Control Surface Effectiveness Research Articles

Effects on Aircraft Performance Due to Geometrical Twist of Wing

This paper provides an overview of the effects of Geometric Twist on Aircraft performance by introducing a washout condition in an aircraft wing. This condition results in an effective Angle of Attack at the wingtip that is lower than the Angle of Attack at the wing root. Using CFD analysis, the variation of aircraft performance factors such as lift, drag coefficients, and aerodynamic efficiency is calculated for different Twist angles. A plot of lift and drag coefficients at varying Angle of Attack angles has been generated based on these analyses. These results illustrate the advantages of wing twist variation, particularly at higher angles of attack. One significant advantage of Geometric Twist is that it causes stall conditions to first occur at the wing root, providing a signal to the pilot to control the aircraft before the stall reaches the wingtip. This ensures the effectiveness of control surfaces, such as Ailerons and Flaps, located at the wing trailing edge. A comparison is made between the lift, drag coefficients, and aerodynamic efficiency of twisted wings and untwisted wings with identical parameters. While twisting the wingtip yields favourable results at higher Angles of Attack (AOA) compared to an untwisted wing, the aerodynamic efficiency of the wings decreases at lower AOA. However, applying the twist angle at high angles of attack, such as 10, 12.5, 15, and 20 degrees, leads to an increase in aerodynamic efficiency.

To study the effectiveness of stern appendages (Cruciform & X Shaped configurations) for maneuverability of autonomous underwater vessel using computational fluid dynamics

Maneuverability is one of the main aspects to be considered during the early design stage of Autonomous Underwater Vehicles. Several captive tests such as Planar Motion Mechanism (PMM), Oblique towing test, and Rotating Arm tests were conducted to obtain the hydrodynamic derivatives of the Autonomous Underwater Vehicle (AUV). In this research, the effectiveness of control surfaces with Cruciform and X configurations has been predicted for the maneuverability of the DARPA SUBOFF model. Computational Fluid Dynamics (CFD) technique was applied and the k- ω turbulence model was used to estimate the flow characteristics in the boundary layer region. The Straight-Line Towing Tank test has been used to calculate the hydrodynamic derivatives due to linear velocity, while the Rotating Arm test with a moving reference frame has been used to estimate the hydrodynamic derivatives. All the tests were conducted at a speed of 6.5 knots and three angles of incidence (10,30,50) for Straight-Line motion and three radii (13m,18m,24m) for Rotating Arm test were adopted. The simulation results of hydrodynamic coefficients have been compared with experimental results for the Cruciform configuration and the error was below 15%. So, the same procedure was adopted for the X-form configuration, and the stability index was calculated for both configurations.

Enhancing aircraft control surface effectiveness by co-flow jet flap at low energy expenditure

This paper conducts Improved Delayed Detached Eddy Simulation (IDDES) to compare applying co-flow jet (CFJ) active flow control on the flap and on the front main part of an aircraft control surface. Trade studies with varying CFJ momentum coefficients (Cμ) and flap deflection angles (δ) indicate that using CFJ on the flap is much more effective and efficient than applying CFJ on the front main part of the control surface. It is attributed to the feature that CFJ is more advantageous to work in severe adverse pressure gradient (APG). With a small Cμ of 0.025, the configuration with CFJ on the flap achieves a 46.1% increase of lift coefficient (CL) with an 80.8% lower power consumption compared with the one using CFJ on the front part. Moreover, the control surface with CFJ on the flap can attach the flow very well at a very high deflection angle of 70∘ and achieves a CL of 2.5 times larger than that of the baseline control surface.

