Low-speed Wind Tunnel Experiments Research Articles

Experiment investigation on secondary flow loss and flow control mechanisms in a variable compressor cascade with penny cavities

Variable Stator Vanes (VSVs) with penny cavities are an important method to improve the efficiency of compressors. However, the annular and radial gaps in VSVs introduce complex flow structures, posing challenges in enhancing compressor efficiency. This article employs low-speed wind tunnel experiments to study the impact mechanisms of VSVs' adjustment angle and radial gap size on secondary flow losses and implements single-hole suction as a means of flow control to reduce total pressure loss. The results indicate that radial gap leakage has a certain inhibitory effect on annular gap leakage at general adjustment angles. Total pressure loss decreases initially as the radial gap increases but once the radial gap further enlarges and dominates the loss components, the total pressure loss increases proportionally with the radial gap. At extreme adjustment angles, where the lateral pressure gradient is significant, radial gap leakage intensity is high and comprises the major part of the loss, leading to an increase in total pressure loss in proportion to radial gap enlargement. Additionally, single-hole suction is effective under various conditions, with the potential to reduce total pressure loss by up to 19.4 %. These findings provide guidance for the application of VSVs in further improving compressor efficiency.

Experimental study on pulsed suction under different momentum coefficients in variable compressor cascade with annular gaps

Variable Stator Vanes (VSVs) with penny cavities are crucial in the compressors of variable cycle engines, where complex interactions between annular and radial gaps can adversely affect aerodynamic performance. Unsteady active flow control techniques have significant potential for reducing secondary flow losses in compressors. This research investigates the internal flow characteristics of a variable compressor cascade with penny cavities using low-speed wind tunnel experiments and flow visualization techniques. This study is the first to experimentally explore the relationship between loss reduction and momentum coefficient using pulsed suction (PS) compared to steady continuous suction (SCS), evaluating the effects of various excitation frequencies. The findings show that at a momentum coefficient of 0.11 and an excitation frequency of 100 Hz, PS achieves optimal aerodynamic performance, reducing total pressure losses by 11.59 %, marking a 92.59 % improvement over SCS. As the momentum coefficient increases, the effectiveness of PS control significantly improves, whereas the changes in SCS effectiveness are less pronounced. The control effects of PS under varying excitation frequencies are complex and influenced by multiple factors. At higher momentum coefficients, the loss reduction effectiveness of PS is strongly correlated with excitation frequency, while at lower coefficients, performance changes due to frequency variations are minimal. This study provides valuable guidance for the design of modern multistage variable compressors.

Experimental design for a novel co-flow jet airfoil

The Co-flow Jet (CFJ) technology holds significant promise for enhancing aerodynamic efficiency and furthering decarbonization in the evolving landscape of air transportation. The aim of this study is to empirically validate an optimized CFJ airfoil through low-speed wind tunnel experiments. The CFJ airfoil is structured in a tri-sectional design, consisting of one experimental segment and two stationary segments. A support rod penetrates the airfoil, fulfilling dual roles: it not only maintains the structural integrity of the overall model but also enables the direct measurement of aerodynamic forces on the test section of the CFJ airfoil within a two-dimensional wind tunnel. In parallel, the stationary segments are designed to effectively minimize the interference from the lateral tunnel walls. The experimental results are compared with numerical simulations, specifically focusing on aerodynamic parameters and flow field distribution. The findings reveal that the experimental framework employed is highly effective in characterizing the aerodynamic behavior of the CFJ airfoil, showing strong agreement with the simulation data.

