Low Angles Of Attack Research Articles

Shock Buffet and Associated Fluid–Structure Interactions of the Benchmark Supercritical Wing

This study focuses on the fluid–structure interactions of the Benchmark Supercritical Wing (BSCW) at buffet conditions (transonic flows and moderate angles of attack). It presents unsteady Reynolds-averaged Navier–Stokes simulations of the fixed BSCW, the BSCW undergoing prescribed pitch motion, and the BSCW suspended on springs in one and two degrees of freedom. The study’s objectives are to 1) establish the BSCW buffet properties and 2) investigate the aeroelastic and prescribed-motion responses of the wing at buffet conditions. The results indicate that the BSCW buffet is predominantly two-dimensional in terms of its characteristics, although it is greatly affected by the tip flow. The response to prescribed pitch motion and the aeroelastic response of the spring-suspended wing display similar phenomena, where the response in the buffet frequency is suppressed once the oscillation amplitude exceeds some threshold, and the structural and aerodynamic responses are solely at the structural frequency. At the studied conditions, the BSCW undergoes stall flutter, dominated by pitch oscillations, compared to a bending-torsion coalescence flutter at lower angles of attack. We conclude that the shock buffet does not drive the BSCW flutter instability. However, both the buffet and stall flutter develop at similar flow conditions of detached boundary layer behind a strong shock at transonic Mach numbers and moderate angles of attack.

HLPW-4: Wall-Modeled Large-Eddy Simulation and Lattice–Boltzmann Technology Focus Group Workshop Summary

A summary of the nine submissions to the Wall-Modeled Large-Eddy Simulation and Lattice–Boltzmann (WMLESLB) Technical Focus Group (TFG) at the 4th High lift Prediction Workshop is provided. The focus of this TFG was to assess the current capabilities of WMLES and LB methods on a complex high-lift configuration across a wide range of angles of attack. Analysis of the submitted data suggests that >250 M spatial degrees of freedom are needed to accurately predict pitching moments at high angles of attack due to large pressure gradients present on the outboard slat and main element for α>17° (corrected for free-air). While some scatter is reported in pitching moment coefficient at the low-angles of attack (α<11°), excellent agreement is observed between submissions near the CL,max state. Objective superiority of WMLES methods over Reynolds-averaged Navier–Stokes (RANS) can be seen in terms of lack of excess outboard separation; a majority of the WMLES and LB submissions predict wedge-shaped separation patterns consistent with the experimental oil flow. The in-tunnel simulations show excellent agreement with the experiment in terms of a) integrated loads, b) surface flow-topology, and c) mechanism for the onset of inboard stall. Further evidence is provided to demonstrate both qualitative and quantitative superiority of the WMLES submissions over RANS.

Design of a Bioinspired Underwater Glider for Oceanographic Research.

The Blue Economy, which is based on the sustainable use of the ocean, is demanding better understanding of marine ecosystems, which provide assets, goods, and services. Such understanding requires the use of modern exploration technologies, including unmanned underwater vehicles, in order to acquire quality information for decision-making processes. This paper addresses the design process for an underwater glider, to be used in oceanographic research, that was inspired by leatherback sea turtles (Dermochelys coriacea), which are known to have a superior diving ability and enhanced hydrodynamic performance. The design process combines elements from Systems Engineering and bioinspired design approaches. The conceptual and preliminary design stages are first described, and they allowed mapping the user's requirements into engineering characteristics, using quality function deployment to generate the functional architecture, which later facilitated the integration of the components and subsystems. Then, we emphasize the shell's bioinspired hydrodynamic design and provide the design solution for the desired vehicle's specifications. The bioinspired shell yielded a lift coefficient increase due to the effect of ridges and a decrease in the drag coefficient at low angles of attack. This led to a greater lift-to-drag ratio, a desirable condition for underwater gliders, since we obtained a greater lift while producing less drag than the shape without longitudinal ridges.