Aeroelastic simulations of a delta wing with a Chimera approach for deflected control surfaces

A numerical tool for the computation of aircraft control surface aerodynamics with flexibility effects is presented. The solution is based on coupled Computational Fluid Dynamics (CFD) and Computational Structural Mechanics (CSM) simulations embedded in the multidisciplinary simulation environment SimServer. In SimServer, the DLR-TAU Code is utilized to obtain the CFD solution by solving the Reynolds-Averaged Navier–Stokes (RANS) equations. Structural displacements are computed with a modal solver. The Chimera implementation of SimServer, suited for hybrid grids, is applied to model the control surfaces. Numerical simulations with the flexible Chimera method are performed for the Model53 wing configuration, which is a generic delta wing with a deployed slat as well as an inboard and outboard trailing edge flap. Aerodynamic and aeroelastic simulations at high dynamic pressure q=45 kPa and transonic speed {text {Ma}} = 0.8 are performed for several angles of attack 10^circ le alpha le 25^circ and flap deflection angles -30^circ le delta le 30^circ. The effect of structural deformations on the flow field and control surface effectiveness are analyzed and compared to computations of components treated fully rigid. At the targeted freestream condition M=0.8 and {text {Re}}=15.1 times 10^7, the flow field around the Model53 configuration is characterized by the interaction of vortices and shock waves. The results of the lift and pitching moment coefficient for the rigid and flexible configuration revealed the importance of taking the structural flexibility into account in order to obtain more accurate results for the considered range of flap deflections. Furthermore, the computational effort of the aerodynamic and aeroelastic simulations are evaluated. The increase in computational effort is shown to be adequate for the given increase in accuracy.

Robust control design based aircraft flat-spin recovery using optimally deflected novel deployable fin

This paper presents novel approach for utilizing deployable fin as a static control surface along with aircraft primary controls for flat-spin recovery. Comparative analysis of with and without deployment of the fin has been thoroughly examined. To demonstrate the present approach, a nonlinear constrained optimization technique based optimal fin deflection is determined, followed by aircraft response in a flat-oscillatory-stable-left-spin flight at optimal fin setting angle, which provides spin attribute compensation. Thereafter, sliding-mode based robust control is designed for commanded departure of the aircraft from flat-spin to steady-straight-level flight. It is shown that finite-time direct spin recovery is achieved without any timescales based separation in dynamics. The present work novelty lies in the proposed strategy of using deployable fin, wherein dynamics of the fin is modeled and incorporated in the standard equations of motion of aircraft, subsequently, robust control law is developed. It is found from closed-loop simulation results that the proposed controller not only reduces the number of turns about spin axis and altitude loss but also minimizes spin recovery time when the fin is deployed. Additionally, rudder control power requirement is reduced, and effectiveness of other aerodynamic control surfaces are regained earlier when the fin is used for spin recovery. The outcome of the present study promises its practical implementation since the proposed solution is simple and compatible to integrate with the existing aircraft.

Investigation of an Underwater Vectored Thruster Based on 3RPS Parallel Manipulator

Autonomous underwater vehicles (AUVs) are important and useful tool platforms in exploring and utilizing ocean resource. However, the effect of control surfaces would decrease even invalid complete in this condition, and it is very hard for conventional AUVs to perform detailed missions at a low forward speed. Therefore, solving this problem of AUVs becomes particularly important to increase the application scope of AUVs. In this paper, we present a design scheme for the vectored thruster AUV based on 3RPS parallel manipulator, which is a kind of parallel manipulator and has advantages of compact structure and reliable performance. To study the performance and characteristics of the proposed thrust-vectoring mechanism, a series of works about corresponding kinematic and dynamic analysis have been performed through the theoretical analysis and numerical simulation. In the part of kinematics, the inverse, forward kinematics, and workspace analysis of the thrust-vectoring mechanism is presented, and the numerical simulations are accomplished to prove the feasibility and effectiveness of this design in AUVs. In order to further verify feasibility of the thrust-vectoring mechanism, based on the considerations of various affecting factors, a dynamic model of the designed thrust-vectoring mechanism is established according to theoretical analysis, and the driving forces of the linear actuator are presented through a series of numerical simulations. In addition, a control scheme based on PID algorithm is proposed for the designed vectored thruster with considering various affecting factors and the application environment. Meanwhile, the control scheme is also established and verified in MATLAB Simscape Mutibody. A series of numerical simulations of the thrust-vectoring mechanism prove the feasibility of the vectored thruster. According to equipping the designed vectored thruster, the AUVs can overcome the limit of weakening the control ability at zero or low forward speeds, and this improvement also expands the application of it, which has been scaled greatly.

