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Three-dimensional (3D) microdisplacement monitoring plays a crucial role in the assembly of large aircraft. This paper presents a broadly applicable high-precision online 3D microdisplacement monitoring method and system based on proximity sensors as well as a corresponding in situ calibration method, which can be applied under various extreme working conditions encountered in the aircraft assembly process, such as compact and obstructed spaces. A 3D monitoring model is first established to achieve 3D microdisplacement monitoring based only on the one-dimensional distances measured by proximity sensors, which concerns the extrinsic sensor parameters, such as the probe base point (PBP) and the unit displacement vector (UDV). Then, a calibration method is employed to obtain these extrinsic parameters with high precision by combining spatial transformation principles and weighted optimization. Finally, calibration and monitoring experiments performed for a tailplane assembly process are reported. The calibration precision for the PBP is better than ±10 μm in the X and Y directions and ±2 μm in the Z direction, and the calibration precision for the UDV is better than 0.07°. Moreover, the accuracy of the 3D microdisplacement monitoring system can reach ±15 μm. In general, this paper provides new insights into the modeling and calibration of 3D microdisplacement monitoring based on proximity sensors and a precise, efficient, and low-cost technical means for performing related measurements in compact spaces during the aircraft assembly process.

Three-dimensional (3D) profile scanning plays a crucial role in the inspection of assembled large aircraft. In this paper, to achieve noncontact automatic measurements of the high-reflective profiles of large-scale curved parts and components, an automated noncontact system and method with high accuracy and high efficiency are presented. First, a hybrid 3D coordinate measurement system based on proximity sensors and cameras is proposed to obtain noncontact measurements while avoiding the influence of high reflection on the measurement accuracy. A hybrid measurement model that combines the one-dimensional distances measured by the proximity sensors and the 3D information obtained by cameras is proposed to determine high-accuracy 3D coordinates of the measured points. Then, a profile-driven 3D automated scanning method and strategy are designed to rapidly scan and reconstruct the profile within the effective range without scratching the profile or exceeding the measurement range of the proposed system. Finally, experiments and accuracy analyses are performed in situ on an assembled tailplane panel (approximately 1760 mm × 460 mm). The automated scanning process is completed in a timeframe of 208 s with an average error of less than 0.121 mm for profile reconstruction. Therefore, the proposed method is promising considering both the high accuracy and high efficiency requirements of profile inspections for large aircraft.

Formation flying techniques, demonstrated by migratory birds, have the potential to significantly save energy when two or more transport aircraft fly the same route at the same time. In fuel-saving...

Any aircraft in flight is subjected to dynamic loads. Following vibration-related accidents, a flow field and vibration analysis can be carried out to analyze the data and study the cause of the accident. When slit airflow enters the cavity between the tailplane structure and the elevator, a mixed vortex is formed. If the vortex-induced vibrational frequency of around 50 Hz happens to be close to the natural frequency of the structure at 46 Hz, it is likely to induce structural vibration (resonance). The resonance can cause excessive fatigue damage which can ultimately lead to structural failure and the loss of the component or the aircraft. Damping methods can be employed to control vibration within the structure by reducing the amplitude of that vibrational motion by 83%. This article details a recreation of one example of structural vibration within an airborne aircraft.

This paper presents the design of an automatic deep stall recovery algorithm for large transport aircraft using optimal trajectory planning. Deep stall is a condition where an aircraft is trapped in a nose-high stall condition and its elevators cannot produce enough nose-down pitching moment to recover the aircraft from the stall. The NASA Generic Transport Model (GTM) is used as the basis for the design and verification of the system. The aerodynamic model of the NASA GTM simulation model is modified to exhibit deep stall behaviour. Simulations are performed to show that the modified aircraft model can be pushed into deep stall, and cannot be recovered using elevator actions only. The deep stall recovery task is formulated as an optimal path planning problem and solved using an A* search algorithm to find the optimal sequence of control actions and the resulting optimal state trajectory to escape from the deep stall. The A* algorithm performs the planning using a simplified, three-degrees-of-freedom (3DOF) aircraft model that models only the fast rotational dynamics. The automatic deep stall recovery is then verified in simulation using the full six-degrees-of-freedom (6DOF) NASA GTM aircraft model. The simulation results show that the system successfully recovers the aircraft from deep stall. The optimal sequence of control actions first uses the rudder to yaw the horizontal tailplane out of the aircraft’s own wake to regain elevator effectiveness, and then uses the elevators to pitch the nose down and recover from the stall.

Aircraft handling qualities may be influenced by wing-tip flow separations and horizontal tail (HT) reduced efficiency caused by loss of local dynamic pressure or local tailplane flow separations in high angle-of-attack manoeuvres. From the flight tester’s perspective, provided that the test aircraft presents sufficient longitudinal control authority to overcome an uncommanded nose-up motion, this characteristic should not be a safety factor. Monitoring and measuring the local airflow in the aircraft’s HT provides information for safe flight-test envelope expansion and data for early aerodynamic knowledge and model validation. This work presents the development, installation and pre-flight calibration using computational fluid dynamics (CFD), flight-test calibration, results and benefits of differential pressure based local angle-of-attack and total pressure measurements through 20 static pressure ports and a Kiel pitot. These sensors were installed in a single-aisle, four-abreast, full fly-by-wire medium-range jet airliner with twin turbofan engines and conventional HT (low vertical position).

