Hover Flight Conditions Research Articles

System identification of flybar-less rotorcraft UAV

Purpose This paper aims to build an accurate mathematical model which is necessary for control design and attitude estimation of a miniature unmanned rotorcraft and its subsequent conversion to an autonomous vehicle. Design/methodology/approach Frequency-domain system identification of a small-size flybar-less remote controlled helicopter is carried out based on the input–output data collected from flight tests of the instrumented vehicle. A complete six degrees of freedom quasi-steady dynamic model is derived for hover and cruise flight conditions. Findings The veracity of the developed model is ascertained by comparing the predicted model responses to the actual responses from flight experiments and from statistical measures. Dynamic stability analysis of the vehicle is carried out using eigenvalues and eigenvectors. The identified model represents the vehicle dynamics very well in the frequency range of interest. Research limitations/implications The model needs to be augmented with additional terms to represent the high-frequency dynamics of the vehicle. Practical implications Control algorithms developed using the first principles model can be easily reconfigured using the identified model, because the model structure is not altered during identification. Originality/value This paper gives a practical solution for model identification and stability analysis of a small-scale flybar-less helicopter. The estimated model can be easily used in developing control algorithms.

Finite State Inflow Flow Model for Coaxial Rotor Configuration

An analytical coaxial rotor inflow model has been developed from potential flow theory using the pressure potential superposition approach. The coaxial rotor pressure potential superposition inflow model (PPSIM) is formulated in statespace form with structure similar to the Peters–He model, except that additional off-diagonal blocks are included in the apparent mass (M-matrix) and influence coefficient matrices (L-matrix). These off-diagonal blocks take into account mutual interference effects present in a coaxial rotor system by relating the rotor's inflows due to other rotor's pressure loadings. Induced inflow distributions on both upper and lower rotors are computed using PPSIM for comparison against predictions from high-fidelity models such as GT-Hybrid and the viscous vortex particle method (VVPM). Good agreement between PPSIM-induced inflow results and GT-hybrid as well as VVPM data has been shown for hover flight condition. At low advance ratio, there are differences in fore-to-aft inflow states between PPSIM and the high-fidelity models. This is because PPSIM assumed rigid, skewed cylindrical wake geometries for both upper and lower rotors during forward flight. But in GT-Hybrid and VVPM, wake structures are allowed to move freely in space and are mainly affected by rotor-induced velocities at low advance ratios. Owing to the close proximity between upper and lower rotors, mutual interference-induced velocities significantly distorted the rotors' wake geometries. The rigid rotor wake geometry assumptions in PPSIM and the distortion captured in higher fidelity models are the reasons behind differences in rotor-induced inflows. At higher advance ratios, wake distortion effects are less prominent since free-stream inflows are significantly larger than rotorinduced velocities. Hence, smaller differences between PPSIM inflow states and those extracted from GT-Hybrid as well as VVPM are observed at high advance ratios.

Rotor interactional effects on aerodynamic and noise characteristics of a small multirotor unmanned aerial vehicle

Small scale unmanned aerial vehicles using multirotor propulsion systems have received considerable attention for a wide range of military and commercial applications in recent years. In the multirotor configuration, the rotor interaction phenomenon occurs severely because the rotors are located in close proximity to each other. Therefore, the separation distance between the adjacent rotor tips has a strong effect on the wake structures and flow fields, which consequently play an important role in determining the aerodynamic performance and noise level of the multirotor vehicle. In the present study, numerical simulations of a quadcopter under hover flight conditions are conducted to investigate the mutual rotor-to-rotor interactional effects on the aerodynamic performance, wake structures, and sound pressure level using the nonlinear vortex lattice method with the vortex particle method and acoustic analogy based on Farassat’s formulation 1A. Calculations for the multirotor configurations with different separation distances show that the average thrust force decreases significantly and force fluctuation is found to increase dramatically as the rotor spacing gets smaller. In addition, the wake geometry and induced flow structure behind the rotor tend to be radially dragged down toward the center of the vehicle due to the existence of the adjoining rotor, which consequently results in strong wake-to-wake interaction and the formation of asymmetric wake structures although the multirotor operates under the hovering condition. It is also observed that unsteady loading introduced by rotor interaction leads to a considerable increase in the sound pressure level, particularly the normal direction of the rotor plane.