A wind tunnel investigation of the unsteady aerodynamic characteristics of bionic flapping wings with three-dimensional complex motion

The bionic flapping wing aircraft realizes flight by imitating the structure and flapping mode of birds. In this paper, a three-dimensional composite motion aerodynamic analysis model of a bionic flapping wing is established. The diverting field and flapping wing grid are divided up using dynamic hybrid grid technology. The flow field of the flapping wing is analyzed by solving the Navier–Stokes (N–S) equation combined with the Boussinesq hypothesis. The lift of the flapping wing under different flutter frequencies and incoming wind speeds is studied under the asymmetric flutter condition. In order to verify the accuracy of the aerodynamic simulation results, a flapping-wing shrinking-ratio model prototype is made, and low-speed wind tunnel experiments are carried out to test the changes in the flight lift of the wing at different flutter frequencies and angles of attack. A comparative analysis of the wind tunnel experiment and the aerodynamic simulation results shows that when the flapping frequency (1–5 Hz) and incoming wind speed (1–5 m/s) increase, the lift force generated by the wing flapping increases. Due to the deviation between the experimental sample airfoil area and the simulated airfoil area, as well as the wing-driven fuselage vibration during the experiment, a sensor error is produced, resulting in a deviation of about 1 N between the experimental result curve and the simulation result curve. However, the aerodynamic characteristics obtained from the aerodynamic simulation analysis are basically consistent with the aerodynamic change law measured in the wind tunnel experiment.

Experimental investigation of hole-type bleed for active flow control in an axial compressor cascade with tip clearance

Boundary layer aspiration using bleeding is a potential technique for active flow control. This study explores the influence mechanism of hole-type bleed on the secondary flow loss through a low-speed wind tunnel experiment. Compared with the suction slot, the hole-type bleed for boundary layer aspiration has the advantages of good flexibility and high strength. In this study, the tip leakage vortex (TLV) and passage vortex (PV) in a compressor cascade with a tip clearance are effectively controlled through the hole-type bleed experimentally. The optimal bleeding hole position is located at 25% cx downstream of the blade leading edge near the blade suction surface. An 8.34% performance benefit in the Cpt with a bleed mass flow rate of 0.66% is obtained. It is found that the optimal bleeding hole position is highly correlated to the position where the TLV and PV originate in the baseline. Moreover, the balance between the controlling of the TLV and PV is considered to strongly determine the performance benefit obtained in the compressor cascade. The findings reveal that the secondary flow loss control mechanism under the hole-type bleed from two aspects: delayed vortex origination and clearance blockage.

Aerodynamic Performance of an Aircraft with Aft-Fuselage Boundary-Layer-Ingestion Propulsion

Boundary-layer ingestion (BLI) is a propulsor–airframe integration technology that promises substantial fuel consumption benefits for future civil aircraft. This paper discusses an experimental study, conducted within the European Union–funded Horizon 2020 CENTRELINE project, on the aerodynamic performance of an aircraft with a BLI propulsor integrated at the aft-fuselage section (known as the Propulsive Fuselage Concept). The low-speed wind-tunnel experiments were carried out at Reynolds and Mach numbers of 460,000 and 0.12, whereas the Reynolds and Mach numbers are 40,000,000 and 0.82 at full-flight scale. Aerodynamic loads measurements show that the BLI propulsor affects the longitudinal and lateral-directional equilibrium of the aircraft in off-cruise conditions. Moreover, velocity and total pressure measurements characterize the flowfield around the BLI propulsor in cruise and off-cruise conditions. The analysis of the momentum and power fluxes in the flowfield shows that, while around 20% of the total aircraft drag is due to the fuselage body, only less than 5% of the total aircraft drag power is dissipated in the fuselage wake. Furthermore, the BLI propulsor recovers around 50% the axial kinetic energy flux in the fuselage boundary layer (the so-called wake-filling effect), suggesting an increased propulsive efficiency.

Low speed aerodynamic characteristics of non-slender delta wing at low angles of attack

Low-speed wind tunnel experiments are conducted to study the aerodynamic performance of a half-span delta wing with 45° leading-edge sweep at subsonic flow regime. The experiments are carried out at a Reynolds number of 8.37 × 105, a free-stream Mach number of 0.1 and angles of attack up to 25°, in steps of 5°. The test model was designed with thirty-two pressure taps fixed on its surfaces (sixteen on each side). Multi-tube manometers were connected to these taps using long tubes to enable recording the pressure readings. Surface pressure distributions and aerodynamic characteristics were calculated at different span-wise locations along the non-dimensional chord-wise distance. Results exhibited that most lift on the studied wing is generated in the region close to the leading edge for all the studied incidence angles. Additional lift is created in the region close to the root chord rather than the tip chord, whereas drag forces increases from tip to root. This can be attributed to the formation of trailing edge vortexes due to the flow separation at the wing leading edge that produces more drag, hence suppressing lift. The study showed also that angle of attack increases the drag coefficient from tip to root, especially at high angle of attack, indicating unfavourable behaviour for manoeuvring. Moreover, the angle of attack increased the pitching moment coefficient up to 10° before it drops sharply until it reaches the tip of the wing model.