In this combined experimental and numerical study, the propeller–airframe aerodynamic interaction is characterized for an aircraft configuration with propellers mounted to the horizontal tailplane. The contributions of the propeller and airframe to the overall loading are distinguished in the experimental analyses by using a combination of external balance and internal load cell data. Validated computational fluid dynamics simulations are then employed to quantify the interaction at a component level. The results show that the propeller installation shifts the neutral point aft with increasing propeller thrust. For the configuration considered herein, the yawing moment due to sideslip is increased by approximately 10%, independent of the propeller thrust coefficient. The changes in propeller loading due to the airframe-induced flowfield are the dominant factor to change the airframe stability and performance. The prominent installation effects occur at high angle of attack, because in that condition the propeller experiences a significant nonuniform inflow that affects the propeller and tailplane. The relatively large propeller diameter compared with tailplane span leads to a change of the tailplane root vortex that causes the tailplane effectiveness to reduce with an inboard-up rotating propeller.

Until 2017, there was no certainty about the fatigue life of Hawk tailplanes in FINAF’s flight conditions. Then full-scale fatigue tests were performed to determine if the FINAF is required to procure more tailplanes, and to extract evidence, which could be used to increase the structural inspection interval times. The tests were executed with two 4000 FH flown tailplanes and the goal was to achieve additional 2000 FH with a scatter factor of 5. Test loads were applied with actuators feeding both buffeting and maneuvering symmetrically at the same time. Test’s spectrum was based on the FINAF OLM strains and on the usage spectrum of the FINAF flights 2014–2015. Limited NDIs were done after every 200–340 EFH and full inspections after every 1000 EFH. Several damages, such as broken rivets and cracks in spars and angles, arose. Following the testing, the tailplanes were subjected to RSTs with the load corresponding the ultimate design load. The tailplanes passed the RSTs without noticeable additional damages. Centre sections were torn down for more detailed inspections. Some fault indications were obtained from the buttstraps, but all the defects were very small. Seven cracks were found on the skins and one location could be determined as the critical location. The centre joint survived the test period. The residual strength was sufficient with a 20 mm crack at the skin rivet hole, which was estimated to be the most loaded. The tests gave solid basis for increasing the TP’s acceptable usage life by 1000 FH. It was possible to determine the crack propagation rate to verify the structural inspection period to be applied. Considerable cost savings will be achieved, because the inspections can now be optimized. In addition, now it is known that the current number of TPs is sufficient with the additional 1000 FH for the targeted HW life cycle, and no additional procurement is required.

Hybrid RANS/LES simulations of the flow around the NASA Common Research Model aircraft configuration were carried out with the focus on understanding the interaction of the separated wake with the tailplane in the presence of massively separated flow on the main wing. Validation of the CFD data using PIV data obtained for the flow conditions at \(\alpha =16^{\circ }\), \(\alpha =18^{\circ }\) and \(\alpha =20^{\circ }\) was carried out, confirming the generally satisfactory performance of the DDES simulations observed in earlier publications. As a next step, the wake characteristics and tailplane forces were evaluated for three angles of attack in order to investigate the flow dynamics in low speed stall. The separation characteristics were found to vary over the span. The wake size and downwash direction varied significantly with higher values of \(\alpha \). The altered wing downwash influenced the tailplane inflow, with the load fluctuations on the latter being significantly affected by the amount of turbulent kinetic energy present in the wake.

Purpose This paper aims to present and discuss the requirements for flying targets which sometimes are contradictory to each other and to perform a trade-off analysis before the design activity is started. It also aims to demonstrate conceptual and preliminary design processes using a practical example of PW-61 configuration and to show how results of experimental flight tests using a scaled flying target will be described and analyzed before manufacturing the full scale flying target. Design/methodology/approach An important part of the paper consists of the selection of tailplane configuration of the flying target UAV to protect some expensive on-board systems against serious damages and to obtain a sufficient dynamic stability, independently of the amount of the petrol in fuel tank. Inverted V-tail, U-tail and H-tail configurations were considered and compared both, theoretically and in-flight experiments. Findings Flight dynamics models and associated computational procedures were useful both in a preliminary design phase and during the final assessment of the configuration after flight tests. Selection of the tailplane configuration for the flying target UAV is very important to protect some expensive on-board systems against serious damages and to obtain a sufficient dynamic stability, independent of the amount of the petrol in fuel tank. Practical implications Flying targets should be speedy, maneuverable, cheap, easy in deployment and multi-recoverable (if not destroyed by live ammunition), must have relatively low take-off weight and an endurance of at least 1 h. This paper can be useful for proper selection of requirements and preliminary design parameters to make the design process more economically effective. Originality/value This paper presents very efficient methods of assessing the design parameters of flying targets, especially in an early stage of the design process. Stability computations are performed based on equations of motion and are supplemented by flight tests using the scaled flying models. It can be considered as an original, not typical, but very practical approach because it delivers lots of data in the early design stages at relatively low cost.

