Lift Increment Research Articles

A comparison of role-related physical fitness between British Army trainees and trained soldiers.

British Army basic training (BT) and initial trade training (ITT) enable personnel to develop role-related physical capability to perform in-service job-roles. The study aimed to compare physical performance of trainees (completing ITT) and trained soldiers, on a series of gym-based fitness tests and representative military tasks. A total of 316 British Army personnel [68 trainees (63 men: 22±3years, 71.6±8.4kg and 1.74±0.07m) and 248 trained soldiers (225 men: 27±6years, 78.7±12.7kg and 1.76±0.08m)] completed two sessions. Session 1; body mass, stature, age and gym-based tests (2km run, broad jump, seated medicine ball throw, hex bar deadlift, 100m shuttle sprints, pull-ups and mid-thigh pull). Session 2; representative military tasks (loaded carriage [stage 1, 4km, 35-40kg and 4.8kmh-1 fixed pace and stage 2, 2km, 20-25kg and individual best-effort speed], tactical movement, casualty drag, stretcher carry, vertical lift, repeated carry and incremental lift). Independent sample t-tests were employed to examine group differences. Compared to trainees, trained soldiers were older (p<0.001), heavier (p<0.001) and scored higher on broad jump (p=0.024), medicine ball throw (p=0.007) and mid-thigh pull (p=0.048), but were slower on 2km run (p=0.047), loaded carriage (p<0.019), tactical movement (p<0.001) and casualty drag (p<0.001). Overall, trainees achieve higher scores on aerobic/anaerobic tests, whereas trained soldiers outperform trainees in strength/power-based tests. Although a cross-sectional comparison does not provide strong evidence, the results may indicate that cardiovascular fitness is developed during BT, whereas muscle strength/power develops post BT/ITT. These findings would need confirming by a longitudinal study and could inform the development/management of role-related fitness during BT, ITT and through career.

Effects of Propeller Distribution on the Aeroelastic Characteristics of Large-Deformation Wings

Considering the large-deformation and multi-propeller characteristics of very flexible aircraft, propeller effects are introduced and accessed in the wing static and dynamic aeroelastic analysis and different propeller distributions are utilized to obtain more aeroelastic benefits. The propeller–wing aeroelastic interactions are innovatively modeled in the paper. For propeller–wing aerodynamic interaction, propeller-induced velocities are considered and added in the nonplanar steady and unsteady vortex lattice methods. For propeller–wing structural interaction, the conversion of loads and displacements between attached propellers and the large-deformation wing is derived. Static aeroelastic cases indicate that thrust can reduce structural deformation and slipstream can cause considerable lift increment. Dynamic cases indicate that thrust can reduce the wing’s maximum response to gust and bring an improvement of 9.4% in the wing’s critical velocity, while slipstream can reduce the gust response amplitude. In addition, using smaller and more propellers is recommended instead of an individual larger propeller. Decreasing and increasing propeller speeds toward the wingtip is more beneficial for cruise status and gust alleviation, respectively.

High-Lift Enhancement of an Airfoil Using Arrays of Discrete Wall Jets

The circulation of a high-lift airfoil is enhanced using a spanwise array of fluidically oscillating wall jets to engender a Coanda-like effect on the suction surface near the cove between the flap and the main element. The actuation jets manipulate the interaction between the main element boundary layer and the cove jet to overcome flow separation over the flap and yield high-lift performance that is comparable to or better than conventional high-lift airfoils. The fluidic actuation enables effective operation at larger flap deflections while diminishing the sensitivity to cove characteristics. The performance of the actuator array varies with both the jet momentum coefficient and the spanwise periodic jet spacing. It is shown that the lift increment per jet, which for a sufficiently sparse jet array is invariant with jet spacing, diminishes as jet spacing is reduced due to spanwise jet interactions. Accordingly, the overall lift increment scales with the number of active jets and jet density, and a characteristic spanwise jet spacing is identified for optimal performance.