Experimental study on an airfoil equipped with an active flow control element

This paper addresses the torque ripple problem of vertical axis wind turbines (VAWT) by applying an oscillating bump which was previously optimised with a numerical model. Since the element primarily affects the drag of an airfoil, a low-speed wind-tunnel experiment using the wake survey method was conducted in order to highlight the effect on the drag of a symmetrical airfoil. Measurements were taken along the centreline of the foil to compare with numerically simulated results on a 2D foil. This study provides an experimental data set for an active flow control device. When the stationary results were compared against numerical simulation they showed an overall good agreement. The moving bump increased the drag coefficient for higher frequencies of oscillation and induced less drag compared to the steady maximum distortion. Uncertainties in the experiments were primarily caused by fluctuations in the test section and the data reduction error. Future work should include the measurement of lift in order to determine the influence of the element on the torque.

Computational Simulations of Fluidic Actuation on Rotor Blades and Their Experimental Validation

Active flow control (AFC) is conventionally used to reduce flow separation at high angles of attack. A novel alternative form of AFC for controlling rotor blade loads in attached flow is examined. The AFC system employs pulsed fluidic jets near the trailing edge of a VR-12 airfoil on the pressure side and suction side. Each jet has two operating modes: tangential blowing and normal blowing. Unsteady Reynolds-averaged Navier–Stokes simulations are performed to characterize the unsteady aerodynamic phenomena induced by actuation at low and moderate angles of attack. The simulation approach is validated using data from recent low-speed () wind-tunnel experiments. The normal blowing modes produce nearly symmetric changes in the lift and moment that are proportional to the relative jet strength. The tangential blowing modes have a minor effect on lift and moment but have a beneficial effect on drag. High-speed simulations, up to , are also performed to determine the effect of freestream Mach number on actuator performance. Finally, simulations representing rotor blade dynamics during vibrations are carried out to illustrate the effect of unsteadiness associated with blade motion. This is the first instance showing that AFC is a viable approach for potentially controlling rotor vibrations in attached flow.

Effects of wing flexibility on aerodynamic performance of an aircraft model

The low-speed wind tunnel experiment is carried out on a simplified aircraft model to explore the influence of wing flexibility on the aircraft aerodynamic performance. The investigation involves the measurements of force, membrane deformation and velocity field at Reynolds number of 5.4 × 104–1.1 × 105. In the lift curves, two peaks are observed. The first peak, corresponding to the stall, is sensitive to the wing flexibility much more than the second peak, which nearly keeps constant. For the optimal case, in comparison with the rigid wing model, the delayed stall of nearly 5° is achieved, and the relative lift increment is about 90%. It is revealed that the lift enhanced region corresponds to the larger deformation and stronger vibration, which leads to stronger flow mixing near the flexible wing surface. Thereby, the leading-edge separation is suppressed, and the aerodynamic performance is improved significantly. Furthermore, the effects of sweep angle and Reynolds number on the aerodynamic characteristics of flexible wing are also presented.

