Delta Wing Model Research Articles

Data-Driven Sparse Sensor Placement Optimization on Wings for Flight-By-Feel: Bioinspired Approach and Application.

Flight-by-feel (FBF) is an approach to flight control that uses dispersed sensors on the wings of aircraft to detect flight state. While biological FBF systems, such as the wings of insects, often contain hundreds of strain and flow sensors, artificial systems are highly constrained by size, weight, and power (SWaP) considerations, especially for small aircraft. An optimization approach is needed to determine how many sensors are required and where they should be placed on the wing. Airflow fields can be highly nonlinear, and many local minima exist for sensor placement, meaning conventional optimization techniques are unreliable for this application. The Sparse Sensor Placement Optimization for Prediction (SSPOP) algorithm extracts information from a dense array of flow data using singular value decomposition and linear discriminant analysis, thereby identifying the most information-rich sparse subset of sensor locations. In this research, the SSPOP algorithm is evaluated for the placement of artificial hair sensors on a 3D delta wing model with a 45° sweep angle and a blunt leading edge. The sensor placement solution, or design point (DP), is shown to rank within the top one percent of all possible solutions by root mean square error in angle of attack prediction. This research is the first to evaluate SSPOP on a 3D model and the first to include variable length hairs for variable velocity sensitivity. A comparison of SSPOP against conventional greedy search and gradient-based optimization shows that SSPOP DP ranks nearest to optimal in over 90 percent of models and is far more robust to model variation. The successful application of SSPOP in complex 3D flows paves the way for experimental sensor placement optimization for artificial hair-cell airflow sensors and is a major step toward biomimetic flight-by-feel.

Analysis of stall characteristics of the cuspidal waverider at subsonic conditions

The stall characteristics of the cuspidal waverider under subsonic conditions were studied at M = 0.4 and H = 0 km incoming flow. The stall characteristics were compared with those of delta wing models with the same leading-edge swept angle and the same aspect ratio. It can be found that the flow field changes of the cuspidal waverider are similar to those of Model 1 and Model 2. The critical angle of attack of the cuspidal waverider is more influenced by the aspect ratio, but less related to the leading-edge swept angle. The lift change of the cuspidal waverider is closely related to the vortex structure on the leeward side. As the angle of attack increases, the vortex structure becomes thicker and stronger, and moves toward the middle of the model under the action of the overflowing high pressure gas on the lower surface. Under the same angle of attack, the vortex structure under the effect of the adverse pressure gradient, the intensity along the flow direction decreases while the vortex continues to move away from the wall. Besides, the possible splitting of large vortex into smaller ones will be another result of the effect of the inverse pressure gradient. The smaller the overall aspect ratio of the model, the greater actual swept angle and the greater effective swept angel of the vortex structure, which allows the vortex to exist at higher angles of attack. When the vortex structure is broken to form a larger recirculation area in the leeward area, the negative pressure at the leeward surface decreases, the lift coefficient decreases, and a stall occurs.

Shock migration on an oscillating straked delta wing using an unsteady Euler solver

Many investigations, including flight tests, wind tunnel tests, and computational methods have attempted to determine the drivers of limit cycle oscillations (LCO) on the F-16 fighter aircraft. One area of interest has focused on shock-induced trailing-edge separation (SITES) as a driver of limit cycle oscillations. The investigation presented herein aims to expand the understanding of unsteady, nonlinear aerodynamics by studying the flow over a wing undergoing pitch oscillations using Euler computational fluid dynamics with a boundary layer coupling scheme capable of estimating viscous flow effects within the boundary layer. A straked NACA 64A204 delta wing model was used for this analysis. Parameters such as trim angle of attack, pitch amplitude, and Mach number were varied to create 71 unique test cases. Results demonstrate that the same single shock systems at low angles of attack and two shock systems at moderate to high angles of attack that were observed by Cunningham in his wind tunnel tests were also present in this investigation. Patterns of significant shock movement and disappearance were observed within the transonic flow region. Hysteresis of the pitching moment was also found to be present. These findings support the theories that shock movement and SITES may play significant roles in driving LCO.

