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

2012 ◽  
Vol 482-484 ◽  
pp. 121-125
Author(s):  
Peng Wu ◽  
Xue Ying Deng ◽  
Yan Kui Wang
Keyword(s):  
Flow Separation ◽  
Deflection Angle ◽  
Leading Edge ◽  
Control Surface ◽  
Control Technique ◽  
High Lift ◽  
Surface Design ◽  
Aerodynamic Efficiency ◽  
Lift Control ◽  
Control Surfaces

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

Transient Response Study on Rolling Effectiveness of Multiple Control Surfaces

2002 ◽  
Author(s):  
Deman Tang ◽  
Aiqin Li ◽  
Earl H. Dowell
Keyword(s):  
Transient Response ◽  
Three Dimensional ◽  
Wind Tunnel Test ◽  
Leading Edge ◽  
Control Surface ◽  
Multiple Control ◽  
Wing Model ◽  
Aerodynamic Model ◽  
Control Surfaces ◽  
Dry Friction Damping

In the present paper, a transient response study of the effectiveness of trailing and leading edge control surfaces has been made for a rolling wing-fuselage model. An experimental model and wind tunnel test are used to assess the theoretical results. The theoretical model includes the inherently nonlinear dry friction damping moment that is present between the spindle support and the experimental aeroelastic wing model. The roll trim equation of motion and the appropriate aeroelastic equations are solved for different combinations of leading and trailing edge control surface rotations using a reduced order aerodynamic model based upon the fluid eigenmodes of three dimensional vortex lattice aerodynamic theory. The present paper provides new insights into the transient dynamic behavior and design of an adaptive aeroelastic wing using trailing and leading edge control surfaces.

scholarly journals Airfoil Selection Procedure, Wind Tunnel Experimentation and Implementation of 6DOF Modeling on a Flying Wing Micro Aerial Vehicle

Micromachines ◽  
2020 ◽  
Vol 11 (6) ◽  
pp. 553 ◽  
Author(s):  
Taimur Ali Shams ◽  
Syed Irtiza Ali Shah ◽  
Ali Javed ◽  
Syed Hossein Raza Hamdani
Keyword(s):  
Wind Tunnel ◽  
Equations Of Motion ◽  
Selection Procedure ◽  
Leading Edge ◽  
Control Surface ◽  
Flight Tests ◽  
Micro Aerial Vehicle ◽  
Aerial Vehicle ◽  
P Factor ◽  
Control Surfaces

Airfoil selection procedure, wind tunnel testing and an implementation of 6-DOF model on flying wing micro aerial vehicle (FWMAV) has been proposed in this research. The selection procedure of airfoil has been developed by considering parameters related to aerodynamic efficiency and flight stability. Airfoil aerodynamic parameters have been calculated using a potential flow solver for ten candidate airfoils. Eppler-387 proved to be the most efficient reflexed airfoil and therefore was selected for fabrication and further flight testing of vehicle. Elevon control surfaces have been designed and evaluated for longitudinal and lateral control. The vehicle was fabricated using hot wire machine with EPP styrofoam of density 50 Kg/ m 3 . Static aerodynamic coefficients were evaluated using wind tunnel tests conducted at cruise velocity of 20 m/s for varying angles of attack. Rate derivatives and elevon control derivatives have also been calculated. Equations of motion for FWMAV have been written in a body axis system yielding a 6-DOF model. It was found during flight tests that vehicle conducted coordinated turns with no appreciable adverse yaw. Since FWMAV was not designed with a vertical stabilizer and rudder control surface, directional stability was therefore augmented through winglets and high wing leading edge sweep. Major problems encountered during flight tests were related to left rolling tendency. The left roll tendency was found inherent to clockwise rotating propeller as ‘P’ factor, gyroscopic precession, torque effect and spiraling slipstream. To achieve successful flights, many actions were required including removal of excessive play from elevon control rods, active actuation of control surfaces, enhanced launch speed during take off, and increased throttle control during initial phase of flight. FWMAV flew many successful stable flights in which intended mission profile was accomplished, thereby validating the proposed airfoil selection procedure, modeling technique and proposed design.

scholarly journals Boundary Layer Control of Airfoil using Rotating Cylinder

2020 ◽  
Vol 8 (6) ◽  
pp. 4742-4750
Keyword(s):  
Leading Edge ◽  
Rotating Cylinder ◽  
Aerodynamic Characteristics ◽  
Favourable Result ◽  
Ansys Fluent ◽  
High Lift ◽  
Freestream Velocity ◽  
Aerodynamic Efficiency ◽  
Two Parameters ◽  
Layer Control

The requirement for improving the aerodynamic efficiency and delaying the formation of stall over the wing has been of prime importance within the field of aviation. The main objective of the project is to further improve upon these two parameters. The configuration used for analysis consists of a NACA 2412 airfoil of chord length 0.982m with a 64mm cylinder at the leading edge. Analysis is completed using ANSYS Fluent, with a freestream velocity of 10m/s. The aerodynamic characteristics of three configuration bare airfoil, Airfoil with static cylinder and Airfoil with rotating cylinder are tabulated and plotted. The comparison is then followed by pressure and velocity contours to visualize the flow over each configuration. The rotating cylinder configuration shows a improvement in the aerodynamics characteristics. The rotating cylinder configuration gives the most favourable result. This study has a potential application in high lift devices and can be used as stall delaying device