Experimental analysis of aircraft directional control effectiveness

Aircraft directional control effectiveness is analyzed through experiments in wind tunnel. Control surfaces on low aspect ratio lifting surfaces exhibit different lifting capabilities compared to those with high aspect ratio. Such behavior must be carefully considered in the preliminary design phase to avoid any overestimation in size, weight, costs and emissions. Traditional sizing methodologies are based on coupling the effects of the wing planform (e.g. aspect ratio) at low angles of control surface deflection with the effects of wing section (e.g. chord ratio) evaluated on section data in the full range of control surface deflection, so that the non-linear aerodynamics is inherited from 2D data, completely neglecting the different aerodynamic behavior of a low aspect ratio wing at high angles of deflection. To fill this gap, an experimental wind tunnel test campaign on a generic regional turboprop aircraft model with a modular vertical tail with rudder has been performed. Results indicate that the aircraft design methodologies present in public literature underestimate the control surface effectiveness at high angle of deflections by 15% to 25%, leading to an average overestimation of control surface size.

Control device effectiveness studies of a 53∘ swept flying wing configuration. Experimental, computational, and modeling considerations

The present investigation covers studies of different control devices on several 53∘ swept flying wing configurations without a vertical tail plane. The wings considered share the same planform but differ in their spanwise profile shapes. The planform, developed under the NATO AVT-251 Task Group, is referred to as the MULDICON (or MULti-DIsciplinary CONfiguration). The objectives of this article are twofold: (1) to design yaw control surfaces for the MULDICON wing using an experimental/computational approach and (2) to develop aerodynamic models that rapidly and accurately predict the effectiveness of control surfaces over a wide range of flight conditions. The yaw control surface design (position and size) should provide sufficient yaw moment with almost no contribution in roll and pitch moment. To identify such concepts, a number of preliminary experiments on a generic flying wing configuration have been conducted. Two promising concepts from wind tunnel tests were then numerically examined for being implemented in the MULDICON baseline wing. Concepts with spoilers and a split flap were specifically considered. For medium to high angles of attack, the flow topology of the baseline wing is dominated by a vortical flow field on the upper outer wing. This leads to interactions between vortex and the control device which influences the flow and the attitude of the control device on the upper wing side. Furthermore, this work considers developing aerodynamic models for predicting stability derivatives of several MULDICON designs over a wide range of flight conditions. Aerodynamic loads models are only developed for normal force and pitch moment coefficients, however, the developed approach can easily being extended to include lateral aerodynamic coefficients as well. Quasi-steady models of this work are power series expansions of traditional linear aerodynamic models to capture nonlinear effects. Additionally, the models can estimate static and dynamic stability derivatives and the control surface powers.

Effect of control surfaces on the aerodynamic database of the Stratofly hypersonic vehicle

This paper aims at studying the control surfaces of the STRATOFLY project reference aircraft, funded by the European Commission, under the framework of Horizon 2020 plan. The values of aerodynamic coefficients in a wide range of flow free-stream conditions are stored in the aircraft aerodynamic database. The research goal is to update a pre-existent database that was developed with fixed control surfaces using the six control surfaces deflection as input. Different Mach numbers determine different flow regimes: subsonic, transonic, supersonic, and hypersonic. In subsonic, transonic and low supersonic regimes a vortex-lattice solver is used to obtain the global coefficients assuming an unviscous flow on a simplified model. In hypersonic flow a build-up approach is applied: the control surfaces deflection contribution is developed by assuming a two-dimensional flow on the airfoil and by applying shock-expansion theory on the geometry. Then the paper analyses results showing stability and L/D results. The final paragraph focuses on trimmability at cruise Mach. No trimmed solution is obtainable to optimize the propulsive system. The solution proposed to solve this issue is to extend the four elevons: larger elevons are found to be able to trim the vehicle at the desired angle of attack.