Design perspectives of a system identification approach of the NATO AVT-316 multi-swept wing aerodynamics

Air supremacy requires enhanced maneuvering capabilities at high angles of attack and even beyond the stall angle. High angle-of-attack (high-α) maneuverability can be achieved using swept wings and tailored leading-edge shapes to replace local and disorganized flow separation with controlled separation-induced leading-edge vortices (LEV). Strong leading-edge vortices delay the stall to higher angles and generate a non-linear lift increment up to the angle of attack where the jet-type flow structure of LEV changes to a reversed-flow bubble (vortex breakdown phenomenon). A multi-swept wing configuration has the potential to delay the vortex breakdown to even higher angles as the vortices formed over the front wing can energize vortices formed over the main wing and lead to vortex merging. This article is focused on understanding the unsteady interactions between multiple leading edge vortices formed over multi-swept wing configurations in the subsonic speed regime. Specifically, this article investigates the aerodynamic characteristics of wings being evaluated under the initiative of the NATO STO AVT-316 Task Group. Highly refined meshes, and the use of hybrid turbulence models such as Detached Eddy Simulations (DES) and Delayed DES are required to accurately resolve the vortical flows over these wings. Simulating all flow conditions of interest is a computationally expensive approach. A system identification method was therefore proposed to rapidly and accurately generate aerodynamic models of these wings. The method uses a novel piece-wise chirp (constant amplitude and increasing frequency) motion as an input signal (training maneuver). The motion computational cost is equivalent to the cost of six static CFD simulations, however it can predict aerodynamic responses of the wings over a wide range of angles of attack. The method has been tested for double and triple delta wings. Some design considerations are provided based on predicted flow features and aerodynamic data. Prediction results with different turbulence models, sting geometries, grid resolutions and an adaptive mesh refinement approach are provided.

The flow control mechanism of trailing-edge Gurney flap on a 50°-swept delta wing in forced pitching

The flow control effect of the trailing-edge Gurney flap (TG) on the dynamic lift characteristics for a 50°-swept delta wing during large-amplitude pitching oscillations at various reduced frequencies (k = 0.072, 0.144, 0.287, and 0.575) was investigated via force, particle image velocity, and dye visualization measurements in a water channel facility. Numerical simulations were carried out to further understand the flow control mechanism of the TG in low and high reduced frequency cases (k = 0.072 and 0.575). It was found that as the reduced frequency increases, the lift increments brought by the TG are magnified and abated during the upstroke and downstroke processes, respectively. The breakdown of the leading-edge vortex (LEV) on the upper surface of the wing is promoted by the TG during the early stage of the pitching cycle. The lift enhancement being benefited by the TG is mainly contributed by the recovery of lower surface pressure along the trailing edge due to the blockage effect of TG, which also stimulates the spanwise flow and strengthens the LEV upon the upper surface. The significant lift increment contribution of the upper surface during the upstroke process can be maintained to higher angle of attack as the reduced frequency increases.

Rapid dynamic aeroelastic response analysis of the highly flexible wing with distributed propellers influence

AbstractA rapid nonlinear aeroelastic framework for the analysis of the highly flexible wing with distributed propellers is presented, validated and applied to investigate the propeller effects on the wing dynamic response and aeroelastic stability. In the framework, nonlinear beam elements based on the co-rotational method are applied for the large-deformation wing structure, and an efficient cylinder coordinate generation method is proposed for attached propellers at different position. By taking advantage of the relatively slow dynamics of the high-aspect-ratio wing, propeller wake is modeled as a quasi-steady skewed vortex cylinder with no updating process to reduce the high computational cost. Axial and tangential induced velocities are derived and included in the unsteady vortex lattice method. For the numerical cases explored, results indicate that large deformation causes thrust to produce wing negative torsion which limits the displacement oscillation, and slipstream mainly increases the response values. In addition, an improvement of flutter boundary is found with the increase of propeller thrust while slipstream brings a destabilising effect as a result of the increment of dynamic pressure and local lift. The great potential of distributed propellers in gust alleviation and flutter suppression of such aircraft is pointed out and the method as well as conclusions in this paper can provide further guidance.

