Induced Rolling Moment Research Articles

Neuroadaptive high-order fully-actuated system approach for roll autopilot with unknown uncertainties

In this paper, a neuroadaptive high-order fully-actuated system approach control scheme incorporating the disturbance observer technique is proposed for the missile roll autopilot, subject to model uncertainties generated by the induced roll moment, along with actuator control efficiency deterioration and external disturbance. To address model uncertainties, the radial basis function neural network is implemented. The external disturbance and approximation error are treated as compound disturbances and estimated by a nonlinear disturbance. To avoid the “differential explosion” inherent in the backstepping technique, the high-order fully-actuated system approach is invoked to track the desired roll angle command. The semi-globally uniformly bounded of the closed-loop system is demonstrated via the Lyapunov method. Numerous simulations under various conditions have been conducted to verify the effectiveness of the proposed roll autopilot.

Wake Vortex Far-Field of a Generic Transport Aircraft Affected by Oscillating Flaps

The influence of oscillating trailing-edge flaps of a generic transport aircraft on the middle and far fields of the wake vortex system is investigated. The aim is to be able to reduce separation distances between aircraft during approach in order to increase capacity at airports. By introducing a certain disturbance by means of oscillating flaps in the near field, the long-wavelength Crow instability should be excited in the further downstream development. The evolution of the wake vortex system is investigated using a large-eddy simulation approach. The wake vortex system is examined up to 133 spans downstream for two high-lift configurations. A nonactuated configuration with statically deflected flaps is compared to a configuration with actuated flaps. The results show that the long-wavelength Crow instability cannot be excited by actuated trailing-edge flaps. However, a faster decrease of the dimensionless circulation over the downstream position can be observed for the actuated configuration compared to the nonactuated configuration, due to a faster diffusion of the rolled-up vortex pair. Furthermore, the averaged induced rolling moment on an aircraft [with the International Civil Aviation Organization’s (ICAO’s) classification of “light”] encountering the wake vortex of the actuated configuration (ICAO’s classification of “heavy”) is up to 32.5% lower than for a follower aircraft, which is entering the wake vortex of the nonactuated configuration.

Wake Vortex Effects due to Horizontal Crossing at Any Angle

The effects of wake crossing at any horizontal angle, airspeed, and altitude are considered, focusing on crossing at cruise speeds at altitudes close to the tropopause. The roll effect is considered in two cases: 1) specification of the aileron deflection as a function of time that exactly compensates for the rolling moment induced on the crossing aircraft by the wake of the leading aircraft so that the crossing aircraft is undisturbed; and 2) in the case when the available roll control power is not sufficient to compensate completely for the wake induced rolling moment, the resulting roll rate and bank angle of the crossing aircraft are calculated as a function of time. The distinction between the two cases is made by a safe separating distance or safe separating time: beyond which, roll compensation 1) is possible; and below which, rolling motion 2) is unavoidable. Besides the roll acceleration, roll rate, and bank angle for the roll axis, the effects of wake crossing are also considered for the vertical axis, including the load factor, vertical velocity, and altitude change. The theory applies to any pair of aircraft and indicates the dependence on aircraft parameters, flight, and atmospheric conditions.

Study on the Influence of a Wake Vortex on an ARJ21 Aircraft Using the Strip Method

A model for the vortex distribution in the wake of an aircraft is elaborated to investigate the wake influence on the behaviour of other aircrafts potentially interacting with it. As a realistic case, the interaction of an ARJ21 aircraft with a (leading) A330-200 aircraft is considered. Different distances are considered, namely, 6 km, 7 km, 8 km, 9.3 km, and 10 km. Simulations based on the used wake dissipation mechanism are used to investigate different conditions, namely, the ARJ21 in take-off and level flight and the changes induced in the related lift by the front aircraft A330-200 during landing. The induced roll moment is also studied and analyzed by means of a strip method. As a result, the roll moment coefficient is determined to quantify the roll degree of the aircraft when it is influenced by the wake vortex. The results show the overall roll moment coefficient of the considered ARJ21 aircraft is less than 0.05, and the wing roll moment coefficient is less than 0.04. Such results are interpreted and discussed according to existing standards.

Design of retractile fins actuator based on hollow piezoelectric motor

In the initial stage of the exterior ballistic, the centroid of the simple guided rocket will change with the fuel consumption of rocket engine, so the flight stability in the initial stage and the manoeuvrability in the whole trajectory correction process must to be comprehensively considered. For this challenge, in this paper, a specific actuator and aerodynamic configuration based on the integration of retractile fins structure and rotor of hollow piezoelectric motor is proposed. And the research on the structure of the actuator and the scheme of the retractile mechanism are carried out. The wind-induced rolling moment of the actuator under different flight conditions (angle of attack, Mach number) is simulated and analyzed by numerical computation method, and the load characteristics of the piezoelectric motor of the actuator is analyzed and tested. The results show that when the flight speed is 0.8 Ma, 1.0 Ma, the angle of attack is 0∼6°, the maximum induced rolling moment of the actuator is 0.30 N.m, and the load speed of the piezoelectric motor is 25r/min when the excitation voltage is 450V and the excitation frequency is 40.2kHz, which can meet the requirements of the projectile flying environment.