Delta wing rocking motion suppression by deep reinforcement learning

The wing rock phenomenon is a self-excited roll motion occurring on the aircraft at high angles of attack (AoAs), which negatively impact safety and maneuverability. The limit cycle oscillation is a typical characteristic of this self-excited roll motion. A controller employing a model-free training approach is constructed on the basis of deep reinforcement learning (DRL) to address this severe nonlinear control problem. For an 80° swept delta wing model, the analytical model describing the nonlinear behavior of wing rock is presented and implemented in the simulation environment. The DRL-based controller is trained with the proximal policy optimization algorithm. A reward function is carefully designed for training to achieve excellent convergence stability. Various simulations were conducted at a series of unlearned initial conditions (roll angles, AoAs) to demonstrate the effectiveness and generalization capability of the proposed method and verify and evaluate the performance of the trained DRL controller. Finally, the scenario that undergoes disturbance is added to confirm the effectiveness and robustness of the proposed DRL-based controller. Results show that the DRL-based controller could effectively regulate the oscillation and retain the capability to suppress the un-wanted behavior with favorable generalization capability and robustness. Therefore, delta wing rocking motion suppression based on DRL has powerful intelligent control performance and provides a new idea for predicting the evolution of aerodynamics in rocking motion suppression.

Bioinspired Aerodynamic Load Estimation for UCAVs in Turbulent Environments

Small aircraft carriers play a crucial role in military navigation and field intelligence. Unmanned Combat Aerial Vehicles (UCAVs) have emerged as indispensable tools for various military operations. However, the efficacy of UCAV operations is often hampered by the turbulent conditions generated by oceanic gusts. To enhance the reliability and autonomy of these systems, this research focuses on advancing aerodynamic load estimation techniques.
 Many biological systems employ sensory systems to collect environmental feedback to improve predation and avoid attack. Examples include the proprioception exhibited in bird feathers and the lateral line function in fish. Drawing inspiration from these biological models, a system of sensors is employed to achieve more precise estimations of the UCAV aerodynamic state and therefore train an existing Multi-Layer Perception Artificial Neural Network (MLP ANN) to respond to the environment.
 A series of experiments were conducted on a non-slender delta-wing model of NACA 0012 mounted on a force/moment sensor and exposed to varying flow conditions in a controlled setup called the WindShaper. To simulate gusty environments, the WindShaper produced unsteady random axial gusts ranging from 5m/s to 15m/s. A total of 100 sets of data, consisting of 60000 data points were collected and used to train the MLP ANN.
 Statistical methods were employed to distinguish valid and edge cases within the datasets. From analyzing the outputs generated by the MLP ANN, it was determined that there exists an importance in collecting data of high quality. These analyses identify factors affecting model performance and highlight specific challenges that hinder accurate gust prediction. The utilization of advanced neural network methodologies as a technique for load estimation has significant implications for naval aviation, aerodynamics, and their related autonomous systems. This research contributes to the development of more robust UCAV operations in challenging environments.

ANALYSIS OF VORTEX ON SHARP-EDGED DELTA WING WITH BLOWING EFFECT

The formation of the primary vortex above the wing is dependent on the Reynolds number, the angle of attack, the Mach number, and the bluntness of the leading edge. This experiment is done in order to assess the impacts of blowing at three sites on the delta wing's flow topology. The blower was placed on the leading edge at the 15%, 50%, and 70% wing sections. The experiment was conducted at a Reynolds number of 8 x 105, which corresponds to a wind speed of 25 m/s and a mean aerodynamic chord. The UTM generic delta wing model with a 55° sweep angle was employed, and the experiment was conducted at the Low-Speed Wind Tunnel at UniversitiTeknologi Malaysia (UTM-LST). During each experiment, the model was subjected to surface pressure monitoring techniques. This project's findings show the influence of blowing on the flow topology above the wing.In specific test settings, the installation of the blower at positions 50 percent significantly increased the size of the primary vortex while delaying its breakdown at greater angles of attack.