Flow separation control of NACA-2412 airfoil with bio-inspired nose

2019 ◽  
Vol 91 (7) ◽  
pp. 1058-1066 ◽  
Author(s):  
Mohamed Arif Raj Mohamed ◽  
Ugur Guven ◽  
Rajesh Yadav
Keyword(s):  
Flow Control ◽  
Flow Separation ◽  
Control Method ◽  
Angle Of Attack ◽  
Leading Edge ◽  
Separation Control ◽  
Cetacean Species ◽  
Aerodynamic Efficiency ◽  
Content Type ◽  
Flow Separation Control

Purpose The purpose of this paper is to achieve an optimum flow separation control over the airfoil using passive flow control method by introducing bio-inspired nose near the leading edge of the NACA 2412 airfoil. Design/methodology/approach Two distinguished methods have been implemented on the leading edge of the airfoil: forward facing step, which induces multiple accelerations at low angle of attack, and cavity/backward facing step, which creates recirculating region (axial vortices) at high angle of attack. Findings The porpoise airfoil (optimum bio-inspired nose airfoil) delays the flow separation and improves the aerodynamic efficiency by increasing the lift and decreasing the parasitic drag. The maximum increase in aerodynamic efficiency is 22.4 per cent, with an average increase of 8.6 per cent at all angles of attack. Research limitations/implications The computational analysis has been done for NACA 2412 airfoil at low subsonic speed. Practical implications This design improves the aerodynamic performance and increases structural strength of the aircraft wing compared to other conventional high-lift devices and flow-control devices. Originality/value Different bio-inspired nose designs which are inspired by the cetacean species have been analysed for NACA 2412 airfoil, and optimum nose design (porpoise airfoil) has been found.

Flow separation control using a bio-inspired nose for NACA 4 and 6 series airfoils

2021 ◽  
Vol 93 (2) ◽  
pp. 251-266
Author(s):  
Mohamed Arif Raj Mohamed ◽  
Rajesh Yadav ◽  
Ugur Guven
Keyword(s):  
Flow Control ◽  
Flow Separation ◽  
Control Method ◽  
Leading Edge ◽  
Aircraft Design ◽  
Separation Control ◽  
Cetacean Species ◽  
Aerodynamic Efficiency ◽  
Content Type ◽  
Flow Separation Control

Purpose This paper aims to achieve an optimum flow separation control over the airfoil using a passive flow control method by introducing a bio-inspired nose near the leading edge of the National Advisory Committee for Aeronautics (NACA) 4 and 6 series airfoil. In addition, to find the optimised leading edge nose design for NACA 4 and 6 series airfoils for flow separation control. Design/methodology/approach Different bio-inspired noses that are inspired by the cetacean species have been analysed for different NACA 4 and 6 series airfoils. Bio-inspired nose with different nose length, nose depth and nose circle diameter have been analysed on airfoils with different thicknesses, camber and camber locations to understand the aerodynamic flow properties such as vortex formation, flow separation, aerodynamic efficiency and moment. Findings The porpoise nose design that has a leading edge with depth = 2.25% of chord, length = 0.75% of chord and nose diameter = 2% of chord, delays the flow separation and improves the aerodynamic efficiency. Average increments of 5.5% to 6° in the lift values and decrements in parasitic drag (without affecting the pitching moment) for all the NACA 4 and 6 series airfoils were observed irrespective of airfoil geometry such as different thicknesses, camber and camber location. Research limitations/implications The two-dimensional computational analysis is done for different NACA 4 and 6 series airfoils at low subsonic speed. Practical implications This design improves aerodynamic performance and increases the structural strength of the aircraft wing compared to other conventional high lift devices and flow control devices. This universal leading edge flow control device can be adapted to aircraft wings incorporated with any NACA 4 and 6 series airfoil. Social implications The results would be of significant interest in the fields of aircraft design and wind turbine design, lowering the cost of energy and air travel for social benefits. Originality/value Different bio-inspired nose designs that are inspired by the cetacean species have been analysed for NACA 4 and 6 series airfoils and universal optimum nose design (porpoise airfoil) is found for NACA 4 and 6 series airfoils.