The behavior of the local correlation-based \(\gamma -Re_\theta \) transition model outside the boundary layer was investigated. It is shown that for perturbed vorticity distributions in the wake the intermittency drops below one. To investigate and quantify this effect, a test case consisting of two NLF(1)-0416 airfoils was simulated imitating a wing tailplane interaction of an aircraft. As a remedy two model modifications are proposed to eliminate the undesired effects. Besides, based on Reynolds-Averaged Navier-Stokes computations a theoretical analyses was performed identifying the stability properties of the original and modified \(\gamma \) equation.

Lightly-damped mechanical systems exhibit strong resonant behaviour which could potentially result in life-threatening situations. To prevent these situations from happening, frequency response function measurements are essential to accurately characterise the resonant modes of the mechanical system. Unfortunately, these measurements are distorted by leakage and (long) transient phenomena. Local modelling techniques have been introduced in the past to resolve these complications but either they do not use the correct model structure or they introduce a bias. This paper proposes a local rational modelling technique which completely removes the bias from the estimation procedure and is applicable to large-scale multiple-input, multiple-output systems. The developed technique uses the bootstrapped total least squares estimator which provides unbiased estimates for the local rational model and generates accurate uncertainty bounds for the obtained non-parametric frequency response function. The proposed technique is successfully verified using a simulation example of a large-scale system which contains 100 resonances and has 100 inputs and 100 outputs. Its practical applicability is illustrated by characterising the resonant behaviour of the tailplane of a glider.

Aircraft performance can be assessed and improved by considering the key variables linked toweight and aerodynamics in the Breguet range equation. In this paper, the authors present a methodfor wing design that allows a reduction in induced drag and minimization of the weight associated withthe aircraft’s vertical tailplane, whilst ensuring desirable lateral-directional flight dynamics. The usecase is a general aviation aircraft for which the wing has been modified using Prandtl’s 1933 approachwhere the span constraint is removed to yield a non-elliptic lift distribution. It is shown that such a liftdistribution also contributes to the aircraft’s lateral-directional stability and as a result, the size andweight of the vertical tail can be reduced. This study was carried out using an analytical frameworkthat combines early design tools such as XFOIL and AVL deemed to be adequate for subsonic flight.Both cruise and approach configurations are considered. Wing twist distribution and span extensionhave been calculated using lifting line theory. The study demonstrates the design trade-off neededbetween flight dynamic modes, such as the Dutch roll mode, and vertical tailplane size when theaircraft is equipped with a wing designed to generate a non-elliptic lift distribution. It is shown thatthis approach allows a 14% improvement in the lift to drag ratio with 44.34% reduction in V-tailweight. These yield a total of 17% improvement in aircraft range. As for the approach phase it shareall the characteristics observed in cruise with the difference that Dutch roll mode is stable for almostall the smaller size of V-tail. Further work requires to focus on the placement of ailerons to removeadverse yaw tendencies.

WOS: 000491169100006, Purpose The purpose of this paper is to rise the autonomous flight performance of the small unmanned aerial vehicle (UAV) using simultaneous tailplane of UAV and autopilot system design. Design/methodology/approach A small UAV is remanufactured in the UAV laboratory. Its tailplane can be changed before the flight. Autopilot parameters and some parameters of tailplane are instantaneously designed to maximize autonomous flight performance using a stochastic optimization method. Results found are applied for simulations. Findings Benefitting simultaneous tailplane of UAV and autopilot system design process, autonomous flight performance is maximized. Research limitations/implications - Authorization of Directorate General of Civil Aviation in Turkey is required for UAV flights. Practical implications - Simultaneous tailplane and autopilot system design process is so useful for refining UAV autonomous flight performance. Social implications - Simultaneous tailplane and autopilot system design process fulfills confidence, high autonomous performance, and easy service demands of UAV users. By that way, UAV users will be able to use better UAVs. Originality/value Creating a novel technique to recover autonomous flight performance (e.g. less overshoot, less settling time and less rise time during trajectory tracking) of UAV and developing a novel procedure performing simultaneous tailplane of UAV and autopilot system design idea., Research Fund of Erciyes University Scientific Research Projects (BAP) Coordination Unit [FBA-2018-6712]; Research Fund of The Scientific and Technological Research Council of Turkey (TUBITAK)Turkiye Bilimsel ve Teknolojik Arastirma Kurumu (TUBITAK) [114M856], This work was supported by the Research Fund of Erciyes University Scientific Research Projects (BAP) Coordination Unit under Project Number: FBA-2018-6712.; This work was also supported by the Research Fund of The Scientific and Technological Research Council of Turkey (TUBITAK) under Project Number: 114M856.; The author would like to also recognize Assoc. Prof. Dr Tugrul OKTAY, Assist. Prof. Dr Mehmet KONAR and PhD Candidate Alpertunga CEYLAN for their supports during UAV manufacturing and UAV flight testing.

