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Purpose The purpose of this study is to design a flight control model for a control surface-less (CSL) tri-tilt-rotor (TTR) unmanned aerial vehicle (UAV) based on a Proportional Integral Derivative (PID) controller to stabilize the altitude and attitude of the UAV subjected to various flying conditions. Design/methodology/approach First, the proposed UAV with a tilting mechanism is designed and analyzed to obtain the aerodynamic parameters. Second, the dynamics of the proposed UAV are mathematically modeled using Newton-Euler formation. Then, the PID controller is implemented in the simulation model to control flight maneuvers. The model parameters were implemented in a mathematical model to find the system’s stability for various flight conditions. The model was linearized to determine the PID gain values for vertical take-off and landing, cruise and transition mode. The PID controller was tuned to obtain the desired altitude and attitude in a short period. The tuned PID gain values were implemented in the PID controller and the model was simulated. Findings The main contribution of this study is the mathematical model and controller for a UAV without any control surface and uses only a thrust vector control mechanism which reduces the complexity of the controller. The simulation has been carried out for various flight conditions. The altitude PID controller and the attitude PID controller for CSL-TTR-UAV were tuned to obtain desired altitude and attitude within the optimum duration of 4 s and deviation in the attitude of 8%, which is within the allowable limit of 14%. The findings obtained from the simulation revels that the altitude and attitude control of the CSL-TTR-UAV was achieved by controlling the rpm of the rotor and tilt angle using the PID controller. Originality/value A novel CSL TTR UAV mathematical model is developed with a dual tilting mechanism for a tail rotor and single axis tilt for the rotors in the wing. The flight control model controls the UAV without a control surface using a PID controller for the thrust vector mechanism.

Reverse pedal operational property in front crosswind flight condition is a potential hazard for accidents involving loss of tail rotor effectiveness (LTE), which is closely related to the main rotor (MR) wake interference on the tail rotor (TR). As understanding of this interaction is vital for the early warning strategy development, the MR wake influence effect on TR thrust and the effect of helicopter yaw stability are examined in this study. For this purpose, the comparison of TR thrust and flow field with wind azimuth and speed in front crosswind environment was performed by experiment and CFD simulation, respectively. Test campaign was performed at a 5.5 m × 4 m wind tunnel in the China Aerodynamics Research and Development Center using a high-position bottom-blade forward-rotating TR and a counterclockwise rotating MR to address the TR thrust under wind speeds of 8–22 m/s with 50°, 60°, and 70° wind azimuths. The influence of MR disc loading was also contrasted. CFD analysis was used to gain insight into the flow physics responsible for the interference effect. It was conducted with unsteady Reynolds-averaged Navier–Stokes simulations, where the MR using the actuator disk approach and the TR blade rotation was modeled via a sliding mesh method. Results indicated that the MR disc vortex has a remarkable interference effect on the TR aerodynamic performance characteristic and that the effect is sensitive to the wind speed, wind direction, and MR disc loading. The observed yaw instability is considered to be related to the lesser inflow introduced by the MR disc vortex due to the change in the relative position of the disc vortex filament and TR with the wind azimuth. The increase in TR thrust at moderate wind speeds is due to the increase in leading edge dynamic pressure caused by the opposite swirl direction of the disc vortex contrasted to the TR. The MR disc loading affects the TR thrust due to the change of disc vortex strength and position.

The complex three-shaft three-reducer structural designs of helicopter transmission systems are prone to changes in the relative positions of shafting under the conditions of main rotor and tail rotor loads. These changes will affect the transmission characteristics of the entire transmission system. In this study, the planetary gear trains of helicopters were examined. Due to the fact that these structures are considered to be the most representative structures of the main reducers of helicopters, they were selected as the study objects for the purpose of examining the meshing characteristics of planetary gear trains when the relative positions of the shafting changed due to the position changes of the main rotor shafts under variable load conditions. It was found that by embedding the comprehensive time-varying meshing stiffness values of the main rotor shafts at different positions, a dynamic model of the relative position changes of the planetary gear trains could be established. Then, combined with the multibody dynamics software, the meshing characteristics of the sun gears, and the planetary gears, the planetary gears and the inner ring gears were simulated and analyzed under different inclinations and offsets of the shafting. The results obtained in this study revealed the following: (1) the average meshing force of the gears increased with the increases in the angle inclinations, and the meshing force between the sun gears and the planetary gears increased faster than the meshing force between the planetary gears and the inner ring gears. It was observed that during the changes in the shafting tilt positions, obvious side frequency signals had appeared around the peak of the meshing frequency in the spectrum. Then, with the continuous increases in the tilt position, the peak was gradually submerged; (2) the average meshing force of the gears increased with the increases in the offset, and the increasing trend of the meshing force between the sun gears and the planetary gears was similar to that observed between the planetary gears and the inner ring gears. It was found that when the shafting offset position changed, there were obvious first and second frequency doubling in the spectrum; (3) the mass center orbit radii of the sun gears increased with the increases in the shafting position changes, and the changes in the angular tilt position were found to have greater influencing effects on the mass center orbit radii of the sun gears than the changes in the offset positions. This study’s research findings will provide a theoretical basis for future operational status monitoring of the main transmission systems of helicopters and are of major significance for improvements in the operational stability of helicopter transmission systems, which will potentially ensure safe and efficient operations.

