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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.

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.

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.

— 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).

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.

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 ...

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.

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 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.

One of the main structural requirements of a leading edge of a tailplane is to ensure that any significant damage caused by foreign object (i.e. birdstrike, etc...) would still allow the aircraft to land safely. In particular, leading edge must be certified for a proven level of bird impact resistance. Since the experimental tests are expensive and difficult to perform, numerical simulations can provide significant help in designing high-efficiency bird-proof structures. The aim of this research paper was to evaluate two different leading edge designs by reducing the testing costs by employing state-of-the-art numerical simulations. The material considered was a sandwich structure made up of aluminium skins and flexcore as core. Before each test was carried out, pre-test numerical analyses of birdstrike were performed adopting a lagrangian approach on a tailplane leading edge of a large scale aircraft using the MSC/Dytran solver code. The numerical and experimental correlation have shown good results both in terms of global behaviour of the test article and local evolution of some measurable parameters confirming the validity of the approach and possible guidelines for structural design including the bird impact requirements.

This work investigates the aerodynamic interference among airplane components caused by rudder deflection for a typical turboprop aircraft geometry through the computational fluid dynamics technique. At no sideslip, an airplane is in symmetric flight conditions. The rudder deflection creates a local sideslip angle close to the vertical tailplane, and this effect is increased by fuselage and horizontal tail. Typical semiempirical methods, such as United States Air Force Stability and Control Data Compendium (USAF DATCOM), do not take into account for these effects, proposing the same corrective parameters both for pure sideslip and rudder deflection conditions. Numerical analyses executed on several aircraft configurations with different wing and horizontal tailplane positions show that the interference factors are smaller than those predicted by the USAF DATCOM procedure, providing guidelines for a more accurate aircraft directional control analysis and hence rudder preliminary design.

The Icing Encounter Flight Simulator is one part of the Smart Icing Systems project at the University of Illinois at Urbana-Champaign. From the Smart Icing Systems project, an ice management system was designed that would sense and characterize ice, notify the pilot, and if necessary take measures to ensure the safety of the aircraft. The icing simulator was used as a platform to integrate and test different components of the Smart Icing System. To create an Icing Encounter Flight Simulator, functionality and Smart Icing System components were added to the FlightGear flight simulator. Functionality added to FlightGear include a reconfigurable aircraft model and an icing model, and Smart Icing System components added include an envelope protection system and a glass cockpit enhanced with ice management system features. To ensure a real-time simulation, computationally intensive processes were distributed over several desktop computers linked together through a local network. In order to demonstrate the effectiveness of the Smart Icing System components, two fictional but historically motivated icing scenarios were developed and tested with a simulated DeHavilland DHC-6 Twin Otter, specifically, a tailplane stall event during a steep descent and a roll upset event during an emergency approach. Language: en

Extreme loads are generated in aircraft flight manoeuvres. Among different manoeuvres considered in this paper are motions following a sudden deflection of elevator and response to a vertical gust. Airplane was assumed to be a rigid body of three degrees of freedom in symmetrical motion. Elevator deflection was either of the step change type, or of the sinusoidal type, gust was assumed to be either of the step change type or harmonic, with a gust cycle time corresponding to the time to travel a distance equal to 25 Mean Aerodynamic Chord. In all cases a jump type elevator deflection was assumed to last for 1 second, whilst the airplane response was observed for 3 seconds. The airplane motion, its velocities, accelerations and load acting on the tailplane were calculated by means of numerical integration of the ordinary differential /of motion.

A summary about the DLR, German Aerospace Center results from the fourth AIAA Computational Fluid Dynamics Drag Prediction Workshop is presented. Compared to the investigations in the previous three workshops, the latest workshop had a stronger focus on drag and trim drag predictions as well as pitching moment calculations. Therefore, the new Common Research Model developed by NASA’s Subsonic Fixed Wing Aerodynamics Technical Working Group and tested in NASA wind tunnels is used. It represents a state-of-the-art transonic transport aircraft configuration, and in contrast to the configurations previously taken, it includes an optional horizontal tailplane with three different tail settings. DLR has defined three objectives for its activities in the fourth drag prediction workshop. At first, investigations should identify solution accuracy and grid convergence behavior using prismatic element dominant grids for boundary-layer resolution in comparison to hexahedral element dominant grids. Second, the influen...