Experimental data are presented for the parasite drag of various helicopter fuselage components, such as skids, external fuel tanks, and tailplane. The experiments were conducted at the Kazan National Research Technical University (Kazan Aviation Institute) T-1K wind tunnel, investigating four versions of a fuselage similar to the Ansat helicopter. It was found that, for the range of pitch angles −10≤α≤10 deg−10≤α≤10 deg, the skids added 80% to the drag of the bare fuselage, whereas the tailplane increased the drag by 20%. At the same conditions, external fuel tanks were found to add 48% to the clean fuselage drag. A simple rotor hub with a tail support added 74% to the bare fuselage in the range of pitch angles −8≤α≤6 deg−8≤α≤6 deg. Streamlining the rear fuselage was found to reduce the drag by 16% over the range of pitch angles −10≤α≤10 deg−10≤α≤10 deg. Apart from the parasite drag, ideas for drag reduction are also discussed.

Tangential blowing over the shoulder of a deflected rudder is applied on a vertical tailplane. For a large rudder deflection angle and without blowing the flow is mostly separated on the rudder. Three configurations are investigated in a numerical study with the aim to increase the side force coefficient. The first one is the baseline without blowing, the second one has a continuous full span slot and for the third one discrete slots are used. With sufficient blowing through the continuous slot the separation on the rudder can be removed completely while the separation extent is greatly reduced when using discrete slots. This is investigated for different sideslip angles. An approximately linear increase in the side force coefficient can be found until a sudden side force breakdown occurs. In a further study, the jet blowing velocity is varied. A smaller jet velocity leads to a smaller increase in the side force coefficient. Comparing the continuous and discrete slot configurations shows that for a similar increase in the side force coefficient a much smaller mass flow rate is needed for the discrete slots. However, for jet velocities below the sonic speed the increase in the side force coefficient is limited. It can only be increased by a larger slot extent in spanwise direction but this comes at the expense of an increase in the required mass flow rate.

This paper provides an overview of current activities of DLR (German Aerospace Center) with respect to stability and control investigations in the context of early stages of aircraft design. For this purpose, DLR follows an interdisciplinary and multi-level design approach. Using an integration framework in combination with a central data exchange format, largely automated process chains are set up that combine calculation and simulation capabilities of the multitude of disciplines required in early aircraft design. Rather than using empirical relations and assumptions based on experience, the underlying methods applied by the tools are mainly based on physical model representations. The major aim of this design approach is to generate all relevant data needed for stability and control investigations, including aerodynamic damping derivatives and to assemble them within a flight dynamics model. Not only does this approach allow for an early consideration of stability and control characteristics, but it also respects interdisciplinary effects and enables automated design changes. This paper describes the infrastructure used for setting up the described process. It presents disciplinary tools used to calculate engine performance maps, calculate aerodynamic performance maps and structural properties, generate flight dynamics models with associated control laws and to assess aircraft handling qualities. Furthermore, this paper provides application examples of early stability and control considerations, using integrated interdisciplinary process chains. This comprises a handling qualities assessment under uncertainty considerations and vertical tailplane sizing for a blended wing body. In addition, engine and split flap sizing processes for an unmanned combat aerial vehicle are shown. The interdisciplinary design approach presented here, serves to find a well justified early configuration and reduces the risk of later design changes.

This paper addresses the effects of propeller installation on the aerodynamic performance of a tailplane featuring tip-mounted propellers. A model of a low aspect ratio tailplane equipped with an elevator and a tip-mounted propeller was installed in a low-speed wind-tunnel. Measurements were taken with an external balance and surface pressure taps to determine the aerodynamic characteristics of the tailplane, while the flowfield in the wake of the model was investigated using particle-image velocimetry. The experimental data are supported by CFD analyses, involving both transient simulations of the full-blade configuration and steady-state simulations the propeller replaced by an actuator-disk model. The upstream effects on the propeller time-average and time-accurate thrust and normal-forces are found to be limited for different tailplane operating conditions. It is shown that for a given propeller rotation direction, the load distribution on the tailplane is highly dependent on the direction of elevator deflection. The rotation direction of the tailplane tip-vortex relative to the propeller swirl therefore significantly affects the integral loads on the tailplane, resulting in differences in the normal-force gradient and elevator effectiveness.

Aircraft flight dynamic characteristics can be greatly changed by ice accretion, which has been considered a considerable threat to aircraft flight safety for a long time. An overview of the studies on several ice accretion effects on aircraft flight dynamics is presented here. Special attention is paid to the following areas: ways to obtain the aerodynamic data of iced aircraft, flight dynamic modeling and simulation for iced aircraft, effects of ice accretion on aircraft stability and control as well as on flight performance and aircraft icing envelope protection and control adaption. Finally based on the progress of existing research in these areas, some key issues which deserve more attention for researchers to resolve are addressed, including obtaining aerodynamic data of iced aircraft through numerical simulation method, consummating the existing calculation models about effects of ice accretion on aircraft aerodynamic derivatives and enhancing the investigation on problems of tailplane ice accretion.

