Main Rotor Research Articles

Method for Determining the Number of Propellers in an Aircraft from a Radar Image Obtained by Inverse Synthesis of the Antenna Aperture

The inverse antenna aperture synthesis technology used to construct a radar image of an aircraft’s propellers has shown high efficiency. The radar image allows to visualize the blades of the rotors included in the functional group (traction rotors of an aircraft, rotors of a twin-rotor helicopter, rotors of a multicopter). In the case of a single rotor in an aircraft (an airplane’s tractor rotor, a single-rotor helicopter’s main rotor), the radar image is simple and clearly perceptible. In the case of several propellers belonging to the same functional group, the ana ysis of the radar image becomes significantly more complicated. This is due to the random relative position of the blades of different propellers at the moment the image begins to be constructed, as well as the possible random coincidence of the spatial position of the blades belonging to different propellers. In this regard, determining the number of propellers in an aircraft is a new urgent task, the solution of which allows us to obtain additional information for recognition. The method under consideration for determining the number of propellers is based on the most common design features – the blades in the propeller follow at the same angular interval, in the propellers of the functional group the number of blades is the same.

Parametric structure optimization of a main rotor blade as an application for FSI analysis in main rotor optimization process

Purpose Modern warfare and modern battlefield are very demanding. The recent conflicts showed that the usage of the helicopters is very limited and only the best constructions are able to provide support for the operations. The purpose of this research is to show the possibilities of new design tool for main rotor aerostructural optimization. It is a next chapter of the research that is aimed at finding new solutions for rotorcraft constructions. Design/methodology/approach This work presents a method of preliminary structure optimization of the main rotor blade using parametric modeling. It is the next step in the main rotor optimization studies. It is the next step after preparing the parametric model for the external shape CFD analysis. As a basis for parametric blade structure calculations, the analytical model is provided in this paper. The equations of rigid blade loads and, as a consequence of the strength elements, stresses are shown. The parametric blade modeling is conducted using the Graphic Integrated Programming language. The parametric design method is shown to be used for various blade planform models and different section airfoils. The structure of a blade is generated automatically after the user enters the parameters. The code-inbuilt analysis systems provide a quick inertia examination of the generated geometry, which is the basis for further optimization. The program calculates the blade loads and verifies them with the given material conditions and proposed safety factors. In the analysis, composite materials for the strength elements were proposed. Findings The results of this research showed the application of parametrization into the main rotor blade design loop. It was presented that the main rotor blade structure can be enhanced using fluid-structure interaction (FSI) methods. The time saving with the implementation the process into design loop is shown. Practical implications This work can be practically used in the main rotor blade design process. It provides the possibilities to check various blade aerodynamic configuration in a structure strength aspect. Originality/value To the best of the authors’ knowledge, there were no published research that combines the main rotor FSI analysis. The method, which is presented in the work, provides a new approach to a rotorcraft design. The application of the parametrization and combining it with the FSI method gives a novel solution for helicopters construction enhancement.

Numerical Investigation of an NACA 13112 Morphing Airfoil.

This article presents a numerical study on the 2D aerodynamic characteristics of an airfoil with a morphed camber. The operational regime of the main rotor blade of the IAR 330 PUMA helicopter was encompassed in CFD simulations, performed over an angle of attack range of α=[-3°; 18°], and a Mach number of M=0.38. Various degrees of camber adjustment were smoothly implemented to the trailing-edge section of the NACA13112 airfoil, with a corresponding chord length of c=600 mm at the Reynolds number, Re=5.138×106, and the resulting changes in static lift and drag were calculated. The study examines the critical parameters that affect the configuration of the morphing airfoil, particularly the length of the trailing edge morphing. This analysis demonstrates that increasing the morphed camber near the trailing edge enhances lift capability and indicates that the maximum lift of the airfoil depends on the morphed chord length. The suggested approach demonstrates potential and can be implemented across various categories of aerodynamic structures, such as propeller blade sections, tails, or wings.