Comparative Study of Machine Learning Modeling for Unsteady Aerodynamics

Modern fighters are designed to fly at high angle of attacks reaching 90 deg as part of their routine maneuvers. These maneuvers generate complex nonlinear and unsteady aerodynamic loading. In this study, different aerodynamic prediction tools are investigated to achieve a model which is highly accurate, less computational, and provides a stable prediction of associated unsteady aerodynamics that results from high angle of attack maneuvers. These prediction tools include Artificial Neural Networks (ANN) model, Adaptive Neuro Fuzzy Logic Inference System (ANFIS), Fourier model, and Polynomial Classifier Networks (PCN). The main aim of the prediction model is to estimate the pitch moment and the normal force data obtained from forced tests of unsteady delta-winged aircrafts performing high angles of attack maneuvers. The investigation includes three delta wing models with 1, 1.5, and 2 aspect ratios with four determined variables: change rate in angle of attack (0 to 90 deg), non-dimensional pitch rate (0 to .06), and angle of attack. Following a comprehensive analysis of the proposed identification methods, it was found that the newly proposed model of PCN showed the least error in modeling and prediction results. Based on prediction capabilities, it is seen that polynomial networks modeling outperformed ANFIS and ANN for the present nonlinear problem.

Identification of Sharp Edge Non-Slender Delta Wing Aerodynamic Coefficient Using Neural Network

Delta wing formed a vortical flow on its surface which produced higher lift compared to conventional wing. The vortical flow is complex and non-linear which requires more studies to understand its flow physics. However, conventional flow analysis (wind tunnel test and computational flow dynamic) comes with several significant drawbacks. In recent times, application of neural network as alternative to conventional flow analysis has increased. This study is about utilization of Multi-Layer Perceptron (MLP) neural network to predict the coefficient of pressure (Cp ) on a delta wing model. The physical model that was used is a sharp edge non-slender delta wing. The training data was taken from wind tunnel tests. 70% of data is used as training, 15% is used as validation and another 15% is used as test set. The wind tunnel test was done at angle of attack from 0°-18° with increment of 3°. The flow velocity was set at 25m/s which correspond to 800,000 Reynolds number. The inputs are angle of attack and location of pressure tube (y/cr) while the output is Cp . The MLP models were fitted with 3 different transfer functions (linear, sigmoid, and tanh) and trained with Lavenberg-Marquadt backpropagation algorithm. The results of the models were compared to determine the best performing model. Results show that large amount of data is required to produce accurate prediction model because the model suffer from condition called overfitting.

An Experimental Investigation of the Vortex Merging over a Cranked-Delta Wing at Subsonic Speed

A series of experiments was carried out to investigate the aerodynamic characteristics of a cranked-delta wing model. It has been noted by various researches that in the vicinity of the take-off angles of attack, an instability in the longitudinal stability of aircraft equipped with cranked delta wings occurs where its origin is not well understood yet. To further study this phenomenon, a semi-span cranked delta wing model was designed and built. Surface pressure data for various angles of attack at low subsonic speeds, to better simulate take-off and landing conditions, were measured.

Influence of Active Flow Control on Blunt-Edged VFE-2 Delta Wing model

This paper highlights the flow topology above blunt-edged delta wing of VFE-2 configuration when an active flow control technique called ‘blower’ is applied in the leading edge of the wing. The flow topology above blunt-edged delta wing is very complex, disorganised and unresolved compared to sharp-edged wing. For the sharp leading-edged wing, the onset of the primary vortex is fixed at the apex of the wing and develops along the entire wing towards the trailing edge. However, the onset of the primary vortex is no longer fixed at the apex of the wing for the blunt-edged case. The onset of the primary vortex develops at a certain chord-wise position and it moved upstream or downstream depending on Reynolds number, angle of attack, Mach number and the leading-edge bluntness. An active flow control namely ‘blower’ technique has been applied in the leading edge of the wing in order to investigate the upstream/downstream progression of the primary vortex. This research has been carried out in order to determine either the flow on blunt-edged delta wing would behave as the flow above sharp-edged delta wing if any active flow control is applied. The experiments were performed at Reynolds number of 0.5×106, 1.0×106 and 2.0×106 corresponding to 9 m/s, 18 m/s and 36 m/s in UTM Low Speed wind Tunnel based on the mean aerodynamic chord of the wing. The results obtained from this research have shown that the blower technique has significant effects on the flow topology above blunt-edged delta wing. The main observation from this study was that the primary vortex has been shifted 20% upstream when the blower technique is applied. Another main observation was the ability of this flow control to delay the formation of the vortex breakdown.