WOS: 000418465900003, In this article, simultaneous longitudinal and lateral flight control systems design for both passive and active morphing tactical unmanned aerial vehicles (TUAVs) is first time applied for autonomous flight performance maximization in the literature. For this purpose longitudinal and lateral dynamics modelling of TUAVs produced in Erciyes University, Faculty of Aeronautics and Astronautics, Model Aircraft Laboratory are considered in order to obtain simulation environments. Our produced TUAV is called as ZANKA-III which has weight of 50 kg, range of around 3000 km, endurance of around 28 hour, and ceiling altitude of around 12500 m. Von-Karman turbulence modelling is used in order to model atmospheric turbulence during flight in both longitudinal and lateral simulation environments. A stochastic optimization method called as simultaneous perturbation stochastic approximation (i.e. SPSA) is also first time applied in order to obtain optimum dimensions of morphing parameters (i.e. extension ratios of wingspan and tailspan, assembly positions of wing and tailplane to fuselage) and optimum magnitudes of longitudinal and lateral controllers' gains (i.e. P, I and D gains) while minimizing cost index capturing terms about both longitudinal and lateral autonomous flight performances and there exists lower and upper constraints on all optimization variables in the literature., Scientific and Technological Research Council of Turkey (TUBITAK) [115M603], This research was funded by a grant (No. 115M603) from the The Scientific and Technological Research Council of Turkey (TUBITAK).

This paper presents a conceptual design study of hybrid electric turboprop aircraft configurations for entry into service in the year 2035. A design mission of 800 nm and 70 passengers was the focus of the study. The hybrid electric architecture was used as an enabler for advanced configurations and improved propulsive efficiency. The main energy source for cruise flight is conventional fuel, batteries are used only to cover peak loads. A design process with four phases covering a wide design space was performed: 1) initial brainstorming phase (48 concepts identified), 2) qualitative down-selection based on expert judgment (reduction to 17 concepts), 3) initial quantitative analysis with simple spreadsheet methods (reduction to four concepts), and 4) evaluation of the four most promising concepts with classical conceptual design methods enhanced by more detailed disciplinary analysis when required. The most promising concept after all phases consisted of a parallel hybrid architecture with a high aspect ratio wing, electrically driven propellers at the wingtip and reduced vertical tailplane. A mission fuel reduction of 4% compared to an advanced EIS 2035 turboprop without hybrid electric propulsion was achieved with this concept. Other promising configurations included a tailless concept and a distributed propulsion configuration.

Numerical simulations are conducted on the DHC-6 Twin Otter wing and horizontal tailplane to explore the aerodynamic penalties that affect airfoil performance in heavy rain conditions. An Eulerian-Lagrangian two-phase flow method is used to simulate the aerodynamic performance of the wing under rainfall conditions. The air is regarded as continuous phase solved by the Reynolds averaged Navier-Stokes (RANS) equations. The rain is regarded as discrete phase solved by the particle equations of motion in the Lagrangian reference frame. Our simulation results agree well with the experimental data and show obvious regularity. The lift decreases and the drag increases under the influence of heavy rain. The wing experiences more severe aerodynamic penalties than the horizontal tailplane at the same rain condition. The influence of rainfall leads to premature stall and separation.

— In this conference article, combined passive and active morphing approach is applied on tactical unmanned aerial vehicles (TUAVs) for autonomous flight performance maximization. For this intention lateral dynamic modeling of TUAVs manufactured in Erciyes University, Faculty of Aeronautics and Astronautics, Model Aircraft Laboratory is investigated in order to obtain lateral state-space model and a simulation model. Our manufactured TUAV is named as ZANKA-III which has weight of 50 kg, range of around 3000 km, endurance of around 28 hour, and ceiling altitude of around 12500 m. Von-Karman turbulence modeling is applied in order to capture atmospheric turbulence in lateral simulation environment. A stochastic optimization method called as simultaneous perturbation stochastic approximation (i.e. SPSA) is used in order to get optimum dimensions morphing parameters (i.e. extension ratios of wingspan and tailspan, assembly positions of wing and tailplane to fuselage).

On a swept vertical tailplane with infinite span tangential blowing over the shoulder of a deflected rudder is applied. For large rudder deflection angles the flow on the rudder is separated without blowing. A numerical study is conducted with the aim to increase the side force coefficient which might be required for a take-off condition if a one-sided engine failure occurs. With a continuous slot and sufficient mass flow rate the separation on the rudder can be reduced or avoided. It is shown that by using discrete slots this can be achieved for a similar side force coefficient gain with a smaller momentum coefficient. In addition the sweep angle of the incoming flow is varied showing a strong impact on the achievable side force coefficient. This is also true for the curvature of the rudder shoulder over which the jet is blown.