Propeller Rotation Effects on Pitching Moment for Transport Aircraft with Conventional Tail

The effects of the propeller rotation direction on the pitching moment characteristics for a typical twin-engine turboprop aircraft with a conventional tail at low and moderate angles of attack under high-power conditions are investigated through unsteady Reynolds-averaged Navier–Stokes simulations. The numerical method is validated using wind-tunnel data for a twin-engine T-tail turboprop aircraft. The computational results of the conventional tail configurations indicate that the corotating propellers significantly degrade the longitudinal stability at low angles of attack in the takeoff configuration. By comparison with corotating propellers, counter-rotating inboard-up propellers not only increase the lift but also lead to a notable improvement in longitudinal stability, especially at low angles of attack. In contrast, the counter-rotating outboard-up propellers further reduce the negative pitching moment curve slope of the airframe. Analysis of the flowfield demonstrates that the effects of the propeller rotation direction significantly alter the downwash gradient at the horizontal tailplane by affecting the vertical location of the high-downwash region. The effectiveness of the horizontal tailplane and the longitudinal stability of the aircraft are therefore changed.

Measurement of aerodynamic force and moment acting on a javelin using a magnetic suspension and balance system

The rules governing the dimensions of the Javelin were substantially changed in 1986. It was considered that this new design guaranteed there was zero pitching moment at 0° angle of attack and that the pitching moment decreased (became negative) with increasing angle of attack. The objective of this study is to investigate if the pitching moment remains always negative (nose-down rotation). To measure accurate aerodynamic forces acting on a Javelin, the world’s largest 1 m magnetic suspension and balance system was used. The magnetic suspension and balance system was able to measure aerodynamic forces without support interference in the wind tunnel. In addition, computational fluid dynamics were carried out to estimate the pitching moment coefficients. It was found that the pitching moment coefficient of a commercially available Javelin becomes positive (nose-up rotation) at lower angles of attack, less than 12°. The pitching moment becomes positive if the upstream side of the center of gravity receives more inflow than the downstream side. This situation can be attained by, for example, increasing the thickness of the upstream side when compared with that of the downstream side.

火星飛行機のための高高度飛行試験MABE-1の機体の運動特性について

In this paper, we describe two issues that have been discovered during the design of a flight control system of Mars airplane and an experimental vehicle. The first issue is that both, the experimental and Mars aircraft, easily tend to entry to the backside region. This has been observed even at higher speeds and lower angle of attack than in conventional aircraft due to different flight environments and aerodynamic characteristics. The second issue arises due to the aircraft being lightweight and able to fly at high cruising speeds. This results in a high natural frequency of short period mode. Due to this high natural frequency, usually negligible time delay effects have a significant effect. Therefore, the phase margin of the system decreases and its control performance tends to deteriorate.

Blunt leading-edge effect on spanwise-varying leading-edge contours of an UCAV configuration

The vortex system on the SACCON model is studied under low Reynolds number by using the flow visualization technique and the surface pressure testing measurement in a low-speed water channel and a wind tunnel, respectively. A main topic is discussed below. The blunt leading-edge section on this SACCON model induces various flow phenomena under different angles of attack. The attached flow appearing around the upstream leading-edge region at a low angle of attack induces different flow phenomena downstream, so does the branched vortices appearing near the trailing-edge. Although this attached flow is vanished at higher angle of attack, the stability of forming an outboard vortex can be shown in the results of pressure coefficients as a function of an angle of attack. Moreover, how the blunt leading-edge contour on SACCON model and other models affecting flow field can be compared with the findings given in the previous studies. The models with only blunt leading-edge contour can prompt an outboard vortex and an inner vortex, compared to the flow fields on the models with spanwise-varying leading-edge contours. This inner vortex is formed due to the attached flow inboard passing downstream and being affected by the outboard vortex. However, this attached flow is different from the attached flow observed on SACCON model. Therefore, flow phenomena occurring on different blunt leading-edge models can be differentiated.

Numerical Study on Flow and Noise Characteristics of an NACA0018 Airfoil with a Porous Trailing Edge

An airfoil with a porous trailing edge has a low noise emission; thus, using a porous medium is a good technique for further reduction of wind turbine noise. In this paper, to reduce airfoil trailing edge noise while minimizing the negative influence of a porous medium on aerodynamic performance, a new filling method is proposed such that a porous medium is only used in the suction side half of the trailing edge, which is more sensitive to the noise generation. The large eddy simulation (LES) technique for flow and the Ffowcs Williams and Hawkings (FW-H) method for acoustics are used. At a Reynolds number of 2.63 × 105 and various angles of attack, an NACA0018 airfoil profile with a porous trailing edge covering 20% of the chord is studied under two porous configurations, namely a fully porous and a suction-side porous trailing edge type. The results show that the flow direction, velocity magnitude, and their distributions along the boundary layer of the two porous airfoils are significantly modified due to the presence of the porous medium. The fluctuation of the pressure coefficient and the increase in the boundary layer thickness are significant at low angles of attack. As compared to the solid airfoil counterpart, the noise radiation from the newly proposed suction-side porous airfoil achieves a noise reduction of 4.3 dB at an angle of attack α = 0°, and a noise reduction of 4.07 dB at an angle of attack α = 2°.

