as an effective high lift device.

Effect of Aspect Ratio on Gurney-Flap Performance

A low-speed wind-tunnel investigation has been undertaken to establish the effect of wing aspect ratio on Gurney-flap performance. Characterization is accomplished using a force balance and flow visualization. The Gurney-flap lift increment due to a shift in the zero-lift angle of attack was observed to scale with that of the lift-curve slope for different aspect ratios. As the aspect ratio reduced, a Gurney flap of greater height was required to maximize aerodynamic efficiency. The dependence of aerodynamic parameters (zero-lift angle of attack, minimum drag coefficient, and lift-curve slope) on the Gurney flap’s height-to-chord ratio was also examined.

Separation and Circulation Control on an Elliptical Airfoil by Steady Blowing

The application of blowing over a thick elliptical airfoil is considered for the purpose of identifying the hierarchy of additional parameters affecting the flow, such as slot width, its location, and orientation, in addition to the mass and momentum flow that were considered as being significant in the past. When a single jet emanates from a narrow slot at a prescribed location, the incremental lift coefficient scales mostly with the momentum coefficient (C_µ). For wider slots, the slot width has to be considered as an independent parameter even when the jet velocity is much higher than the freestream. The lift increment depends strongly on the slot location when the latter is near the natural separation location of the flow, but it is independent of Reynolds number. Deleterious effects of steady blowing on C_L were observed for low momentum inputs. For wider slots located upstream of the natural separation location, the drag is sensitive to Reynolds number and it is not solely affected by C_µ. Flowfield measurements are used to explain some of the observations made.

Aerodynamic performance of a wing with a deflected tip-mounted reverse half-delta wing

The impact of a tip-mounted 65°-sweep reverse half-delta wing (RHDW), set at different deflections, on the aerodynamic performance of a rectangular NACA 0012 wing was investigated experimentally at Re = 2.45 × 105. This study is a continuation of the work of Lee and Su (Exp Fluids 52(6):1593–1609, 2012) on the passive control of wing tip vortex by the use of a reverse half-delta wing. The present results show that for RHDW deflection with −5° ≤ δ ≤ +15°, the lift was found to increase nonlinearly with increasing δ compared to the baseline wing. The lift increment was accompanied by an increased total drag. For negative RHDW deflection with δ < −5°, the RHDW-induced lift decrement was, however, accompanied by an improved drag. The deflected RHDW also significantly modified and weakened the tip vortex, leading to a persistently lowered lift-induced drag, regardless of its deflection, compared to the baseline wing. Physical mechanisms responsible for the observed RHDW-induced phenomenon were also discussed.

Experimental investigation of the methods of energy supply to the external surface of models of flying vehicles

The results of investigation of the effects of external supply of energy to the wind tunnel models at the Mach number M=5 are presented. The energy supply has been modeled by a light gas (helium) supply to the models. The loads have been measured by wind-tunnel balance. It is obtained that at small angles of attack the lift increment under the external supply of energy occurs without an increase in drag and under some conditions is more efficient than the lift increase by means of higher engine thrust under the same energy supply to it. The similarity criterion (“influence coefficient”) is proposed to relate the wind tunnel results to full-scale conditions. Based on the fact that the application of hot hydrogen is the most reasonable for the external supply of energy under full-scale conditions, the aspects of accelerating its ignition and combustion at low temperatures are investigated. It is shown that addition into the hydrogen of its atoms and activated molecules by passing the hydrogen through a glow-discharge zone makes it possible to implement ignition at a temperature of ∼300 K. Without such a treatment, ignition occurs at a temperature of ∼900 K.