A new calculation method of helicopter rotor/fuselage acoustic scattering is developed. Firstly, a CFD analysis model is developed to simulate flow field of the rotor, which is based on the motional embedded grid system and RANS equations, and provides aerodynamic data for rotor noise calculation. Then, FW-H equations are employed to calculate the aeroacoustic characteristics of isolated rotor, and G 1A formulas are applied to calculate the rotor acoustic gradient to provide boundary condition for acoustic scattering. Based on these, the time-domain equivalent source method is applied to calculate acoustic scatter field, and the total acoustic field that considered the fuselage scatter is superposed by isolated rotor acoustics and the scatter one. Finally, the numerical simulations of helicopter main-rotor/fuselage and tail-rotor/fuselage scatter effect are conducted by using the developed models. The results indicate that the helicopter fuselage has important scatter effect on the high frequency acoustics of main rotor and tail rotor, and the acoustic scatter effect become more obvious with the smaller space between the main rotor (tail rotor) and fuselage.

The tail rotor of a helicopter operating under low-speed crosswind undergoes highly complex flow due to the interaction between the main rotor, fuselage, and tail rotor system. In this study, numerical simulations have been conducted on the complete configuration of a helicopter with a ducted fan tail rotor system (comprising a main rotor, ducted fan tail rotor, fuselage, and empennage) to analyze the wake interaction in hovering flight under various crosswind directions. The flow characteristics around the tail rotor, the tail rotor thrust, and the yawing moment of the helicopter are investigated and evaluated. The aerodynamic forces are compared with those of a helicopter with an open-type tail rotor. The results indicate that the aerodynamic performance of the ducted fan tail rotor is highly affected by the wakes of both the main rotor and port wing. Nevertheless, the helicopter with a ducted fan tail rotor is observed to be much more directionally stable under various crosswind directions, than that with an open-type tail rotor. This is because the rotor is protected by the fixed part of the tail rotor system in the former case.

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.

Lightning is one of the most important threats to the safe operation of an aircraft. Rotorcraft can experience lightning strikes on the main rotor blades and tail rotor blades which can seriously affect the rotorcraft and its components in several ways. When lightning strikes a rotor blade made of composite material, electrical–thermal multi-physical phenomena such as the Joule heating effect occur. In this study, an electrical-thermal computational simulation of lightning strikes on composite rotor blades was performed. The simulation analyzed various situations including the presence of a lightning protection system on the composite rotor blade. Emphasis was placed on the effects of geometric parameters of expanded metal foil on the lightning protection performance. In addition, the effects of curvature on thermal and electrical response were investigated. The thermally damaged area was found to substantially increase in inner layers compared to the flat plate case.

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.

The insistence of this paper is to design Fractional order-proportional Integral Derivative (FO-PID) and Integral order –Proportional Integral Derivative (IO-PID) controllers for trajectory tracking control of Twin rotor Multiple Input Multiple Output (MIMO) system (TRMS). The aim of these designed controllers is to move the main and tail actuator at the desired angles accurately to fulfill the demanding attributes so that TRMS should stabilize while hovering. The controller parameters of both the controllers are optimized using fmincon function as well as Multiobjective genetic algorithm. Both the controller are tested in real time and their performance are compared for control of yaw and pitch angles of TRMS. Manual disturbances are given at after 40 seconds of first given disturbance, and the designed controllers are highly capable in diminishing quickly the effect of manual perturbations given to the main and tail rotor of TRMS. From experiment it is observed from comparative performance that FO-PID controller tuned using multiobjective genetic algorithm give best performance in term of time taken to stabilization after the given disturbance.

For the integrated stealth issue of ducted tail rotor noise radiation and radar scattering, a comprehensive optimization method based on high frequency electromagnetic calculation theory and boundary element method is introduced. The radar cross section of the rotor, radiated noise and radar cross section of the ducted tail rotor are designed as optimization goals under the constraints of geometric parameters and aerodynamic force. The model of the ducted tail rotor is established by using the full factorial design and the flow field is constructed by high-precision unstructured grid technology. The aerodynamic characteristics of the ducted tail rotor is simulated by the computational fluid dynamics method based on Navier–Stokes equations and k–e standard viscous model. The noise radiation is solved by boundary element method and the radar cross section value is calculated by physical optics method and physical theory of diffraction. On the basis of these calculation results, the optimization model of the ducted tail rotor is obtained and generated by comprehensive optimization method based on Pareto solution. The final scheme has been satisfactorily improved in terms of noise suppression, radar cross section reduction and aerodynamic lifting. The proposed approach is very effective and efficient for the acoustic/radar comprehensive stealth design of the ducted tail rotor.

The present work is aimed at illustrating a hybrid analytical/numerical methodology that could be used at the early-design stage to predict the tonal noise of an axial-flow fan surrounded by solid surfaces. A model shrouded helicopter tail-rotor architecture is selected as a typical example of unconventional rotor-stator stage. The advanced design involves uneven rotor-blade spacing and leant stator vanes, one of which is replaced by the transmission shaft. Both the rotor and the stator generate tonal noise at the same frequencies, because of the potential distortion around the shaft and because of the impingement of the blade wakes on the vanes, respectively. Source-modes made of circular distributions of stationary sources properly phase-shifted are introduced to reproduce all sources in a unified way, based on analytical expressions. This part is based on simplified unsteady aerodynamic theories. The source-modes are then used as input in a numerical software solving the Helmholtz equation to assess the sound scattering by the shroud. The approach is applied only to the stator noise for illustration and typical results are discussed.