Influence of passive deformation in the lift coefficient of a NACA0012 wing model

The extensive use of lightweight materials in aerial vehicle wings involves structural flexibility phenomena that generate non-negligible deformation effects. This influence is not restricted to big aircraft but also plays a role in smaller aeroplanes and Unmanned Aerial Vehicles (UAVs). Here, we conduct wind tunnel experiments to analyze the effect of passive deformation on the wing model lift slopes. To isolate the deformation effect, we compare rigid wings with a NACA0012 airfoil imposing a prescribed spanwise deformation. We study three levels of deformation: non-deformed, around 2% and 4.5% of tip deflection. Also, we consider the effect of the wing length by using three different semi-aspect ratios (1, 2, and 4), so a total of nine rigid wing models have been analyzed for a range of Reynolds number from 80×103 to 160×103. Deformed wing models show an increase in lift coefficient compared to non-deformed wing cases. Both deformation levels exhibit a qualitatively similar lift increment. A correlation to predict lift coefficient slope in a flat plate is adapted for a NACA0012 airfoil and validated using our experimental results and literature data. The adjusted correlation can quantify the deformation effect on the lift slope, which is comparable to using a slightly longer wing model.

A Numerical Study of the Impact of Bending and Torsional Stiffness on the Static Aeroelastic Characteristics of a Large Aspect Ratio Composite Wing

The bending and torsional stiffness influence structural deformation to varying degrees. Utilizing the coupled numerical method of Computational Fluid Dynamics and Computational Structural Dynamics (CFD/CSD), a comprehensive study was conducted on the impact of bending and torsional stiffness on the static aeroelastic behavior of a high-aspect-ratio wing, employing FLUENT software across diverse flight conditions. The study summarized this influence by comparing the computational outcomes. The results indicate that as the angle of attack rises, the lift increment diminishes gradually until the angle reaches 14 degrees, rendering the strategy of enhancing bending and torsional stiffness to gain more lift ineffective. Between 2 and 14 degrees of angle of attack, the lift difference between the scenarios decreases from 23.98% to 7.31%. At higher Mach numbers and lower angles of attack, augmenting wing stiffness significantly boosts wing lift. Optimal lift-to-drag characteristics are achieved at approximately 6 degrees of angle of attack. By averagely increasing the wing’s bending and torsional stiffness by 8.28% and 5.22%, respectively, the lift-to-drag characteristics can be enhanced by 5.27% at a low angle of attack of 0.75 Mach. The disparity in maximum deflection between the two stiffness wings is most pronounced at higher flight speeds and smaller angles of attack, with the opposite trend observed for the difference in maximum torsion angle. The key findings presented in this paper can expedite the integrated design of stiffness for this type of wing structure by providing vital technical insights

Effect of Tip-Mounted Propeller and Trailing-Edge Flap on Wing's Lift and Drag Coefficients and Vortical Wake

Abstract The lift and drag forces and vortical wake of a rectangular semiwing equipped with a tip-mounted propeller and a plain trailing-edge flap both outside and in ground effect were investigated experimentally at Re = 2.49 × 105. Outside the ground effect, the slipstream of the inboard-up propeller rotation attenuated the wingtip vortex, significantly decreasing the lift-induced drag. The propeller rotation also led to an increased maximum lift and stall angle. At a fixed lift condition, the propeller rotation led to a reduced drag. The deployment of the flap further increased the lift curve slope and maximum lift but at the expense of an increased drag and earlier stall as compared to the unflapped wing. In ground effect, the lift augmentation of the unflapped wing increased pronouncedly with propeller rotation as the ground was approached. For the flapped wing, the propeller rotation, however, produced a smaller lift increment than the unflapped wing due to the recirculation of the deflected slipstream into the propeller disk. The extent of this recirculation increased with flap deflection and propeller rotation, contributing to the inferior lift production of the flapped wing with a tip-mounted propeller in ground effect.