Dynamic Wind-Tunnel Testing of a Sixty-Degree Delta-Wing Model Without Support Interference

The longitudinal dynamic characteristics of a 60 deg sweptback delta-wing model at angles of attack of up to 21 deg were investigated in this study using a magnetic suspension and balance system. The magnetic suspension and balance system can obtain wind-tunnel data without support interference by suspending a model using magnetic forces. Wind-tunnel tests during accurate pitch oscillations were conducted, and the dynamic longitudinal forces and moment data were obtained. We proposed a method to accurately identify the dynamic pitching moment during pitch motions from our data using a dynamic derivative model. In our model, the dynamic derivative due to the pitch rate and pitch acceleration were considered, and these dynamic derivatives were treated as linear functions of angle of attack. Multiple regression analysis was used to identify the coefficients in the model. Our model can separately describe the static and dynamic part of the total pitching moment. The dynamic pitching moment coefficients were precisely identified by our dynamic derivative model with consideration of the acceleration term for low angles of attack. In addition, the pitch acceleration term could be considered to have the same effect as virtually changing the moment of inertia of the delta-wing model.

Tuft Flow Visualisation on UTM-LST VFE-2 Delta Wing Model Configuration at High Angle of Attacks

This paper reports on flow visualisation and surface pressure measurements over the upper surface of a blunt-edged delta wing model at high angles of attack. The flow structure above the upper surface of the blunt-edged delta wing was found to be different compared to delta wing with sharp leading edge. The flow becomes more complicated especially in the leading edge region of the wing. Currently, there is no data available to verify if the primary vortex could reach the apex of the wing when the angle of attack is further increased. Most prior experiments were performed at the angles of attack, α, below 23° with only a few experiments that had gone to α = 27°. These prior experiments and some CFD works stipulated that the attached flow continue to exist in the apex region of the delta wing even at very high angles of attack above 23°. In order to verify this hypothesis, several experiments at high angles of attack were conducted in Universiti Teknologi Malaysia Low Speed wind Tunnel (UTM–LST), using a specially constructed VFE2 wing model equipped with blunt leading edges. This series of experiments employed two measurement techniques; the first was the long tuft flow visualisation method, followed by surface pressure measurements. The experiments were performed at Reynolds numbers of 1.0×106 and 1.5×106. During these experiments, several interesting flow characteristics were observed at high angles of attack, mainly that the flow became more sensitive to changes in Reynolds number and the angles of attack of the wing. When the Reynolds number increased from 1×106 to 1.5×106, the upstream progression of the initial point of the main vortex was relatively delayed compared to the sharp-edged delta wing. The experiments also showed that the flow continued to be attached in the apex region up to α = 27º.

Wall shear stress measurements by white-light oil-film interferometry

Wall shear stress and wall streamline direction measurements were conducted on a generic delta wing model to demonstrate the potential of the white-light oil-film interferometry in three-dimensional flows. With the proposed algorithm, an automatic oil-film thickness calculation is possible and yields reproducible and plausible wall shear stress distributions on the leeward side of the delta wing for various angles of attack.Graphic abstract

Pulsed Blowing Interacting with a Leading-Edge Vortex

Manipulation of vortex instabilities for aerodynamic performance increase is of great interest in numerous aeronautical applications. With increasing angle of attack, the leading-edge vortex of a semi-slender delta wing becomes unsteady and eventually collapses, endangering the flight stability. Hence, active flow control by pulsed blowing stabilizes the vortex system, enlarging the flight envelope for such wing configurations. The most beneficial outcome is the reattachment of the separated shear layer during post-stall, contributing to a lift increase of more than 50%. In contrast to high power consuming brute-force actuation, manipulating the flow instabilities offers a more efficient alternative for mean flow field control, which has direct repercussions on the aerodynamic characteristics. However, the flow mechanisms involving jet–vortex and vortex–vortex interactions and the disturbance convection through the flow field are little understood. This paper reports on the unsteady flow field above a generic half delta wing model with a 65 ° sweep angle and its response to periodic blowing. Numerical and experimental results are presented and discussed in a synergistic manner.

