Aircraft Flap Research Articles

A new active asymmetry monitoring and control technique applied to critical aircraft flap control system failures

Asymmetry limitation requirements between left and right wing flap surfaces play an important role in the design of the implementation of the secondary flight control system of modern airplanes. In fact, especially in the case of sudden breaking of one of the torsion bars of the flap transmission line, the huge asymmetries that can rapidly develop could compromise the lateral-directional controllability of the whole aircraft (up to cause catastrophic occurrences). Therefore, in order to guarantee the aircraft safety (especially during take-off and landing flight phase in which the effects of asymmetries could generate uncontrollable aircraft attitudes), it is mandatory to timely detect and neutralize these asymmetries. The current monitoring techniques generally evaluate the differential angular position between left and right surfaces and, in most the events, limit the Flaps Control System (FCS) asymmetries, but in severe fault conditions (e.g. under very high aerodynamic loads), unacceptable asymmetries could be generated, compromising the controllability of the aircraft. To this purpose, in this paper the authors propose a new active monitoring and control technique capable of detecting the increasing angular error between the different flap surfaces and that, after stopping the whole actuation system, acts on the portion of the actuation line still connected to the PDU to minimize the FCS asymmetries.

Technique for Reducing the Effects of Nonlinear Terms on Electric Field Measurements of Electric Field Sensor Arrays on Aircraft Platforms

AbstractA generalized technique has been developed that reduces the contributions of nonlinear effects that occur during measurements of natural electric fields around thunderstorms by an array of field mills on an aircraft. The nonlinear effects can be due to nearby charge emitted by the aircraft as it acquires and sheds charge, but the nonlinear effects are not limited to such sources. The generalized technique uses the multiple independent measurements of the external electric field obtained during flight to determine and remove nonlinear contaminations in the external vector electric field. To demonstrate the technique, a simulated case with nonlinear contaminations was created and then corrected for the nonlinear components. In addition, data from two different field programs utilizing two different aircraft and field mill configurations, each containing observable and different nonlinear effects, were also corrected for the significant nonlinear effects found in the field mill outputs. The expanded independent measurements in this new technique allow for the determination and correction of components in the field mill outputs from almost any measurable source. Alternate utilization of the technique can include removing effects in the aircraft charging such as aircraft altitude, cloud properties, engine power settings, or aircraft flap deployment. This technique provides a way to make more precise measurements of the true external electric field for scientific studies of cloud electrification.

Degradation of high loaded oscillating bearings: Numerical analysis and comparison with experimental observations

Nowadays mechanical systems are expected to sustain extreme working conditions, due to the increase of the involved power and the optimized design. Moreover special applications like aeronautics, space engineering and robotics require the reduction of the contact area in the contact pairs, which allow the relative motion between components, by increasing the power transmitted per unit of contact area.As a consequence the joint degradations are between the main issues, resulting in possible failures or increase of consumption and maintenance costs. This paper addresses the analysis of the degradation mechanism of oscillating ball bearings subjected to high loads. These bearings, needed to assure the repetitive relative rotation between two members of the mechanical system (e.g. repetitive motions of assembling and manufacturing robots, ship helms, aircraft flaps, etc.), can reach extreme contact pressures at the ball-race contact surfaces; the oscillations of the bearings provide a fatigue loading of the contact area due to the repetitive rotation of the balls between the races. This work is aimed to calculate the contact stress distributions due to the specific boundary conditions, in order to relate them with the bearing degradation. A numerical model is presented to reproduce the loading conditions and calculate the contact stress and strain distributions in order to correlate them with the damages observed experimentally; results obtained by 2D modelling and 3D modelling, for both elastic and plastic properties of the bearing material, are reported and compared. The numerical results and the comparison with tribological observations of a degraded bearing, after cyclic load application, suggest a possible degradation scenario that brings to the failure of the bearing due to subsurface damages.

Flap side edge noise modeling and prediction

This paper presents a model for aircraft flap side edge noise prediction. The flap side edge noise is modeled as two components, respectively for the low and high frequency domain and corresponding to two noise generation mechanisms in the flap side edge flows. The high frequency noise mostly comes from the flow separation near the sharp corners of the flap, mainly in the forward half of the flap chord, and the low frequency component is largely due to the interactions between the large scale vortex structure and the sharp corners of the flap, occurring in the mid and aft half chord region of the flap where the rollup vortex is well established with significant vertical energy to be scattered into sound by the corners. The prediction model is developed from the theory of aerodynamic sound generation, relating the noise spectrum to the statistics of the surface pressure fluctuations in the vicinity of the flap side edges, the characteristic length and time scales in the flow, and the Green's function that accounts for the sound-flow coupling and propagation effects. For both the low and high frequency components, the individual physical mechanisms for the noise generation and propagation are analyzed to develop prediction models for the features of the radiated noise, such as the spectral shapes, the Mach number dependencies and the far field directivities. The individual models are validated by experimental data, showing good agreements between the developed prediction models and data.