Comparative Study and Aerodynamic Analysis of Rectangular Wing Using High-Lift Systems

High-lift systems are designed to expand the flight envelope and have a most important effect on the size of the wing, economy, and safety of many airliner configurations. Even a small increment of lift using a high-lift system can significantly impact an aircraft’s profitability. The effective design of the airfoil shape with the required aerodynamic performance is still difficult. In the early days, the designs of airfoils were randomly set up and tested in the flow section, and then, the Wright brothers emerged with a cambered section. NACA has provided an appropriate airfoil definition that supports us in making airfoil designs using formulas, not randomly. This paper describes the influence of aerodynamic analysis of wing with flaps at various deflection angles. Aerodynamic variables for the aircraft wing, which is made up of the NACA airfoil 6412 model with and without flaps, have been studied at various angle of attack (AOA) (i.e., -4, -2, 0, 2, 4, 6, 8, 10, 12, and 16) and different Mach number at 0.2, 0.3, and 0.4. Also, the analysis was done for the 15000 ft altitude to check the density effects for the real-time applications. The coefficients of lift and drag are gained by examining the pressure distribution over the surface of the wing. Lift increases as the approach ascends from a low to a high angle of attack, and the most extreme lift is produced at a specific point. After that, when the angle of attack increases further, the drag component increases, so the stall occurs at that point in time. The results showed that the NACA 6412 airfoil obtained the maximum lift at 14°, and the lift value started to decrease. The CFD computations are performed in Ansys Fluent by performing hybrid mesh using ICEM-CFD. The analysis is performed for various configurations of the wing section, and the effects of flow parameters like angle of attack, altitude, and the gap distance between the main wing and slotted flap were compared to identify the better configurations.

Effectives of Different Shaped Dimples on a NACA Airfoil

This research addresses changes in the aerodynamic properties of the airfoil caused by modifying the surface in the form of dimples. The first surface changes mentioned here the wing model has dimples that face inside and outward. Airfoil models that have been modified to illustrate variation lifting and dragging at varying angles of attack (AOA). Surface adjustments are made here by taking into account the various types and forms of dimples. When an airfoil reaches a certain angle of attack, dimples serve to reduce pressure drag because wake production begins owing to boundary layer separation. Dimples on an aeroplane wing function similarly to vortex producers. They cause turbulence, which slows down boundary layer splitting, decrease wake, and reduces pressure drag. It has helped to improve the lift and raise the AOA of the stall. Dimples on the surface of an aircraft wing do not considerably increase pressure drag due to the wing already has an aerodynamic form, but they may impact its aerodynamic properties when the airfoil is at AOA, which is one of the issues raised in this research. The outcomes legitimize the increment in the general lift and decrease in drag by the airfoil

A Computational Fluid Dynamics Investigation of a Flapping Hydrofoil as a Thruster

The paper features a computational fluid dynamics study of a flapping NACA0015 hydrofoil moving with a combination of sinusoidal heaving and pitching. Several kinematic configurations are explored, varying sequentially pitch and heave amplitude, Strouhal number and phase angle, in an attempt to determine the influence of each parameter on the propulsive performance. To optimize efficiency the angle of attack should assume the highest value that also avoids the arise of the leading edge vortex generated in the dynamic stall state. At low Strouhal number optimum is reached at high heave amplitudes, which correspond to the configurations minimizing the hysteresis in the (Cy,Cx) plane. The same outcome in terms of hysteresis minimization has been verified to occur when optimal phase shift was considered. Differently, when the Strouhal number and the angle of attack become higher, to exploit efficiently the lift increment owed to dynamic stall it emerged the necessity of adopting low heave amplitude to improve separation resistance, avoiding the occurrence of deep stall.

Aero-Propulsive Interactions between UAV Wing and Distributed Propellers Due to Their Relative Position

The purpose of this paper is the evaluation of the aero-propulsive effects on a UAV wing model with distributed propulsion. An array of three propellers is placed ahead of the leading edge of a rectangular wing with flap. The investigation was performed with high-fidelity numerical analyses to provide insights into the phenomenology and to screen the interesting positions to be validated in the wind tunnel. The propellers’ array is moved into twelve different positions, allowing longitudinal and vertical translations. The wing has an untwisted and constant section profile, with a single slot trailing-edge flap that is deflected into three positions. The flap span is entirely covered by the propellers’ blowing. Results show an increment of lift, drag, and pitching moment coefficients with distributed propellers enabled. For a given thrust level, the magnitude of such increments depends on the propellers’ positions, the flap configuration, and the angle of attack. The lift enhancement sought in distributed propulsion applications comes at the expense of a significant increase in drag and pitching moment magnitude. In some combinations, the wing’s contribution to the aircraft longitudinal stability is severely affected. Conversely, the propellers’ inflow is altered such that thrust is increased in all the investigated configurations, with a small reduction of propulsive efficiency.