Experimental and Theoretical Investigation of Ground Effect at Low Reynolds Numbers

An experimental study is presented investigating the effect of ground proximity on an airfoil at a low Reynolds number. A S8036 airfoil was evaluated atRe 120; 000 and 180,000. The groundwas simulated using a thin splitter plate. Ground clearance was varied from 1 to 0.1 of the chord, measured from the airfoil’s quarter-chord. Force balance as well as surface pressure and ground plane pressures were recorded. Two semiempirical prediction methods were developed to estimate the two-dimensional ground effect. Results indicate that ground proximity increases the lift coefficientwith a peak occurring at a gap of 0.2 of the chord.Basedupon lift coefficient, ground effect may reduce drag, while for a given airfoil incidence, it increases the drag coefficient. Surface pressure evaluation at 0 deg incidence indicated that ground effect increases suction on both airfoil surfaces. The theoretical methods showed close accord with the experiment and predicted the location of the lift increment maxima. Theoretical decomposition suggests that ground effect has the net impact of increasing airfoil camber and reducing the effective freestream.

Wingtip vortex control via the use of a reverse half-delta wing

The effect of a 65° sweep reverse half-delta wing (RHDW), mounted at the squared tip of a rectangular NACA 0012 wing, on the tip vortex was investigated experimentally at Re = 2.45 × 105. The RHDW was found to produce a weaker tip vortex with a lower vorticity level and, more importantly, a reduced lift-induced drag compared to the baseline wing. In addition to the lift increment, the RHDW also produced a large separated wake flow and subsequently an increased profile drag. The reduction in lift-induced drag, however, outperformed the increase in profile drag and resulted in a virtually unchanged total drag in comparison with the baseline wing. Physical mechanisms responsible for the RHDW-induced appealing aerodynamics and vortex flow modifications were discussed.

Application of Pulsed Blowing Technique in High-Lift Control Surface Design

Because the flight performance of aircraft is so dependent on aerodynamic efficiency of control surfaces, it is very important to eliminate the flow separation over the control surfaces at high deflection angle in order to keep the aircraft having good flight capability, especially for the modern aircraft with tailless aerodynamic configuration. A novel flow control technique to eliminate flow separation of control surface at high deflection angle and creat high lift increment by pulsed blowing at leading edge of control surface is discussed in this paper. The performance of lift enhancment of control surface which used this technique is investigated, and based on the zonal analysis of pulsed frequency, the control characteristic of this technique is also discussed.

Modeling of Cross-Coupling Responses on Hingeless Helicopters via Gyroscopic Effect

The perplexing cross-coupled responses between the control axes on hingeless helicopters have puzzled researchers for years. Unlike previous studies, which introduced a physically meaningless phase-lag to account for the cross-couplings, this paper proposes a new method to relate both on-axis and off-axis responses by the gyroscopic moments through the actuation mechanism of hingeless helicopters. This new method allows investigators to directly and analytically quantify the debatable cross-coupling due to the complex actuation of the rotor. This method is based on the fact that when the angular momentum of the spinning rotor is disturbed by an incremental lift along the main blades due to the varying cyclic pitch angle controlled by the servo mechanisms, off-axis moments are induced to counteract the changes in angular momentum according to the principles of the gyroscope, and hence these gyroscopic moments directly exhibit on-axis responses. This new method yields a parametric framework for examining the previously unexplained cross-coupled responses of hingeless helicopters and demonstrates that, in addition to aerodynamics, intricate dynamic nonlinearities also occur due to non-intuitive actuation mechanisms on the rotor of hingeless helicopters. Finally, simulations and experiments were performed to validate the proposed modeling method.

The Effect of Canard on Aerodynamics of Blended Wing Body

This paper describes the wind tunnel testing of a Blended Wing Body (BWB) with rectangular canards and twisted wing (Baseline II E2 configuration) developed in Universiti Teknologi MARA (UiTM). The experiment work was carried out in UiTM low speed wind tunnel using 1:6 scaled model of BWB at Mach 0.1. The testing is conducted for canard’s deflection angle between 0° to +20. The results show by adding the canard surface to the BWB’s body, at 12 degree and higher angles of attack,α there will be a slight increment in lift. Maximum lift-to-drag ratio decreases with increasing canard surface deflection. Also, by adding the canard surface, the value of moment at zero lift,CM,0 is increased.