Implementation of Multidisciplinary and Multifidelity Uncertainty Quantification Methods for Sonic Boom Prediction

To advance a supersonic aircraft design process, an uncertainty quantification study is conducted for sonic boom prediction while considering uncertainties associated with flight and atmospheric conditions. The uncertainty quantification process is implemented within a multidisciplinary analyses framework and assisted with a multifidelity surrogate model based approach. The sonic boom prediction framework requires input from the flowfield pressure distribution solution to generate the near-field pressure signature of the aircraft, which is then propagated throughout the atmosphere to the ground by using aeroacoustic methods. The open-source SU2 suite is employed as a high-fidelity flow solver tool to obtain the aerodynamic solution, while in-house postprocessing scripts are developed to generate the required near-field pressure signature. For low-fidelity flow analysis, A502 PAN AIR, a higher-order panel code which solves flows around slender bodies in low angles of attack for subsonic and supersonic regimes, is used. For nonlinear aeroacoustic propagation, NASA Langley Research Center’s code sBOOM exploits the near-field pressure signature for both high-fidelity and low-fidelity sonic boom calculations. Efficient uncertainty quantification tools are developed in house by implementing multifidelity polynomial chaos expansion and multifidelity Monte Carlo methods. Several flight and atmospheric parameters are selected to include randomness where these uncertainties are propagated into the sonic boom loudness prediction of the JAXA wing-body model, which is a low boom aircraft. Finally, an overall assessment of the multifidelity uncertainty quantification methods is presented in terms of efficiency and numerical accuracy.

Bio-inspired optimization of leading edge slat

PurposeThe high lift devices are effective at high angle of attack to increase the coefficient of lift by increasing the camber. But it affects the low angle of attack aerodynamic performance by increasing the drag. Hence, they have made as a movable device to deploy only at high angles of attack, which increases the design and installation complexities. This study aims to focus on the comparison of aerodynamic efficiency of different conventional leading edge (LE) slat configurations with simple fixed bioinspired slat design.Design/methodology/approachThis research analyzes the effect of LE slat on aerodynamic performance of CLARK Y airfoil at low and high angles of attack. Different geometrical parameters such as slat chord, cutoff, gap, width and depth of LE slat have been considered for the analysis.FindingsIt has been found that the LE slat configuration with slat chord 30% of airfoil chord, forward extension 8% of chord, dip 3% of chord and gap 0.75% of chord gives higher aerodynamic efficiency (Cl/Cd) than other LE slat configurations, but it affects the low angles of attack aerodynamic performance with the deployed condition. Hence, this optimum slat configuration is further modified by closing the gap between LE slat and the main airfoil, which is inspired by the marine mammal’s nose. Thus increases the coefficient of lift at high angles of attack due to better acceleration over the airfoil nose and as well enhances the aerodynamic efficiency at low angles of attack.Research limitations/implicationsThe two-dimensional computational analysis has been done for different LE slat’s geometrical parameters at low subsonic speed.Practical implicationsThis bio-inspired nose design improves aerodynamic performance and increases the structural strength of aircraft wing compared to the conventional LE slat. This fixed design avoids the complex design and installation difficulties of conventional movable slats.Social implicationsThe findings will have significant impact on the fields of aircraft wing and wind turbine designs, which reduces the design and manufacturing complexities.Originality/valueDifferent conventional slat configurations have been analyzed and compared with a simple fixed bioinspired slat nose design at low subsonic speed.