Modelling and dynamic analysis of a synchropter

This paper addresses the flight dynamics modeling, trim, and dynamic analysis of an intermeshing-rotor helicopter, indicated as synchropter. This configuration has gained a great interest for its suitability within heavy load lifting and transportation in extreme high temperature and altitude, and other harsh environments. The paper presents some relevant features related to synchropter's flight dynamics modeling of the interference between its two tilted main rotors. Trim results show the advantage of the synchropter in forward flight where the yawing moment is naturally balanced at almost all speeds and no lateral-directional compensation is needed. The synchropter's dynamic stability shows similarity to a conventional helicopter in the longitudinal phugoid. However, in the lateral phugoid, the synchropter is unstable at all flying speeds and therefore its vertical fin needs to be carefully designed.

SIMULATION OF TORQUE VARIATIONS IN A DIESEL ENGINE FOR LIGHT HELICOPTERS USING PI CONTROL ALGORITHMS

This article presents the results of simulation research of a diesel engine for a light helicopter. The simulations were performed using the 1D software AVL Boost RT. The engine model includes elements such as cylinders, turbine, compressor, inlet and outlet valves, ambient environment definition, and fuel injection control strategy. The simulations aimed to evaluate the engine's response to step changes in the main rotor load, both increasing and decreasing power demands. Parameters analyzed included power deviation, torque, engine rotational speed, and stabilization time of the main rotor rotational speed. All tests were conducted using a single set of PI controller settings. The results demonstrate that these parameters are dependent on the magnitude of the step change in the main rotor load demand. The study compares the maximum engine rotational speed deviation from the nominal value for both increasing and decreasing main rotor load demands. The findings indicate that using PI regulator to control rotational speed in the diesel engine in a light helicopter significantly depends on the change in the load torque on the rotor.

Handling Qualities Assessment and Discussion for Helicopter with Slung Load Systems Utilizing Various Sling Configurations

Sling configurations significantly influence the coupled dynamics of the helicopter with slung load system (HSLS), resulting in alterations to handling qualities (HQs) that remain inadequately understood. This study introduces a computer-oriented, generalized method for constructing the HSLS model with various sling configurations. To evaluate the HQs of 1-point, 2-point, and 4-point sling configurations, both the stability and response criteria outlined in ADS-33E and a newly proposed criterion for slung loads towards the updated ADS-33F were employed. Modal analysis was conducted to elucidate the coupled mechanisms of the HSLS under different sling configurations. The findings reveal that the dynamics of the main rotor can attenuate the lateral swing motions of the load in the 4-point sling configuration. While multiple-point sling configurations can enhance the helicopter’s bandwidth, they also amplify the magnitude notch in the helicopter’s response. Nevertheless, when a larger hook distance is employed, the notch frequency is sufficiently distant from the load swing bandwidth, leading to a reduced degradation in HQs. A 4-point configuration with lateral and longitudinal hook distances equal to twice the width and length of the slung load is recommended in practice to achieve sufficient swing stability and mitigate HQ degradation.

Metaheuristic Algorithm-Based Proportional–Integrative–Derivative Control of a Twin Rotor Multi Input Multi Output System

Metaheuristic algorithms are computational techniques based on the collective behavior of swarms and the study of organisms acting in communities. These algorithms involve different types of organisms. Finding controller values for nonlinear systems is a challenging task using classical approaches. Hence, using metaheuristics to find the controller values of a twin rotor multi-input multi-output system (TRMS), one of the nonlinear systems studied in the literature, seems to be more appropriate than using classical methods. In this study, different types of metaheuristic algorithms were used to find the PID controller values for a TRMS, including a genetic algorithm (GA), a dragonfly algorithm, a cuckoo algorithm, a particle swarm optimization (PSO) algorithm, and a coronavirus optimization algorithm (COVIDOA). The obtained graphs were analyzed based on certain criteria for the main rotor and tail rotor angles to reach the reference value in the TRMS. The experimental results show that when the rise and settlement times of the TRMS are compared in terms of performance, the GA took 1.5040 s (seconds) and the COVIDOA took 9.59 s to increase the pitch angle to the reference value, with the GA taking 0.7845 s and the COVIDOA taking 2.4950 s to increase the yaw angle to the reference value. For the settling time, the GA took 11.67 s and the COVIDOA took 28.01 s for the pitch angle, while the GA took 14.97 s and the COVIDOA took 26.69 s for the yaw angle. With these values, the GA and COVIDOA emerge as the foremost algorithms in this context.

