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Higher harmonic control (HHC) is an approach for achieving reduced helicopter vibration by controlling the vibratory rotor airloads in such a way that the fuselage excitation is minimized. This paper is a historical look at how a program aimed at helicopter vibration reduction started as an outgrowth of fixed wing flutter suppression at NASA Langley Research Center, proved the HHC concept on aeroelastically scaled wind tunnel models and went on to demonstrate viability in full-scale flight testing on the OH-6A helicopter in 1982. Following the OH-6A flight tests, the helicopter research community was stimulated to prove the effectiveness of HHC on different configurations through analysis, wind tunnel tests, and flight tests. All of these investigations have shown HHC to be effective in reducing vibration to levels not attainable with conventional vibration control methods and without any detrimental side effects. HHC development has progressed to the point that the technology provides one more option to address the ever-present vibration problem in helicopters. The literature demonstrates that helicopter ride quality equivalent to that of fixed wing aircraft is attainable with application of HHC. [ABSTRACT FROM AUTHOR]

Purpose: The paper aims at developing a novel algorithm to estimate high-order derivatives of rotorcraft angular rates to break the contradiction between bandwidth and filtering performance because high-order derivatives of angular rates are crucial to rotorcraft control. Traditional causal estimation algorithms such as digital differential filtering or various tracking differentiators cannot balance phase-lead angle loss and high-frequency attenuation performance of the estimated differentials under the circumstance of strong vibration from the rotor system and the rather low update rate of angular rates. Design/methodology/approach: The algorithm, capable of estimating angular rate derivatives to maximal second order, fuses multiple attitude signal sources through a first-proposed randomized angular motion maneuvering model independent of platform dynamics with observations generated by cascaded tracking differentiators. Findings: The maneuvering flight test on 5-kg-level helicopter and the ferry flight test on 230-kg-level helicopter prove such algorithm is feasible to generate higher signal to noise ratio derivative estimation of angular rates than traditional differentiators in regular flight states with enough bandwidth for flight control. Research limitations/implications: The decrease of update rate of input attitude signals will weaken the bandwidth performance of the algorithm and higher sampling rate setting is recommended. Practical implications: Rotorcraft flight control researchers and engineers would benefit from the estimation method when implementing flight control laws requiring angular rate derivatives. Originality/value: A purely kinematic randomized angular motion model for flight vehicle is first established, combining rigid-body Euler kinematics. Such fusion algorithm with observations generated by cascaded tracking differentiators to estimate angular rate derivatives is first proposed, realized and flight tested. [ABSTRACT FROM AUTHOR]

Purpose: This paper aims to present the novel methodology of computational simulation of a helicopter flight, developed especially to investigate the vortex ring state (VRS) – a dangerous phenomenon that may occur in helicopter vertical or steep descent. Therefore, the methodology has to enable modelling of fast manoeuvres of a helicopter such as the entrance in and safe escape from the VRS. The additional purpose of the paper is to discuss the results of conducted simulations of such manoeuvres. Design/methodology/approach: The developed methodology joins several methods of computational fluid dynamics and flight dynamic. The approach consists of calculation of aerodynamic forces acting on rotorcraft, by solution of the unsteady Reynold-averaged Navier–Stokes (URANS) equations using the finite volume method. In parallel, the equations of motion of the helicopter and the fluid–structure-interaction equations are solved. To reduce computational costs, the flow effects caused by rotating blades are modelled using a simplified approach based on the virtual blade model. Findings: The developed methodology of computational simulation of fast manoeuvres of a helicopter may be a valuable and reliable tool, useful when investigating the VRS. The presented results of conducted simulations of helicopter manoeuvres qualitatively comply with both the results of known experimental studies and flight tests. Research limitations/implications: The continuation of the presented research will primarily include quantitative validation of the developed methodology, with respect to well-documented flight tests of real helicopters. Practical implications: The VRS is a very dangerous phenomenon that usually causes a sudden decrease of rotor thrust, an increase of the descent rate, deterioration of manoeuvrability and deficit of power. Because of this, it is difficult and risky to test the VRS during the real flight tests. Therefore, the reliable computer simulations performed using the developed methodology can significantly contribute to increase helicopter flight safety. Originality/value: The paper presents the innovative and original methodology for simulating fast helicopter manoeuvres, distinguished by the original approach to flight control as well as the fact that the aerodynamic forces acting on the rotorcraft are calculated during the simulation based on the solution of URANS equations. [ABSTRACT FROM AUTHOR]