Computational Analysis of Rotor-Blown-Wing for Electric Rotorcraft Applications

This Paper examines the performance of a rotor-blown-wing in hover and forward flight conditions. Flow solutions were generated using Helios, with a delayed detached eddy simulation model. The configuration is based on one set of rotors/wing of a quadrotor biplane tail-sitter aircraft. Simulations were conducted to assess the effect of varying angle of attack, rotor revolutions per minute with rotor speed, and rotor diameter. Two configurations were examined, the rotor-blown-wing baseline with two 60.96 cm diameter rotors and the configuration with four 30.48 cm diameter rotors. The smaller rotors were analyzed at the revolutions per minute corresponding to producing half the thrust of the rotor-blown-wing baseline rotor. The presence of the wing in hover resulted in the rotor-blown-wing baseline rotor operating at 1.9% lower power loading than the same rotor in isolation. In comparison, the configuration reduced power loading by 13.6% compared to the isolated rotor. In forward flight (airplane mode), an increasing angle of attack was shown to increase thrust produced on the downstroke side of the rotor. Furthermore, the presence of the wing increases the thrust produced on the bottom half of the disk. Both the blown wings increase the ratio from 3.2 for the isolated wing to 3.6 and 4.3 for the rotor-blown-wing baseline and configurations, respectively.

Aerodynamic characteristics analysis of electrically controlled rotor based on viscous vortex particle method

An electrically controlled rotor (ECR), also called as swashplateless rotor, replaces a swashplate with a trailing edge flap system to implement primary rotor control. To investigate the aerodynamic characteristics of the ECR, an analysis model based on the viscous vortex particle method, ECR blade pitch equation and Weissinger-L lifting surface model is established. In this model, the ECR wake flow field vorticity is discretized as multiple vortex particles, and the vorticity-velocity form Navier-Stokes equation is solved to simulate the transport diffusion of the vorticity. The impact of flap deflection on the blade bound vortex circulation is calculated via the equivalent incidence angle. The flap deflection induced blade pitch movement is obtained by solving the ECR blade pitch movement equation via the Runge-Kutta fourth-order method. On this basis, to address the low computing efficiency of the viscous vortex wake model, a “relaxation difference”-based trim strategy is developed to reduce the calculation times of the viscous vortex wake model during trim iteration. Based on this model, the aerodynamic characteristics of a sample ECR in hovering and forward flight conditions are analysed. The results show that in hovering and forward flight conditions, when the rotor trim requires large flap deflection, in the rotor wake vorticity field, apart from blade tip vortex, there are intense vortices at both ends of the flap. This additional intense vortex results in an ECR wake vorticity distribution that is more complex than that of a conventional rotor, significantly impacting the distribution of the rotor disc inflow and load. Additionally, in the forward flight condition, the ECR flap collective deflection results in cyclic blade pitch movement, while flap cyclic deflection results in a significant portion of the 2Ω harmonic component in the blade pitch movement when the advance ratio is high.

Entire Flow Field Modeling Including Wake Region through Improved Finite-State Method

In the recent decades, the accuracy and robust of finite state inflow model to provide the velocity on the disk have been greatly improved. It is adequate to compute the thrust, pitch and roll moment real time for normal flight conditions. For lifting rotor flying close to a non-penetrable surface, the wake will interact with the inflow of the lifting rotor. Additionally, for co-axial rotors the wake of the upper rotor influences the lower rotor. Therefore, it is necessary to develop a model to compute the flow field in the wake since the Peters-He model cannot give the induced velocity below the rotor disk. In this paper, the improved finite state model will be introduced to compute the axial induced velocity everywhere in the flow field including wake and downstreamregions. The adjoint theorem will be introduced, and the computational efficiency of the convolution method and numerical algorithm for computing the adjoint varibales with time delay will be compared. With the parameters of a Harrington rotor, the results in the time domain for such a rotor in hover and forward flight conditions will also be illustrated.