Wind-tunnel experimental investigation on rotor-rotor aerodynamic interaction in compound helicopter configuration

In recent years, the compound helicopter configuration, featuring the addition of lateral propellers to a helicopter main rotor, has gained renewed interest for its ability to achieve higher flight speed. The aerodynamic behaviour of compound rotorcraft is dominated by mutual interactions between the rotors and the wakes they generate, which can affect their performance and the handling qualities of the aircraft. In order to study the effects of the rotor-wake and wake-wake interactions, a test rig was developed and an extensive wind-tunnel test campaign was carried out measuring the performance of rotor and propellers under different flight conditions. Hovering flight and forward flight at different advance ratios were considered, as well as crosswind flight under various wind directions. In general, an increase in thrust for both main rotor and propellers is found due to the interaction. In particular, the significant increase in the propeller thrust is attributed to the influence of the main rotor downwash, which impacts edgewise on the propellers. In some specific conditions, a decrease in the propeller's thrust is observed, which might be related to blade-vortex interaction (BVI) effects. To aid in the interpretation of the experimental results, numerical simulations with a mid-fidelity aerodynamic code were also performed.

Optimal control for helicopter/turboshaft engine system based on hybrid variable speed

In order to address the limitation of fixed-ratio transmission (FRT), which compromises the attainment of both optimal main rotor speed and optimal power turbine speed, an optimal speed control method based on hybrid variable speed (HVS) is proposed. Firstly, based on the integrated performance calculation model of helicopter/turboshaft engine system, the distribution factors of variable speed are applied, and the integrated optimization method of optimal speed is proposed based on the minimum engine fuel flow. Subsequently, an online estimation technique employing a high-order filter is devised and engineered to achieve superior cascaded control of turboshaft engines. Finally, a novel real-time optimal speed control method based on hybrid variable speed is proposed. The simulation results under different operation conditions demonstrate that regardless of whether it is FRT or HVS, the optimal main rotor speed increases with forward velocity. In the case of HVS, turboshaft engine degradations have a significant impact on the optimal power turbine speed rather than optimal main rotor speed. Adopting an estimation method based on high-order filtering for gas turbine rotational acceleration proves more advantageous in mitigating high-frequency oscillation and continuous saltation of estimated values. Moreover, in comparison with the optimal speed control method of FRT, HVS-based approach enables simultaneous attainment of the optimal main rotor speed and power turbine speed, thereby enhancing the overall efficiency of the integrated helicopter/turboshaft engine system and significantly decreasing engine fuel consumption by over 2%. Consequently, there has been a remarkable enhancement in the overall performance of the integrated helicopter/turboshaft engine system.

Control of Helicopter Using Virtual Swashplate

This article presents a virtual swashplate mechanism for a mini helicopter in classic configuration. The propeller bases are part of a passive mechanism driven by main rotor torque modulaton, this mechanism generates a synchronous and opposite change in the propellers angle of attack, then the thrust vector tilts. This approach proposes to control the 6 degrees of freedom of the aircraft using two rotors. The main rotor controls vertical displacement and uses torque modulation and swing-hinged propellers to generate pitch and roll moments and the horizontal displacement while the yaw moment is controlled by the tail rotor. The dynamic model is obtained using the Newton-Euler approach and robust control algorithms are proposed. Experimental results are presented to show the performance of the proposed virtual swashplate in real-time outdoor hover flights.

Aeroelastic fluid dynamics assessment of performance and vibration on active twisting rotors

The effects of passive, and higher-harmonic rotor blade twist on helicopter main rotors are investigated. High-fidelity CFD is used for calculations due to the demanding flow conditions considered, including high blade loading, and high-speed cases. A strong structural coupling scheme is employed for enhanced performance and vibration predictions. Quantitative predictions are made for the effects of blade twist in hover, and the results showed: figure of merit improvements of up to 1 count per degree of twist, and diminishing returns with increasing twist. In high-speed level flight, the higher twist blade showed 18% increased vibration. A 2/rev active twist input shows vibration reductions of 22% in high-speed and 8% at high load cases. The active twist achieved a marginally improved aerodynamic efficiency at high-speed, and reduced the aerodynamic rolling moment in the dynamic stall conditions.