Effects of the Propeller Advance Ratio on Delta Wing UAV Leading Edge Vortex

Delta wing is a triangular-shaped platform that can be applied into the unmanned aerial vehicle (UAV) or drone applications. However, the flow above the delta wing is governed by complex leading-edge vortex structures which result in complicated aerodynamics behaviour. At higher angles of attack, the vortex burst can take place when the swirling flow is unable to sustain the adverse pressure gradient. More studies are needed to understand these vortex phenomena. This paper addresses an experimental study of active flow control called propeller on a generic 55° swept angle sharp-edged delta wing model. In this experiment, a propeller was placed at two different locations. The first location was at the apex of the wing while the second position was at the rear of the wing. The experiments were conducted in a 1.5 × 2.0 m2 closed-loop wind tunnel facility at Universiti Teknologi Malaysia. The freestream velocities were set at 20 m/s and 25 m/s. The research consisted of an intensive surface pressure measurement above the wing surface to investigate the effects of rotating propeller towards the leading-edge vortex. The experiments were divided into four configurations. The clean wing configuration was performed without the propeller and followed by pusher-propeller configuration using 10-inch 9-inch propellers. The final configuration was the tractor-propeller with a 10-inch propeller. The results emphasise the influences of the propeller size and its location corresponding to vortex properties above the delta-winged UAV model. The findings had indicated that the vortex peak is increased when the propeller is installed for both pusher and tractor configurations. The results also indicate that the pressure coefficient is increased when the propeller advance ratio increases.

Effect of Shapes and Turbulent Inlet Flow to Vortices on Delta Wings

This paper aims to estimate the effect of turbulent inlet flow to vortices on Delta wing with four different turbulence intensity from 0.5% to 15% and the effect of taper ratios to aerodynamic characteristics of Delta wings with four taper ratios: 0, 0.1, 0.2, 0.3, 0.4, 0.5, 0.6, 0.7. The main purpose of this paper is to find out the formation, development, and breakdown of vortices on Delta wings when changing taper ratios and turbulence intensity thence determining the center of vortices with the range of attack angles from 5o to 40o in low velocities about 2.5 m/s. This research uses Delta wing models with a 40o swept-back leading edge, the root chord length 150 mm, and a thickness 5 mm. The problem is simulated by using ANSYS fluent and experiment in the subsonic wind tunnel to compare and validate results. The Delta wing models are meshed by using ICEM to improve the mesh quality and using the turbulence model for low Reynolds number flows Transition SST (4 equations) to calculate aerodynamic characteristics such as lift coefficient, drag coefficient, pressure coefficient... find the paths which connect centers of the vortices, and show the contours of pressures and velocities to evaluate the change of centers of the vortices. The results showed that the two vortices grow up and tend to move inward when the attack angle increase, the vortices are broken strongly in high attack angles, the aerodynamic quality of Delta wings change insignificantly when changing turbulent intensity at inlet. This research also carried out that the stall angle increase when increasing the taper ratio.

Magnetic Suspension of a Moving Aircraft Model at High Angles of Attack

The magnetic suspension and balance system (MSBS) suspends a model without physical support systems. With the use of a pair of two-dimensional CMOS image sensors, a position sensing system that detects the position of the model by dynamically tracking the model attitude was developed with the aim of applying an MSBS for high angles of attack (AOAs). The position sensing system developed in this study was able to detect the position and attitude of the model within the range of AOAs of -40 to 40 degrees. The position sensing system was installed to an existing 60 cm MSBS at JAXA, and static and dynamic wind tunnel tests were conducted to investigate the ability of the MSBS and the dynamic characteristics of a delta wing aircraft model at various AOAs. The MSBS could suspend and oscillate the model in pitch direction at AOA of 15 degrees with the amplitude of 6 degrees in the wind. The delta wing model was dynamically stable at each tested AOAs.