High Reliability Permanent Magnet Brushless Motor Drive for Aircraft Application

Reliability is a fundamental requirement in aircraft safety-critical equipments. Its pursuing involves the adoption of protective design concepts such as fault-tolerant or redundant approaches, aiming to minimize mission failure probabilities. Multi-phase motor drives are gaining a growing interest to this extent, because they permit a boost in torque and power density, allowing the design of very compact high efficiency drives with intrinsic fault-tolerant capabilities. This paper presents a five-phase permanent magnet brushless motor drive developed for an aircraft flap actuator application. The motor is designed to satisfy the load specifications with one or two phases open or with a phase short circuited, while a failure in the rotor position sensors is remedied through a sensorless strategy. Design studies aiming to predict the faulty mode performance in case of different remedial strategies are presented. Experimental tests on the drive prototype are included, which confirm its capability to satisfy the planned degraded modes of operation.

Research on Fault Diagnosis Method of the Aircraft Flap System Based on Weighted Fuzzy Petri Net

Accordingly to the problem that during the fault diagnosis of the flap system, the existing experience and knowledge are always uncertainty, inconsistency, incompleteness, what’s more, the different fault cause has different effect on the fault. So the Weighted Fuzzy Petri Net is adopted, and fuzzy Petri net and matrix operations are combined, then the formal process of fuzzy inference reasoning algorithm is researched in order to obtain the fault probability of each event in the fault tree of flap system. The presented method effectively makes up the deficiency of the traditional FTA and FMEA, and provides a new fault diagnosis method for complex systems, so the fault diagnosis method based on Weighted Fuzzy Petri Net has a great practical value.

A Study of Contact Stress on Coatings of Wheel-Slide Using ANSYS

The main purpose of this research is to find a solution for the coating/substrate of the wheel-slide which is the key component of civil aircraft flaps and slats, firstly, the simulation test is designed based on wheel-slide’ s working environment. Further, stress analysis of the coating/substrate is researched, and the results are as the input parameters of the finite element model. Then the finite element software, ANSYS, is applied to simulate the contact stress of coating/substrate in wheel-slide, and the stress’s distribution of coating/substrate are analyzed under different load, different coating thickness, different elastic modulus ratio of coating and other parameters. So the stress field characteristics are gained to guide the conduct of modeling experiments. The results can be applied to explore the failure of mechanism coating/substrate system, and provide theoretical guidance to design the coatings.

Importance Analysis of the Aircraft Flap Mechanism Movement Failure

basic events cannot be properly ranked in the case of the uncertain probability. Contrarily, the new importance measure is on the assumption of the uncertain probability of the basic event. By considering the distribution of the traditional importance measure corresponding to the actual uncertainty of the basic event probability, and using a characteristic value, usually the expectation of the importance measure distribution, to describe the effects, the new importance measure can properly rank the basic events with the uncertainty probability. After the probability density evolution method is employed to solve the new importance measure, the basic event importance is analyzed for the unilateral asymmetric movement failure of the aircraft flap mechanism. The results show that the new importance measure solution based on probability density evolution method is efficient under the acceptable precision. Neglect of the actual uncertainty of the basic event probability may introduce error rank, which is demonstrated by the comparison of the illustrations.

Use of a Novel Fiber Optical Strain Sensor for Monitoring the Vertical Deflection of an Aircraft Flap

The present paper reports the use of a plastic optical fiber-based sensor for elongation measurements in an aircraft flap subjected to different types of flexural loading conditions. The sensor, bonded to the surface of the aircraft structure, relies on measuring the phase shift that occurs between two sinusoidally modulated light signals when the aircraft structure is bent. The light signals are guided through two optical fibers, one of them fixed to the top surface of the flap, and the other one to the bottom surface. The sensor offers good signal stability and repeatability and represents a cost-effective alternative to other more sophisticated health-monitoring systems currently used.