Purpose The purpose of this paper is to determine dependencies between a rotor-blade shape and a rotor performance as well as to search for optimal shapes of blades dedicated for helicopter main and tail rotors. Design/methodology/approach The research is conducted based on computational methodology, using the parametric-design approach. The developed parametric model takes into account several typical blade-shape parameters. The rotor aerodynamic characteristics are evaluated using the unsteady Reynolds-averaged Navier–Stokes solver. Flow effects caused by rotating blades are modelled based on both simplified approach and truly 3D simulations. Findings The computational studies have shown that the helicopter-rotor performance may be significantly improved even through relatively simple aerodynamic redesigning of its blades. The research results confirm high potential of the developed methodology of rotor-blade optimisation. Developed families of helicopter-rotor-blade airfoils are competitive compared to the best airfoils cited in literature. The finally designed rotors, compared to the baselines, for the same driving power, are characterised by 5 and 32% higher thrust, in case of main and tail rotor, respectively. Practical implications The developed and implemented methodology of parametric design and optimisation of helicopter-rotor blades may be used in future studies on performance improvement of rotorcraft rotors. Some of presented results concern the redesigning of main and tail rotors of existing helicopters. These results may be used directly in modernisation processes of these helicopters. Originality/value The presented study is original in relation to the developed methodology of optimisation of helicopter-rotor blades, families of modern helicopter airfoils and innovative solutions in rotor-blade-design area.

Control design of a TRMS system such as helicopter is very complicated task. Since the TRMS are very high non-linear system, so its non-linearity too high and the control design is very difficult problem for TRMS system. It has two rotors, the first one is the main rotor and the second one is tail rotor. The TRMS System are based on the beam and support by a counter balance. Horizontal and vertical plane are included in the TRMS System for PID and LQR controller both are discuss in this paper. PID and LQR controller both are discuss in this paper. By Simulink result for vertical and horizontal plane, LQR controller better the PID control. All the simulation work is done by the MATLAB/Simulink of the results, environment and the working of simulation results have been done at the end of this paper.

Twin-rotor two-input two-output (TITO) aerodynamical system (TRAS) is a highly non-linear system with a cross-coupling effect. The objective of the controller design in TRAS is to track the trajectory of azimuth and pitch angles, reach the desired azimuth and pitch angle positions, and stabilize TRAS in the presence of significant cross-coupling. The coupling effect between the main rotor and tail rotor is a severe issue for the controller design. So, this paper proposes a proportional integral derivative double derivative (PIDD2) controller design for TRAS, first time. Further, the tuning of PIDD2 controller is carried out via a modified grey wolf optimizer (MGWO). Besides, the real-time hardware-in-loop (HIL) implementation of the PIDD2 controller with the TRAS laboratory prototype setup is carried out, first-time. Moreover, It is tested under external disturbances for robustness and effectiveness analyses. In addition, this paper performs both simulations as well as hardware results analyses. These analyses reveal that the experimental results are found to match closely with simulation results. Finally, the efficacy of the proposed control design approach is presented by comparing it with the recently published controldesign schemes.

Helicopter tail rotors adopt a segmented driveline connected by flexible couplings, and dry friction dampers to suppress resonance. Modeling for this system can provide a basic foundation for parameter analysis. In this work, the lateral-torsional vibration equation of the shaft with continuous internal damping is established. The static and dynamic effects caused by flexible diaphragm couplings subject to parallel and angular misalignment is derived. A novel dual rub-impact model between the shaft and dry friction damper with multiple stages is proposed. Finally, a model of a helicopter tail rotor driveline incorporating all the above elements is formulated. Numerical simulations are carried out by an improved Adams–Bashforth method following the design flowchart. The dynamics of multiple vibration suppression, and the static and dynamic misalignment are analyzed to illustrate the accuracy and characteristics of the model. The coeffect of the rub impact and the misalignment on shafts and dampers are presented through the results of simulation and experiment. It provides an accurate and comprehensive mathematical model for the helicopter driveline. Response characteristics of multiple damping stages, static and dynamic misalignment, and their interaction are revealed.

There is currently interest in the design of small electric vertical take-off and landing aircraft to alleviate ground traffic and congestion in major urban areas. To support progress in this area, a conceptual design method for single-main-rotor and lift-augmented compound electric helicopters has been developed. The design method was used to investigate the feasible design space for electric helicopters based on varying mission profiles and technology assumptions. Within the feasible design space, it was found that a crossover boundary exists as a function of cruise distance and hover time where the most efficient configuration changes from a single-main-rotor helicopter to a lift-augmented compound helicopter. In general, for longer cruise distances and shorter hover times, the lift-augmented compound helicopter is the more efficient configuration. An additional study was conducted to investigate the potential benefits of decoupling the main rotor from the tail rotor. This study showed that decoupling the main rotor and tail rotor has the potential to reduce the total mission energy required in all cases, allowing for increases in mission distances and hover times on the order of 5% for a given battery size.

The Twin Rotor Mimo System (TRMS) consists of two rotors namely tail rotor (∅) and main rotor (ψ). The main objective of this article is improving the stability of the TRMS system in rotation and elevation. To achieve that, a predictive controller based on the model (MPC), and a sliding mode controller (SMC) were implemented. The analytical modelling is based on the study of the system dynamics to obtain a linear model. Additionally, concepts of control involved in the design of the MPC, PID and SMC controllers are considered.

