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High requirements have been placed on the stealth performance of armed helicopters during wartime missions，in order to reduce the probability of the helicopter being found during the task，based on geometric optics method and consistent diffraction theory，the effects of different tail rotor profile thickness，curvature and tail rotor scissors angle on RCS are studied. By comparing different parameters of the tail rotor and the RCS peak value and mean value of the coated absorbing material，the conditions which are more favorable to the stealth of the helicopter tail rotor are determined. When viewed from the ground，the RCS with small tail rotor thickness and small curvature will be reduced appropriately，and the change of scissors angle will lead to the change of RCS peak phase. When the radar is irradiated in parallel，the RCS of tail rotor with smaller thickness is smaller，and the RCS peak value of tail rotor with smaller curvature is smaller. Applying wave absorbing materials on the leading edge and tip of the blade can effectively reduce the overall RCS of the blade. Comprehensive analysis shows that by using the smaller thickness and curvature of the tail rotor，it is beneficial to reduce the RCS of the helicopter tail rotor，and the blade tip surface and the blade leading edge of the absorbing material coating can effectively reduce the overall RCS.

Finding out the reason of the fracture fault of tail rotor on a helicopter is of practical significance for using and maintaining the helicopter and is of significant reference for improving safety of new type helicopter. The fracture fault of tail rotor blade on a helicopter is analyzed by fault tree analysis method, and the bottom event analysis results are obtained. Then, the macro and micro fractographic analysis of the tail rotor blade is carried out. Finally,the simulation analysis of the fractographic damage evolution of flexible beam is performed. The results show that the fracture property of the blade is low-stress and high-cycle bidirectional bending fatigue fracture. When the strength value of the fiber compression direction of the flexible beam is less than 700 MPa, the fractographic type is similar to that of the faulty tail rotor flexible beam. The tail rotor will produce large vibration under large fatigue load condition, and the increase of vibration can intensify the load of the tail rotor, which leads to the increase of load at the root of the flexible beam and exceeds the designed bearing capacity of the flexible beam, then fatigue fracture will occur at the root of the flexible beam of the blade.

To investigate the potential of Statically and Dynamically Extendable Chord (SEC/DEC) in performance improvement of helicopter tail rotors, a prediction model of helicopter flight performance is developed and validated. It is suitable to deploy the SEC and DEC close to the blade tip for power reduction of tail rotor in hover and high-speed flight. The deployment locations of SEC and DEC for tail rotors are similar to that of main rotors. In hover, the deployment of the SEC and DEC results in little power reduction. From low to medium speed flight, the deployment of the DEC and SEC increases the power. In high-speed flight, the SEC and DEC can significantly decrease the power, especially at a large take-off weight. At 300 km/h, when the chord extension is 20% of blade chord, the SEC and DEC with 1/rev input decrease the power by 14.44% and 20.47% at a take-off weight coefficient of 0.0091. The DEC shows greater potential in power reduction than the SEC. There is an optimal phase input for the maximum power reduction, which does not vary with forward speed. For the DEC with 1/rev input, the optimal phase is 0 °. The optimal phase input for the maximum power reduction of tail rotors is different from that of main rotors. [ABSTRACT FROM AUTHOR]

Dynamically extendable chords with controlled non-harmonic motions are studied for improving the performance of variable-speed tail rotors. A validated helicopter performance model is utilized, which consists of a rotor, a fuselage, a tail rotor and a propulsive trim method. The results show that the optimal azimuth angles for the deployment of extendable chords are 40°-50° and 130°-140° at the nominal tail rotor speed. In order to improve the efficiency of the extendable chord, a non-harmonic motion of the extendable chord is proposed. Extendable chords can reduce the blade area with high drag and lower the angle-of-attack, which can lead to significant power savings during high-speed flight. Extendable chords are more suitable to reduce the power of variable-speed tail rotors. When the tail rotor speed is reduced by 20%, the power can be reduced up to 19.4%. [ABSTRACT FROM AUTHOR]

The paper is focused on the numerical study of the aerodynamic characteristics of 3-bladed and scissors tail rotors used on Mi-8/17 series helicopters. The hovering and axial flow modes, including the Vortex Ring State (VRS) area are considered. The free wake model developed by authors is used in the research. The study has been performed for two approaches: with a fixed blade pitch angle and with a fixed time-average thrust value. The dependences of the rotor thrust and torque coefficients vs. axial flow velocity and vs. number of revolutions of a rotor have been obtained. The vortex wake shapes and flow images via streamlines have been analyzed. The comparative analysis of the results has been performed. Scissors rotor demonstrates better aerodynamic characteristics in comparison with 3-bladed rotor. Among them: increased hovering efficiency, lower level of thrust and torque pulsations in the VRS modes, smaller power consumption while constant thrust. The authors came to the conclusion about the greater efficiency of the aerodynamic design of scissors tail rotor compared to 3-bladed rotor both in hover and in the VRS modes area. [ABSTRACT FROM AUTHOR]