A computational investigation was conducted to determine the effect of simulated upper-surface spanwise ice shapes on airfoil aerodynamics. These shapes are representative of supercooled large droplet shapes on aircraft with active de-icing boots. The numerical investigation employed steady-state simulations with a high-resolution full Navier‐Stokes solver using a solution-adaptive unstructured grid for both non-iced and iced cone gurations. Thestudy investigated a modie ed NACA23012 (with and without e ap dee ection )and airfoilsoftheNASAModern Airfoil program. Arangeof protuberancelocation, size, and shapewere examined, and comparisons weremade to availableexperimental data. In general, the performance of thecomputational methodology was particularly good for pressure and hinge-moment distributions (including the nonlinear break points ), whereas lift was predicted reasonably well up to (but not past) fully separated e ow conditions. The airfoil shape sensitivity studies indicated that the NACA 23012m exhibited the most detrimental performance with respect to lift loss, which tended to be greatest around x/c ofabout0.1 that also correspondsto the location of minimum Cp. The more evenly loaded NLF 0414 airfoil tended to have less separation for equivalent clean-airfoil lift conditions and did not exhibit a unique critical ice-shape location. The forward-loaded airfoils of the business jet main wing model and the commercial transport horizontal tailplane model exhibited critical ice-shape locations close to the leading edge (x/c = 0.02), which was close to the minimum Cp location.

Modal approaches like eigen- structure assignment have shown them- selves to be efficient for flight control design. Performance requirements are easily met using this approach. How- ever, generally, robustness is not satisfac- tory. This paper presents a technique that can be viewed as an improvement over traditional eigenstructure assignment as it produces systems wich meet robust- ness requirements (multimodel approach). The proposed technique reduces to solv- ing a quadratic problem under linear con- straints. The application treated concerns the landing phase of a large transport air- craft. It is shown that standard gain scheduling can be replaced by a single low dimensional dynamic feedback.

The out-of-plane vibration of a panel on an aircraft tailplane, excited by random noise in the 100 – 500 Hz range, was suppressed by an active controller that used four piezoceramic patches, two as strain actuators and two as strain sensors. The strain energy levels at these sensor locations were signiŽ cantly reduced. Three accelerometers at arbitrarily chosen locations elsewhere on the panel demonstrated that the out-of-plane vibrational energy levels were also reduced there. The multi-input, multi-output digital controller, designed by a physically meaningful technique, also resulted in a system with good robustness properties.

The paper presents a deep investigation on the aerodynamics of the vertical tailplane and the correct estimation of its contribution to aircraft directional stability and control, especially during the preliminary design phase. Nowadays the most used methodologies in preliminary design to estimate the contribution of vertical tailplane on aircraft directional stability and control are (i) the classical method proposed by USAF DATCOM (also presented in several aeronautics textbooks) and (ii) the method presented in ESDU reports. Both methodologies derive from NACA World War II reports of the first half of the ʼ900, based on obsolete geometries that do not represent the typical shape of a transport aircraft. The other limit is that these methods give quite different results for certain configurations, e.g. in the case of horizontal stabilizer mounted in fuselage. As shown in literature, the main effects on the sideforce coefficient of the vertical tail are due to the interactions among the aircraft components. In order to better highlight these effects, a different approach using the RANS equations has been adopted. Several CFD calculations have been performed on some test cases (used as experimental database) described in NACA reports to verify the compliance of CFD results with available experimental data. The CFD calculations (performed through the use of a parallel supercomputing platform) have shown a good agreement between numerical and experimental data. Subsequently the above mentioned effects have been deeply investigated on a new set of aircraft configurations. The configurations that have been prepared differ among them for wing aspect ratio, wing–fuselage relative position (high-wing/low-wing), vertical tailplane aspect ratio (vertical tail span versus fuselage height) and horizontal tailplane position respect to the vertical tailplane (with the aim of investigating the effect of fin-mounted T configuration, typical of regional turboprop transport aircraft). All the CFD analyses have been carefully post-processed and have been useful to obtain new curves to predict the above mentioned effects and thus to have a more accurate estimation of vertical tailplane contribution to aircraft directional stability and control.