Wake Turbulence Evolution and Hazard Analysis for General Aviation Takeoff Accident

Wake turbulence is commonly a commercial aviation topic involving relatively heavy transport aircraft. However, wake turbulence is also safety relevant for business and general aviation flight operations. A detailed analysis has been conducted for a specific accident involving two aircraft of the wake turbulence category light. A four-seat aircraft with 1 ton maximum takeoff weight took off behind a biplane with 5.5 ton maximum takeoff weight. This led to a fatal accident. Based on available information and several assumptions, the wake turbulence aspects for the specific accident scenario have been assessed. The roll axis was the dominating factor in this case. Accordingly, the wake induced rolling moment was analyzed and related to the assumed available roll control power. Furthermore, the wake vortex behavior was simulated and analyzed in detail. The results indicate that wake turbulence can be considered to be the cause of the accident. The theoretical results and the conclusions regarding the accident are also supported by the results of flight tests, which were conducted in the context of the accident investigation.

Roll Control of Low-Aspect-Ratio Wings Using Articulated Winglet Control Surfaces

The use of tip mounted winglets with independently variable cant angles was investigated as a means of roll control on wings with an aspect ratio of one. Wind-tunnel testing was performed in which a six-axis force balance was used to measure the total aerodynamic load on wings with winglet control surfaces. Stereoscopic digital particle image velocimetry of the near-wake plane was used to show how the topology of the tip vortices changed with winglet deflection. Shifts in the location of the right tip vortex core are considered to be responsible for roll moment generation because they indicate changes in the symmetry of suction-side flow structures. All winglet deflections were observed to shift the right tip vortex core inboard, and thereby shorten the effective span of the wing. The effect of a winglet deflection may be approximated as a change in the wing aspect ratio and a lateral shift in the wing aerodynamic center. Prandtl’s lifting line theory provides a closed-form estimate for the reduction in lift caused by a winglet deflection. A geometrical argument was made to account for the induced roll moment. The right tip vortex core also shifts vertically, following the deflected wing tip. Vertical shifts in the right tip vortex result in an angle between the wing span line and a line connecting the two tip vortices. A positive angle is defined as the right tip vortex higher over the wing than the left, and it is accompanied by a positive roll moment. While in sideslip, the wing with no winglet deflection experiences a considerable roll moment as a result of a vertical and lateral shift in the two tip vortices. The articulated winglets are observed to partially mitigate these effects when the upstream winglet is actuated, and thus show promise as a direct means of disturbance rejection.

Aerodynamic Aspects of Unmanned Aerial Vehicle Aerial Refueling

The feasibility of refueling a high-endurance unmanned aerial vehicle was investigated numerically, focusing on the aerodynamic effects. For this study, a single-aisle transport aircraft-based tanker was proposed (the Airbus A320) and, for the much smaller unmanned aerial vehicle, the MQ-9 Reaper was selected (and the weight ratio between these two airplanes was less than 6%). This refueling scenario was different from most previous studies where the tanker and the refueling airplanes were of similar weight. Computed results showed that the aerodynamic coefficients changed significantly with lateral and vertical separation between the two airplanes, and they changed only marginally when changing the following distance. In spite of the size and performance differences, it was found that, when the refueling plane approached from reasonably below and behind the tanker airplane, the induced rolling moments were within the unmanned aerial vehicle aileron control limit and a refueling maneuver was possible.

Influence of Differential Spoiler Settings on the Wake Vortex Characterization and Alleviation

Experimental investigations using particle image velocimetry technique have been carried out for the evaluation of the differential spoiler setting capabilities in modifying the spanwise wing load and further reduce the wake vortex hazard. The aircraft half model (at high lift configuration) was investigated for two differential spoiler settings. Results reveal a noticeable inboard shift of spanwise wing loading. Implementation of a differential spoiler setting results in a substantial redistribution of the flap tip vortex circulation with an increase in the diameter of the merged vortex by a factor of up to 2.72 relative to the undisturbed flap tip vortex. Inspection of the cross-stream distribution of axial vorticity shows a reduction by a factor of up to 2.33 in the peak vorticity value. A 44% decrease of the maximum crossflow velocity relative to the undisturbed flap tip vortex crossflow velocity was recorded for the case of deployed spoilers. Finally assessment of the differential spoiler setting capabilities as a wake vortex attenuation device reveals that, while position of the maximum induced rolling moments in the flap tip area is little influenced by the differential spoiler setting, the maximum induced rolling moment coefficient was reduced to nearly one third of the undisturbed flap tip vortex value.

