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This paper aims to provide the international aeronautical community with details of the development of a new disruptive technology for aircraft propulsion.

This paper describes the results achieved by a small Austrian aeronautical innovations company in developing a cyclogyro propulsion system capable of vertical launch and efficient forward flight. The research team progressed from concept definition and simulation (2004-2006), through experimental validation and concept demonstration (2006), component optimization (2006-2012), full system demonstration (2012-2014) and examination of ability to scale (both larger and smaller) (2015 onwards). This paper provides details of the results of each of these stages.

The research team proved that cyclogyro propulsion can be used for the vertical launch, and that, in forward flight, it has the potential to achieve efficiency beyond the range of conventional fixed wing and rotorcraft.

This research indicates that the efficiency increases with forward speed within the range achieved in standard wind tunnels (up to 35 m/s). This efficiency appears to be caused by a unique chamber effect within the cyclogyro rotor assembly. Future research should be conducted to analyse this chamber effect in greater detail and to test the cyclogyro rotor for speeds beyond 35 m/s.

This work indicates that cyclogyro propulsion could have the potential to provide vertical launch, high speed and highly efficient aircraft that have reduced wing span, no external rotors and exceptional agility. This technology could therefore be feasible for vertical take-off and landing aircraft that can safely form densely packed swarms.

It could be researched as an efficiency increase in forward flight completely different to existing propulsion systems. This could open a way for a more efficient air traffic in future and faster reduction of CO₂ and NOX emission an allow an environment-friendlier air travelling.

This paper provides the details of the first cyclogyro aircraft to have flown and will serve the aeronautical community by stimulating the debate on this new disruptive technology.

This paper aims to present the design of a stability augmentation system (SAS) in the longitudinal and lateral axes for an unstable helicopter.

The feedback controller is designed using linear quadratic regulator (LQR) control with full state feedback and LQR with output feedback approaches. SAS is designed to meet the handling qualities specification known as Aeronautical Design Standard (ADS‐33E‐PRF). A helicopter having a soft inplane four‐bladed hingeless main rotor and a four‐bladed tail rotor with conventional mechanical controls is used for the simulation studies. In the simulation studies, the helicopter is trimmed at hover, low speeds and forward speeds flight conditions. The performance of the helicopter SAS schemes are assessed with respect to the requirements of ADS‐33E‐PRF.

The SAS in the longitudinal axis meets the requirement of the Level 1 handling quality specifications in hover and low speed as well as for forward speed flight conditions. The SAS in the lateral axis meets the requirement of the Level 2 handling quality specifications in both hover and low speed as well as for forward speed flight conditions. The requirements of the inter axis coupling is also met and shown for the coupled dynamics case. The SAS in lateral axis may require an additional control augmentation system or adaptive control to meet the Level 1 requirements.

The study shows that the design of a SAS using LQR control algorithm with full state and output feedbacks can be used to meet ADS‐33 handling quality specifications.

The XH‐59A ABC™ demonstrator aircraft is a research vehicle designed to investigate the unique characteristics of the Advanced Blade Concept. The aircraft, shown in Fig. 1, has now completed an extensive flight test programme investigating its full airspeed, altitude, and manoeuvring envelope. Aircraft design was initiated in 1972, with first flight in July 1973, with certain interruptions to review and analyse flight data and make modifications to the aircraft. The culmination of the flight test programme was a 12.7 hour evaluation of the aircraft by the US Army Aviation Development and Test Activity at Ft. Rucker, Alabama. This evaluation is the subject of this technical report. Reference 1 is the Army report documenting their evaluation. The XH‐59A programme has been jointly funded by the US Army, Navy and Air Force, the National Aeronautics and Space Administration, and Sikorsky Aircraft. The majority of funding has come from the Army.

Computational fluid dynamics (CFD)/computational structural dynamics (CSD) coupling analysis is an important method in the research of helicopter aeroelasticity due to its high precision. However, this method still suffers from some problems, such as wake dissipation and large computational cost. In this study, a new coupling method and a new air load correction method that combine the free wake model with the CFD/CSD method are proposed to maintain computational efficiency whilst solving the wake dissipation problem of the prior coupling methods.

A new coupling method and a new air load correction method that combine the free wake model with the CFD/CSD method are proposed. With the introduction of the free wake model, the CFD solver can adopt two-order accuracy schemes and fewer aerodynamic grids, thus maintaining computational efficiency whilst solving the wake dissipation problem of the prior coupling methods.

Compared with the predictions of the prior methods and flight test data, those of the proposed method are more accurate and closer to the test data. The difference between the two methods in high-speed forward flight is minimal.

Because of the chosen research approach, the research results may lack generalisability. Therefore, researchers are encouraged to test the proposed method further.