Design and Modeling of the Fowler Flap With Adaptive Elements

Aerospace ◽  
2006 ◽  
Author(s):  
Oleksandr Kozlov
Keyword(s):  
Finite Element ◽  
Finite Element Model ◽  
Deformation Behaviour ◽  
Leading Edge ◽  
Element Model ◽  
High Lift ◽  
Aerodynamic Efficiency ◽  
Transport Aircraft ◽  
Sma Actuators ◽  
Active Flap

This paper describes the process of the designing of the Fowler flap with adaptive elements. Modern passenger and transport aircraft use high lift devices for take off and landing. Of great importance for the creation of high lift during take off and landing is the shape and size of the gap between wing and extended Fowler flap. To improve the deformation behaviour of the leading edge of this flap, and to improve the geometry of the gap, it was decided to use shape memory alloy (SMA) actuators in the flap structure. First, the complex finite element model of the passive flap was created and then this model was extended with active elements, which are modelling the SMA actuators. As a result, the complex finite element model of the active flap was obtained. This modelling was done using software MATLAB and a finite element model was created using software ANSYS. The main result of this work is that with help of SMA actuators integrated in the flap it was possible to influence and improve the geometry of the gap between the wing and the extended flap, resulting in the aerodynamic efficiency of this flap being increased.

scholarly journals Control surface design for radio-controlled aircraft. Case: SAE Aero Design Micro-class prototype

2021 ◽  
Author(s):  
Rafael A. Márquez ◽  
Miguel A. Martínez ◽  
Manuel J. Martínez
Keyword(s):  
Weight Ratio ◽  
Aircraft Design ◽  
Control Surface ◽  
Primary Control ◽  
Surface Design ◽  
C Language ◽  
Design Variables ◽  
Surface Sizing ◽  
Design Competition ◽  
Control Surfaces

This research article presents a design methodology for primary control surfaces (Ailerons, Rudder and Elevator) for experimental unmanned radio-controlled aircraft. The methodology is based on the proposal and standardization of the required mechanical and aerodynamic analysis for each control surface sizing, considering the SAE Aero Design competition objectives within Micro Class. It is used on empirical results previously described in references about aeronautical design, computerized fluids dynamics (CFD) software, and aircraft controllability regulations in order to obtain the design variables. Based on this information, the iteration sequences required for design were automated by a C++ language code to obtain the optimal characteristics for each surface, thereby reducing the possibility of calculation errors, overall time, and workload invested in the design process. The application of the methodology to the latest aircraft design reduced the total control systems weight to the aircraft’s empty weight ratio to a minimum of 3.4%.

scholarly journals Hybrid Deflection of Spoiler Influencing Radar Cross-Section of Tailless Fighter

Sensors ◽  
2021 ◽  
Vol 21 (24) ◽  
pp. 8459
Author(s):  
Zeyang Zhou ◽  
Jun Huang
Keyword(s):  
Cross Section ◽  
Electromagnetic Scattering ◽  
Deflection Angle ◽  
Elevation Angle ◽  
Radar Cross Section ◽  
Control Surface ◽  
Linear Mode ◽  
Linear Dynamic ◽  
The Given ◽  
Control Surfaces

With the continuous development of advanced fighters towards tailless and flying wing layouts, diverse control surfaces have become the mainstream design. To study the influence of spoiler control surface on the radar cross-section (RCS) of a tailless fighter, a calculation method is presented. The deflection angle of the spoiler is controlled by the fixed mode, linear mode, and smooth mode. The results show that the opening action of the spoiler will break the original stealth characteristics of the aircraft at the key azimuth angles of the head and tail. As the elevation angle increases, this adverse effect will spread to the side. The influence of the different dynamic deflection modes of the spoiler on the aircraft RCS is analyzed. Compared with the linear dynamic deflection mode, the smooth dynamic deflection mode is conducive to the reduction in the average RCS at the given head azimuth. The presented method is effective to study the influence of the spoiler deflection on the electromagnetic scattering characteristics of the tailless aircraft.

scholarly journals A trim problem formulation for maximum control authority using the Attainable Moment Set geometry

2021 ◽  
Author(s):  
Carmine Varriale ◽  
Mark Voskuijl
Keyword(s):  
Inequality Constraints ◽  
Problem Formulation ◽  
Control Surface ◽  
Control Effort ◽  
Decision Variables ◽  
Control Authority ◽  
Control Forces ◽  
Lift Control ◽  
Moment Set ◽  
Control Surfaces

AbstractThis paper presents a generic trim problem formulation, in the form of a constrained optimization problem, which employs forces and moments due to the aircraft control surfaces as decision variables. The geometry of the Attainable Moment Set (AMS), i.e. the set of all control forces and moments attainable by the control surfaces, is used to define linear equality and inequality constraints for the control forces decision variables. Trim control forces and moments are mapped to control surface deflections at every solver iteration through a linear programming formulation of the direct Control Allocation algorithm. The methodology is applied to an innovative box-wing aircraft configuration with redundant control surfaces, which can partially decouple lift and pitch control, and allow direct lift control. Novel trim applications are presented to maximize control authority about the lift and pitch axes, and a “balanced” control authority. The latter can be intended as equivalent to the classic concept of minimum control effort. Control authority is defined on the basis of control forces and moments, and interpreted geometrically as a distance within the AMS. Results show that the method is able to capitalize on the angle of attack or the throttle setting to obtain the control surfaces deflections which maximize control authority in the assigned direction. More conventional trim applications for minimum total drag and for assigned angle of elevation are also explored.

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