Finite-time sliding mode and super-twisting control of fighter aircraft

The development of two nonlinear robust higher-order flight control systems for roll-coupled maneuvers of fighter aircraft with uncertain parameters is discussed in this article. The objective is to independently control the output variables (roll angle, pitch angle and sideslip angle) using aileron, elevator and rudder control surfaces. For a nominal model of aircraft, first a finite time stabilizing (FTS) control law, based on the notion of geometric homogeneity, is designed. Then for robust control in the presence of parameter uncertainties, (i) a discontinuous sliding mode (DSM) control law and (ii) a super-twisting (STW) continuous control law is designed. It is shown that in the composite closed-loop system consisting of either (a) the FTS and DSM control laws or (b) the FTS and STW control systems, the output trajectory tracking error and its first-order derivative converge to the origin in finite time. Digital simulation results for a swept-wing fighter aircraft, including the two composite control systems, are obtained. These results show that each of the designed flight controllers accomplishes precise simultaneous large longitudinal and lateral maneuvers, despite uncertainties in the aerodynamic and inertia parameters, turbulence, and partial loss of control surface effectiveness.

Flight Test of GL-1 Glider Half Scale Prototype

GL-1 is a single-seat mid-performance glider, designed to be Indonesian National Glider. The Glider have been developing since 2014. The development produced a half scale prototype called BL-1, which had accomplished static test in 2016, then followed by first flight test at April 20th 2017, and second flight test at May 21st 2017. The purpose of the flight test was to obtain familiarization of the aircraft, aerodynamics characteristics and flow visualization, with data from flight recorded in FDR. The flight test resulted in two flights with total length of 21 minutes. The data from FDR and flight test documents extracted to analyze the characteristics and behavior of the aircraft during flight test. The aerodynamics characteristic was close to analytical results. The control was good; however, the effectiveness of control surface may need to be further analyzed. The result of the flight test will be used as a reference for further improvements and may need further testing.

Numerical simulation of vortex generators on a winglet control surface

The objective of the present paper is improving the effectiveness of a winglet control surface at deflections greater than ≈5° through passive vortex generators (VG) placed on the control surface. Without deflection, the VGs are aligned with the flow direction and thus have almost no impact. The conceptual idea is that a control surface deflection changes the flow direction in such a way that the VGs become active and delay flow separation from the trailing edge. To decide on an appropriate initial Vortex Generator (VG) configuration the aerodynamic characteristics of the winglet are investigated first with Reynolds averaged Navier–Stokes (RANS) simulations. The functional principle of the concept is then demonstrated. A parametric study of different VG dimensions investigates the effectiveness for a range of flow conditions and deflection angles.

Аналитическая реконфигурация системы управления полетом воздушного судна с сохранением заданных характеристик устойчивости

The paper sets out the conceptual issues of methods of analytical solution of the problem of defined aircraft stability characteristics conservation during flight control system reconfiguration under actuating subsystem failure. It is shown that the use of the reconfiguration technique with analytic set of stabilizing control laws synthesized by the decomposition method taking into account the cross-linking between the control channels allows to fully compensate the effects of failures of the actuating subsystem in each of the control channels under the conditions of complete controllability of aircraft as a dynamic system. It thus makes it possible to reconfigure the flight control system that always ensures the stability of the motion of the aircraft with a precisely prescribed poles. During study of failures in the form of a gradual reduction of the effectiveness of control surfaces it found that the synthesized analytical control law is a corresponding automatic increase feedback gain serviceable channel actuating subsystem, as well as the reallocation of control signals between the drives within the existing cross-connections between the control channels. Some of the solutions obtained by reconfiguring the control laws are not satisfactory, because of significantly limited possibility of maneuvering. The explanation for this is the relatively small quantities of system controllability measures, which are a functions of controllability gramian, and directly related to the severity of failures actuating subsystem. The results can be used at perspective control systems design stage to optimize the characteristics of the control actuators to meet the increasing requirements for fail-safe flight.