A comprehensive helicopter flight dynamics code is developed based on the UH-60 helicopter and named Texas A\&M University Rotorcraft Analysis Code (TRAC). This is a complete software package, which could perform trim analysis to autonomous flight simulation and the capability to model any helicopter configuration. Different components of the helicopter such as the main rotor, tail rotor, fuselage, vertical tail, and horizontal tail are modeled individually as different modules in the code and integrated to develop a complete UH-60 model. Since the code is developed on a module basis, it can be easily modified to adopt another component or configure a different helicopter. TRAC can predict the dynamic responses of both the articulated rotor blades and the helicopter fuselage and yields the required pilot control inputs to achieve trim condition for different flight regimes such as hover, forward flight, coordinated turn, climb/descent, etc. These trim results are validated with the test data obtained from the UH-60 flight tests conducted by the US Army. Beyond trim analysis, TRAC can also generate linearized models at various flight conditions based on a first-order Taylor series expansion. The extracted linear models show realistic helicopter dynamic behavior and were used to simulate a fully autonomous flight that involves a UH-60 helicopter approaching a ship and landing on the deck by implementing a Linear Quadratic Regulator (LQR) optimal controller., For software package, contact author via bochan.lee@tamu.edu

Calculating the unsteady convective heat transfer on helicopter blades is the first step in the prediction of ice accretion and the design of ice-protection systems. Simulations using Computational Fluid Dynamics (CFD) successfully model the complex aerodynamics of rotors as well as the heat transfer on blade surfaces, but for a conceptual design, faster calculation methods may be favorable. In the recent literature, classical methods such as the blade element momentum theory (BEMT) and the unsteady vortex lattice method (UVLM) were used to produce higher fidelity aerodynamic results by coupling them to viscous CFD databases. The novelty of this research originates from the introduction of an added layer of the coupling technique to predict rotor blade heat transfer using the BEMT and UVLM. The new approach implements the viscous coupling of the two methods from one hand and introduces a link to a new airfoil CFD-determined heat transfer correlation. This way, the convective heat transfer on ice-clean rotor blades is estimated while benefiting from the viscous extension of the BEMT and UVLM. The CFD heat transfer prediction is verified using existing correlations for a flat plate test case. Thrust predictions by the implemented UVLM and BEMT agree within 2% and 80% compared to experimental data. Tip vortex locations by the UVLM are predicted within 90% but fail in extreme ground effect. The end results present as an estimate of the heat transfer for a typical lightweight helicopter tail rotor for four test cases in hover, ground effect, axial, and forward flight.

With the continuous improvement and development of armed helicopters, the research on their stealth characteristics has become more and more in depth. In order to obtain the complex effect of stealth characteristics caused by the high-speed rotation of rotor-like components, a dynamic scattering method (DSM) is presented. Rotation speed, azimuth, elevation angle, pitching angle, and rolling angle are studied and discussed in detail. The results show that the electromagnetic scattering characteristics of the main rotor and tail rotor are dynamic and periodic. This period characteristic is related to the rotation speed and attitude angle of the rotor. The radar cross-section (RCS) of the helicopter varies greatly at different observation angles and attitude angles, but the dynamic electromagnetic scattering effect caused by the main rotor and tail rotor cannot be ignored. The presented DSM is effective and efficient for studying the dynamic RCS of the rotor-type parts of a helicopter or the whole machine.

Condition monitoring and quick fault detection and diagnostics of machinery problems can lead to significant reduction in operational and maintenance costs in addition to improving the safety level. In this paper, a new method is proposed for fault detection and diagnoses of the AH-64D helicopter tail rotor drive shaft problems. The proposed method depends on extracting minimum number of indicative features using search algorithm based on wavelet decomposition and power spectrum of vibration signals, then feeding them to a Neuro-fuzzy classification algorithm. The proposed method is verified by using experimental vibration data collected using an AH-64D helicopter test bed. The proposed algorithm shows superior performance in finding a set indicative features and classifying different shaft faults.

Tilt rotor aircrafts combine the advantages of the long mileage of fixed-wing airplanes and the vertical take-off and landing (VTOL) capabilities of rotorcrafts by using tiltable rotors. However, tilt rotor aircrafts suffer from poor autorotation performance and landing stabilitiy. This work proposes an alternative actuation configuration of a tricopter tilt rotor aircraft that aims to optimize the flight duration, VTOL capabilities and stability control. The proposed configuration employs two front tilt rotors that are able to provide thrust-vectoring control with the addition of a fixed tail rotor that produces upward thrust. Dynamic modeling of the proposed actuation configuration is developed to provide insights on how different motions (hovering, ascending, descending, rotating and airplane-like flying) are achieved by manipulating the amounts of thrusts and rotor tilt angles. Preliminary prototype as well as the electronic and control systems are developed and discussed. Proposed configuration has the potential to achieve more sustainable flying.

With the development of electric helicopters’ motor technology and the widespread use of electric drive rotors, more aircraft use electric rotors to provide thrust and directional control. For a helicopter tail rotor, the wake of the main rotor influences the tail rotor’s inflow and wake. In the procedure of controlling, crosswind will also cause changes to the tail disk load. This paper describes requirements and design principles of an electric motor drive and variable pitch tail rotor system. A particular spoke-type architecture of the motor is designed, and the performance of blades is analyzed by the CFD method. The demand for simplicity of moving parts and strict constraints on the weight of a helicopter makes the design of electrical and mechanical components challenging. Different solutions have been investigated to propose an effective alternative to the mechanical actuation system. A test platform is constructed which can collect the dynamic response of the thrust control. The enhancement of the response speed due to an individual motor speed control and variable-pitch system is validated. Keywords: Electric tail rotor, Helicopter, More Electric Aircraft (MEA), Propulsion optimization, Tail rotor

SUMMARYA generally accepted helicopter model used for control includes the effect of Small Body Forces (SBF) which couple the vehicle's rotational subsystem inputs to its translational dynamics. SBF result from tail rotor thrust and lateral forces due to main rotor flapping. It is well-known that SBF lead to a theoretically challenging stabilization problem for the tracking error dynamics. Hence, much of the existing work has neglected SBF in order to simplify control design. We design a controller that directly compensates the influence of the tail rotor component of the SBF. The design is validated in simulation and flight tests.