Lift Enhancement by Trapped Vortex at Ultra-low Reynolds Numbers

A means of passively trapping a vortex to enhance the aerodynamic characteristics of an airfoil is explored in this paper. The flow over an airfoil with a cavity on the upper surface, an NACA 0004 airfoil, and a flat plate airfoil is studied computationally in the ultra-low Reynolds number regime. At low angles of attack a vortex is seen to be trapped in the cavity. With increasing angle of attack the trapped vortex grows in size and occupies the entire cavity and the effective geometry resembles a smooth airfoil. The vortex circulation also increases and enhances the lift characteristics of the airfoil. The drag coefficient of the cavity airfoil is comparable to the NACA 0004 airfoil. The lift-to-drag ratio of the cavity airfoil is found to be higher than the smooth airfoils in the entire range of angles of attack considered in the present study. The cavity airfoil, with a trapped vortex, offers an interesting prospect for engineering nano-air vehicles.

Minimizing airfoil drag at low angles of attack with DBD-based turbulent drag reduction methods

Dielectric Barrier Discharge (DBD) based turbulent drag reduction methods are used to reduce the total drag on a NACA 0012 airfoil at low angels of attack. The interaction of DBD with turbulent boundary layer was investigated, based on which the drag reduction experiments were conducted. The results show that unidirectional steady discharge is more effective than oscillating discharge in terms of drag reduction, while steady impinging discharge fails to finish the mission (i.e. drag increase). In the best scenario, a maximum relative drag reduction as high as 64 % is achieved at the freestream velocity of 5 m/s, and a drag reduction of 13.7 % keeps existing at the freestream velocity of 20 m/s. For unidirectional discharge, the jet velocity ratio and the dimensionless actuator spacing are the two key parameters affecting the effectiveness. The drag reduction magnitude varies inversely with the dimensionless spacing, and a threshold value of the dimensionless actuator spacing of 540 (approximately five times of the low-speed streak spacing) exists, above which the drag increases. When the jet velocity ratio smaller than 0.05, marginal drag variation is observed. In contrast, when the jet velocity ratio larger than 0.05, the experimental data bifurcates, one into the drag increase zone and the other into the drag reduction zone, depending on the value of dimensionless actuator spacing. In both zones, the drag variation magnitude increases with the jet velocity ratio. The total drag reduction can be divided into the reduction in pressure drag and turbulent friction drag, as well as the increase in friction drag brought by transition promotion. The reduction in turbulent friction drag plays an important role in the total drag reduction.

Experimental flow control investigation over suction surface of turbine blade with local surface passive oscillation

Impact of the local flexible membrane (LFM) on aerodynamic phenomena including the formation of a laminar separation bubble (LSB) and transition to turbulence was experimentally investigated over the suction surface of a Clark-Y airfoil first time in literature. The experiments such as aerodynamic force measurement, smoke-wire flow visualization and hot-film tests were carried out at the free-stream velocity of U∞ = 3.2 m/s, U∞ = 6.4 m/s, U∞ = 9.6 m/s, U∞ = 12.8 m/s, and Reynolds number based upon on the chord length was Rec = 3.5 × 104, Rec = 7.0 × 104, Rec = 1.05 × 105 and Rec = 1.4 × 105, respectively. The experimental angle of attack was set at 0° = α ≤ 20°. In detailed intermittency analysis by the hot-film sensor over the uncontrolled airfoil, it was seen that the LSB and transition to turbulence formed close to the trailing edge at a lower angle of attack, and it moved towards the leading edge when increasing the angle of attack simultaneously. Employing LFM on the suction surface obviously affected the progress of these flow phenomena. In the results of smoke-wire flow visualization, either the size of the laminar separation bubble (LSB) was reduced or its presence was suppressed at lower incidences. The aerodynamic force measurement results also supported those behaviors. In particular, at lower incidences, the negative effects of LSB were mitigated, resulting in the presence of a more stable lift curve. Additionally, it was clearly observed that utilizing LFM ensured positive effects, especially at the pre- and the post-stall regions in terms of fewer fluctuations at the CL curve, meaning that less aerodynamic vibration and noise on wind/hydro turbine could be obtained.