Induced rolling moment for NACA4412 plain and flapped wing

PurposeThe paper aims to compute rolling moments on a follower aircraft wing due to vortices generated by a plain and flapped NACA4412 wing using experimental particle image velocimetery (PIV) data.Design/methodology/approachThis paper describes the detailed variation of the induced rolling moment on a follower aircraft wing derived from a PIV velocity field measurement. A rectangular wing of a subsonic wall interference model is used as a vortex generator in two different configurations: plain wing of NACA4412 cross‐section profile; and flapped wing with trailing edge flap of NACA0012 cross‐section profile extended at 20°.FindingsResults obtained showed that the maximum induced rolling moment coefficient depends on the strength of the vortex produced by the generating aircraft wing and increases linearly with the increment of the angle of attack.Originality/valueThis paper provides an insight on the effects of different angles of attacks for plain and flapped wings on the induced rolling moment coefficient and therefore, the hazard imposed on the following aircraft can be estimated.

Dynamics of Vortical Flows Induced by Swept-Notched Wings in Aircraft Wake Simulations

The motivation behind the present investigations is the alleviation of vortex wake hazard. The influence of wing-tip geometry modifications on the near-field wake behind swept-notched wings is investigated up to 24 span lengths, using computational fluid dynamics. Three-dimensional parallel simulations are run at a chord-based Reynolds number of 5.34 x 10 4 . The effectiveness of notched wing modifications on wake properties is assessed via comparison with conventional and original notched wings. Current results are promising: the tip and flap vortices produced by the swept-notched wings remain distinct and unmerged, unlike those associated with the conventional wing. In all cases, the wakes of the novel wings have circulations and vorticity values that are substantially lower than the original notched wing. Numerical data clearly indicate that such geometrical combinations break apart the concentrated area of streamwise vorticity shed by the wing into three or more less intense regions, significantly reducing the wake strength and induced rolling moment, and affecting the location and merging process of the overall wake system.

Analytic Model of Catastrophic Yaw

A quantitative model of the non-linear motion leading to catastrophic yaw of finned missiles is presented. The coupled roll-yaw dynamics of the missile, acted on by the trim angle of attack due to slight configurational asymmetries, is represented by pitch-yaw and roll equations of motion that include cubic aerodynamic coefficients and roll orientation-dependent induced moments. The steady-state equilibrium points of the system are found and their stability is determined by linearization. The solutions are evaluated by comparison with numerical integration of the equations of motion, proving the capability of the model to predict catastrophic yaw. Nomenclature CD = drag force coefficient CLα = angle of attack coefficient of the lift force CNα = angle of attack coefficient of the normal force CMα = angle of attack coefficient of the pitch moment Cli = induced roll moment coefficient Clp = roll damping moment coefficient Clδ = roll effectiveness moment coefficient CMpα = angle of attack coefficient of the Magnus moment

Excitation of Instabilities in the Wake of a Wing with Winglets

This paper describes experimental investigations of possibilities to destabilize the vortex wake behind an airfoil by using winglets with integrated rudders. These rudders can be deflected statically, as well as oscillated periodically around their static deflection. The oscillation is intended to excite short-wavelength instabilities inherent in the vortex system to provoke an accelerated decay of the vortices. The used model also comprises deflectable ailerons and can be equipped with a pair of trailing edge flaps to modify the vortex system. Quantitative wake investigations by Particle Image Velocimetry have been conducted up to 43 spans behind the wing model in a water towing tank. Results show a markedly earlier decay of the vortex system and abatement of induced rolling moments in the wake for certain rudder configurations in comparison to the clean airfoil configuration. However, the oscillation of the winglet rudders did not lead to a further acceleration of vortex decay compared to the statical case in the frame of the experiments conducted here.

Tip Vortex Wake Destabilization with and without Winglets

AbstractWake vortices are an inevitable by‐product of lift. Another airplane following too closely can be dangerously influenced by induced rolling moments. Consistent adherence to the prescribed safety margins ensures safe flight, however it also obstructs the capacity of large airports. To remedy the expected future capacity problems, research is being carried out world wide. This paper describes some results of the research work to decrease the hazard of the vortex wake accomplished at the ILR. (© 2006 WILEY‐VCH Verlag GmbH & Co. KGaA, Weinheim)