For the planned optimization of the vertical tailplane (VTP) of a Megaliner transport aircraft configuration a prior basic study is undertaken where the influence of simplifying the aircraft geometry on the characteristics of the VTP is examined. Four different configurations with varying simplification level were investigated. These are: a VTP on a base plate, a VTP on a fuselage, a VTP on a fuselage with horizontal tailplane and a VTP on a fuselage with horizontal tailplane, wing and belly fairing. Several flight cases were identified where the characteristics of the VTP are most important and then analyzed by means of numerical flow analysis using the flow solver TAU developed at the German Aerospace Center DLR. Determined are the side force coefficient generated at maximum rudder deflection, the high speed drag, the gradient of the side force coefficient due to rudder deflection and the gradient of the side force coefficient due to sideslip angle. The latter two were evaluated at low and high speed conditions. The simplified configurations were then compared to the most complete one as reference case.

The present study aims at assessing the influence of the interference effects caused by the horizontal tail plane (HTP) and vertical tail plane (VTP) on the aerodynamic behaviour of a transport aircraft in high-lift configuration. This aspect is of particular interest close to aircraft stalling conditions. The present study therefore assesses the effects caused by the tail plane during onflow conditions that represent stall manoeuvres. In this regard numerical simulations at extreme points of flight envelopes are analysed, by considering effects of discretization, physical turbulence models and incidence stall settings.

In order to achieve the required level of system safety, it is often necessary to specify scheduled single system check periods for duplicated or triplicated systems, where failure of one system is neither indicated nor evident to the pilot or ground crew. This paper presents the techniques used in the analysis of the A300 tailplane trim actuator service data, for the purpose of identifying a check period and overhaul life, consistent with continued safe economic operation.

This chapter deals with the static longitudinal stability of fixed-wing aircraft having a conventional configuration consisting of a wing and a horizontal tailplane with elevator and trim tab. The aircraft is considered as a rigid body. We provide a definition for the mean aerodynamic chord and show its role in the longitudinal stability problem. Then the pitching moment equations are written for the wing alone and the full aircraft. Trim equations are derived for stick-fixed and stick-free conditions; the corresponding stability points are calculated. Finally, the pull-up up maneuver is described under steady-state conditions; stick-free and stick-fixed maneuver points are derived. Keywords: aircraft flight; longitudinal stability; center of gravity; pull-up; longitudinal trim; trim drag; stick-fixed; stick-free stability

Computations using the DLR TAU code are presented to predict the effect of the wind tunnel on models which are directly attached to the wind tunnel side-wall. Results show that the effect of the sidewalls is similar to an offset in the angle of attack (but can only partially be corrected by a change in α), and that this effect is proportional to the lift and Mach number. These results were used in the design phase for the iGREEN wing-tailplane interference experiment.

The rotating shaft of an aircraft tailplane can be considered as a structural system with several significant failure modes. Each significant failure mode is constituted by one critical cross-section of the shaft. The method of enumerating critical cross-sections is explained. All the significant failure modes are considered as series and correlative for computing the system reliability of the shaft.

