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A low-cost but credible method of low-subsonic flutter analysis based on ground vibration test (GVT) results is presented. The purpose of this paper is a comparison of two methods of immediate flutter problem solution: JG2 – low cost software based on the strip theory in aerodynamics (STA) and V-g method of the flutter problem solution and ZAERO I commercial software with doublet lattice method (DLM) aerodynamic model and G method of the flutter problem solution. In both cases, the same sets of measured normal modes are used.

Before flutter computation, resonant modes are supplied by some non-measurable but existing modes and processed using the author’s own procedure. For flutter computation, the modes are normalized using the aircraft mass model. The measured mode orthogonalization is possible. The flutter calculation made by means of both methods are performed for the MP-02 Czajka UL aircraft and the Virus SW 121 aircraft of LSA category.

In most cases, both compared flutter computation results are similar, especially in the case of high aspect wing flutter. The Czajka T-tail flutter analysis using JG2 software is more conservative than the one made by ZAERO, especially in the case of rudder flutter. The differences can be reduced if the proposed rudder effectiveness coefficients are introduced.

The low-cost methods are attractive for flutter analysis of UL and light aircraft. The paper presents the scope of the low-cost JG2 method and its limitations.

In comparison with other works, the measured generalized masses are not used. Additionally, the rudder effectiveness reduction was implemented into the STA. However, Niedbal (1997) introduced corrections of control surface hinge moments, but the present work contains results in comparison with the outcome obtained by means of the more credible software.

Reports on the MSc group design project of students at the College of Aeronautics, aerospace vehicle design in 1995. The students worked on advanced short take‐off and vertical landing of a combat aircraft. Details the project showing aircraft dimensions and design. Full assessment of the results is pending, but outlines a number of problems faced by the students.

The purpose of this paper is to investigate the flow around Parastoo UAV's wing, with the aim of improving its aerodynamic performance. A major source of concern is the use of relatively large flaps in the original design. This unmanned aerial vehicle (UAV) operates at low Reynolds numbers of below 500,000 and was designed for short‐range reconnaissance.

A finite volume solver is utilized to investigate the flow over different wing designs to find a replacement for the current one. To check the accuracy of this numerical modeling, the authors first duplicate the conditions of available relevant experiments. The numerical results are in good agreement with the wind tunnel experiments. Here, the aerodynamic performances of Parastoo's wing at different flight conditions with and without the proposed modification are studied and compared.

As the original design of Parastoo uses relatively large flaps, it is found that the aerodynamic performance of Parastoo is significantly hampered due to their existence. The use of a new wing cross‐section can improve the aerodynamics efficiency of Parastoo. It is recommended that FX‐63137 airfoil is a more suitable cross section instead of Parastoo's original NACA‐63215 airfoil. It is shown that this change improves the aerodynamic performance of the UAV and with the use of smaller flaps (changing the flaperons to only ailerons), the existing payload weight can be increased by 90 per cent.

The issues discussed for this UAV may be of use for other small unmanned plane designers. The numerical data generated for this study are useful for other design teams, both as in direct uses of the data in their own designs and/or for the validation of their numerical methods before investigating other wing designs.

The present study aimed to demonstrate different computational models, data and stability results obtained in a wide number of projects of various aircrafts such as unmanned aerial vehicles (UAVs), general aviation and big passenger flying airliners in blended wing body (BWB) configurations. Many details of modeling and computing are shown for unconventional configurations, namely, for a BWB aircraft and for tailless UAVs.

Mathematical models for analysis of static and dynamic stability were built and investigated based on equations of motion in the linearized form using the so-called state variable model for a steady-state disturbed, generally asymmetric, flight.

Flight dynamics models and associated computational procedures appeared to be useful, both in a preliminary design phase and during the final assessment of the configuration at flight tests. It was also found that the difference between thresholds for static and dynamic stability conditions was equal to 9 per cent of mean aerodynamic chord (MAC) in the case of BWB and 3 per cent of MAC in the case of tailless UAVs.

Many useful information about aircraft dynamics can be easily obtained from computational analyses including time to half/double and periods of oscillation, undamped frequencies, damping ratio and many others. Stability analysis of different unconventional configurations will be easier and faster if an access to such configurations is available.

This paper presents a very efficient method of assessment of the designing parameters, especially in an early stage of the design process. In open literature, there are a great number of datasets for classical configurations, but it is hard to find anything for passenger BWB and tailless UAVs. Stability computations are performed based on equations of motion derived in the stability frame of the reference fixed with one-quarter of MAC. It can be considered as an original, not typical but a very practical approach because values of stability and control derivatives do not change even if the centre of gravity is travelling.