Helicopter rotor in a propeller slipstream: Aerodynamic and rotor blade flapping responses

During helicopter air-to-air refueling the helicopter's rotor can engage the tanker aircraft's wing and flap end tip vortices and the slipstream of its propellers. The tip vortices represent a swirl around an axis that is essentially parallel to the rotor longitudinal axis. The slipstream of the propellers includes a strong air jet downstream that can cover a significant amount of the helicopter's main rotor. This phenomenon was recently treated analytically for a rigid rotor by means of blade-element momentum theory and the rotor control angles required to reject the disturbance were computed. In this article, the rotor blade flapping is introduced as degree of freedom. It is shown that the rotor coning is affected by the slipstream position with respect to the rotor hub. Without retrim, the largest amount of coning and cyclic flapping occurs for slipstream positions on the retreating side of the rotor.

Helicopter vibration control employing superelastic pitch link with shape memory alloy

One of the main applications of shape memory alloy (SMA) concerns vibration control, especially superelastic SMA (SMA-SE), which presents a significant stress hysteresis and can work as a damper while adding minimal weight to the structure. In helicopters, vibration is a phenomenon inherent to their operation, and although cannot be eliminated, must be minimized to acceptable levels of safety and comfort. In this context, this paper aims to present the design, testing, and dynamic behavior of a new device entitled smart pitch link, a device with superelastic material for passive vibration control. The paper depicts the design process, assembly, and device testing using a whirl tower that simulates a helicopter rotor in hover. The approach adopts a comparative analysis of the dynamic response on the prototype in reference condition, with a stiff pitch link, with the modified prototype, and with the superelastic pitch link installation. In addition, beyond hover tests, cyclic commands and an asymmetrical lateral gust were employed to observe the output response of the device under transient loadings. Contrary to the literature data about the low time response of SMA, the device presents satisfactory time response and attenuation in frequencies close to main rotor frequencies, with signal attenuations between 2 to 4 dB without any external energy consumption, which, in a comparative analysis, may exceed more than 50% of vibrational amplitude attenuation.

Wind Tunnel Experiments on Parallel Blade–Vortex Interaction with Static and Oscillating Airfoil

This study aims to experimentally investigate the effects of parallel blade–vortex interaction (BVI) on the aerodynamic performances of an airfoil, in particular as a possible cause of blade stall, since similar effects have been observed in literature in the case of perpendicular BVI. A wind tunnel test campaign was conducted reproducing parallel BVI on a NACA 23012 blade model at a Reynolds number of 300,000. The vortex was generated by impulsively pitching a second airfoil model, placed upstream. Measurements of the aerodynamic loads acting on the blade were performed by means of unsteady Kulite pressure transducers, while particle image velocimetry (PIV) techniques were employed to study the flow field over the blade model. After a first phase of vortex characterisation, different test cases were investigated with the blade model both kept fixed at different incidences and oscillating sinusoidally in pitch, with the latter case, a novelty in available research on parallel BVI, representing the pitching motion of a helicopter main rotor blade. The results show that parallel BVI produces a thickening of the boundary layer and can induce local flow separation at incidences close to the stall condition of the airfoil. The aerodynamic loads, both lift and drag, suffer important impulsive variations, in agreement with literature on BVI, the effects of which are extended in time. In the case of the oscillating airfoil, BVI introduces hysteresis cycles in the loads, which are generally reduced. In conclusion, parallel BVI can have a detrimental impact on the aerodynamic performances of the blade and even cause flow separation, which, while not being as catastrophic as in the case of dynamic stall, has relatively long-lasting effects.