Wind Tunnel Experiment for Low Wind Speed Wind Turbine Blade

Environment and green energy awareness are two main factors why this study has been carried out. This research is focused on aerodynamics study for airfoil structure modification based on NACA 0044 and NACA 0063 by using wind tunnel experiment. Aerodynamic characteristics such as lift coefficient, CL, drag coefficient, CD, lift to drag ratio and cell relative velocity has been investigated in this study. CFD simulation has been carried out at the early stage of the investigation (for NACA 0044 and NACA 0063), and a new airfoil profile had been created (0044-63) by modified the chord length and the location of maximum thickness of the airfoil by using the modified NACA Four-Digit Series. Wind tunnel experiment has been take place for three different wind speeds from 25m/s, 35m/s and 45m/s at various angles of attack from 0oto 40owith 5oincremental for the respective airfoil. The results show that the modified 0044-63 produced the better lift coefficient and this airfoil has been fabricated and tested in the wind tunnel experiment in order to validate the CFD result. This paper reports the result of aerodynamics characteristics for respective new airfoil and it shows that at angle of attack between 5oto 15o, this airfoil produced good lift to drag ratio value. Also, by modified the location of maximum thickness 30% to the trailing edge give the increment of lift to drag ratio produced approximately 15% and at the same time, give insignificant changes to the drag coefficient value.

Effects of a trapped vortex cell on a thick wing airfoil

The effects of a trapped vortex cell (TVC) on the aerodynamic performance of a NACA0024 wing model were investigated experimentally at Re = 106 and $$6.67\times 10^{5}$$ . The static pressure distributions around the model and the wake velocity profiles were measured to obtain lift and drag coefficients, for both the clean airfoil and the controlled configurations. Suction was applied in the cavity region to stabilize the trapped vortex. For comparison, a classical boundary layer suction configuration was also tested. The drag coefficient curve of the TVC-controlled airfoil showed sharp discontinuities and bifurcative behavior, generating two drag modes. A strong influence of the angle of attack, the suction rate and the Reynolds number on the drag coefficient was observed. With respect to the clean airfoil, the control led to a drag reduction only if the suction was high enough. Compared to the classical boundary layer suction configuration, the drag reduction was higher for the same amount of suction only in a specific range of incidence, i.e., α = −2° to α = 6° and only for the higher Reynolds number. For all the other conditions, the classical boundary layer suction configuration gave better drag performances. Moderate increments of lift were observed for the TVC-controlled airfoil at low incidence, while a 20% lift enhancement was observed in the stall region with respect to the baseline. However, the same lift increments were also observed for the classical boundary layer suction configuration. Pressure fluctuation measurements in the cavity region suggested a very complex interaction of several flow features. The two drag modes were characterized by typical unsteady phenomena observed in rectangular cavity flows, namely the shear layer mode and the wake mode.

Integrated Design of an Active Flow Control System Using a Time-Dependent Adjoint Method

An exploratory study is performed to investigate the use of a time-dependent discrete adjoint methodology for design optimization of a high-lift wing configuration augmented with an active flow control system. The location and blowing parameters associated with a series of jet actuation orifices are used as design variables. In addition, a geometric parameterization scheme is developed to provide a compact set of design variables describing the wing shape. The scaling of the implementation is studied using several thousand processors and it is found that asynchronous file operations can greatly improve the overall performance of the approach in such massively parallel environments. Three design examples are presented which seek to maximize the mean value of the lift coefficient for the coupled system, and results demonstrate improvements as high as 27% relative to the lift obtained with non-optimized actuation. This lift gain is more than three times the incremental lift provided by the non-optimized actuation.