Aviation accidents related to unintentional rotation may periodically occur while flying single-rotor helicopters. On-time and correct actions may help the pilot to find the way out of this hazardous situation. But it is also important to understand the situation which contributes to the unanticipated yaw occurrence, and whether there are any factors which can stop the pilot from preventing such unintentional rotation, in order to avoid these conditions. Literature analysis shows that researchers studying this phenomenon don’t have the shared vision on unanticipated yaw occurrence conditions. In regards to this fact the decision to carry out a series of wind tunnel experiments using helicopter model and propeller was taken. The main object of research was a radio-controlled model of the Blade 130 x helicopter, mounted on a platform rotating around a vertical axis, which was installed on a vertical strut. Research-laboratory aerodynamic complex belonging to the Aerodynamics, Design and Aircraft Strength Chair of Moscow State Technical University of Civil Aviation was used to generate airflow. A set of dynamic experiments was carried out to determine the conditions contributing to unanticipated yaw occurrence. The analysis of the experiments has shown that there is a range of sliding angles at a certain speed of the incoming air flow which makes the helicopter yaw balancing impossible, and if the helicopter occasionally gets into this range, it inevitably leads to the unintended rotation of the helicopter on the yaw occurrence. Helicopter yaw trim inability occurs at negative sideslip angles because of tail rotor thrust decrease due to the incoming airflow blowing which decreases the blades angles of attack and worsens helicopter airframe aerodynamic moment that coincides in direction with main rotor torque if helicopter airframe possesses directional stability. In these conditions the required tail rotor pitch is greater than the available pitch so the pilot is not able to counteract the initiated unanticipated yaw rotation of the helicopter that has begun. The possibility of helicopter unanticipated yaw rotation caused by the impact of the main rotor on the tail rotor was not experimentally confirmed. It was impossible to create the conditions of unanticipated yaw occurrence during the experiments because of the tail rotor vortex ring state.

With the promotion and enhancement of stealth technology of helicopter rotor components, the research on the dynamic radar cross section (RCS) of helicopter rotor is becoming more and more important and imminent. In order to facilitate the calculation and analysis of the electromagnetic scattering characteristics during rotor rotation, a dynamic scattering calculation (DSC) method based on quasi-static principle (QSP) and grid coordinate transformation is presented. After analyzing the advantages and disadvantages of QSP, the dynamic principle is used to describe the rotation process of the rotor. Combined with the grid coordinate transformation method, the RCS of the rotor is accurately calculated by physical optics (PO) and physical theory of diffraction (PTD). Then the influence of azimuth, elevator angle and observation distance on rotor dynamic RCS is analyzed. The results show RCS of the tail rotor is indeed dynamic and periodic and its main influencing factors include azimuth and elevation angle. The proposed DSC method is efficient and effective for studying the dynamic RCS of tail rotor. [ABSTRACT FROM AUTHOR]

The radar image construction algorithms for propellers of aerial vehicles belonging to the classes «helicopter» and «propeller-driven aircraft» were proposed in [1]. The algorithms are designed to be implemented in a radar sensor with inverse synthesis of antenna aperture. The fundamental condition for the qualitative construction of radar images of the propellers is accurate measurement of their rotation frequency. The measurement method for the repetition frequency of the propeller blades on the base of the convolution of the spectrum of the «secondary» modulation of the signal reflected from propeller was proposed in [2]. This paper discusses the results of experimental studies of the measurement method for repetition frequency of the propeller blades and of the radar images construction algorithms of propellers of typical low-altitude aircraft (helicopters and propeller-driven aircrafts).

A novel measurement method for repetition frequencies of the blades of an aerial vehicle (in part, helicopter's main and tail rotor) for adaptation circuits of the radar sensor with an inverse synthesis of antenna aperture is proposed. The method based on the convolution of the spectrum of the «secondary» modulation of the signal with simultaneous compensation of the Doppler frequency of the reflected signal. This allows in typical conditions for the radar sensor rapid change of Doppler frequency to provide a long coherent accumulation and, accordingly, high accuracy of measurement of the rotational frequencies of the rotors.