The article addresses investigating the possibilities of making a parameterized flight dynamics model in a mixed reality environment. The purpose was to obtain a model for a Mi-17 helicopter. The Mi-17 VI helicopter flight test characteristics were used to evaluate the above model. As a comparative criterion, compatibility between the model's selected characteristics and the real object was adopted. Simulation tests were conducted on a cabin simulator set made via MR (mixed reality) technology. Test results confirmed the possibility of reconstructing the selected helicopter type's flight dynamics using the parameterized model. [ABSTRACT FROM AUTHOR]

Regime recognition is an important tool in the creation of usage spectra and component lifetime prediction. During development of regime recognition codes, it is necessary to establish baseline performance using scripted flight-test data. Various problems arise when using scripted flight data for verification and validation that may be mitigated by using simulated data. This paper presents a virtual pilot algorithm designed to generate scripted flight-test data from a helicopter simulator. The virtual pilot receives a maneuver script as input and streams control inputs to a flight simulator to perform the required maneuvers in an accurate and reliable manner. The algorithm is formulated as a variable structure controller where each maneuver is mapped to a single feedback law. Several techniques are employed to validate the virtual pilot approach, and results show that synthetic Health and Usage Monitoring data generated by the virtual pilot compare favorably against actual flight data. [ABSTRACT FROM AUTHOR]

This paper describes the experimental design studies conducted to maximize the hover endurance of a micro air vehiclescale quadrotor helicopter (gross weight<50 g) by careful design and selection of the individual subsystems while understanding their interdependencies. The different subsystems of the quadrotor that were examined include the rotors, motors, electronic speed controllers, gear transmission systems, battery, and airframe structure. An optimized micro air vehicle-scale rotor with a diameter of 110mm and a figure of merit of 0.67 served as the basis for the vehicle design. Brushed and brushless motor systems were paired with the optimized rotor and were examined for maximum system level efficiency and capability to produce the required thrust. Propulsion systems using brushed motors with speed controllers and the optimum gear ratio showed up to 35% improvement in efficiency over brushless outrunner motors with speed controllers. Lithium-ion polymer batteries were compared and evaluated on voltage drop-off characteristics and achievable endurance versus weight. Quadrotor airframe weight has been minimized to 7.4% of the gross weight, representing a 30-50% decrease compared existing similarly sized quadrotors. The highest endurance was obtained using brushed motors with a 5.33:1 gear ratio and a 650 mAh battery. Presently, the longest continuous hover endurance achieved with this design is 31 min. [ABSTRACT FROM AUTHOR]

Piloted simulation tests were conducted to evaluate advanced response types for ship-based rotorcraft. The simulation used the GENHEL-PSU model of an UH-60 operating from a FFG-7 frigate. The Control Equivalent Turbulence Input (CETI) model was used for airwake turbulence. A full-authority nonlinear dynamic inversion control law was designed to provide attitude command/attitude hold (ACAH), acceleration command/velocity hold (ACVH), and translational rate command/position hold (TRC/PH) response types. Ship deck motion measurements were fed into these control laws to provide ship-relative control. Piloted evaluations were performed for each response type as well as the baseline mechanical control system. A maritime mission task element was performed with and without ship motion and airwake effects. Pilot comments, handling qualities ratings, and workload metrics indicated that the best performance was achieved using an ACVH response type on approach, which then automatically transitions to TRC/PH over the ship deck. [ABSTRACT FROM AUTHOR]

This work focuses on targeted flight of rotary wing micro aerial vehicles, in previously unexplored, global positioning system-denied environments. Targeted flight is defined as navigating to a goal position, defined by relative distance from a known initial position. It differs from the more common task of mapping, as it may not rely on loop closure opportunities to smooth out errors and optimize the generated map. Therefore, position estimates' accuracy becomes a necessity. A refined position and mapping algorithm is hereby developed, and its accuracy is quantitatively assessed using a dedicated metric, showing significant advantages over current methods. Repeatability and robustness are shown using a set of 12 repeated experiments in a benchmark scenario. The algorithm is then coupled with a global A* path planner and applied on a single rotor helicopter, performing targeted flight missions using a pilot-in-the-loop implementation. The pilot is provided with only heading information, and artificial (invisible) obstacles are introduced, apparent only to the path planner, to prevent the pilot from interfering with the obstacle avoidance task. The resulting path shows the helicopter successfully avoids both real and artificial obstacles, while following the planned path to the goal. [ABSTRACT FROM AUTHOR]