Pressure Feedback Control of Aerodynamic Loads on a Delta Wing in Transverse Gusts

A feedback active flow control (AFC) scheme is studied for control of unsteady aerodynamic loads that act on a generic tailless delta wing during transverse gust encounters. The transverse gust and the roll angle of the wing create an unsteady roll moment. The trailing edge actuators generate unbalanced lift increments on the two sides of the model to control the roll moment. Spatially distributed pressure sensors measure surface pressure on the suction side of the wing, and the real-time aerodynamic loads are estimated through models that are identified by the Sparse Identification of Nonlinear Dynamics algorithm. The aerodynamic load estimation is then used as a surrogate to provide the AFC system with a feedback signal to alleviate the unsteady roll moment due to the gust effect. A physical interpretation of the model identification results is made to locate the source of the lift and roll moment fluctuations in gust encounters. Experimental results show that the AFC effect with a momentum coefficient input of 2% is equivalent to 26° elevon deflection in terms of the roll moment change, and the AFC doubles the control authority relative to elevons. The real-time control results show that the pressure feedback control system reduces the mean roll moment error and suppresses the fluctuations to less than 20% of the uncontrolled fluctuation level at a high angle of attack.

Numerical Evaluation on Large Ground-Effect Seaplane at Takeoff with Power Augmented Ram

Power-augmented ram has been considered as a promising solution for increasing lift coefficients for large ground-effect seaplanes that typically are required to take off at sea. To date, however, a quantitative evaluation of the power-augmented ram (PAR) technique is still lacking, except for a few conceptual analyses; and its possible technical barriers are unclear. In this study, a typical large-scale ground-effect vehicle is employed to assess the effect of the PAR on lift enhancement at takeoff using numerical simulation. The results show that high-lift enhancement can be obtained as front-installed engine jets ram into the open cavity between a wing and the ground. The jet flow produces high pressure on the lower surface of the wing, significantly increasing the lift. However, the high pressure is primarily imposed on the trailing-edge flap that is deflected down at takeoff, thus yielding a large nosedown pitching moment that is basically proportional to the lift increment. The extra pitching moment will pose a challenge to trimming the seaplane with conventional elevators. Reasonable designs (including multiple engine layouts, increasing nozzle area, and decreasing nozzle velocity) can improve the undesired pitching moment, thus achieving better aerodynamic performances.

Flutter and discrete gust load alleviation characteristics of multi-segment folding wings with constant cross section

Flutter and discrete gust load alleviation characteristics of multi-segment folding wings with constant cross section are studied. A bending-torsional beam element is developed to simulate the bending and pitching modes. Aerodynamic lift data of a folding wing at different folding angles are obtained from the modified aerodynamic strip theory. Firstly, a structural dynamic model including folding angle parameters is established based on finite element method, and modal truncation technology is introduced to decouple the corresponding dynamic model. Secondly, aerodynamic lifts and pitching moments are obtained in the frequency domain according to Theodorsen unsteady aerodynamic theory. Then V – g method for flutter calculation is introduced to get flutter velocities and flutter frequencies at different folding angles. Thirdly, aerodynamic lift increments caused by discrete gust are applied to the folding wing and a set of integro-differential equations which can be solved by Laplace transform method is derived. Finally, a three-segment folding wing of a near ground UAV is investigated and the results show that existences of folding angles can effectively improve the flutter velocity of the folding wing. And the gust loads can be effectively alleviated by reasonably selecting folding modes and folding angles, especially, the peak values of bending moment at wing root can be alleviated up to about 62%.