This paper describes the effect of different rotor blades on the X-shaped tail rotor of the Mi-171 LL, observed conducting flight tests. The tests were carried out on the same helicopter in the similar atmospheric conditions.The objective of the tests was the comparison of flight performance of two sets of rotor blades of the helicopter Mi-171 LL. However, materials test revealed a difference in the angles of the tail rotor at different MRs with the same takeoff weight.The authors are grateful to I.G. Peskov, S.R. Zamula and A.I. Orlov for assistance in carrying out this work and the preparation of this article.Noted that the helicopter takeoff weight when hovering out of ground effect in ISA with blades from polymer composite materials (PCM) exceeds the takeoff weight of the helicopter with the serial blades in the nominal mode of the engine operation at ~ 750kg, in the takeoff mode at ~ 700kg.Knowing the altitude and climatic characteristics of the engine, the obtained dependence allows to determine the balancing value of jрв on hovering at different combinations of pressure altitude and outside air temperature for a given speed of the main rotor (MR).It follows from the work that when the same value Nпр(95/nнвпр)3 or Nfact the balancing values of jрв for the helicopter with the main rotor blades from the PCM is less than for the helicopters with serial blades by 0.5…0.9°. The difference in the angles of the tail rotor increases with growing of Nепр(95/nнвпр)3 (Nfact). Perhaps this is caused by different induction effect of the main rotor on the tail rotor to the MR from PCM and the serial ones.As follows from the materials, the thrust of the main rotor with blades from PCM with the same engine power is more in comparison with the serial blades. Consequently inductive speeds of the main rotor are more and the angles of the tail rotor are less. It can be assumed that a large induced velocity of the main rotor increases the thrust of X-shaped tail rotor.

For single-rotor helicopters there are special flight modes, when tail rotor (TR) is under significant inductive influence of vortical wake of main rotor (MR). Inductive influence of vortical wake of MR can provoke essential changes in flowing of TR and its aerodynamic characteristics comparing to isolated rotor. In this case increase of tail rotor pitch, necessary for helicopter controlling, is possible.The article contains computational modelling of TR work with vortical wake of MR at the example of MIL Mi-171 helicopter. The modelling has been made on the base of non-linear blade (free wake) vortical model of rotor, produced at Helicopter Design Department of MAI.Helicopter hovering modes with crosswind of various intensity Vz was considered. Thrust-time lationship forisolated TR and TR with vortical wake of MR for equal flight modes was obtained. Flow around the rotors was analyzed, its vortical wake was considered.The results make it possible to clarify the peculiarities of TR work on considered modes and MR influence on its work. It was found out that vortical wake of MR has a more significant impact on TR work with crosswind on the right, when TR falls into vortex ring state mode. Inductive influence of vortical wake of MR leads to vortexring state mode for TR on lower speeds (Vz :: 5 m/s) than in case of isolated work of TR (Vz :: 12,5 m/s). In that case,the required tail rotor pitch has increased by 13% for Vz = 5 m/s. The results of modelling and flight tests led to good agreement.

This article is concentrated on modeling and researching of aerodynamic characteristics of main and tail rotors’ combination of a single-rotor helicopter scheme. Non-linear free wake vortical model with diffusion of free wake is used. The model is constructed on the base of Helicopter Design Faculty, MAI. There was analyzed a number of special flight regimes of a helicopter. There were counted aerodynamic characteristics of rotors.

The article presents the results of numerical studies of aerodynamic characteristics of isolated three-bladed tail rotor, and X-shaped tail rotor at "vortex ring". Calculations of aerodynamic characteristics of rotors are based on nonlinear blade vortex model developed at the Department of "The helicopter design" MAI. Considered hover and blasting at positive angles of attack αH, including the area of "vortex ring". Built "vortex ring" area in the coordinates of the horizontal Vx and vertical velocity Vy of airflow at different angles of attack of the rotor by condition drop rotor thrust against hover thrust at fixed angles of rotor collective pitch.

Variable tail rotor speed is investigated as a method for reducing tail rotor power, and improving helicopter performance. A helicopter model able to predict the main rotor and tail rotor powers is presented, and the flight test data of the UH-60A helicopter is used for validation. The predictions of the main and tail rotor powers are generally in good agreement with flight tests, which justifies the use of the present method in analysing main and tail rotors. Reducing the main rotor speed can result in lower main rotor power at certain flight conditions. However, it increases the main rotor torque and the corresponding required tail rotor thrust to trim, which then decreases the yaw control margin of the tail rotor. In hover, the tail rotor may not be able to provide enough thrust to counter the main rotor torque, if it is slowed to follow the main rotor speed. The main rotor speed corresponding to the minimum main rotor power increases, if the change of tail rotor power in hover is considered. As a helicopter translates to cruise, the induced power decreases, and the profile power increases, with the profile power dominating the tail rotor. Reducing the tail rotor speed in cruise reduces the profile power to give a 37% reduction in total tail rotor power and a 1.4% reduction to total helicopter power. In high-speed flight, varying the tail rotor speed is ineffective for power reduction. The power reduction obtained by the variable tail rotor speed is reduced for increased helicopter weight. [ABSTRACT FROM PUBLISHER]