Modern articulated rotors utilize an elastomeric bearing on the root of each blade to allow multiaxis articulation through a single component. The bearing forms a spherical joint that makes angular displacements in specific blade axes both coupled and difficult to measure. A smart sensing concept is developed to uncouple and separately measure the blade flapping, feathering, and lead-lag position. The sensor concept projects a magnetic field from the blade root to the rotor hub such that flap, lag, and pitch displacements can be derived from a hub-mounted pickup array. The device utilizes permanent magnets on the blade root and an array of anisotropic magnetoresitive sensors on the hub. The sensor array is integrated with other electronic components, including a microprocessor in a printed circuit board, forming a self-contained sensing system. The "smart" sensor is software driven, self-checking, and can be networked with other sensors for application flexibility. The sensor accuracy is ensured through a process that develops calibration coefficients that are loaded into system nonvolatile memory. The sensor system is designed for initial application on an articulated tail rotor for flight-test purposes. Risk reduction testing is conducted in simulated tail rotor environmental conditions including vibration and shock, elevated centrifugal force, temperature extremes, and electromagnetic emissions/susceptibility. All development and test results indicate the new sensor technology is valid and ready for use in its initial flight-test application. [ABSTRACT FROM AUTHOR]

There is a critical lack of quantitative data regarding the mechanism of brownout cloud formation. Recognizing this, tests were conducted during the U.S. Air Force Research Laboratory 3D-LZ Brownout Test at the U.S. Army Yuma Proving Ground. Photogrammetry was utilized during two rounds of flight tests with an instrumented EH-60L Black Hawk to determine whether this technique could quantitatively measure the formation and evolution of a brownout cloud. Specific areas of interest included the location, size, and average convective velocity of the cloud, along with the characteristics of any defined structures within the cloud. Following the first flight test, photogrammetric data were validated through comparison with onboard vehicle data. Lessons learned from this test were applied to the development of an improved photogrammetry system. A second flight test, using the improved system, demonstrated that obtaining quantitative measurements of the brownout cloud is possible. Results from these measurements are presented. Flow visualization with chalk dust seeding was also tested to gain additional insight on the particle pickup process. Based on the amount of chalk removed during the landings, the pickup forces of the brownout cloud appear to be very low. Overall, these tests demonstrate the viability of photogrammetry as a means for quantifying brownout cloud formation and evolution. [ABSTRACT FROM AUTHOR]

This article offers a new perspective in defining comfortable helicopter flight along with two solutions based on passive and active main rotor morphing. Constrained optimization design problems aimed at minimizing flight control energy while satisfying variance constraints on flight parameters that are considered important in passenger comfort and noise reduction are formulated and solved. Output variance constrained (OVC) control is used for control system design and simultaneous perturbation stochastic approximation (SPSA) is employed to solve the resulting constrained optimization problems. Details on the computation of the control energy are given. Closed-loop responses of designs that satisfy prescribed variance constraints with a very small margin on the achievable variance bounds are compared with responses of designs that satisfy such constraints with a larger margin both for passively and actively morphing helicopters. [ABSTRACT FROM PUBLISHER]

Several U.S. Navy rotorcraft models are equipped with Health and Usage Monitoring Systems. On-board or on-ground Regime Recognition (RR) programs are used to develop individual rotorcraft usage spectrums to compute component and airframe fatigue life expended (FLE). FLE is a function of rotorcraft gross weight (GW) and center of gravity (CG), but neither are monitored by sensors. Thus GW and CG computation algorithms based on measured engine torque curves and trim equations during hover are developed and incorporated into RR. The algorithms are validated using numerous fleet flights with known GWs and CG positions. Excellent correlation between predicted and actual values is observed with low standard error of estimate. The in-flight variation of GW and CG is determined with the help of recorded fuel weight. RR is used to analyze 32,000 h of data from 111 rotorcraft to develop GW and CG utilization distributions. Furthermore, GW variation is modeled with a beta probability distribution and the joint probability distribution of GWand CG utilization is determined to assess airframe and dynamic component fatigue lives. [ABSTRACT FROM AUTHOR]