Numerical Simulation of Ground Effect on Circulation Control Airfoil

Circulation control (CC) is used extensively to control the attitude of rudderless aircraft experiencing ground effect in take-off and landing phases. The investigation of ground effect on airfoil CC is necessary to improve flight performance and quality in proximity to the ground. The aerodynamics and flow field of a modified NACA0012 airfoil with CC in ground effect are investigated with numerical simulations. The compressible Reynolds-averaged Navier–Stokes equations with the shear-stress transport k − ω turbulence model equations are solved using the Finite Volume Method (FVM). Simulation results show that the ground effect changes the lift increment per unit jet momentum coefficient, and CC can reverse the polarity ground effect. The effective angle of attack α E and the downwash space downstream airfoil are altered by the ground effect resulting in variations of airfoil surface pressure and lift. Unlike the unbounded flow field, the jet attachment distance is not only determined by the jet momentum coefficient but it can change with the ride height, which is the distance from the ground to the center of the semicircular Coanda surface, for the same jet momentum coefficient.

Experimental Investigation of Flow Separation Control over Airfoil by Upper Surface Flap with a Gap

This paper focused on the effects of the flap with a tiny gap on the flow separation over the NPU-WA-180 airfoil. The effects of the geometric parameters of the flaps, such as the flap gap height, angle, and position, were investigated. The study showed that the flap can significantly improve the stall features of airfoil in a limited phase of angle of attack (AoA), and increase lift and reduce drag at a high AoA. It can increase the lift coefficient and drag coefficient in the case of high AoAs, and the angle range of the lift augmentation and drag reduction can reach more than 9°. Furthermore, an excessively large gap is not conducive to the improvement of the airfoil stall performance. The flap angle plays a key role in the airfoil stall characteristic. As the flap angle decreases, the angle range of improving airfoil stall characteristics becomes larger, the pitching moment increment becomes smaller. However, the maximum lift increment and the effect of the drag reduction will decrease. And the effects of the position of the flap on the airfoil performance were also studied. Considering the maximum lift coefficient and drag coefficients in large AoA, the Type1 installed at the 0.7c position has the best effect; from the perspective of delayed stall, the Type1 installed at the 0.6c position has the best delay effect. These results can provide the data and theoretical support for the flap application in engineering.

Experimental research on aero-propulsion coupling characteristics of a distributed electric propulsion aircraft

Distributed Electric Propulsion (DEP) aircraft use multiple electric motors to drive the propulsors, which gives potential benefits to aerodynamic-propulsion interaction. To investigate and quantify the aerodynamic-propulsion interaction effect of the wing section, we built a DEP demonstrator with 24 “high-lift” Electric Ducted Fans (EDFs) distributed along the wing’s trailing edge. This paper explores and compares the aero-propulsion coupling characteristics under various upstream speed, throttle, and EDF mounting surface deflection angles using a series of wind tunnel tests. We compare various lift-augmentation power conditions to the clean configuration without propulsion unit under the experiment condition of 15–25 m/s freestream flow and angles of attack from −4° to 16°. The comparison of computational results to the experimental results verifies the effectiveness of the computational fluid dynamic analysis method and the modeling method for the DEP configuration. The results show that the EDFs can produce significant lift increment and drag reduction simultaneously, which is accordant with the potential benefit of Boundary Layer Ingestion (BLI) at low airspeed.

Numerical investigation of the aerodynamic interference in 2-aircraft formation flight

Due to obvious reduction in drag, formation flight attracts wide attention of major powers in the world. Based on wild geese flying in a ‘V’ formation, a formation composed of two flying-wing aircrafts is designed. The aerodynamic interference in 2-aircraft formation flight is studied by CFD. Results show that 2-aircraft formation flight has little impact on leading aircraft, but has an evident impact on following aircraft. At the same AoA, the effect of increment in lift and reduction in drag of following aircraft is large. As considering the balance of gravity and lift, the reduction in drag of following aircraft is much larger, that is up to 25.1%. Under the aerodynamic interference effect, the following aircraft has a nosedown pitching moment increment, and becomes lateral static instability, which is opposite to the characteristic of solo flying-wing aircraft. The aerodynamic interference mainly comes from upwash flow induced by wingtip vortex of leading aircraft, which reduces the actual AoA and induced drag of following aircraft, and changes the lift distribution of both sides of following aircraft.