Some helicopter manufacturers are exploring the compound helicopter design as it could potentially satisfy the new emerging requirements placed on the next generation of rotorcraft. It is well understood that the main benefit of the compound helicopter is its ability to reach speeds that significantly surpass the conventional helicopter. However, it is possible that the introduction of compounding may lead to a vehicle with significantly different flight characteristics when compared to a conventional helicopter. One method to examine the flight dynamics of an aircraft is to create a linearised mathematical model of the aircraft and to investigate the stability derivatives of the vehicle. The aim of this paper is to examine the stability derivatives of a compound helicopter through a comparison with a conventional helicopter. By taking this approach, some stability, handling qualities and design issues associated with the compound helicopter can be identified. The paper features a conventional helicopter and a compound helicopter. The conventional helicopter is a standard design, featuring a main rotor and a tail-rotor. The compound helicopter configuration features both lift and thrust compounding. The wing offloads the main rotor at high speeds, whereas two propellers provide additional propulsive thrust as well as yaw control. The results highlight that the bare airframe compound helicopter would require a larger tailplane surface to ensure acceptable longitudinal handling qualities in forward flight. In addition, without increasing the size of the bare airframe compound helicopter’s vertical fin, the Dutch roll mode satisfies the ADS-33 level 1 handling qualities category for the majority of the flight envelope.

© 2014 Elsevier Ltd.Flutter of T-tail configurations is caused by the aeroelastic coupling between the vertical fin and the horizontal stabiliser. The latter is mounted on the fin instead of the fuselage, and hence the arrangement presents distinct characteristics compared to other typical empennage setups; specifically, T-tail aeroelasticity is governed by inplane dynamics and steady aerodynamic loading, which are typically not included in flutter clearance methodologies based on the doublet lattice method. As the number of new aircraft featuring this tail configuration increases, there is a need for precise understanding of the phenomenon, appropriate tools for its prediction, and reliable benchmarking data. This paper addresses this triple challenge by providing a detailed explanation of T-tail flutter physics, describing potential-flow modelling alternatives, and presenting detailed numerical and experimental results to compensate for the shortage of reproducible data in the literature. A historical account of the main milestones in T-tail aircraft development is included, followed by a T-tail flutter research review that emphasises the latest contributions from industry as well as academia. The physical problem is dissected next, highlighting the individual and combined effects that drive the phenomenon. Three different methodologies, all based on potential-flow aerodynamics, are considered for T-tail subsonic flutter prediction: (i) direct incorporation of supplementary T-tail effects as additional terms in the flutter equations; (ii) a generalisation of the boundary conditions and air loads calculation on the double lattice; and (iii) a linearisation of the unsteady vortex lattice method with arbitrary kinematics. Comparison with wind-tunnel experimental results evidences that all three approaches are consistent and capture the key characteristics in the T-tail dynamics. The validated numerical models are then exercised in easy-to-duplicate canonical test cases. These parametric studies illustrate the impact of well-known factors in T-tail flutter, namely horizontal tailplane dihedral, flexibility and static deformations. In addition, scenarios are exposed in which the stability behaviour is dictated by typically second-order effects, such as chordwise forces and quadratic modes, revealing drastically different qualitative flutter curves. It is also shown that there is a distinction between angle of attack of the whole tail assembly and incidence of the horizontal tailplane relative to the fin, which might yield very counterintuitive trends depending on the configuration parameters. The paper concludes with flight test results of the Airbus A400M, epitome of modern T-tail aircraft. Tests performed in a wake-vortex encounter campaign complement the virtually nonexistent literature in the topic, demonstrate how T-tail effects can be measured in flight and restate the adequacy of potential-flow models for T-tail flutter prediction.

The present work is concerned with the accurate modeling of transport airplanes. This is of primary importance to reduce aircraft development risks and because multi-disciplinary design and optimization (MDO) frameworks require an accurate airplane modeling to carry out realistic optimization tasks. However, most of them still make use of tail volume coefficients approach for sizing horizontal and vertical tail areas. The tail-volume coefficient method is based on historical aircraft data and it does not consider configuration particularities like wing sweepback angle and tail topology. A methodology based on static stability and controllability criteria was elaborated and integrated into a MATLAB application for airplane design. Immediate advantages with the present methodology are the design of realistic tail surfaces and properly sized airplanes. Its validation was performed against data of five airliners ranging from the regional jet CRJ-100 to the Boeing 747-100 intercontinental airplane. An existing airplane calculator application incorporated the present tail-sizing methodology. In order to validate the updated application, the Fokker 100 airliner was fully conceptually designed using it.

This article presents aerodynamic modelling and real-time control of a 1-degree-of-freedom free to pitch Tailplane. The system is designed to serve as an experimental test facility for investigating flight dynamics principles, model validation and different feedback control paradigms. A high-fidelity plant model is an important first step in many flight-related applications such as control system design, analyses and pilot training. To achieve these objectives, a detailed study is conducted employing analytical as well as system identification (SI) techniques. Analytical approach although less accurate complements SI process. This synergy is exploited along with statistical and time domain tests to arrive at a high-fidelity model. It is demonstrated that such an integrated approach is suitable for modelling a class of unmanned air vehicles. The SI model is then employed for controller synthesis. Finally, real-time pitch control under stick command is demonstrated utilizing classical proportional integrato...