Development of a smart wing

PurposeThe objective of this paper is to develop an actuation system utilizing smart materials such as shape memory alloys (SMA) to control the position of an aircraft's flaps.Design/methodology/approachThe proposed smart wing consisted of SMA springs that were fixed at one end to the wing box toward the leading edge of the airfoil. The other end of each spring was attached tangentially to a rotating cylinder fixed to the flap. The springs were arranged in an upper and a lower layer to cause rotation of the flap in both the upward and downward directions. The spring actuators were controlled by the introduction of heat resulting from the applied current. A prototype of the smart wing was developed and tested to demonstrate the design concept.FindingsA prototype of a smart actuation system for controlling the flaps of an aircraft was successfully developed. Through the experimental and theoretical analyses conducted, the design was validated and showed strong potential for future application.Practical implicationsThe proposed concept can be applied to other aircraft systems such as ailerons, slats, rudders and elevators.Originality/valueThe prototype of a smart wing is unique. It utilizes smart materials for aircraft flap actuation. The concept can be applied on ailerons, slats, rudders and elevators.

Robust control of a shape memory alloy wire actuated flap

In this paper, shape memory alloy (SMA) wire actuators are used to control the flapmovement of a model airplane wing. Conventionally, the flap of an aircraft wing is drivenby electric motors or hydraulic actuators. The use of SMA actuators has the advantage ofsignificant weight reduction. Two SMA actuators are used: one to move the flap up and theother to move the flap down. A sliding mode based nonlinear robust controller isdesigned and implemented on a real-time data acquisition and control platform tocontrol the position of the flap. Feedback control experiments of both positionregulation and tracking control are conducted. To demonstrate the controller’srobustness to uncertainties and disturbances, experiments are conducted withadditional mass on the flap, changing thermodynamic conditions and time varyingaerodynamic loads. Experiments in all cases show that the actual position of theflap closely follows the desired command during experiments. In conclusion, thispaper shows the feasibility of using SMA wire actuators for aircraft flap control.

2.2.1 Model‐based Safety Analysis of a Flap Control System

AbstractFault tree analysis is a widely adopted technique to systematically analyze causes for a given failure of a complex system. Traditionally, a fault tree is constructed top‐down based on knowledge about the structure of the system and the interaction of subsystems. With the increasing system complexity and the accompanying introduction of model‐based development techniques in the industrial process, a substantial amount of this knowledge is laid down in the system models. The main focus of the presented techniques and tools is to automatically exploit this knowledge by extracting a fault tree suitable for FaulTree+ directly from a given design modeled in Statemate. The resulting fault tree is complete wrt. the specified failure, i.e. the analysis considers every possible causal failure combination which is guaranteed by applying model checking techniques. Using an aircraft Flap control system this paper shows how to smoothly integrate the technique into an existing model‐based process.

On noise reduction by flap side edge fences

This paper discusses the effects of fences on side edge cross flows. Experimental data are reviewed to show that fences can modify the cross flow in the side edge region of aircraft flaps in such a way that noise reduction can be achieved. Near field pressure data reveal that fences reduce the characteristic frequency of the flow fluctuations. This leads us to postulate that the fences make the far field noise spectra shift downward in frequency. Because airframe noise spectra typically have a negative slope in the most important mid to high frequency domain, this downward shift in frequency leads to an equivalent reduction of noise levels, when measured by conventional metric such as the effective perceived noise level in aircraft certification. To explain the decrease of the characteristic frequency of the flow fluctuations due to the fences, an analytical model is developed, based on vortex dynamics. The predictions from the model are shown to be in very good agreement with measured data, both for the dominant frequency in the baseline case without the fences and for the trends of the decreasing frequency with increasing fence height.

Surface pressure fluctuations on aircraft flaps and their correlation with far-field noise

This paper discusses unsteady surface pressures on aircraft flaps and their correlation with far-field noise. Analyses are made of data from a 4.7% DC-10 aircraft model test, conducted in the 40 × 80 feet wind tunnel at NASA Ames Research Center. Results for various slat/wing/flap configurations and various flow conditions are discussed in detail to reveal major trends in surface pressure fluctuations. Spectral analysis, including cross-correlation/coherence, both among unsteady surface pressures and between far-field noise and near-field fluctuations, is used to reveal the most coherent motions in the near field and identify potential sources of noise related to flap flows. Dependencies of surface pressure fluctuations on mean flow Mach numbers, flap settings and slat angles are discussed. Dominant flow features in flap side edge regions, such as the formation of double-vortex structures, are shown to manifest themselves in the unsteady surface pressures as a series of spectral humps. The spectral humps are shown to correlate well with the radiated noise, indicating the existence of major noise sources in flap side edge regions. Strouhal number scaling is used to collapse the data with satisfactory results. The effects of flap side edge fences on surface pressures are also discussed. It is shown that the application of fences effectively increases the thickness of the flaps so that the double-vortex structures have more time to evolve. As a result, the characteristic timescale of the unsteady sources increases, which in turn leads to a decrease in the dominant frequency of the source process. Based on this, an explanation is proposed for the noise reduction mechanism of flap side edge fences.