Wake Vortex Alleviation Flow Field Studies

Wake-vortex-allevialion research was conducted in the far-field vortex wake of a generic wing-tail aircraft model. The goats were to achieve accelerated vortex strength reduction and to map the conditions at which this reduction would occur. The wing-tail model was run in a water tow tank to generate a pair of unequal-strength counterrotating vortices on each side of centerline. Dye flow visualization provided physical insight into the nature of the vortex interactions, and three-component particle image velocimetry allowed quantification of key characteristics of the flowfield, including circulation, vorticity, vortex trajectory, and induced rolling moments. Experiments were conducted for a variety of model angles of attack, tail incidence angles, and tail spans

Study of Four-Vortex Aircraft Wakes and Layout of Corresponding Aircraft Configurations

A sequence of numerical simulations for the efficient determination of four-vortex wakes, consisting of two counter-rotating vortex pairs, with low induced rolling moments after the unstable merging of inner and outer vortices, is proposed. After a systematic investigation of generic four-vortex systems in a second step, an airliner geometry in high-lift configuration is modified such that it produces a vortex wake similar to the identified optimal four-vortex system. The configuration modification is done using simple design rules. For verification of the existence of a similar four-vortex system, the wake vortex of the modified airliner configuration is simulated up to the extended near field in a final step. Two four-vortex configurations are considered in the present paper, different only in the vortex spacing ratio

On wake vortex response for all combinations of five classes of aircraft

AbstractThe present paper concerns the response of a following airplane to a pair of wing tip vortices left by a leading aircraft, represented by the Hallock-Burnham model, including the effect of vorticity decay between the two aircraft. The effect of vorticity is evaluated in terms of the induced rolling moment and also the lift loss; these specify the roll acceleration and the downward acceleration, respectively. The corresponding two response equations can be put into the same dimensionless form, and integrated using exponential integrals. This specifies the roll rate and sink rate as a function of time; besides the latter, the bank angle and altitude loss, are also plotted, all also as a function of time, for all combinations of leading and following aircraft in five classes. These are the three ICAO weight categories of light, medium and heavy, plus two other cases, viz the special case of the Boeing 757, which requires larger separations distances, and the case of a future very large transport aircraft (VLTA) exceeding significantly the size of a Boeing 747.

Studies on the influence of outboard flaps on the vortex wake of a rectangular wing

Aircraft trailing vortices constitute a hazard to following aircraft, and are therefore one of the main concerns for airport capacity constraints. At the Institute of Aerospace Engineering (ILR) experiments on wake vortices up to a distance of 60 spans behind the model of a rectangular wing are conducted in a towing tank. The motivation behind the presented experiments is the alleviation of the rolling moment induced on following aircraft by constructive means at wings and flaps of the preceeding aircraft. Following the approach of Ortega et al., the influence of the geometry of outboard flap extensions on the wake of a rectangular wing is investigated. Towing tank experiments using Particle Image Velocimetry are conducted, and the induced rolling moment is calculated from the experimentally determined downwash. In addition a stability analysis is performed for the experimentally investigated four-vortex configurations utilizing the linearized Biot–Savart law as proposed by Crow. It is assumed that the vortex system attunes to the most unstable eigenform. The theoretical results help giving an interpretation of the experimental data. It is shown that a significant reduction of the induced rolling moment by more than 50% can be achieved by the use of outboard flap extensions. This is due to an increase of the effective vortex size resulting from the counter-rotating vortex pairs produced by the flaps.

On the Structure and Attenuation of an Aircraft Wake

The results of an experimental investigation of the wake-vortex structure behind a swept and tapered wing with slats and e aps arepresented. The wind-tunnel test is part of a project funded by the German Research Association to investigate the e ow structure of wings with different e ap extensions. The investigation aims at studying the characteristics of lift-generated vortices in order to e nd ways to attenuate dangerous wake vortex encounters for following aircraft during takeoff and landing such that an increase in airport capacity can be achieved. The wake structureismeasuredfordifferente ap settingsbymeansofhot-wireanemometry atdifferent streamwisepositions in the near e eld behind the wing. The measured data are evaluated to examine vortex parameters such as core radius, maximum tangential velocities, circulation distributions, turbulence levels, and maximum induced rolling moments on a following aircraft. In addition, for one e ap setting means of alleviation with a wing e n at different angles of incidence mounted at the outboard e ap edge region are examined. The e n causes a large increase of the vortex core and a reduction of the maximum induced rolling moment within the measured region.

Numerical simulation of roll induced moment of cruciform tactical missiles

This paper presents a numerical evaluation of the induced roll moment of cruciform missiles by a turbulent Navier–Stokes solver. The numerical results have been computed with a high resolution implicit upwind scheme. Comparisons with experiments performed at ONERA show a good agreement as well as a decisive improvement over numerical simulations based on Euler equations. Results are presented for subsonic and supersonic onflow conditions and two different configurations: fuselage-fins and fuselage-wing-fins. Comparative detailed analysis results in the origin of the rolling moment as well as the contributions of the different lifting surfaces.