A generic wind tunnel model for tailplane stall research was designed and experimentally investigated to establish a database for code validation. The configuration is numerically optimised to obtain large Reynolds numbers at the horizontal tailplane in a wind tunnel of limited size. It consists of a fuselage, a detachable horizontal tailplane and a tip-truncated wing, that mounts the model to the turntables of the closed test section. The wing was designed to reproduce a representative downwash in the tailplane region. The tests were conducted at a freestream Mach number of 0.16 and at a Reynolds number of 0.72 wx 106. Tailplane stall in case of natural and fixed transition was visualised by the oil-flow technique and quantified by pressure and force measurements and by the Particle Image Velocimetry of the turbulent trailing-edge separation. Numerical simulations, using the unstructured Reynoldsaveraged Navier Stokes Code TAU, are in good agreement with the experiments. They show a separation of the boundary layer starting at the trailing edge with high crossflow velocities at the outer tailplane. Depending on the boundary-layer transition, the stall occurs abruptly for natural transition, initialised by a burst of a laminar separation bubble, or gradually for fixed transition.

This paper presents aerodynamic modelling and realtime control of a 1 degree-of-freedom (DOF) 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 this objective a detailed study is conducted employing analytical as well as system identification techniques. Analytical approach although less accurate complements system identification process. This synergy is exploited along with statistical and time domain tests to arrive at a high fidelity model. It is argued 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 utilising classical PI control law.

*† ‡ § The Zeppelin NT is a semi rigid airship, produced by Zeppelin Luftschifftechnik GmbH (ZLT) in Friedrichshafen, Germany. It uses three identical tailplanes in an inverted Yarrangement for steering. These tailplanes are built in a stiffened, thin and light CFRPsandwich shell construction. For this tailplane concept the Institute for Lightweight Structures (LLB) at Technical University of Munich, Germany, has developed an In-Situ monitoring method, based on LAMB-waves, to detect damages due to hailstone impacts. Alternatively, the LLB has researched advanced tailplane designs – to improve resistance to hailstones and to reduce mass versus the current tailplanes – as well as repair mechanisms.

When designing flight control laws using linearisations of an aircraft model about dierent flight conditions, some form of scheduling of the resultant gains will be required to implement the controller over wide operating regions. In practice, the controller gains are often scheduled against slowly-varying system states such as altitude or velocity. However, it may also be desirable to schedule gains against rapidly-varying states such as angle-of-attack or pitch rate. To obtain optimum performance it may be necessary to schedule a state feedback gain with that same state, creating a cyclic dependence. This paper presents a numerical method of accounting for this dependence when scheduling state feedback gains against themselves. The resulting ‘dynamic gain schedule’ is shown to overcome some limitations normally imposed on conventionally scheduled controllers, and to improve transient response in nonlinear regions. The controller design process, using eigenstructure assignment, is demonstrated on a 2 nd order longitudinal aircraft model in incidence and pitch rate. The controller was implemented experimentally on an approximate BAe Hawk wind tunnel model mounted on a dynamic pendulum support test rig. Despite the significantly nonlinear characteristics of the model and tailplane actuation system, the results show that satisfactory control could be achieved using dynamic scheduling over the entire operating region.

The conceptual design is presented of a Mach 1.6 supersonic commercial transport aircraft carrying 250 passengers in mixed class seating over a distance of 5,500 nm. The design has been used as a baseline for a European MDO study emphasizing wing planform optimization. Prediction methods for aerodynamic and mass properties requiring low computational efiort were derived from the open literature, and have been validated with the aid of a Mach 2.0 reference aircraft design. The selected four mid tandem fan engines enable the aircraft to comply with Chapter 3 airfleld noise requirements. Several studies identifying major technological, geometrical and sizing sensitivities provided arguments for the choice of a tailplane conflguration and an aspect ratio 2.8 modifled arrow wing, featuring leading- and trailing-edge ∞aps. The three-leg main landing gear retracts into and behind the wing center section structure. Performance analysis indicates that the Mach 1.6 design has a signiflcantly longer range than the reference Mach 2.0 design with the same take-ofi weight. Field lengths are similar to those of present-day long-range transonic transports, and ∞ight characteristics at subsonic speeds are more conventional than those of a Mach 2.0 airliner.