Horizontal Wind Effect on the Aerodynamic Performance of Coaxial Tri-Rotor MAV

The coaxial Tri-rotor micro air vehicle (MAV) is composed of three coaxial rotors where the aerodynamic characteristics of is complicated in flight especially when the wind effect is introduced. In this paper, the hovering performance of a full-scale coaxial Tri-rotor MAV is analyzed with both the simulations and wind tunnel experiments. Firstly, the wind effect on the aerodynamic performance of coaxial Tri-rotor MAV is established with different rotor speed (1500–2300 rpm) and horizontal wind (0–4 m/s). Secondly, the thrust and power consumption of coaxial Tri-rotor (L/D = 1.6) were obtained with low-speed wind tunnel experiments. Furthermore, the streamline distribution, pressure distribution, velocity contour and vortex distribution with different horizontal wind conditions are obtained by numerical simulations. Finally, combining the experiment results and simulation results, it is noted that the horizontal wind may accelerate the aerodynamic coupling, which resulting in the greater thrust variation up to 9% of the coaxial Tri-rotor MAV at a lower rotor speed. Moreover, the aerodynamic performance is decreased with more power consumption at higher rotor speed where the wind and the downwash flow are interacted with each other. Compared with no wind flow, the shape of the downwash flow and the deformation of the vortex affect the power loading and figure of metric accordingly.

Wind loading of a finite prism: aspect ratio, incidence and boundary layer thickness effects

A systematic set of low-speed wind tunnel experiments was performed at Re = 6.5x104 and 1.1x105 to study the mean wind loading experienced by surface-mounted finite-height square prisms for different aspect ratios, incidence angles, and boundary layer thicknesses. The aspect ratio of the prism was varied from AR = 1 to 11 in small increments and the incidence angle was changed from α = 0° to 45° in increments of 1°. Two different boundary layer thicknesses were used: a thin boundary layer with δ/D = 0.8 and a thick boundary layer with δ/D = 2.0–2.2. The mean drag and lift coefficients were strong functions of AR, α, and δ/D, while the Strouhal number was mostly influenced by α. The critical incidence angle, at which the prism experiences minimum drag, maximum lift, and highest vortex shedding frequency, increased with AR, converged to a value of αc = 18° ± 2° once AR was sufficiently high, and was relatively insensitive to changes in δ/D. A local maximum value of mean drag coefficient was identified for higher-AR prisms at low α. The overall behaviour of the force coefficients and Strouhal number with AR suggests the possibility of three flow regimes.

Yaw control of a flying-wing unmanned aerial vehicle based on reverse jet control

Due to its layout, there are difficulties in realizing heading attitude control of a flying-wing unmanned aerial vehicle. In this paper, a reverse jet control scheme has been designed: (1) to replace the resistance rudders that are used for the yaw control of a conventional flying-wing unmanned aerial vehicle, (2) to assist and optimize heading attitude control, eliminate the adverse effects of the control surface and enhance stealth performance, and (3) to promote the use of rudderless flight for flying-wing unmanned aerial vehicles. To explore the control mechanism and the flow field of the reverse jet scheme, three-dimensional numerical simulations and low-speed wind tunnel experiments were carried out. First, the numerical simulations evaluated the feasibility and effectiveness of the reverse jet control scheme and explored and optimized the excitation parameters for the scheme. Then the forces were measured in a wind tunnel, and particle image velocimetry experiments were carried out. A reverse jet actuator was independently designed to verify the results of the numerical simulation. The results show that when the reverse jet excitation is applied, the jet obstructs the mainstream, destroys the flow field at the excitation position, and causes early separation of the flow, which increases the pressure drag on the wings and produces a control effect. The control effect mainly depends on the separation degree of the leeward surface. The larger the jet momentum coefficient is, the smaller the jet angle is, and the closer the excitation position is to the leading edge, the greater the separation degree of the leeward surface is, the better the heading attitude control effect is.