Computational investigations of Ka-226T helicopter coaxial rotor aerodynamics in the vortex ring state

The vortex ring states are observed when the rotor is flowed at positive angles of attack. For the main rotor, these conditions are realized during a steep descent of the helicopter at low speeds. The rotor vortex ring states are affected by significant phenomena related to the behavior of its aerodynamics, including negative phenomena. The latter is, firstly, referred to decrease in rotor thrust, increase in the required power, pulsations of thrust and torque, unsteady flapping blade motion, etc. In terms of helicopter piloting, it means a sharp loss of altitude, increase in the control force, a high level of vibration, defocusing attenuation of rotor spinning cone, as well as controllability deterioration. All these factors determine the relevance of research on these modes and the importance of solving a problem of defining their boundaries. Recently, due to the rapid development of computer technologies and computational models, it has become practical to perform numerical research of the rotor aerodynamics in vortex ring states. The paper presents the study results of Ka-226T helicopter coaxial rotor aerodynamics in steep descent modes in the field of vortex ring states. The angles of rotor attack αВ = 90...30° and the range of vertical descent velocities Vу = 0...26 m/s are considered. The original nonlinear bladed vortex model of the rotor developed at the Moscow Aviation Institute (MAI) was used. The total and distributed aerodynamic rotor characteristics were calculated. The shapes of the vortex wake and the rotor flow patterns were analyzed. The boundaries of the vortex ring states in velocity coordinates “Vx − Vу” were constructed according to various criteria reflecting the known features of these states. The results obtained significantly complement the existing experience of experimental and numerical research in this field.

Performance Improvement of Winged Compound Helicopter Due to Differential Flap Deflections

A lift-offset system for a single-rotor-type compound helicopter is proposed to improve the aerodynamic performance in high-speed flight. The proposed system utilizes the differential flap deflections on the fixed wings to produce a rolling moment, which is counteracted by the single main rotor, causing the rotor to operate in a lift-offset state. The performance of the proposed system is evaluated through numerical simulations. At first, the effect of lift offset with regard to lift-share ratio on the rotor performance is investigated. Then, the impact of lift offset due to the differential flaps on the overall effective lift-to-drag ratio is studied. The results show that the lift offset significantly improves the rotor performance and the overall effective lift-to-drag ratios, especially at larger rotor lift-share ratios. The overall effective lift-to-drag ratio increases by 10% due to the differential flaps compared to the zero-flap deflections. It is concluded that the lift offset due to the differential flaps achieves more efficient cruising flight for a single-rotor-type compound helicopter.

Autonomous flight performance maximization for slung load carrying rotary wing mini unmanned aerial vehicle

PurposeThis research study aims to minimize autonomous flight cost and maximize autonomous flight performance of a slung load carrying rotary wing mini unmanned aerial vehicle (i.e. UAV) by stochastically optimizing autonomous flight control system (AFCS) parameters. For minimizing autonomous flight cost and maximizing autonomous flight performance, a stochastic design approach is benefitted over certain parameters (i.e. gains of longitudinal PID controller of a hierarchical autopilot system) meanwhile lower and upper constraints exist on these design parameters.Design/methodology/approachA rotary wing mini UAV is produced in drone Laboratory of Iskenderun Technical University. This rotary wing UAV has three blades main rotor, fuselage, landing gear and tail rotor. It is also able to carry slung loads. AFCS variables (i.e. gains of longitudinal PID controller of hierarchical autopilot system) are stochastically optimized to minimize autonomous flight cost capturing rise time, settling time and overshoot during longitudinal flight and to maximize autonomous flight performance. Found outcomes are applied during composing rotary wing mini UAV autonomous flight simulations.FindingsBy using stochastic optimization of AFCS for rotary wing mini UAVs carrying slung loads over previously mentioned gains longitudinal PID controller when there are lower and upper constraints on these variables, a high autonomous performance having rotary wing mini UAV is obtained.Research limitations/implicationsApproval of Directorate General of Civil Aviation in Republic of Türkiye is essential for real-time rotary wing mini UAV autonomous flights.Practical implicationsStochastic optimization of AFCS for rotary wing mini UAVs carrying slung loads is properly valuable for recovering autonomous flight performance cost of any rotary wing mini UAV.Originality/valueEstablishing a novel procedure for improving autonomous flight performance cost of a rotary wing mini UAV carrying slung loads and introducing a new process performing stochastic optimization of AFCS for rotary wing mini UAVs carrying slung loads meanwhile there exists upper and lower bounds on design variables.