Over the past decade, NASA, under a succession of rotary-wing programs, has been moving towards coupling multiple discipline analyses to evaluate rotorcraft conceptual designs. Handling qualities is one of the component analyses to be included in such a future Multidisciplinary Analysis and Optimization framework for conceptual design of Vertical Take-Off and Landing (VTOL) aircraft. Similarly, the future vision for the capability of the Concept Design and Assessment Technology Area of the U.S Army Aviation Development Directorate also includes a handling qualities component. SIMPLI-FLYD is a tool jointly developed by NASA and the U.S. Army to perform modelling and analysis for the assessment of the handling qualities of rotorcraft conceptual designs. Illustrative scenarios of a tiltrotor in forward flight and a single-main rotor helicopter at hover are analysed using a combined process of SIMPLI-FLYD integrated with the conceptual design sizing tool NDARC. The effects of variations of input parameters such as horizontal tail and tail rotor geometry were evaluated in the form of margins to fixed- and rotary-wing handling qualities metrics and the computed vehicle empty weight. The handling qualities Design Margins are shown to vary across the flight envelope due to both changing flight dynamics and control characteristics and changing handling qualities specification requirements. The current SIMPLI-FLYD capability, lessons learned from its use and future developments are discussed.

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 j рв 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 N пр (95/n нвпр ) 3 or Nfact the balancing values of j рв 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 N епр (95/n нвпр ) 3 (N fact ). 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.

Sparse signal representation is used to coherently separate the main rotor blade components of a helicopter from the composite data of complex time series radar returns. Received signals of this type typically consist of returns from the rotating main and tail rotor blades, the helicopter body, and other residual components which may overlap in time and frequency, thus making it difficult for conventional time–frequency separation techniques to be applied for the signal separation. In the proposed algorithm, a sparse signal representation is applied with the use of tunable Q wavelet transform to construct the dictionary, and basis pursuit denoising is used for the signal reconstruction of the component of interest. The algorithm shows very effective separation of the main rotor blade component from the composite radar returns, which is demonstrated using both simulated and real radar data at X -band.

This paper considers the Mi-8 helicopter main rotor aerodynamics when flying near the infinite ground surface. The research is based on the free wake model developed by authors at Moscow Aviation Institute. The distance from the rotor's hub to the ground surface in the range of H = 6−16 m and the values of free stream (flight) velocity in the range of V = 0−15 m/s are considered. The results of the visualization for both rotor wake shapes and streamlines are obtained. The influence of the ground proximity on the rotor wake shape, including the formation of "supervortex" and "ground vortex" structures are analyzed. The induced velocity fields of the main rotor in the area of the tail rotor location for various azimuth positions relative to the main rotor axis are studied. The conclusion is made about the significant influence of the ground effect on the rotor wake structure and induced velocities field, including the area of the tail rotor location. Particularly, at slip flight with speed V = 10 m/s, when the distance to the ground surface H increases from 6 to 12 m, the value of the average induced velocity at the tail rotor plane is growing up to four times. The obtained data have allowed taking into account such effects of aerodynamic interference in the simplified mathematical model of the research flight simulator of JSC Helicopters Mil and Kamov.

In order to improve the aerodynamic performance of the ducted tail rotor, parametric investigations and optimization of the interior duct surface are performed based on the Computational Fluid Dynamics (CFD) method. Firstly, to predict the aerodynamic characteristics of the ducted tail rotor with high precision and efficiency, the Reynolds-Averaged Navier-Stokes (RANS) solver coupled with body-fitted mesh around rotor blade is established based on the rotating reference frame. Then, novel parametric researches on the aerodynamic characteristics of the ducted tail rotor are conducted with different lip radii, transition section lengths, and duct outlet areas respectively, and the flow separation mechanism on the different interior duct surfaces are obtained. Finally, a new curve lip shape parameterized by the Class Shape Transformation method is employed to replace the conventional circular lip, besides, the transition length and the outlet area which have great influences on the flow separation phenomenon are combined to conduct the optimization of the interior duct surface. Attribute to the suppression of the flow separation, the required power and the FM of the optimized ducted tail rotor show a maximum decrement of 2.49% and a maximum increment of 2.44% with respect to the baseline SA365N1 ducted tail rotor respectively.

Rotorcraft drivelines typically utilize a multi-segmented metallic system to transmit power between the engine and tail rotor. The typical arrangement of metal driveshaft segments, hanger bearings, and flexible couplers contribute to a significant logistical footprint, maintenance downtime, and life-cycle costs. Thus, an innovative flexible matrix composite driveshaft design alternative is presented in this paper, intended to simultaneously reduce the number of couplers and bearings, as well as, provide high fatigue strain capacity. Through reduction in number of parts, the maintenance cost and time as well as weight of the system are reduced. Composite driveshafts, representing those used in utility helicopters, were designed using an optimization process that considers: (1) damping-induced self-heating, (2) whirling stability, (3) torsional buckling stability, and (4) lamina strength. The paper provides a ballistic comparison study between a baseline carbon/epoxy composite and flexible carbon/polyurethane composite driveshaft segments. One driveshaft of each material was torsionally loaded to failure without ballistic impact. Additionally, two driveshafts were impacted obliquely at zero torque with 7.62 and 12.7 mm armor piercing/incendiary (API) rounds. After impact, the driveshafts were loaded in torsion to failure. Residual torsional strengths were 17–21% and 13% of un-impacted strengths for the 7.62 mm and 12.7 mm rounds, respectively. For the small sample size, flexible driveshafts had a marginally higher residual strength compared to the carbon/epoxy counterpart. Residual torsional stiffness values were 83–86% and 52–59% for the 7.62 mm and 12.7 mm rounds, respectively.