The Analytic Solutions to the Task of Maintaining the Specified Characteristics Stability in Aircraft Flight Control System Reconfiguration

The paper sets out the conceptual issues of methods of analytical solution of the problem of defined aircraft stability characteristics conservation during flight control system reconfiguration under actuating subsystem failure. It is shown that the use of the reconfiguration technique with analytic set of stabilizing control laws synthesized by the decomposition method taking into account the cross-linking between the control channels allows to fully compensate the effects of failures of the actuating subsystem in each of the control channels under the conditions of complete controllability of aircraft as a dynamic system. It thus makes it possible to reconfigure the flight control system that always ensures the stability of the motion of the aircraft with a precisely prescribed poles. During study of failures in the form of a gradual reduction of the effectiveness of control surfaces it found that the synthesized analytical control law is a corresponding automatic increase feedback gain serviceable channel actuating subsystem, as well as the reallocation of control signals between the drives within the existing cross-connections between the control channels. Some of the solutions obtained by reconfiguring the control laws are not satisfactory, because of significantly limited possibility of maneuvering.

Aeroelastic Tailoring and Active Aeroelastic Wing Impact on a Lambda Wing Configuration

Two aeroelastic design approaches are applied to a supersonic lambda wing configuration with the goal of improving the wing’s aeroelastic performance. Aeroelastic tailoring and active aeroelastic wing technologies are investigated in a sequential manner to quantify their benefits. Industry-standard tools are used to perform these design studies. First, a finite element model of a lambda wing is developed. Second, eight structural variants of the finite element model are created to study variations in wing stiffness. Finally, two sizing optimization processes are applied. The first is an aeroelastic sizing approach standard to MSC NASTRAN, and the second is an active aeroelastic wing design process, which is presented in this paper. The latter process integrates control and structural sizing design variables together. Both of these processes are used to size all eight structural variants. The optimized variants are compared based on weight and control surface effectiveness. Additionally, the optimal variations in percent ply utilization over the wing for the lightest weight concept are explored. It is found that applying aeroelastic tailoring simultaneously with active aeroelastic wing has the potential to reduce the aircraft gross takeoff weight by 7% in lambda wing configurations.

Experimental aerodynamic assessment and evaluation of an agile highly swept aircraft configuration

Extensive experimental and numerical investigations on a highly swept generic unmanned combat aerial vehicle (UCAV) configuration of lambda type with a variable leading edge contour have been conducted. Within these investigations, it was shown that the flow field is dominated by complex vortex systems including vortex-to-vortex and vortex-to-boundary layer interactions. The vortex-dominated flow field has a strong nonlinear influence on the aerodynamic behavior of the configuration. Hence, controllability is demanding and poses a real challenge in the design of these kinds of configurations. Especially, the dimensioning of the control surfaces, for the lateral- and longitudinal stability of tailless configurations of low aspect ratio and high leading-edge sweep, poses a challenging task which is not yet solved. To understand the problem of lacking lateral- and longitudinal stability for these kinds of configurations, experiments in the subsonic and transonic flow regime have been conducted for the Stability and Control Configuration (SACCON), which has a leading edge sweep of 53°, to assess the control surface effectiveness of conventional trailing-edge control devices. The present study reviews the experimental investigations conducted with the highly swept generic UCAV configuration SACCON.

Computational Fluid Dynamics Predictions of Control-Surface Effects for F-16XL Aircraft

Computational fluid dynamics methods play an increasingly important role in aircraft design and development. Some examples are conditions that cannot be tested before a flight test. To rely on these methods, it is essential that they are assessed and evaluated with a state-of-the-art wind tunnel test and/or in-flight data. In a previous project, Cranked-Arrow Wing Aerodynamics Project International, it was reported that all computational fluid dynamics methods failed to some degree in the transonic regime where shock–vortex interaction phenomena were present. Detailed analysis of surface pressure distribution showed that computational fluid dynamics was not able replicate the correct flowfield and produce acceptable results. This paper aims to restudy those transonic flight conditions for which computational fluid dynamics underperforms and to try to determine or shed light on the extent to which surface effects contribute to the computational fluid dynamics predictions.