Wing sweep effects on laminar separated flows

We reveal the effects of sweep on the wake dynamics around NACA 0015 wings at high angles of attack using direct numerical simulations and resolvent analysis. The influence of sweep on the wake dynamics is considered for sweep angles from $0^\circ$ to $45^\circ$ and angles of attack from $16^\circ$ to $30^\circ$ for a spanwise periodic wing at a chord-based Reynolds number of $400$ and a Mach number of $0.1$ . Wing sweep affects the wake dynamics, especially in terms of stability and spanwise fluctuations with implications on the development of three-dimensional (3-D) wakes. We observe that wing sweep attenuates spanwise fluctuations. Even as the sweep angle influences the wake, force and pressure coefficients can be collapsed for low angles of attack when examined in wall-normal and wingspan-normal independent flow components. Some small deviations at high sweep and incidence angles are attributed to vortical wake structures that impose secondary aerodynamic loads, revealed through the force element analysis. Furthermore, we conduct global resolvent analysis to uncover oblique modes with high disturbance amplification. The resolvent analysis also reveals the presence of wavemakers in the shear-dominated region associated with the emergence of 3-D wakes at high angles of attack. For flows at high sweep angles, the optimal convection speed of the response modes is shown to be faster than the optimal wavemakers speed suggesting a mechanism for the attenuation of perturbations. The present findings serve as a fundamental stepping stone to understanding separated flows at higher Reynolds numbers.

The impact of the ground on flow structure and aerodynamic characteristics of a double delta wing

The ground effect (GE) on the formation of the vortical flow over a double delta wing having sweep angles of Λ=70°/40° was examined using the Particle Image Velocimetry (PIV) and dye flow visualization methods in this study. Qualitatively and quantitatively measured results were demonstrated for angles of attack, α=15° and 25° cases in the side and cross-flow planes to reveal the detail of the flow structures. The distance, h between surfaces of the ground and the delta wing normalized with the chord length of the double delta wing, c, was selected as h/c=0.1, 0.25 0.55, and the ground-free case (GFC). The results demonstrated that the magnitudes of the strake and wing vortices at α=15° and the coiled-up vortex at α=25° progressively diminish when the delta wing moves from the free-stream region into the boundary layer region close to the ground surface, h/c=0.1. Also, the rate of the inboard movement of leading edge vortices towards the wing central axis increases on the pressure surface of the wing as the altitude of the double delta wing is gradually lowered towards the ground surface. The lift coefficient, CL of the delta wing enhances due to the development of air stagnation and ram pressure between wing and ground surfaces by lowering the level of the double delta wing from the free-stream region into the boundary layer flow region which is strongly influenced by the presence of the ground surface. Also, the drag coefficient, CD increases more than the lift coefficient, CL resulting in a deterioration of the lift to drag ratio, CL/CD under the ground effect (GE). Finally, an increment in CL/CD is much more prominent at lower angles of attack, α regardless of ground effect.

Preset angle, aspect ratio, and stall representations effects on the aeroelastic responses of small-scale wind turbine blades

This work provides an inclusive nonlinear aeroelastic study on the effects of horizontal axis Small-Scale Wind Turbines (SSWTs)’s preset angle of attack as well as aspect ratio on its aerodynamic stability. For this purpose, an enhanced nonlinear reduced-order model is established and a state–space formulation for the aerodynamics of a comprehensive wing-based aeroelastic system is constructed. This representation captures the influence of the preset angle of attack where the blade no longer displays an oscillatory motion around its zero-equilibrium point in the post-flutter region, instead, a new equilibrium point is introduced to the system through the incorporation of the preset angle. The effect of the preset angle on the blade stability is investigated by jointly performing linear analyses that provide insights on this parameter’s influence on the onset of flutter, accompanied with nonlinear analyses that evaluate the latter design parameter’s effect on the blade bending and torsion amplitudes in the post-flutter regime. Additionally, the Duhamel’s principle specific to low angles of attack is extended to represent unsteady flows with non-conventional lift curves of low aspect ratio blades while considering different representations of the aerodynamic stall. Results show the significant effect of the preset angle on both the onset of flutter as well as the post flutter amplitudes where the higher the preset angle, the higher the flutter speed. Furthermore, this research effort also demonstrates the importance of accurate lift modeling since ill-evaluated lift characteristics leads to a potential over- or under-estimation of the blade’s torsional and bending amplitudes post-flutter, and the conventional 2π slope assumption frequently used in various reduced-order-models within the literature seldom predicts erroneous onset of flutter values as well as inaccurate amplitudes in the post-flutter regime especially for the case of low aspect-ratio systems.