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.

A distributed atmospheric turbulence model is developed for helicopter flight simulation. The impacts of different aerodynamic surfaces on handling qualities with helicopter in turbulence are analyzed. Both helicopter flight dynamic model and turbulence model are validated against the flight test data, which shows that the proposed model improves the accuracy of the simulation of rotor blade element turbulence model on the vertical axis within the interested frequency range for handling-quality analysis. A quantitative index is developed to study the impact of turbulence on helicopter handling qualities by analyzing the effects of different aerodynamic surfaces on helicopter response to turbulence. The results show that the main rotor has the dominant effects on both the heave and roll handling qualities, the airframe has the primary effect on the pitch handling quality, and the tail rotor has an important effect on the yaw handling quality. Furthermore, the impacts of the main rotor in turbulence on the ...

The aim of this paper was to demonstrate the improved functionality and performance of an electromechanical brake for a helicopter main rotor, which to date has been hydraulically actuated using a disc brake and caliper arrangement. Increasingly, designers seek higher performing solutions to traditional problems through the integration of modern actuation and control strategies. This electromechanical device is required to constrain the helicopter tail rotor shaft protruding from the main rotor gearbox to allow safe taxiing and storage of the helicopter. A systematic and rigorous design methodology was used to converge on an effective solution which satisfied a very demanding specification. The design was further detailed and optimized, leading to the development of a prototype at a high technology readiness level that was tested within a bespoke rig, simulating the torque requirements found on a helicopter main rotor using the torque and position control. The design was shown to meet the required holding torque whilst providing additional functionality of continuous holding capability and meeting the challenging volumetric constraints.

Control design of a TRMS system resembling helicopter is very complicated task. Since it is highly non-liner system, which makes the controller design for TRMS system very complex problem. TRMS system is connected with two rotors, in which first one is main rotor and the other is tail rotor. This is based on a beam, supported by a counter balance. Two degrees of freedom (2DOF) PID controller and FOPID controller model of TRMS involving horizontal and vertical plane has been obtained for controller design. 2-DOF PID and FOPID controller both have been discussed in this paper. With the Simulation results obtained for vertical and horizontal plane, factional order PID controller has been found to be better than 2-DOF controller. The results of all the simulation work has been performed on the MATLAB/Simulink environment.

A cyclocopter is a vertical takeoff and landing aircraft that has cycloidal blade systems that consist of several blades rotating about a horizontal axis. The cycloidal blade system allows changes in direction and magnitude of the generated thrust. This paper describes the design and experimental studies of a 110 kg cyclocopter having two cycloidal blade system rotors and one tail propeller rotor powered by a 300-cm3-class rotary engine. The tail rotor generates pitching motion and compensates for the torque generated by the front two rotors. The upward thrust direction of the tail rotor provides 10% of the lifting force. The design thrust of the cycloidal blade system rotor is estimated in analytical and numerical analyses. An accurate analytical aerodynamic model is developed. The analytically estimated results have a trend similar to that of computational fluid dynamics analysis. Both analysis results match well with the results of a ground test. The onboard flight control computer system has two embed...

Helicopter with the advantage of light, flexible and hovering in the air, is widely used. In both military and civilian domain, it gets more and more attention. It becomes the main subject of research in the world. In the power transmission system of helicopter, the tail rotor shaft is an important part. It is one of the key technologies to predict the dynamic characteristics of the shaft accurately for improving the overall performance of the helicopter. Based on the theory of elastic mechanics, rotor dynamics and the finite element analysis, the former main drive shaft is taken as an example. The dynamic characteristics of the transmission shaft is calculated by the multi/few degrees of freedom finite element method. The validity of the finite element method with few degrees of freedom is determined. Then the structure characteristics of the helicopter tail drive shaft are analyzed, and the dynamic characteristics of the helicopter tail rotor shaft are calculated by using the finite element method with few degrees of freedom. The research results provide the basis and guidance for the design of the helicopter tail drive system.

Brownout dust cloud simulations were conducted for rotorcraft undergoing representative landing maneuvers, primarily to better understand the effects of different rotor placement and rotor/airframe configurations. A time-dependent free-vortex wake for the rotors and surface singularity method for the airframe were employed to represent the carrier phase of the flow. A coupling strategy for the free-vortex method was developed to include the effects of rotors operating at different rotational speeds, such as a tail rotor. For the dispersed phase of the flow, particle tracking was used to model the dust cloud based on solutions to a decoupled form of Basset–Boussinesq–Oseen equations appropriate to dilute gas/particle suspensions of low Reynolds number Stokes flow. The predicted flow fields near the ground and resulting dust clouds during the landing maneuvers were analyzed to better understand the nature and severity of their development and to examine differences produced by various rotor and airframe con...