On the high-lift characteristics of a bio-inspired, slotted delta wing

A bio-inspired, slotted delta wing was abstracted from a multi-vane propulsor geometry ubiquitous in nature, and analysed to investigate aerodynamic performance during acceleratory and steady-state motions. Evolutionary convergence of slotted geometries in nature suggests an aerodynamic benefit in manoeuvrability, as exemplified in the fins and wings of a broad range of extant and extinct swimmers and flyers, respectively. Through direct force measurements and stereoscopic particle image velocimetry, it was found that the abstracted, slotted geometry exhibited a region of steady-state lift and drag enhancement at angles of attack greater than 25° when compared to a reference profile based on a delta-wing plate. At an angle of attack of 30°, the lift and drag measured on the abstracted model were 15.3% and 17.0% higher than the delta-wing model, respectively. In contrast, these shapes showed little difference in performance during an acceleration-from-rest manoeuvre. It was found that the secondary and tertiary vanes of the abstraction encouraged the formation of additional leading-edge vorticity. The formation of these additional leading-edge vortices was confirmed by an increase in streamwise circulation measured near each effective leading edge along the length of the chord. As such, this configuration provides lift augmentation appropriate for the development of high-performance control surfaces.

Experimental Investigation on Blunt-Edged UTM Delta Wing VFE-2 Configurations at Low Reynolds Number

The flow behaviour over the upper surface of a blunt-edged delta wing is mainly governed by a complicated leading edge vortex. This paper presents an experimental investigation on blunt-edged UTM VFE-2 delta wing model at low Reynolds numbers. The primary vortex for sharp-edged delta wing develops in the apex region even at low angles of attack. However, this situation does not happen for blunt-edged wings; the primary vortex in the leading edge region develops at a certain chordwise position on the wing. The primary vortex progresses upstream or downstream depending on the leading edge profile, the angle of attack and Reynolds number. A decrease in Reynolds number will accelerate the upstream progression of the primary vortex towards the wing apex. In the VFE-2 research group, many experiments were conducted at Reynolds number of 1×10^6 and above. Thus, the main objective of this study was to investigate the upstream progression of leading edge primary vortex on a blunt-edged delta wing at Reynolds numbers of 0.5×10^6, 0.75×10^6 and 1.0×10^6. Wind tunnel experiments were performed at three different flow velocities, 8.7 m/s, 13.1 m/s and 17.5 m/s corresponding to Reynolds numbers of 0.5×10^6, 0.75×10^6 and 1.0×10^6. Two measurement techniques were employed on the upper surface of the wing, i.e. experimental surface pressure and tuft techniques. Experiments were also conducted on two different leading edge profiles namely medium and large radius profiles with different leading edge radius-to-wing chord ratio. Pressure coefficients on the upper surface of the wing were plotted to observe the characteristic primary vortex, the upstream progression of the primary vortex and also the characteristics of vortex breakdown. The results obtained indicate that the primary vortex rapidly formed when the Reynolds number was decreased. Analysis in the apex region showed that the attached flow still existed even at a very low Reynolds number of 0.5 × 10^6.

Experimental study of riblet effects on the aerodynamic performance and flow characteristics of a delta wing

The present study investigates the effects of riblet on the aerodynamic performance and flow characteristics of a delta wing. Flow visualizations and measurements of the aerodynamic forces are performed on a smooth-surface as well as few textured-surface models of the delta wing. Flow visualizations were undertaken at flow speed of 2.5 m/s and various angles of attack in a vertical wind tunnel. The effects of riblet are investigated on the diameter of the vortices, location of the vortex breakdown, distance of the vortices from the wing surface, pattern of flow lines around the wing, and separation phenomenon. The results show that the riblet with a height-to-chord ratio of 0.013 has the most positive effect on the aforementioned parameters as far as drag reduction is concerned. In the second part of the study, lift and drag forces of the models are measured in a horizontal wind tunnel for a smooth-surface model as well as the other two textured models with height-to-chord ratios of 0.006 and 0.013, height-to-distance ratio of 1, sweep angle of 63.5°, and at Reynolds number of 2 × 105 and 0 to 35° angles of attack. The riblet-surface delta-wing model with height-to-chord ratio of 0.013 shows an increased lift-to-drag ratio at the whole range of angles of attack from 0 to 35°.