Aerodynamic Performance of Hex-Rotor UAV Considering the Horizontal Airflow

In this paper, the aerodynamic performance of a Hex-rotor unmanned aerial vehicle (UAV) with different rotational speeds (1500–2300 RPM) considering the horizontal airflow conditions is analyzed by both simulations and experiments. A low-speed wind tunnel experiments platform is applied to measure the thrust, torque, and power consumption of a Hex-rotor UAV with different rotational speeds in horizontal airflow, which varied from 0 m/s–4 m/s. First, this paper introduces the effect of horizontal airflow on a UAV. Then, the low-speed wind tunnel experiments were carried out on a Hex-rotor UAV (D/L = 0.56) with different horizontal velocities to determine the hover performance. Finally, numerical simulations were obtained with the streamline distributions, pressure distributions, velocity contour, and vortex distributions at different horizontal airflow conditions to describe the aerodynamic interference effect of different horizontal airflows. Combined with the experimental results and numerical simulations results, the horizontal airflow proved to have a significant influence on the aerodynamic performance of the Hex-rotor UAV with an increase in thrust and power. Indeed, the streamlines in the flow field were coupled to each other at the presence of the incoming airflow. Especially when the incoming airflow was larger, the Hex-rotor UAV could properly use low-speed flight to maintain high power loading. Finally, it is inferred that the aerodynamic performance of the Hex-rotor UAV is also related to the movement and deformation of the vortex at the tip of the rotor.

Alleviation of Propeller-Slipstream-Induced Unsteady Pylon Loading by a Flow-Permeable Leading Edge

The impingement of a propeller slipstream on a downstream surface causes unsteady loading, which may lead to vibrations responsible for structure-borne noise. A low-speed wind-tunnel experiment was performed to quantify the potential of a flow-permeable leading edge to alleviate the slipstream-induced unsteady loading. For this purpose, a tractor propeller was installed at the tip of a pylon featuring a replaceable leading-edge insert in the region of slipstream impingement. Tests were carried out with four flow-permeable inserts, with different hole diameters and cavity depths, and a baseline solid insert. Particle-image-velocimetry measurements showed that the flow through the permeable surface caused an increase in boundary-layer thickness on the pylon’s suction side. This led to a local drag increase and reduced lift, especially for angles of attack above 6 deg. Furthermore, it amplified the viscous interaction with the propeller tip-vortex cores, reducing the velocity fluctuations near the pylon surface by up to 35%. Consequently, lower tonal noise emissions from the pylon were measured in the far field. This suggests that the desired reduction in surface pressure fluctuations was achieved by application of the flow-permeable leading edge.

Yaw-control efficiency analysis for a diamond wing configuration with outboard split flaps

The yaw-control device of a low-aspect ratio flying wing with diamond-shaped wing planform is investigated. Extensive low-speed wind tunnel experiments have been carried out to obtain surface pressure data and the aerodynamic forces and moments of the configuration for six different flap deflection angles at varying angles of attack and sideslip. Complementary unsteady Reynolds-averaged Navier–Stokes simulations are performed for selected configurations. The experimental data is used to examine the validity of the numerical results. The analysis is focused on the aerodynamic coefficients and derivatives. Yaw-control effectiveness, yaw-control efficiency, crosswind landing capabilities and coupling effects are discussed. The results show sufficient yaw-control effectiveness and efficiency for a wide range of considered freestream conditions. The outboard flap exhibits a non-linear characteristic with respect to the flap deflection angle and freestream conditions. The efficiency is considerably reduced at high angles of attack due to large-scale flow separation in the wing outboard section. Non-linear coupling effects with the rolling moment become obvious for moderate to large flap deflections over the whole angle of attack polar. The numerical results show good agreement with the experimental data in the surface pressure distributions and longitudinal aerodynamic coefficients. The yawing moment is overpredicted by numerical simulations for large flap deflection angles.

Velocity field measurements in the wake of a propeller model

Turboprop configurations are being revisited for the modern-day regional transport aircrafts for their fuel efficiency. The use of laminar flow wings is an effort in this direction. One way to further improve their efficiency is by optimizing the flow over the wing in the propeller wake. Previous studies have focused on improving the gross aerodynamic characteristics of the wing. It is known that the propeller slipstream causes early transition of the boundary layer on the wing. However, an optimized design of the propeller and wing combination could delay this transition and decrease the skin friction drag. Such a wing design would require the detailed knowledge of the development of the slipstream in isolated conditions. There are very few studies in the literature addressing the requirements of transport aircraft having six-bladed propeller and cruising at a high propeller advance ratio. Low-speed wind tunnel experiments have been conducted on a powered propeller model in isolated conditions, measuring the velocity field in the vertical plane behind the propeller using two-component hot-wire anemometry. The data obtained clearly resolved the mean velocity, the turbulence, the ensemble phase averages and the structure and development of the tip vortex. The turbulence in the slipstream showed that transition could be close to the leading edge of the wing, making it a fine case for optimization. The development of the wake with distance shows some interesting flow features, and the data are valuable for flow computation and optimization.