The present article gives an overview of numerical investigations performed during a research project aimed at in-depth understanding of noise generation mechanisms of a shrouded helicopter tail rotor such as Airbus Helicopters Fenestron $$^\mathrm{TM}$$ . Both aerodynamic and aeroacoustic studies are performed based on a full-scale lightweight transport helicopter configuration with detailed geometry of the Fenestron $$^\mathrm{TM}$$ and by neglecting of the main rotor downwash effect. Analysis is carried out by using a hybrid method, combining Unsteady Reynolds-Averaged Navier-Stokes (URANS) simulation with Ffowcs Williams and Hawkings (FW-H) acoustic analogy. Two representative helicopter flight conditions, namely fast forward flight and hovering, are considered. Thereby, it is intended to assess the influence of cross flow in forward flight condition on the Fenestron $$^\mathrm{TM}$$ acoustic property. Effect of turbulence model on the accuracy of noise prediction is also investigated by comparing the Shear Stress Transport (SST) turbulence model with the Scale Adaptive Simulation (SAS). The capability of the applied hybrid methodology is estimated by means of flight test measurements. A good agreement is found between the predicted and measured sound spectra in terms of blade passing frequency (BPF) and its corresponding sidebands for both flight conditions investigated.

The compound helicopter design could potentially satisfy the new emerging requirements placed on the next generation of rotorcraft. The main benefit of the compound helicopter is its ability to reach speeds that significantly surpass those of the conventional helicopter. However, it is possible that the compound helicopter design can provide additional benefits in terms of maneuverability. The paper features a conventional helicopter and a compound helicopter. The conventional helicopter features a standard helicopter design with a main rotor providing the propulsive and lifting forces, while a tail rotor, mounted at the rear of the aircraft, provides the yaw control. The compound helicopter configuration features both lift and thrust compounding. The wing offloads the main rotor at high speeds, and two propellers provide additional axial thrust as well as yaw control. This study investigates the maneuverability of these two helicopter configurations using inverse simulation. The results predict that a compound helicopter configuration is capable of attaining greater load factors than its conventional counterpart, when flying a pullup–pushover maneuver. In terms of the accel–decel maneuver, the compound helicopter configuration is capable of completing the maneuver in a shorter time than the conventional helicopter, but at the expense of greater installed engine power. The addition of thrust compounding to the compound helicopter design reduces the pitch attitude required throughout the acceleration stage of the maneuver.

We describe a new configuration of fixed-pitch miniature robot rotorcraft that combines the energetic efficiency of a helicopter and the mechanical simplicity of a quadrotor. The large power required to hover is proportional to the inverse of the rotor radius; thus, for a given diameter footprint, a single large rotor will energetically outperform several smaller rotors within the same boundary. However, smaller rotors are able to respond more quickly than large rotors, which require complex actuation to provide control. Our “triangular quadrotor” configuration uses a single large rotor for lift and three small rotors for control, gaining the benefits of both. The small rotors are canted slightly to also provide the same service as a conventional helicopter's tail rotor. Momentum theory analysis shows that a triangular quadrotor may provide a 20% reduction in required hover power, compared with a quadrotor of the same mass and footprint. This is particularly valuable for flying robots working indoors where maximum rotor size is constrained. Using conventional quadrotor and a triangular quadrotors constructed to be a similar as possible, we demonstrate that the triangular quadrotor uses 15% less power, without optimization. A power efficiency budget is provided, and the influence of drive system efficiency is explored. We present a dynamic model and demonstrate experimentally that the aircraft can be stabilized in flight with simple PID control.

Insight into the effects of motions and waves on a ship airwake can be used to provide information regarding safe conditions for at sea flight operations and pilot training. We present a study on the effects of waves and motions on a ship's airwake and a helicopter operating above the flight deck using full scale computational fluid dynamics simulations. The ONR Tumblehome Ship (ONRT) geometry is analyzed in wind and regular head waves corresponding to nominal sea states 3 and 6. In order to separate the effects of waves and motions, simulations are conducted for both sea states with combinations of: waves and motions, waves and no motions, no waves with motions, and no motions or waves. A triple velocity decomposition is conducted in order to quantify changes in the airwake due to motions and waves. The aerodynamic loads on a helicopter hovering in the airwake are studied with one-way and two-way coupling approaches. The one-way coupled analysis uses the velocity field data from the full scale airwake simulations along with disk actuator theory to calculate thrust fluctuations for three different rotor sizes operating above the flight deck. In the two-way coupled study a helicopter loosely based on the Sikorsky SH-60 Seahawk hovering above the flight deck of the ONRT is simulated, including dynamic overset grids of the main rotor, tail rotor and fuselage. This approach captures the interaction between the rotor downwash and the ONRT airwake. The study shows that for the lower sea state 3 the motions and waves have little effect on the airwake behavior, exhibiting little change with respect to the baseline case without motions and waves. At sea state 6 the flow field is modified significantly, which also affects the forces on the helicopter. Force fluctuations show main peaks at the wave encounter frequency and the blade passage frequency, as well as higher harmonics. The one-way coupling approach shows that larger rotors effectively filter out small scale fluctuations while smaller diameter rotors are affected by smaller flow structures.

This paper presents and discusses the results of the transient numerical simulation carried out for an isolated main rotor model and a helicopter in vertical descent autorotation by developing the 6-degree of freedom, rigid-body equations of helicopter motion. The Unsteady Reynolds Averaged Navier-Stokes model is used to simulate the unsteady flow field surrounding four-bladed BO-105 isolated main rotor and complete helicopter which includes the fuselage and the tail rotor. The principal autorotation characteristics for the helicopter such as induced velocity, rate of descent, main rotor angular speed, thrust and torque have been calculated. The simulation results show the creation of a spanwise flow on the upper surface of the blade, and moving fluid in the direction of the blade tip in a region downstream and beneath the leading edge vortex. The aerodynamic characteristics on the complete helicopter are similar to the ones attained for the isolated rotor simulation.