음향상사법을 이용한 멀티콥터형 소형무인기의 공력 소음 예측

The rotor interaction phenomenon occurs frequently in unmanned aerial vehicles (UAVs) with multi-rotor configuration. In this study, we investigated the rotor interaction effects of a small-size quadcopter UAV on the aerodynamic performance and acoustic characteristics. The numerical simulation of DJI Phantom 2 model in the hover flight condition was conducted using the nonlinear vortex lattice method (NVLM) with the vortex particle method (VPM) and acoustic analogy based on the Farassat-1A formula. NVLM can consider the nonlinear aerodynamic characteristics that are mainly associated with viscous and low Reynolds number effects by incorporating airfoil look-up table and vortex strength correction. VPM is well suitable for simulating complex wake structures of quadcopter UAV because maintaining the connectivity between adjacent vortex particles is not required. Calculations for an isolated rotor were compared against measurements to validate the predicted results, including thrust force and tonal noise at 1st blade passing frequency (BPF). The calculations for a quadcopter configuration show that the average thrust coefficient decreases and thrust coefficient begins to fluctuate dramatically due to severe rotor interaction. In addition, the sound pressure level of the quadcopter is much higher than that of the isolated rotor and a significant difference in noise directivity between the quadcopter and isolated rotor is clearly captured. It is observed that unsteady loading introduced by rotor interaction leads to a considerable increase in noise. This study is a preliminary study to the aeroacoustic problem in multicopter UAV. The discussion in this work gives scope for further research on noise reductions of multicopter UAV by controlling a separation distance between the rotors.

Systematic design methodology for development and flight testing of a variable pitch quadrotor biplane VTOL UAV for payload delivery

This paper discusses the conceptual design and proof-of-concept flight demonstration of a novel variable pitch quadrotor biplane Unmanned Aerial Vehicle concept for payload delivery. The proposed design combines vertical take-off and landing (VTOL), precise hover capabilities of a quadrotor helicopter and high range, endurance and high forward cruise speed characteristics of a fixed wing aircraft. The proposed methodology is described by taking an example of a VTOL Unmanned Aerial Vehicle (UAV) for a mission requirement of carrying and delivering 6 kg payload to a destination at 16 km from the point of origin. First, the design of proprotors is carried out using a physics based modified Blade Element Momentum Theory (BEMT) analysis, which is validated using experimental data generated for the purpose. Proprotors have conflicting requirement for optimal hover and forward flight performance and a compromise solution has to be chosen. Next, the biplane wings are designed using simple lifting line theory. The airframe design is followed by power plant selection and transmission design. Finally, weight estimation is carried out to complete the design process. The proprotor design with 24° preset angle and −24∘ twist is found to be suitable for the current mission profile, based on 70% weightage to forward flight and 30% weightage to hovering flight conditions. A reduction in operating RPM of the proprotors from 3200 during hover to 2000 during forward flight results in 27% increase in proprotor efficiency. The VTOL UAV design proposed in this paper is also predicted to consume 64% less power in airplane mode compared to the quadrotor mode, thereby establishing the benefit of this hybrid concept. A proof-of-concept scaled prototype is fabricated using commercial-off-the-shelf parts and a Proportional Integral Derivative (PID) controller is developed and implemented on open source PixHawk autopilot board to enable stable hovering flight and attitude tracking during hover flight tests.

Investigation of the Rotor–Obstacle Aerodynamic Interaction in Hovering Flight

In this paper, a comprehensive experimental survey of the aerodynamic interaction of a hovering rotor in the proximity of a ground obstacle is described, taking advantage of multiple experimental techniques. Load measurements on the rotor were carried out to assess the change in the rotor performance for different positions with respect to the cubic obstacle, thus simulating a set of possible hovering flight conditions around the obstacle. Laser Doppler anemometry measurements of the rotor inflow were used to investigate how the aerodynamic interaction affected the rotor performance. Stereoscopic particle image velocimetry measurements in the region between the rotor and the obstacle were carried out to gain a better insight of the interacting flow field. The investigation showed two main regions of interest. The first region is the one above the edge of the obstacle, where the rotor experiences a gradual ground effect as it is positioned over the obstacle. The second region, probably of more interest, is the one just beside the obstacle where a recirculation region between the rotor and the obstacle develops, causing both a significant reduction in the thrust augmentation experienced in unobstructed hover in ground effect and significant pitching and rolling moments, due to the nonsymmetrical inflow pattern on the rotor.

UH-1H 로터 블레이드의 제자리 비행 시 투과면을 이용한 원방 소음 해석

본 연구에서는 투과면을 이용하는 음향상사법으로 제자리 비행하는 UH-1H 로터 블레이드 주위의 원방 소음을 예측하였다. 두께 소음과 하중 소음, 그리고 충격파 및 끝단 후류 등에 의해 발생하는 유동 소음을 예측하기 위해 블레이드 표면을 포함하는 투과면을 구성하였다. 3차원 압축성 Euler 방정식 및 Navier-Stokes 방정식을 적용하여 공력 해석을 수행하고 비교하였다. 투과면의 위치에 따라 High Speed Impulsive 소음을 예측 및 검증하였다. 블레이드 끝단에서 발생하는 충격파에 의한 소음원이 지배적인 요소임을 확인하였으며, 충격파를 온전히 포함하도록 투과면을 구성하는 것이 중요함을 보였다.