NASA is developing and validating technology to incorporate aircraft icing effects into a flight training device concept demonstrator. Flight simulation models of a DHC-6 Twin Otter were developed from wind tunnel data using a subscale, complete aircraft model with and without simulated ice, and from previously acquired flight data. The validation of the simulation models required additional aircraft response time histories of the airplane configured with simulated ice similar to the subscale model testing. Therefore, a flight test was conducted using the NASA Twin Otter Icing Research Aircraft. Over 500 maneuvers of various types were conducted in this flight test. The validation data consisted of aircraft state parameters, pilot inputs, propulsion, weight, center of gravity, and moments of inertia with the airplane configured with different amounts of simulated ice. Emphasis was made to acquire data at wing stall and tailplane stall since these events are of primary interest to model accurately in the flight training device. Analyses of several datasets are described regarding wing and tailplane stall. Key findings from these analyses are that the simulated wing ice shapes significantly reduced the C , max, while the simulated tail ice caused elevator control force anomalies and tailplane stall when flaps were deflected 30 deg or greater. This effectively reduced the safe operating margins between iced wing and iced tail stall as flap deflection and thrust were increased. This flight test demonstrated that the critical aspects to be modeled in the icing effects flight training device include: iced wing and tail stall speeds, flap and thrust effects, control forces, and control effectiveness.

This chapter focuses on the design and construction of aircrafts. Large civil airliners have a low wing design, in which the wing structure passes through the freight area beneath the passenger cabin. Small airliners may use the high wing design, with a bulge over the top line of the fuselage so as not to restrict passenger headroom. Having a continuous upper surface to the wing (as in the high-wing layout) can improve the L/D ratio and keeps the engines at a higher distance from the ground, thereby avoiding debris from poor or unpaved runways. Tailplane configuration is matched to the wing type and includes high tail, low tail, flat, vee, and dihedral types. Low tails increase stability at high angles of attack but can also result in buffeting and non-linear control response during normal flight. High tails are generally necessary with rear-fuselage mounted engines and are restricted to high-speed military aircraft use. The rear-engine configuration has generally been superseded by under-wing mounted engines, which optimizes bending moments and enables the engine thrust loads to be fed directly into the wing spars. In contrast, rear-fuselage mounted engines decrease cabin noise. Aircraft cabin design is constrained by the need to provide passenger areas and an underfloor cargo space within the confines of the standard tube-shaped fuselage. Fuselage design is influenced by the ground servicing needs of an aircraft. Ground servicing represents commercial downtime, therefore it is essential to ensure that maximum ground servicing activities be carried out simultaneously.

This paper focusses on the stability analysis of a nonlinearly scheduled fault tolerant control system with varying structure. An aircraft problem is considered in which the tailplane suffers from degraded control effectivity, including the possibility of complete failure of the tailplane actuator, leaving the engines as only input to the system. A controller synthesis approach is followed that gives the designer more freedom in changing the controller strategy and scheduling of the local controllers, while providing fault tolerance for the specified failure. Since controller synthesis and global stability analysis is decoupled, special effort is put on proving global stability. Conservatism in the stability analysis, introduced by nonlinear scheduling, is effectively reduced. Problems in the stability analysis with varying structure of the closed-loop system, caused by input reduction, could be prevented and the method is a good alternative in provision of fault tolerance.

This chapter describes how multivariable control has been applied to the Defence Evaluation and Research Agency (DERA) Bedford Harrier and flight tested. The H∞ method is a robust multivariable design method. The term multivariable means that it is suited to systems with several inputs and outputs where each input strongly affects more than one output. The longitudinal control of the Harrier is such a system, the three primary motivators being nozzle angle, engine thrust and blended tailplane/reaction control system (RCS). Moving any of these three inputs affects forward, normal and pitching motions. The collaborative programme set up between DERA and Bedford and Cambridge University had two objectives, namely: to show how H∞ methods could be used in a practical application; and to demonstrate implementation of multivariable control on an aircraft.

