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The effects of rotor blade design variables and their mutual interactions on aerodynamic efficiency of helicopters are investigated. The aerodynamic efficiency is defined based on figure of merit (FM) and lift-to-drag responses developed for hover and forward flight, respectively.

The approach is to couple a general flight dynamic simulation code, previously validated in the time domain, with design of experiment (DOE) required for the response surface development. DOE includes I-optimality criteria to preselect the data and improve data acquisition process. Desirability approach is also implemented for a better understanding of the optimum rotor blade planform in both hover and forward flight.

The resulting system provides a systematic manner to examine the rotor blade design variables and their interactions, thus reducing the time and cost of designing rotor blades. The obtained results show that the blade taper ratio of 0.3, the point of taper initiation of about 0.64 R within a SC1095R8 airfoil satisfy the maximum FM of 0.73 and the maximum lift-to-drag ratio of about 5.5 in hover and forward flight.

The work shows the practical possibility to implement the proposed optimization process that can be used for the advanced rotor blade design.

The work presents the rapid and reliable optimization process efficiently used for designing advanced rotor blades in hover and forward flight.

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.

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.

The purpose of this paper is to explore the effects of the camber on gliding and hovering performance of two-dimensional corrugated airfoils. While the flying mechanism of natural flyers remains a myth up to nowadays, the simulation serves as a minor step toward understanding the steady and unsteady aerodynamics of the dragonfly flight.

The lattice Boltzmann method is used to simulate the flow past the cambered corrugated dragonfly airfoil at low Reynolds numbers. For gliding flight, the maximum camber, the distance of the location of maximum camber point from the leading edge and Reynolds number are regarded as control variables; for hovering flight, the maximum camber, the flapping amplitude and trajectory are considered as control variables. Then corresponding simulations are performed to evaluate the implications of these factors.

Greater gliding ratio can be reached by increasing the maximum camber of the dragonfly wing section. When the location of the maximum camber moves backward along the wing chord, large scale flow separation can be delayed. These two effects result in better gliding performances. For hovering performances, it is found that for different flapping amplitudes along an inclined plane, the horizontal force exerted on the airfoils increases with the camber, and the drag growths first but then drops. It is also found that the elliptic flapping trajectory is most sensitive to the camber of the cambered corrugated dragonfly wing section.

The effects of the camber on gliding and hovering performance of the cambered dragonfly wing section are explored in detail. The data obtained can be helpful when designing micro aerial vehicles.

The experience gained since 1959 at MBB, Military Aircraft Division, in the development and flight testing of V/STOL combat aircraft having the capability to reach Mach 2 and to take off with after‐burning temperatures is described. The German project VJ 101 C and the US/FRG project AVS as well as the joint US/FRG V/STOL Technology Programme conducted during the years 1967 through 1970 serve as examples. The paper consists of two main sections:

The results of preliminary tests concerning estimation and widening of helicopter limiting manoeuvre abilities are presented. Research space applies to super‐ and hipermanoeuvrability problems that are especially important for helicopters, because of better manoeuvrability influence on higher safety level and effectiveness in special applications. In airplane engineering, these types of tests are advanced and aerodynamic system improvements are introduced as well as thrust vector control. There are also new manoeuvres recognized for advanced manoevrability airplanes: Cobra, Kulbit, Hook, Bell, Herbst manoeuvre. Although helicopter is “originally” thrust controlled, systematic researches on this field are still not conducted. The paper deals with the problem of helicopter flight mechanics at low flight speeds. The purpose of performed analysis is to achieve possibility of helicopter angular position control within wide range of angular displacements. This is performed by linear and centrifugal acceleration control. Rotor thrust vector control makes those accelerations appear.

RECENT developments in the field of convertible aircraft have shown the practicability of numerous methods of achieving vertical take‐off, combined with forward flight that is largely, or completely, supported by fixed wings. The aerodynamics of aircraft which derive their hovering lift from jet thrust are straightforward since there is no down‐wash to complicate the airflow over the wings; if a rotor is used, however, the wash of the rotor will vary from a relatively minor effect in forward flight to the only source of airflow over the wings in hovering.

The purpose of this paper is to develop a real‐time simulation environment for the validation of controller for an autonomous small‐scale helicopter.

The real‐time simulation platform is developed based on the nonlinear model of a series of small‐scale helicopters. Dynamics of small‐scale helicopter is analyzed through simulation. The controller is designed based on the extracted linear model.

The model‐based linear controller can be effectively designed and tested using real‐time simulation platform. The hover controller is demonstrated to be robust against wind disturbance.

To use the real‐time simulation environment to test and validate controllers for small‐scale helicopters, basic helicopter parameters need to be measured, calculated or estimated.

The real‐time simulation environment can be used generically to test and validate controllers for small‐scale helicopters.

The paper presents the design and development of a low‐cost hardware in the loop simulation environment using xPC target critical for validating controllers for small‐scale helicopters.

This paper aims to build an accurate mathematical model which is necessary for control design and attitude estimation of a miniature unmanned rotorcraft and its subsequent conversion to an autonomous vehicle.

Frequency-domain system identification of a small-size flybar-less remote controlled helicopter is carried out based on the input–output data collected from flight tests of the instrumented vehicle. A complete six degrees of freedom quasi-steady dynamic model is derived for hover and cruise flight conditions.

The veracity of the developed model is ascertained by comparing the predicted model responses to the actual responses from flight experiments and from statistical measures. Dynamic stability analysis of the vehicle is carried out using eigenvalues and eigenvectors. The identified model represents the vehicle dynamics very well in the frequency range of interest.

The model needs to be augmented with additional terms to represent the high-frequency dynamics of the vehicle.

Control algorithms developed using the first principles model can be easily reconfigured using the identified model, because the model structure is not altered during identification.

This paper gives a practical solution for model identification and stability analysis of a small-scale flybar-less helicopter. The estimated model can be easily used in developing control algorithms.

A NUMBER of approaches to the calculation of rotor downwash have already been discussed. Broadly spsaking, the methods of Castles and DeLeeuw and Squire and Mangler are the same. In both methods the downwash at the rotor disk is assumed to be perpetrated in a helical downwash sheet which, as the slipstream, extends below the rotor to infinity. The downwash in the disk due to the bound vortices, and the additional downwash in the disk which is induced by the helical sheets in the slipstream (Castles and DeLeeuw substitute downwash rings for helices, in the interest of mathematical simplicity) is calculated, on the assumption of an infinite number of lightly loaded blades. The final results of Castles and DeLeeuw on the one hand, and Squire and Mangier on the other, are in very wide disagresment. This disagreement is principally due to the fact that, whereas the first investigation assumes constant circulation along the blade (ideal twist and taper), Mangier and Squire assume a ‘practical’ variation of the form likely to be encountered on an untwisted untapered blade. We conclude that the radial distribution of lift on a helicopter blade will have a profound effect on the downwash pattern: which in turn will affect the calculated lift.