In this paper, a novel aeroservoelastic effector configuration that is actuated by piezoelectric tabs is presented. The effector exploits trailing-edge tabs installed on free-floating flaps (FFFs). These flaps are used to prevent flutter from occurring and to alleviate loads originating from external excitations such as gusts. A vertical tailplane wind-tunnel model with two free-floating rudders and a flutter control mechanism were designed, and the aeroelastic stability and response characteristics have been modeled numerically. The controller uses the tailplane tip acceleration as a sensor and sends control signals to the piezoelectrically actuated tabs. Wind-tunnel experiments were performed to demonstrate the feasibility of the technology. It was demonstrated experimentally that the flutter speed associated with the free rudders could be increased by 80%. The same controller, applied to the external rudder, was used to alleviate the aeroelastic response of the tailplane to the excitation of the other ...

Marques P, Maligno A, Dierks S and Bachouche A. (2013). Flight dynamics principles of canard aircraft: Implications for UAV configuration decision. International Journal of Unmanned Systems Engineering. 1(2): 12-30. The canard configuration is an appealing alternative to the conventional aft tailplane stabiliser in the design of novel unmanned aerial vehicle (UAV) concepts. A canard foreplane increases the maximum lift coefficient and reduces the trim drag of the aircraft, making the canard set up particularly suitable for high-altitude long-endurance (HALE) UAVs. The inherent flight instability of canard-configured aircraft can be exploited to augment the manoeuvrability of advanced military UAVs at high angles of attack. Canard-wing arrangements result in complex strongly-coupled aerodynamic flow fields that include upwash-downwash effects, vortex-vortex interactions and vortex-surface interactions. This review paper provides a historical overview of the development of the canard aircraft, compares different canard airframes, explores the aerodynamics and aeroelasticity of canard aircraft, assesses longitudinal stability in conventional and canard arrangements, evaluates the aerodynamic efficiency of the tandem wing configuration, and gives examples of canard UAV platform development. The aerodynamic features of the Buraq HALE UAV canard prototype by Tunisia Aero Technologies Industries – UK are outlined. The paper also discusses the advantages and disadvantages of the canard configuration that inform the decision process in the development of next-generation UAV configurations. © Marques Aviation Ltd – Press.

On a swept vertical tailplane with infinite span tangential blowing over the shoulder of a deflected rudder is applied. For large rudder deflection angles the flow on the rudder is separated without blowing. A numerical study is conducted with the aim to increase the side force coefficient. This could, for example, be required during take-off if the engine on one side fails, necessitating the compensation of a large yawing moment. If this criterion is critical for the sizing of the vertical tailplane, active flow control like tangential blowing could help to reduce the size of the vertical tailplane and thus save weight and fuel. With a continuous slot it is demonstrated that the separation on the rudder can be reduced or avoided. It is shown that using discrete slots this can be achieved with a smaller momentum coefficient. To analyze the effects of the discrete slots and their jets on each other and their effectivity with regard to the gain in side force coefficient a parameter study is conducted. The number of slots as well as the size of the slots in spanwise direction is varied and the impact of Jet velocity changes is also studied. In comparing the results for a constant increase in side force and constant slot size in spanwise direction the configuration with the smaller number of slots but a higher jet velocity proved to be the most effective one.

Stability is as essential to flying as lift itself, but previous discussions of how flying animals maintain stability have been limited in both number and scope. By developing the pitching moment equations for gliding animals and by discussing potential sources of roll and yaw stability, we consider the various sources of static stability used by gliding animals. We find that gliding animals differ markedly from aircraft in how they maintain stability. In particular, the pendulum stability provided when the centre of gravity lies below the wings is a much more important source of stability in flying animals than in most conventional aircraft. Drag-based stability also appears to be important for many gliding animals, whereas in aircraft, drag is usually kept to a minimum. One unexpected consequence of these differences is that the golden measure of static pitching stability in aircraft—the static margin—can only strictly be applied to flying animals if the equilibrium angle of attack is specified. We also derive several rules of thumb by which stable fliers can be identified. Stable fliers are expected to exhibit one or more of the following features: (1) Wings that are swept forward in slow flight. (2) Wings that are twisted down at the tips when swept back (wash-out) and twisted up at the tips when swept forwards (wash-in). (3) Additional lifting surfaces (canard, hindwings or a tail) inclined nose-up to the main wing if they lie forward of it, and nose-down if they lie behind it (longitudinal dihedral). Each of these predictions is directional—the opposite is expected to apply in unstable animals. In addition, animals with reduced stability are expected to display direct flight patterns in turbulent conditions, in contrast to the erratic flight patterns predicted for stable animals, in which large restoring forces are generated. Using these predictions, we find that flying animals possess a far higher degree of inherent stability than has generally been recognized. This conclusion is reinforced by measurements of the relative positions of the centres of gravity and lift in birds, which suggest that the wings alone may be sufficient to provide longitudinal static stability. Birds may therefore resemble tailless aircraft more closely than conventional aircraft with a tailplane.