Numerical investigation on the use of various blade tips for the helicopter's main rotor blade in forward flight regarding aerodynamic performance and HSI noise considerations

The goal of this research is to investigate aerodynamic performance and high-speed impulsive (HSI) noise of various blade tips for the helicopter's main rotor in forward flight. For this purpose, three-dimensional compressible flow field around the rotor blades is numerically simulated by the solution of the unsteady Reynolds averaged Navier-Stokes (URANS) equations with the "SST k-ω" turbulence model. Solution is quantified by the calculation of torque, drag, and thrust coefficients of CQ, CD, CT, and L/D ratio. HSI noise on the rotor plane is assessed by the calculation of sound pressure level at certain points of the flow field using the Ffowcs Williams-Hawking (FW-H) equation. Blade tips are eleven and include anhedral, dihedral, Eagle and combinations of them, and also Blue Edge and rectangular. The combinations are inspired by the tip configurations of the fixed wings. In this regard effects of the tip geometry parameters including curvature and height of its parts have been considered as well. All of the tips are examined for two flight conditions of Caradonna and Tung two-blade rotor with fixed pitch angle and time-varying pitch angles. Based on the present results, it is found that blade tips of dihedral family do not improve the aerodynamic performance of the helicopter blade. Also, addition of anhedral part to the blade tip improves the L/D ratio. According to the present acoustic analysis, it was seen that after the Blue Edge, the Eagle tip stands in order of priority. For the condition of Caradonna and Tung two-blade rotor with tip Mach number of 0.8 and advance ratio of 0.2, HSI noise of the Eagle-tip blade decreases by 2.39 % compared to that of the two-blade rotor with rectangular tip. However, the aerodynamics performance of the Eagle tip is much better than that of the Blue Edge.

Flight dynamic modeling and stability of a small-scale side-by-side helicopter for Urban Air Mobility

This paper aims to explore the development of a flight dynamics model for a small-scale side-by-side helicopter and describe its trim and stability characteristics. The helicopter is considered a suitable candidate for Urban Air Mobility (UAM) solutions, because of its reliable design and low noise characteristics, but still very small knowledge is present on the mathematical modeling approaches and dynamic properties. A 14 degrees of freedom nonlinear mathematical model is developed and semi-analytical models are employed to account for the presence of the shrouds. An iterative trim routine is developed and applied with a suitable control mix that allows the use of classic helicopter controls. To control the vertical speed and roll rate, the paper assumes an equal collective pitch and lateral cyclic in the two rotors, while a uniform plus a differential longitudinal cyclic is adopted for pitch and yaw maneuvers. The paper discusses unique characteristics of the side-by-side configuration as obtained from the stability analysis: an unstable high-frequency mode, governed by the vertical velocity and pitch angle arises when the center of gravity (CG) of the vehicle is aligned or placed in front of the two main rotors. Similarly, both lateral phugoid and roll subsidence modes are sensitive to the CG location. The side-by-side configuration presents also a stable spiral mode which needs to be carefully designed.

Development and flight-test verification of two-dimensional rotational low-airspeed sensor for small helicopters

This paper describes the development and the verification of flight test results of a differential pressure-based, two-dimensional low-airspeed sensor designed for the navigation or disturbance detection in small helicopters. The compact and lightweight sensor is integrated with the main rotor of a small helicopter and comprises two probes at both arm ends, a differential pressure sensor, rotary encoder with one magnet and two sensors, microcomputer, a wireless data link, and battery. It measures the differential pressure between the total pressures captured by two total-pressure probes at each rotor angle, instead of using static pressure probes. Thus, the airspeed of the fuselage can be evaluated from the low speed. Flight tests were conducted employing a reference ultrasonic two-dimensional airspeed sensor for comparison. The results demonstrated that the magnitude error of the airspeed is less than 2 m/s for low-airspeed flights (<∼\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$<\\sim$$\\end{document}23 m/s) when utilizing Pitot-type probes. The error in wind angle approximated 30∘\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$^\\circ$$\\end{document}, and the delay was less than or equal to that observed with a global navigation satellite system sensor.