Experimental Investigations of a Generic Swept Unmanned Combat Air Vehicle with Controls

Several static and dynamic forced-motion wind-tunnel tests on a generic unmanned combat air vehicle configuration with a 53 deg swept leading edge are reviewed. The static tests include both subsonic and transonic measurements. These tests are part of an international research effort to assess and advance the state-of-the-art of computational-fluid-dynamics methods to predict the static and dynamic stability and control characteristics for this type of configuration. The tests included analysis of trailing-edge control surface effectiveness. This paper describes the data collected from three different models of this unmanned combat air vehicle configuration in four different wind tunnels. A subset of this large experimental dataset was selected as focus test cases for computational-fluid-dynamics comparison. The focus test case data are presented in detail.

Aerodynamics of delta wings with flaps at hypersonic speeds

Control surface effectiveness is an important parameter for any aeroplane. For a hypersonic aircraft, though the power required to operate the flaps is determined by low speed flying conditions, it is imperative to know the effect of flaps at hypersonic speeds. Hence, studies have been done on this topic by aerodynamicists for over 40 years. In spite of this, only a limited data is available in the literature on this subject. This paper discusses the experimental study of the effect of sweep on the aerodynamic characteristics of thin slab delta wings with flaps at hypersonic speeds. For the purpose of this investigation, a novel special thin six-component balance, which has a thickness of 4mm and can be housed inside wings with 8mm thickness, has been designed. The wings had a sweep of 76degrees, 70degrees and 65degrees, t/c of 0.053 and flaps with 12% of wing area and 12% of wing chord. Testing were done at Mach 8.2, Re number of 2.13 x 10(6) (based on chord), from alpha = -12degrees to 12degrees and flap angle of 20degrees, 30degrees and 40degrees. Separation lengths, measured from Schlieren pictures, clearly show that there is 'no appreciable' effect of sweep on them. Also, using a simple local flow field calculation, the separation has been identified to be transitional in nature. These features of separation reflect in the force data. Because of the small separation length, the flaps (inspite of their small size) were very effective in generating additional C-N, C-M and C-l, which increased with increase in flap angle. In general, the C-N, C-M and X-CP were unaffected by sweep for symmetric flap deflection at positive incidences and asymmetric flap case, For symmetric flap case at negative incidences, only C-N was not influenced by the sweep but C-M decreased and X-CP moved upstream as the sweep is decreased, The wing with lower sweep produces higher CA and lower (L/D)(max) for both symmetric and asymmetric flaps. The rolling moment and adverse yaw increased with decrease in sweep for asymmetric flap deflection. Newtonian theory is shown to be incapable of predicting the effect of sweep on C-l, C-n and on the incremental values of C-N, C-M and C-A. In conclusion, it can be said that a small flap is generally adequate for hypersonic aeroplanes provided they operate at altitudes where transitional and turbulent separation can be expected to occur. This would make the flaps effective and thus enable ample control authority.

Investigation of Missile Control Surface Effects on F-16 Limit-Cycle Oscillation

Tactical aircraft with thin wings and heavy external wing stores are susceptible to aeroelastic limit-cycle oscillations. This investigation applies both frequency-domain and time-domain solutions using aeroelastic computational models to predict limit-cycle oscillation onset, frequency, and amplitude on an F-16. The test configuration uses missile shapes with reconfigurable aerodynamic and mass properties. Statistical analysis of flight-test results identifies significant experimental variables of missile aerodynamic configuration, flight condition, and aircraft fuel state. Furthermore, statistical analysis permits the normalization of flight-test results for direct comparison to aeroelastic models. Results of the comparison show that flight-test data analysis mostly supports the prediction trends, but the magnitude of the aerodynamic effects due to the canards and fins is much less significant in the test data.