Experimental study of surface curvature effects on aerodynamic performance of a low Reynolds number airfoil for use in small wind turbines

This paper presents the wind tunnel experimental results to investigate the effects of surface gradient-of-curvature on aerodynamic performance of a low Reynolds number airfoil Eppler 387 for use in small-scale wind turbines. The prescribed surface curvature distribution blade design method is applied to the airfoil E387 to remove the gradient-of-curvature discontinuities and the redesigned airfoil is denoted as A7. Both airfoils are manufactured with high precision to reflect the design. Low-speed wind tunnel experiments are conducted to both airfoils at chord based Reynolds numbers 100 000, 200 000, and 300 000. Surface pressure measurements are used to calculate the lift and pitching-moment data, and the wake survey method is applied to obtain the drag data. The experimental results of E387 are compared with NASA Low Turbulence Pressure Tunnel (LTPT) results for validation. The gradient-of-curvature discontinuities of E387 result in a larger laminar separation bubble which causes higher drag at lower angles of attack. As the angle of attack increases the separation bubble of the airfoil E387 moves faster towards the leading edge than that of A7, resulting in a premature bubble bursting and earlier stall on E387. The impact of the gradient-of-curvature distribution on the airfoil performance is more profound at higher angles of attack and lower Reynolds number. The aerodynamic improvements are integrated over the 3D geometry of a 3 kW small wind turbine, resulting in up to 10% increase in instantaneous power and 1.6% increase in annual energy production. It is experimentally concluded that an improved curvature distribution results in a better airfoil performance, leading to higher energy output efficiency.

An Auto-Tuning PI Control System for an Open-Circuit Low-Speed Wind Tunnel Designed for Greenhouse Technology.

Wind tunnels are a key experimental tool for the analysis of airflow parameters in many fields of application. Despite their great potential impact on agricultural research, few contributions have dealt with the development of automatic control systems for wind tunnels in the field of greenhouse technology. The objective of this paper is to present an automatic control system that provides precision and speed of measurement, as well as efficient data processing in low-speed wind tunnel experiments for greenhouse engineering applications. The system is based on an algorithm that identifies the system model and calculates the optimum PI controller. The validation of the system was performed on a cellulose evaporative cooling pad and on insect-proof screens to assess its response to perturbations. The control system provided an accuracy of <0.06 m·s−1 for airflow speed and <0.50 Pa for pressure drop, thus permitting the reproducibility and standardization of the tests. The proposed control system also incorporates a fully-integrated software unit that manages the tests in terms of airflow speed and pressure drop set points.

Combustion-Powered Actuation for Dynamic-Stall Suppression: High-Mach Simulations and Low-Mach Experiments

An investigation on the dynamic-stall-suppression capabilities of combustion-powered actuation applied to a tabbed VR-12 airfoil is presented. In the first section, the results from the computational fluid dynamics simulations carried out at Mach numbers of 0.3 and 0.4 are presented. Several geometric parameters are varied, including the slot chordwise location and angle. Actuation-pulse amplitude, frequency, and timing are also varied. The two-dimensional simulations suggest that cycle-averaged lift increases of approximately 4 and 8% with respect to the baseline airfoil are possible at Mach numbers of 0.4 and 0.3 for deep and near-deep dynamic-stall conditions. In the second section, the static-stall results from low-speed wind-tunnel experiments are presented. The low-speed experiments and high-speed computational fluid dynamics suggest that slots oriented tangential to the airfoil surface produce stronger benefits than slots oriented normal to the chord line. The low-speed experiments confirm that chordwise slot locations suitable for Mach 0.3–0.4 stall suppression (based on computational fluid dynamics) will also be effective at lower Mach numbers.