In this paper, a CFD/CSD/free wake coupling method is proposed to improve the computational accuracy of the traditional CFD/CSD coupled method and ensure the computational efficiency.

This paper aims to improve the computational efficiency and to achieve high-order accuracy for the computation of helicopter rotor unsteady flows in forward flight during the industrial preliminary design stage.

The integral arbitrary Lagrangian–Eulerian form of unsteady compressible Navier–Stokes equations with low Mach number preconditioned pseudo time terms based on non-inertial frame of reference undergoing rotating and translating was derived and discretized in the framework of multi-block structured finite volume grid using three types of spatial reconstruction schemes, i.e. the third-order accurate monotonic upwind scheme for conservation laws, the fifth-order accurate weighted essentially non-oscillatory and the fifth-order accurate weighted compact nonlinear schemes.

The results show that the present non-inertial computational method can obtain comparable results with other methods, such as the dynamic overset method, and make sure that the higher-order spatial schemes can significantly improve the tip vortex resolution.

The computational grid used by the present method remained static during the whole unsteady computation process, with only local deformations induced by blade cyclic pitch and other operating motions, which greatly reduced the complexity of grid motion and enhanced the efficiency and robustness.

This study aims to investigate the effects of exhaust direction on exhaust plume and helicopter infrared radiation in hover and cruise status.

Four exhaust modes are concerned, and the external flow field and fuselage temperature field are calculated by numerical simulation. The infrared radiation intensity distributions of the four models in hovering and cruising states are computed by the ray-tracing method.

Under the hover status, the exhaust plume is deflected to flow downward after it exhausts from the nozzle exit, upon the impact of the main-rotor downwash. Besides, the exhaust plume shows a “swirling” movement following the main-rotor rotational direction. The forward-flight flow helps prevent the hot exhaust plume from a collision with the helicopter fuselage generally for the cruise status. In general, the oblique-upward exhaust mode provides moderate infrared radiation intensities in all of the viewing directions, either under the hover or the cruise status. Compared with the hover status, the infrared radiation intensity distribution alters somewhat in cruise.

Illustrating the influences of exhaust direction on plume flow and helicopter infrared radiation and the differences of helicopter infrared radiation under hover and cruise statuses are identified. Finally, an appropriate exhaust mode is proposed to provide a better IR signature distribution.

To investigate the use of centre of gravity location on reducing cyclic pitch control for helicopter UAV's (unmanned air vehicles) and MAV's (micro air vehicles). Low cyclic pitch is a necessity to implement the swashplateless rotor concept using trailing edge flaps or active twist using current generation low authority piezoceramic actuators.

An aeroelastic analysis of the helicopter rotor with elastic blades is used to perform parametric and sensitivity studies of the effects of longitudinal and lateral center of gravity (cg) movements on the main rotor cyclic pitch. An optimization approach is then used to find cg locations which reduce the cyclic pitch at a given forward speed.

It is found that the longitudinal cyclic pitch and lateral cyclic pitch can be driven to zero at a given forward speed by shifting the cg forward and to the port side, respectively. There also exist pairs of numbers for the longitudinal and lateral cg locations which drive both the cyclic pitch components to zero at a given forward speed. Based on these results, a compromise optimal cg location is obtained such that the cyclic pitch is bounded within ±5° for a BO105 helicopter rotor.

The reduction in the cyclic pitch due to helicopter cg location is found to significantly reduce the maximum magnitudes of the control angles in flight, facilitating the swashplateless rotor concept. In addition, the existence of cg locations which drive the cyclic pitches to zero allows for the use of active cg movement as a way to replace the cyclic pitch control for helicopter MAV's.

THE simple actuator disk theory, first postulated by Froude over sixty years ago, is the basis of most helicopter induced flow theory. This disk is an idealization of a rotor which uniformly accelerates the air with no loss of thrust at the blade tips. It can therefore be regarded as the limit case of a rotor with an infinite number of blades. It is also assumed to be infinitely thin so that no discontinuities in velocity occur on the two sides of the disk.

In this paper results arc discussed of wind‐tunnel smoke tests with a small helicopter rotor. The test apparatus includes a specially developed hot‐wire smoke‐generator. An attempt is made to describe the flow pattern in the neighbourhood of a helicopter rotor in the vortex ring state by introducing the spread of the slipstream, causing an airbody around the rotor. By considering the proportions of this airbody the unpleasant behaviour of helicopters in the partial power‐off descent at low forward speeds can be more clearly understood.

THE purpose of this paper is to examine the part that metal fatigue plays in the engineering of the helicopter, and to outline the methods used at present to estimate the safe fatigue life of the component parts of the helicopter.