Streamline Effect Improvement of Additive Manufactured Airfoil Utilizing Dynamic Stream Control Procedure

In the era of fast transport, to create inventive stream flow management solutions that are capable of diminishing the aerodynamic drag of the vehicles, there is a need to modify the flow characteristics over the vehicle by deferring or expelling the position of the flow partition. The objective of this study involves the parameterized design of an airfoil utilizing the Bezier curve technique with the assistance of the simulation program. For flow regulations, synthetic jet modules are ingrained at different percentages of the chord to manage the stall characteristics. The parametrization system, combined with the stream control method, can give a much better insight into flow re-energization and pave some way for the reduction of the wake. Digital fabrication technique (3d printing or Rapid Prototyping method) is used to fabricate the end product for aerodynamic testing. The comparative outcome showed a reduction in drag at certain angles of attack due to the surface finish obtained. By comparing the results, the aerodynamic efficiency showed a significant rise of 13.05% at lower angles of attack when compressed gas was used in the synthetic jet closer to the frontier edge of the airfoil. Near the stall angle of attack, the coefficient of lift (Cl) and coefficient of drag (Cd) values showed no progress.

Generation of a wingtip vortex pair using a pressure-side fin

We present results from a combined experimental and numerical investigation of the flow around a rectangular NACA 0012 wing, whose tip geometry is modified by the addition of a perpendicular fin. The purpose of this modification is to generate a system of two closely-spaced concentrated tip vortices, instead of the familiar single trailing vortex. Experiments were carried out in a water tunnel, where velocity measurements were made in the near wake of the wing using Stereoscopic Particle Image Velocimetry, in order to determine the characteristics of the tip vortex pair. Various configurations were tested by changing the position, dimensions and angle of attack of the fin. It was found that the fin could generate a strong and concentrated secondary vortex only when placed on the pressure side of the wing. Visualisations of surface streamlines using oil painting in a wind tunnel, as well as numerical simulations of the flow in the wingtip region, show that separation occurs with the fin on the suction side, even at low angles of attack. Both experiments and simulations were used to determine the influence of fin parameter variations on the tip vortex pair, with respect to its circulation ratio and separation distance. Whereas co-rotating vortex pairs of similar circulation could be generated easily, the fin's ability to produce counter-rotating secondary vortices is rather limited. The current results have relevance in the context of reducing the impact of concentrated tip vortices, which for rotor configurations can lead to strong blade-vortex interactions, causing fatigue and aerodynamic noise.

Technique Used to Perform Rocket-Plane Free-Flight Tests

Abstract This paper includes description of the technique that was applied for free-flight (drop) tests of the rocket-plane scaled model. The main aim of the experiment was to validate the numerical approach to be used to simulate the gliding flight of the rocket-plane, especially the transition between high to low angles of attack and the rocket-plane response to control. The primary goal of this paper is to show what kind of challenges must be addressed when planning the flight test campaign. This paper includes description of how the rocket-plane model was scaled and built, the model preparation, experimental design and flight procedure. This paper shows an overview of how the experiment can be planned for different scenarios and the lessons learned during the deep stall free-flight tests.

The control mechanism of the soft trailing fringe on the flow characteristics over an airfoil

Inspired by the owl’s silent flight, we experimentally investigated the flow control mechanism of the soft trailing fringes (STFs) on the wake of the S833 airfoil at the Reynolds number of Re = 2 × 104. A high-speed Particle Image Velocimetry (PIV) system is employed to visualize and analyze the flow structures in the wake of the airfoil at different angles of attack (AOA). Furthermore, spectral proper orthogonal decomposition and bispectral mode decomposition are carried out to identify the coherent flow structures and reveal the control mechanism from the perspective of simplified models. PIV measurements’ results demonstrate that the STFs evidently suppress the turbulent quantities including turbulent kinetic energy and Reynolds shear stress in the airfoil wake. On the one hand, the STFs at low AOAs prevent the interaction between the upper and lower shear layers, and the leading- and trailing-edge vortices (TEVs) are significantly suppressed, thus destructing the von Karman vortex streets. On the other hand, the STFs at high AOAs divide the lower shear layer into two parts, markedly attenuating the TEVs and modifying the vortical structures in the wake. Besides, the quadrant analysis reveals that the STFs can mitigate the high-amplitude wall-pressure peaks, indicating that the STFs may manipulate the trailing-edge noise. However, the control effect is limited at median AOAs because the region with high triadic interactions moves upward in the interaction maps, which limits the impact of the STFs.