Modelling of Hovering Helicopter and its Stability analysis using Participation factor

A helicopter during its autonomous missions is considered to be in hover flight condition for very long extent compared to any other flight conditions and they have to hover before any complex maneuvering mode. Hence the analysis of stability and design of a controller for this flight condition is necessary to facilitate the autonomous flight operation. This paper covers the stability analysis of a hovering helicopter and evaluates the influence of each state variable on different flight modes particularly on unstable ones using participation factor. It can be utilized further for stability augmentation purpose as well as for controller design to meet other performance measures.

System identification and robust stabilization using structured controller for a novel ducted fan flying robot

This paper proposes complete procedures for system identification and control of a novel ducted fan flying robot on near hover flight condition. A control-oriented model structure is derived from a comprehensive mechanism model and identified by ground tests and flight experiments. In order to achieve state decoupling and reference tracking, a two-loop control architecture is employed. Non-smooth optimization algorithm is applied to efficiently tune controller parameters against multiple control requirements. Experimental results show that designed structured controller can provide better performance than classical proportional -integral -derivative (PID) controller.

A New Robust Controller Formulation for the Full–State Feedback Position Tracking of a Small–Scaled Unmanned Model Helicopter

This work focuses on the robust attitude tracking control problem for a small– scaled unmanned helicopter where the actual system inputs, namely the elevator servo input, the aileron servo input and the rudder servo input, are used in the controller formulation. The design process is divided into two parts. Initially the problem is transformed into a second order system with an uncertain non–symmetric input gain matrix by utilizing some reasonable simplifications for the rotor model under the hovering flight conditions. Then a novel robust control methodology is utilized via a matrix decomposition method. The stability of the overall system is ensured by Lyapunov type analysis where asymptotic position tracking is ensured. Numerical simulation results are presented to demonstrate the efficiency of the proposed method.

Computational Analysis and Flow Physics of a Ducted Rotor in Edgewise Flight

This study examines the performance of the ducted rotor in hover and edgewise flight conditions. The flow over a threedimensional model of a ducted rotor was simulated using the Spalart–Allmaras Reynolds-averaged Navier–Stokes model implemented in a stabilized finite element method. A sliding mesh was used to conveniently account for the large-scale motion associated with rotor revolutions. The simulation results were analyzed to understand the flow physics and quantify the contributions of the rotor and various sections of the duct interior surfaces on the total aerodynamic forces (thrust, drag, and side force) and moments (pitching and rolling). In edgewise flight, freestream flow separates off the front of the duct inlet, causing a region of recirculating flow and upwash in the rotor plane. The upwash region biases rotor thrust production to the front of the disk. The swirl velocity further biases the region of flow separation over the inlet and upwash at the front of the rotor toward the retreating side of the disk. The shift of thrust production on the rotor and duct toward the front produces a strong nose-up pitching moment on the ducted rotor. The rear of the diffuser is a significant contributor to the total drag; this force includes a nose-down pitch moment, which partially negates the moment from the duct inlet. The rotor is the primary source of vertical vibratory forces as well as vibratory pitching and rolling moments. The small tip clearance of the rotor causes a local interaction between the blade tip and duct that is the dominant contributor to in-plane vibratory forces on the ducted rotor.

System Identification of Quadrotor Aircrafts Based on an Improved PSO

This paper presents a novel PSO designed to accurately model quadrotor dynamic system. Initially,this paper eatablished the quadrotor model under hover or low speed small angle flight conditions. Then, this paper employed an improved PSO for quadrotor system identification to obtain system parameters. Finally, the proposed system produced vertical acceleration and the attitude angle acceleration,this paper compared these data with the real flight data. The results show that the proposed method is effective and reliable.