APART from a short hop take‐off and landing up to a height of about 20 ft, this was the first flight of this Denney Kitfox Mk3. A passenger was carried because the pilot considered that the extra weight would be an extra advantage in what was a relatively high performance aircraft. His two hour familiarisation in the Kitfox were dual. The aircraft took off and the engine was set to 6,000 RPM and the throttle was adjusted during the ground run to maintain this figure. The initial climb was as expected until about 150 ft agl when the indicated airspeed was noted to have been about 45 mph.

The purpose of this paper is to find an influence of the reduced stiffness of actuators, located on the most outer parts of ailerons, flaperons, rudders, elevators and elevons on the excitation of flutter. This phenomenon is especially important for unmanned aerial vehicles because they continuously use all these control surfaces for trimming and stabilisation and on the other hand, the numerous statistics show that failure of elements of flight control systems are still the most probable reasons of aircraft critical failure.

Flutter calculations were performed by use of the classical modal approach. The normal vibrations of the free aircraft were measured in the ground vibration test (GVT). Test results were used either for verification of the FEM model of the structure – in this case for flutter calculation the MSC.Nastran software was used, or directly for flutter calculation. Based on the flutter analysis, the control surfaces critical for flutter were determined.

These so‐called critical control surfaces –, i.e. surfaces responsible for flutter excitation at first – are localized on outer parts of wing and empennage. It was found that the critical surfaces should have been mass balanced or should be irreversible. In the second case, i.e. when the control surfaces are irreversible, the actuators and drivers should have been of a high reliability, because disconnection of these elements could involve flutter.

This approach within the computational analysis is limited to linear case, otherwise NASTRAN software cannot be used for flutter analysis. GVTs could be performed successfully independently if the structure has linear or non‐linear properties.

It was found that before any flight the stiffness in the flight control system of all control surfaces must carefully be checked and kept above the critical stiffness value. Safety level strongly depends on the reliability of actuators used on such unmanned aerial vehicles. The simulation of disconnection (as a result of damage) of selected control surfaces is possible even if the GVT were provided on undamaged vehicle. To do it, the rotational mode of so‐called “free control surface” should be prepared (as an artificial resonant mode) for all deflected control surfaces; next all the resonant modes should be orthogonalized, relative to this artificial control surfaces mode.

This paper was based on two big European and national projects, and all presented results are original and were never published before. Some selected graphs were shown during the EASN Workshop, Paris, September 2010 at the presentation entitled: “Aeroelastic analysis of remotely controlled research vehicles with numerous control surfaces”.

Currently undergoing NASA flight testing is the Grumman X‐29 forward swept wing (FSW) demonstrator aircraft which was built under a contract sponsored by the Defence Advanced Research Projects Agency (DARPA) and funded through the Air Force. The FSW aircraft first took to the air in December, 1984 and after four flights with the manufacturer was handed over to NASA for a verification programme involving all the benefits of this design. Advanced technology features which will be demonstrated are the forward swept wing for improved aerodynamic efficiency and good control at high angles of attack; tailored composite wing structure that resists the tendency towards structural divergence inherent in this configuration; thin supercritical aerofoil for improved transonic performance at high lift coefficients; variable camber trailing edge to reduce drag at all lift coefficients and avoid supersonic drag associated with aerofoil camber; canard longitudinal control for efficient trimming of variable camber pitching moments and favourable canard‐wing interactions; highly relaxed static stability for very low trim drag at all Mach numbers; and a digital fly‐by‐wire flight control system.

Lockheed‐Georgia Company is using a special tool to form a composite assembly for the new V‐22 Osprey tilt rotor aircraft. The tool is a mould fixture for forming a relatively large composite structure. It is used to build the cove spar assembly for the trailing edge of the V‐22's wings.

ROLLED out in May and scheduled to make its first flight in November is the V‐22 Osprey tiltrotor multi‐mission aircraft that has been ordered by four United States Services. Technological advances incorporated in the V‐22 had their beginnings in the XV‐3 in the mid 1950's which proved the feasibility of the tiltrotor concept and in the 1970's the XV‐15 programme explored and confirmed its validity. The latter aircraft which were jointly sponsored by NASA and the Army and Navy, have completed 600 hours of test flying and have demonstrated the suitability of this configuration for all kinds of military purposes.

Outlines the development work on the tiltrotor undertaken during past decades and the more recent progress on the V‐22 Osprey now in production for the US Armed Forces. Also describes the first civil tiltrotor, the Bell Agusta 609, which is due to make its first flight in 2000.