The bird-strike statistics given in Chapter 2 indicated that aircraft have been struck by birds on several parts and that some of these parts are more vulnerable than others. The ability of aero-engine critical structures to withstand bird-strike events must be validated by reliable methods. The parts of an aircraft that are likely to strike a flying bird are obviously the forward-facing areas. Considerable research has been conducted on the resistance of different aircraft components, e.g. fuselage, wing leading edge, tailplane leading edge, empennage, transparent components, fan blades, cockpit, etc., against bird-strike events. A review of the relevant numerical and experimental researches is presented in this chapter.

A numerical 2D investigation of a vertical tailplane airfoil using active flow control with tangential blowing over the rudder shoulder is conducted. The aim of the flow control application is to increase the maximum lift or side force, which can be created by the vertical tailplane, at critical flight conditions like the one-engine-inoperative failure case. In this case the rudder is highly deflected, leading to a large separation on the rudder without blowing. With constant tangential blowing at the rudder shoulder it was shown that fully attached flow can be achieved. To increase the efficiency, pulsed blowing is applied in the present study, leading to a similar increase in the lift coefficient at a reduced actuation mass flow rate requirement. Parameters like the blowing momentum coefficient and the dimensionless frequency are varied and the results are compared to those of the flow calculations with constant blowing. It is observed that pulsed blowing with a small momentum coefficient leads to a strong increase in the lift coefficient compared to constant blowing. The resulting lift increment depends on the dimensionless frequency selected.

For the electrothermal de-icing system of an aircraft tailplane, firstly, an unsteady numerical simulation was conducted for understanding of the heating of the skin. Then ground tests in dry air conditions for the electrothermal de-icing system were designed and conducted, where the ambient temperatures were under the freezing point. The results of theoretical calculation show that, under the given conditions, the highest surface temperature of the tailplane electrothermal de-icing system can reach 67°Cand the lowest surface temperature is close to the ambient temperature. Starting from the third heating cycle, the fluctuation of the surface temperature reaches a stationary cycle. The ground tests show that the designed electrothermal de-icing system can meet the design requirements. It takes less than 25 seconds to raise the surface temperature from -20°C to 0°C. The peak surface temperature in the periodical heating cycles is below 80°C, which is lower than the tolerance limit of the skin material.

Based on the maximum likelihood method, this paper analyzed the influence of tailplane icing on aerodynamic parameters by parameter identification. A nonlinear longitudinal flight dynamics model for aircraft was built, and an identification system was constructed using maximum likelihood method. According to the flight test data of DHC-6 aircraft, the aerodynamic parameters of clean aircraft and two different cases of iced aircraft with particular ice shapes on the tailplane were identified. Eventually, the results of the identification show that the tailplane icing has several adverse effects on the aircraft flight characteristics, including flight performance, elevator effectiveness, stability and safety.

The objective of this paper is to describe the procedure of optimizing the leading-edge structure of a tailplane in an industrial environment. The paper also investigates the effect of implementing a foam core between aluminum sheets in a tailplane leading-edge structure. Bird strike against two types of leading-edge structures, one with and the other without a foam core, was investigated and then the optimum design for each case was determined. The results indicate that if a foam core is embedded between the aluminum sheets instead of increasing the thicknesses of aluminum sheets, the skin overall weight can be reduced by 32%.

The results of investigations related to aerodynamic performances of a high-speed aircraft model, the body of which is designed in the form of rotary solids consisting of different combinations of conical bodies, for example, with permeable conical blunt-nosed body, if there is an active thermal protection, central cylindrical or poorly conical segment and with conical tailplane have been presented in the work. The profiles of pressure distribution over the body’s surface under different angles of attack have been obtained. These profiles make it possible to obtain information about the dynamic performance of the aircraft and reactive drag.

The aerodynamics and acoustics of a generic coaxial helicopter with a stiff main rotor system and a tail-mounted propulsor are investigated using Brown’s Vorticity Transport Model. In particular, the model is used to capture the aerodynamic interactions that arise between the various components of the configuration. By comparing the aerodynamics of the full configuration of the helicopter to the aerodynamics of various combinations of its sub-components, the influence of these aerodynamic interactions on the behaviour of the system can be isolated. Many of the interactions follow a simple relationship between cause and effect. For instance, ingestion of the main rotor wake produces a direct effect on the unsteadiness in the thrust produced by the propulsor. The causal relationship for other interdependencies within the system is found to be more obscure. For instance, a dependence of the acoustic signature of the aircraft on the tailplane design originates in the changes in loading on the main rotor that arise from the requirement to trim the load on the tailplane that is induced by its interaction with the main rotor wake. The traditional approach to the analysis of interactional effects on the performance of the helicopter relies on characterising the system in terms of a network of possible interactions between the separate components of its configuration. This approach, although conceptually appealing, may obscure the closed-loop nature of some of the aerodynamic interactions within the helicopter system. It is suggested that modern numerical simulation techniques may be ready to supplant any overt reliance on this reductionist type approach and hence may help to forestall future repetition of the long history of unforeseen, interaction-induced dynamic problems that have arisen in various new helicopter designs.