The effects of tailplane icing on aircraft dynamics and tailplane aerodynamics were investigated using, NASA's modified DHC-6 Twin Otter icing research aircraft. This flight program was a major element of the four-year NASA/FAA research program that also included icing wind tunnel testing, dry-air aerodynamic wind tunnel testing, and analytical code development. Flight tests were conducted to obtain aircraft dynamics and tailplane aerodynamics of the DHC-6 with four tailplane leading-edge configurations. These configurations included a clean (baseline) and three different artificial ice shapes. Quasi-steady and various dynamic flight maneuvers were performed over the full range of angles of attack and wing flap settings with each iced tailplane configuration. This paper presents results from the quasi-steady state flight conditions and describes the range of flow fields at the horizontal tailplane, the aeroperformance effect of various ice shapes on tailplane lift and elevator hinge moment, and suggests three paths that can lead toward ice-contaminated tailplane stall. It was found that wing, flap deflection was the most significant factor in driving the tailplane angle of attack toward alpha(tail stall). However, within a given flap setting, an increase in airspeed also drove the tailplane angle of attack toward alpha(tail stall). Moreover, increasing engine thrust setting also pushed the tailplane to critical performance limits, which resulted in premature tailplane stall.

The effects of tailplane icing were investigated in a four-year NASA/FAA Tailplane Icing, Program (TIP). This research program was developed to improve the understanding, of iced tailplane aeroperformance and aircraft aerodynamics, and to develop design and training aides to help reduce the number of incidents and accidents caused by tailplane icing. To do this, the TIP was constructed with elements that included icing, wind tunnel testing, dry-air aerodynamic wind tunnel testing, flight tests, and analytical code development. This paper provides an overview of the entire program demonstrating the interconnectivity of the program elements and reports on current accomplishments.

A detailed analysis of two of the dynamic maneuvers, the pushover and elevator doublet, from the NASA/FAA Tailplane Icing Program are discussed. For this series of flight tests, artificial ice shapes were attached to the leading edge of the horizontal stabilizer of the NASA Lewis Research Center icing aircraft, a DHC-6 Twin Otter. The purpose of these tests was to learn more about ice-contaminated tailplane stall (ICTS), the known cause of 16 accidents resulting in 139 fatalities. The pushover has been employed by the FAA, JAA and Transport Canada for tailplane icing certification. This research analyzes the pushover and reports on the maneuver performance degradation due to ice shape severity and flap deflection. A repeatability analysis suggests tolerances for meeting the required targets of the maneuver. A second maneuver, the elevator doublet, is also studied.

The improvement of an integrated aircraft predesign code and its application to future aircraft concepts is reported. For the simulation of a Megaliner canard configuration as well as a supersonic commercial transport aircraft (SCT) the predesign code PrADO was extended by the higher-order panel method HISSS. The improvement of the aerodynamic model allows the simulation of interference effects between wing, tailplane and canard as well as the extension of the simulated flight regime to supersonic speeds. The application of the calibrated predesign code to a Megaliner configuration leads to differences between published and predicted design weights below 6%. Preliminary results of the simulation of a canard configuration show that a canard may not automatically increase the aerodynamic efficiency L/D. A reduction of the induced drag is overcompensated by an increase of the minimum drag due to additional friction drag of the canard. The SCT design studies also provide converged solutions, however, a redesign of the Concorde shows significant discrepancies with respect to the operational empty weight, fields for future investigations.

STOVL aircraft have several force "motivators" controlling the pitch plane. These include tailplane angle, flap position, throttle and nozzle angle. Future configurations are likely to have these motivators, possibly adding other new features such as independent nozzle thrust vectoring, and canard foreplanes. The trimmap based pitch plane control law, known as Control Law 001, is a solution to the major challenge of optimally controlling these many motivators with just two pilot inceptors. This paper describes design, development and initial flight trials of the VAAC 001 control law.