Reliability analysis of structure and control mechanism of aircraft flap

The aircraft flap structure and its control mechanism are considered as an entire system for analysis of the system's reliability. The significant failure modes of a flap structure are based on the following engineering criteria: the magnitude of the computational safety factors of different structural members and the possible patterns that compose failure modes. According to these significant failure modes, the failure probability of a single failure mode can be computed, and the failure probability of the flap structural system can then be solved based on these single failure probabilities. For the flap control mechanism, besides the static strength reliability, it is necessary to compute the function reliability. The reliability of the control mechanism can be synthesized from the above two kinds of reliabilities. Finally, the entire flap system failure probability can be solved by synthesis of the flap structural system failure probability and the flap control mechanism failure probability.

A comparison of honeycomb-core and foam-core carbon-fibre/epoxy sandwich panels

Cellular core sandwich panels of carbon-fibre/epoxy fabric laminate skins, simulating the construction of an aircraft flap, were cured and bonded in a single-step autoclave operation. Nomex honeycomb and Rohacell WF foam of different densities were employed as the core material. The panels were examined to identify voids in the laminate skins and cell collapse and coalescence in the foam core. Test-pieces were subjected to low-energy impact and the induced damage was determined by ultrasonic C-scan. The maximum damage area in the face-skin was comparable to the projectile cross-sectional area. Residual compressive capacity showed an asymptotic decrease with increasing impact levels, most panels gave similar values but the modes of failure were different depending on the type of core. The resistance of the separate panel components to full impact puncture was assessed using instrumented impact load-deflection traces.

Airframe noise using experimental, cross‐correlation techniques and a theoretical, vortex model

Correlation measurements have been made on large scale models placed in wind tunnels. The correlation has been taken between surface pressures and farfield radiated sound. The technique permits one to find the amount of sound radiated by a surface even when that sound is up to 20 dB and more below the ambient noise in the operating wind tunnel. These techniques have strongly suggested that the major source of radiated airframe noise is at the outboard edges of deployed aircraft flaps. Measurements have been made to determine the dependence upon flow Mach number, upon angle, and upon the various flap‐surface positions. A theoretical model has been devised using a vortex (and its image) embedded in a potential flow about a corner. The theory has but one adjustable parameter; all computations can be made on a hand calculator. The radiated sound is found using the Curie theory. The effects of many vortices are included. The predicted, radiated sound matches the correlation‐measured sound to within 3 dB for mo...

On the sonic fatigue life estimation of skin structures at room and elevated temperatures

Fighter/trainer empennages and STOL (Short Take-Off and Landing) aircraft flap systems are subjected to severe acoustic pressure levels as high as 150–170 dB. As a result, acoustic fatigue has become one of the major factors in design. Empennages and flap systems are also subject to high temperatures and thus the influence of thermal buckling on fatigue life must be taken into consideration. To estimate the sonic fatigue life of skin structures, combined use is made of the Monte Carlo method of non-linear panel response analysis and local stress-strain simulations with rheological models. Calculations were conducted and comparison of the results with experimental data shows that the method estimates the sonic fatigue life accurately when adequate values of fatigue notch factors are chosen. Example calculations with thermal as well as static pressure effects were also carried out and the effect of thermal buckling on sonic fatigue life is clarified.

Application of Electronic Data Processing Airport Analysis in Airlines Operations and for Manufacturers

An analysis of the factors affecting aircraft takeoff and landing weights is presented. This analysis, made with the aid of a computer, considers all factors, including: type of aircraft; weight; temperature; wind; engines; payload; policies; limits; runway slope, conditions, and length; obstructions; and aircraft flap settings.

Structural Design

THE structural design of the BAC One‐Eleven generally follows closely that of the Vickcrs Vanguard and the VC10†—involving the use of a considerable number of integrally‐machined components. As a short‐haul aircraft the average time per flight of the BAC One‐Eleven is expected to be of the order of 45 min. during which period full cabin pressure differential will be attained, speeds of the order of its design cruising speeds will be achieved and the undercarriage and flaps will be operated for take‐off and landing. Based on current estimations this involves a design aim of a minimum crack‐free life of 40,000 flights, landings and take‐offs —a much more severe requirement than that for the long‐range subsonic jets. Critical areas of the aircraft (undercarriage, flaps, tailplane and cabin pressure skins) are thus designed on fatigue considerations related principally to the number of flights made. The accent has therefore been placed on building a rugged structure which is easy to maintain and has a long service life. Small amounts of additional weight, properly disposed, can effect large improvements in the service life, particularly necessary on a short‐haul aircraft, and although weight saving is always of prime importance it must be balanced by other factors—especially in the primary structure.