The rotating structures on an airborne target cause additional Doppler modulation in the return signal which is known as the micro-Doppler effect. Information regarding the rotating structures can be extracted from this effect. A technique based on time-frequency transform and Inverse Radon Transform (IRT) to extract helicopter main and tail rotor blades parameter for the purpose of target recognition is investigated. Inverse Radon transform is performed on the spectrogram of simulated signals for different frequencies. A mathematical model to simulate the micro-Doppler effects from the helicopter considering both main and tail rotor blades was developed. The proposed algorithm shows that transformed images corresponding to the main and tail blade’s rotation rates are well focused. The focused image is chosen by calculating the concentration measure of the inverse Radon transformed images. Also blade parameters like the number of rotor blades, the blade length and the rotation rate of the helicopter’s main and tail rotor can be obtained from IRT image.

We demonstrate a quasi-distributed sensor for cantilever health inspection measurements using a fiber Bragg grating (FBG) array inscribed in a polymer optical fiber. The FBGs were characterized and calibrated for axial strain, temperature, and relative humidity prior to their mounting on a carbon cantilever beam, the tail rotor of a helicopter. By using the zero-crossing demodulation algorithm, we recovered the time-dependent, wavelength response from each Bragg grating sensor and the vibration response of the beam was extracted. We used the response of the beam to study how the addition of masses at different positions on the beam influences the vibrational behavior and mimics the location of “damage” through the time-dependent results. We show that health inspection measurements are feasible with polymer-based fiber Bragg gratings, offering accurate and rapid detection of damage points on a structural beam.

In the present work, H infinity \(\left( {H_{\infty } } \right)\) controller on the basis of linear matrix inequality is designed to control TRMS which is a nonlinear highly coupled system. A decoupler is appended to nullify the cross-coupling. The \(H_{\infty }\) controller designed for the decoupled system achieves robust performance even in presence of parametric uncertainty. Perturbations in the moment of inertia of main and tail rotor are considered as parametric uncertainties. The performance of the designed controller is inspected through simulation results in MATLAB/Simulink environment. The responses of the main and tail rotors are found to be adequate.

The development of high-performance computing and computational fluid dynamics methods have evolved to the point where it is possible to simulate complete helicopter configurations with good accuracy. Computational fluid dynamics methods have also been applied to problems such as rotor/fuselage and main/tail rotor interactions, performance studies in hover and forward flight, rotor design, and so on. The GOAHEAD project is a good example of a coordinated effort to validate computational fluid dynamics for complex helicopter configurations. Nevertheless, current efforts are limited to steady flight and focus mainly on expanding the edges of the flight envelope. The present work tackles the problem of simulating manoeuvring flight in a computational fluid dynamics environment by integrating a moving grid method and the helicopter flight mechanics solver with computational fluid dynamics. After a discussion of previous works carried out on the subject and a description of the methods used, validation of the computational fluid dynamics for ship airwake flow and rotorcraft flight at low advance ratio are presented. Finally, the results obtained for manoeuvring flight cases are presented and discussed.

The paper proposes a simulation approach to evaluate the power required by a rotorcraft in standard flight missions and in emergency landing maneuvers, and the corresponding fuel consumption, in order to compare the feasibility and potential fuel savings for different hybrid power systems. More in detail, three options are analyzed, namely electrification of the tail rotor, fully hybrid electric propulsion and electric emergency landing. Weight penalty and potential fuel saving for the proposed hybridization schemes are evaluated for an Agusta-Westland A109 twin engine helicopter model. Nonetheless the discussed methods of analysis have general validity for single main rotor helicopter configurations. Two different scenarios are considered in this investigation: current technologies for batteries and motors and improved electrical components, with performance projections as of 2040. According to this analysis, electrification of the tail rotor and parallel hybridization are feasible with available technology, whereas a fully electrical power system for emergency landing could be developed only in the future. Finally, a parallel hybrid electric power system is sized according to the analysis of power request over four different missions. Fuel savings are evaluated for different energy management strategies. According to the results of this investigation, the parallel hybrid electric power system with present-day and future technologies can save fuel up to 5% and 12%, respectively, with an appropriate energy management strategy.

Control of a small type helicopter is an interesting research area in unmanned aerial vehicle development. This study aims to detect a more typical helicopter unequipped with markers as a means by which to resolve the various issues of the prior studies. Accordingly, we propose a method of detecting the helicopter location and pose through using an infrastructure camera to recognize its in-air natural features such as ellipse traced by the rotation of the helicopter's propellers. A single-rotor system helicopter was used as the controlled airframe in our experiments. Here, helicopter location is measured by detecting the main rotor ellipse center and pose is measured following relationship between the main rotor ellipse and the tail rotor ellipse. Following these detection results we confirmed the hovering control possibility of the helicopter through experiments.

An inverse kinematic solution for the main rotor actuating mechanism of a small-scale helicopter that includes a four-point swashplate system and a Bell–Hiller mixer and the tail rotor actuating mechanism are derived using an approach that is suitable for real-time control applications. A closed-form solution is obtained for the inverse kinematics of the swashplate mechanism and for the forward and inverse kinematics of the tail rotor. Then, a computationally efficient solution is obtained for the inverse kinematics of the Bell–Hiller mixer by converting the nonlinear kinematic equations into a generalized eigenvalue problem. The nonlinear kinematic model is compared to a linear approximation and is validated using experiments.