Validation on Conceptual Design and Performance Analyses for Compound Rotorcrafts Considering Lift-offset

This work conducts a validation study for the XH-59A helicopter using a rigid coaxial rotor system in order to establish the techniques of the conceptual design and performance analysis for the lift-offset compound rotorcraft. As a tool for conceptual design and performance analysis, NDARC (NASA Design and Analysis of Rotorcraft) is used for the present study. An assumed mission profile is considered for the conceptual design of the XH-59A. As a validation result of the design, the dimensions and weight of the XH-59A are appropriately designed when compared to the target values since the relative error is less than 0.5%. Then, performance analyses are conducted for the designed XH-59A model with and without auxiliary propulsion in hover and forward flight conditions. The present analyses show good validity since the prediction results compare well with both the flight test and previous analyses. Therefore, the techniques for the conceptual design and performance analysis of the lift-offset compound helicopter are overall considered to be appropriately established. In addition, this study investigates the influence of the lift-offset on the rotor effective lift-to-drag ratio of the XH-59A helicopter with auxiliary propulsion. As a result, the improvement of the rotor effective lift-to-drag ratio can be obtained by appropriately increasing the lift-offset in high-speed flight.

Numerical investigation of the vortex roll-up from a helicopter blade tip using a novel fixed-wing adaptation method

This contribution relates to the simulation of the flow around the tip of a helicopter rotor blade in hovering flight conditions. We here propose a new methodology of framework adaptation, using a comprehensive rotor code and high-fidelity numerical simulations. We construct an equivalent fixed-wing configuration from a rotating blade, in which centrifugal and Coriolis forces are neglected. The effect of this approximation on the solution is analyzed. The method is validated by a detailed comparison with wind tunnel data from the literature, concerning aerodynamic properties and tip vortex roll-up. This validation also includes variations of the pitch angle and rotational speed, up to transonic tip velocities. Compared to previously published methods of framework adaptation, the new hybrid method is found to reproduce more accurately the flow around a rotating-blade tip.

Dynamic Mesh Analyses of Helicopter Rotor–Fuselage Flow Interaction in Forward Flight

AbstractUnsteady compressible flow analyses of helicopter rotor–fuselage interaction in hover and forward flight conditions are carried out using commercially available computational fluid dynamics (CFD) solver FLUENT. The individual effects of each component on the flow are investigated by simulating the isolated fuselage and the isolated rotor configurations. Then, the rotor-fuselage interaction problem is analyzed. Azimuthal variations of the flap and pitch motions of the blades are prescribed a priori as a first-order Fourier series through a user-defined function feature of the code. The prescribed blade motion may result in meshes with undesirable grid qualities, which may lead to unphysical solutions. A non-overset dynamic mesh motion method that applies volume mesh deformation and cell remeshing within a priori organized block mesh structure is used to accommodate the rigid blade motion. The remeshing is performed when the grid deformation is more than a predefined skewness value of 0.95. The near...

Robust Flight Control Using Nonsmooth Optimization for a Tandem Ducted Fan Vehicle

AbstractThis work addresses the aerodynamic modeling and near‐hover‐flight control design for an unconventional aerial robot of the tandem ducted fan configuration, which is intended to be prototypical of a flight service vehicle. The main model elements of this novel unmanned vehicle, which exhibit highly nonlinear and unstable open‐loop modes, are presented. A frequency‐domain controllability analysis concerning the plant's behavior around the hovering flight condition is then adopted to determine the expected control performance, which is of important practical significance to controllability improvement through vehicle design changes. A robust controller that stabilizes the unmanned vehicle under wind disturbances is designed using a newly developed nonsmooth optimization algorithm, which rigorously and efficiently tunes the arbitrarily predefined structured controller against multiple control requirements. A successive two‐loop architecture is employed in the designed controller. In this architecture, the inner loop provides stability augmentation and decoupling, and the outer loop guarantees the desired velocity tracking performance. Simulation results under stochastic wind gusts are presented to verify the performance of the proposed controllers. Preliminary flight tests are also carried out to demonstrate the performance of the system.

Aerodynamic Interaction between Rotor and Tilting Wing in Hovering Flight Condition

The hovering performance and the lifting capability of tiltrotor aircraft are strongly affected by the aerodynamic interaction between wing and rotors. The tiltwing concept represents an interesting technology to increase the hover performance by reducing the wing–rotor interference. The present work investigates the aerodynamic interaction between wing and rotor in hover for a scaled tiltwing aircraft half-span model. A comprehensive experimental campaign, including force measurements and particle image velocimetry surveys, was performed together with computational fluid dynamics simulations. Numerical predictions were validated using experimental data and were used to describe the flow field.