Flap Gap Oscillatory Blowing on 2D and 2.5D Wing

Here we present preliminary results obtained in developing an active flow control system for high lift systems at advanced TRL level. The work is based on theoretical and experimental work performed in AVERT EU FP6 project where the oscillatory flap gap blowing system was designed and tested on a INCAS F15 2D wing model. Pressure data and global loads have been recorded for a complex evaluation of the basic flow control mechanism. In 2.5D test cases this work has been extended so that the proposed system may be selected as a mature technology in the JTI Clean Sky, Smart Fixed Wing Aircraft ITD. For this goal, new experimental setup was used and also updated electronics for the blowing system have been introduces. This was complemented by a new extension for the data acquisition system and visualization tools. Finally global correlations for basic lift increments have been compared with the reference 2D case and analysed with respect to the system efficiency. INTRODUCTION The overall objective of AVERT Project is to deliver upstream aerodynamics research that will enable breakthrough technology development and innovative aircraft configuration development leading to a step change in aircraft performance. The project objective is to be achieved through the evaluation of selected types of sensor, actuator and control systems, the assessment of these devices against baseline aircraft configurations and the evaluation of the most promising technologies in a medium/large scale wind tunnel test. Figure 1 – AVERT F15 model with end plates (basic 2D setup) The considered flow control technologies are focussed on low-speed applications, mainly investigations with AFC for standard (e.g. DLR F15) and non-standard (e.g. ONERA DND A310) high-lift configurations. Several technologies are investigated and developed INCAS BULLETIN No. 2/ 2009 122 DOI: 10.13111/2066-8201.2009.1.2.17 through both numerical simulations and experiments in wind tunnel. INCAS activities are related to separation control by flap gap oscillatory blowing on a high lift configuration. A high-lift, 2-D model geometry suitable for the application of oscillatory blowing in the flap gap has been selected from previous investigations (DLR F15 model) and manufactured from new at INCAS (Figure 1). A number of test cases with different slot and blowing parameters have been investigated numerically for wind tunnel conditions by DLR. From the numerical results, recommendations for the model design were given (blowing angle and slot widths). A new F15 model has been designed and manufactured for wind tunnel testing at INCAS subsonic wind tunnel. This model was designed so that 19 TU Berlin actuators could be integrated in a 2m span model and tested in a wide range of blowing conditions for the F15 proposed geometry (Figure 2). TU Berlin has designed and manufactured dedicated actuators suitable for flap integration and flap gap oscillatory blowing experimental investigation on a high lift configuration in INCAS facility. Figure 2 – Flap setup (external view) and actuators integration Wind tunnel tests results for the selected test cases have been compared and validated against numerical results and existing experiments (with previous existing WT results at DLR for the F15 high-lift airfoil). EXPERIMENTAL SETUP The experiments in the Subsonic Wind Tunnel at INCAS have been conducted so that we could benefit from the medium size of the facility and to enable large model validation. 1. The maximum wind tunnel test room is of 2.5m width & 2 m height; the maximum permissible span of the model is aprox. 2.0m, in order to make room for end plates and to enable popper distance to the side walls of the test section. 2. In order to achieve a high Reynolds number in the range of 3 million (for wt speed close to 90 m/s), the basic chord length was selected as 600 mm. However, the basic experiments were performed at Reynolds 2 million. INCAS BULLETIN No. 2/ 2009

Aerodynamic Modification of NACA 0012 Airfoil by Trailing-Edge Plasma Gurney Flap

The effect of a novel plasma Gurney flap on the aerodynamic characteristics of a NACA 0012 airfoil is studied by solving the Reynolds-averaged Navier–Stokes equations. The plasma actuator is simulatedwith a phenomenological model. The results indicate that the plasma Gurney flap can increase the lift and nose-down pitching moment of the airfoil, and the mechanism is the same as that of the conventional Gurney flap. The flowfield presents that the vonKarman vortex street disappears near the trailing edge of the airfoil with a plasmaGurney flap, which decreases the airfoil’s drag and thus increases the lift-to-drag ratio before stall. By comparing the lift and nose-down pitching moment increments with the conventional Gurney flap and the jet Gurney flap, the equivalent height or jet blowing momentum coefficient of the plasma Gurney flap are estimated.