The stabilization of "difficult" loads that become aerodynamically unstable at airspeeds well below the power-limited speed of the helicopter-load configuration has been studied since the 1960s. This paper looks at the possibility of stabilizing slung loads in forward flight by imposing a slow steady rotation in yaw (spin stabilization). Slow rotations of 100-150 deg/s suffice to suppress the pendulum motions of the load. A swivel is required at the hook, and only a few foot-pounds of yaw moment are needed to overcome swivel friction and impose the desired yaw rate. The approach is limited to single-point suspensions. A stabilizer design consisting of a one-shaft anemometer-like device with hemispherical cups at the ends was developed in wind tunnel tests. The shaft angle can be controlled to vary the applied yaw moment and allow feedback regulation of the load yaw rate. Flight tests with two cargo containers demonstrated that a simple linear control law with fixed gains was effective in maintaining the desired yaw rate in forward flight over the range of configurations of the test loads. Wind tunnel data were obtained at all stages of the development and testing and proved to be an accurate source of design data and an accurate predictor of performance in flight. [ABSTRACT FROM AUTHOR]

To measure the noise footprint of a helicopter approach procedure, microphones are usually distributed over a large area. In this paper, however, a method is utilized where the measurement of the noise footprint is conducted in parts by distributing the microphones over a limited area and executing the same approach procedure, while the position of the landing point is shifted. The flight tests are conducted with the EC135-ACT/FHS helicopter. Guidance to the pilots is provided by use of a "tunnel-in-the-sky" head down pilot display. Since the same flight procedure is executed multiple times, it is possible to conduct a statistical analysis on the accuracy and reproducibility of the approach procedure in terms of position, velocity, and generated noise. This statistical analysis shows that the geometric accuracy is on the order of 6 m in the horizontal plane and 12 m in the vertical plane. The variations in the measured sound exposure level from one approach to another are on the order of 3-7.5 dB(A). The approach procedure is also executed with excessive cyclic pilot control inputs to identify an upper limit of the extra noise generated by unsteady flight. However, these approaches are indistinguishable, in terms of the sound exposure level or sound pressure level, from the approacheswith normal pilot control inputs. [ABSTRACT FROM AUTHOR]

An ongoing concern with helicopter simulations is that they often have an incorrect off-axis response to cyclic control inputs when compared with the corresponding flight-test data. The more commonly suggested contributing factors for this discrepancy include the influence of dynamic wake distortion, rotor blade elasticity, and unsteady blade airfoil aerodynamics. A rotor model was developed using a Ritz expansion approach with constrained elastic modes to account for the blade elasticity, as it is computationally compact and efficient and therefore suitable for use in a real-time simulation. The effect of including this blade flexibility on the dynamic response, and in particular the on-axis and off-axis response, of the University of Toronto Institute for Aerospace Studies helicoptermodels is examined. In addition, the combined effects of dynamic wake distortion, unsteady blade section aerodynamics, and blade flexibility on the dynamic response are examined. The various features were successful in altering the off-axis response, with improvements in some areas, while not disrupting the on-axis response. In some conditions, the magnitude of the resulting change due to flexibility was greater than the differences noted due to the addition of dynamic wake distortion or unsteady aerodynamics. [ABSTRACT FROM AUTHOR]

This work developed and verified a computational model to predict self-excited limit cycle yaw oscillation (SELCYO) instability of external sling payloads carried under aircraft with dual-point suspension. Inverted-V and inverted-Y slings during steady-state level flight are discussed. The primary goal was to provide a design tool for comparison of the onset of SELCYO between alternate sling geometries. The computational model incorporates steady-state aerodynamic loading during level flight based on scale-model wind tunnel testing. Scale-model sling tests of the onset of SELCYO in the same wind tunnel were used for validation. Predictions of cargo hook load for a full-size HMMWV-M1025 payload carried by inverted-V slings are compared to V-22 Osprey flight-test data. Predictions of stability indicate that stiffer slings are generally more stable, and inverted-V slings are significantly more stable than inverted-Y slings. Small differences between right and left front sling leg lengths caused by rigging error can significantly reduce stability. [ABSTRACT FROM AUTHOR]

This paper presents the design, development and construction of a flight test stand for a quadrotor UAV. As opposed to alternate forms of UAV, the power plant in the case of the quadrotor serves a dual purpose of control and propulsion. Since control and propulsion are coupled, the power plant (BLDC motor coupled with propeller) was studied in detail using a black box structure. Extractions of motor parameters in previous studies used traditional BLDC motor equations and propeller theory however the accuracy achievable and confidence in the extracted parameters remained questionable. The developed data acquisition process served to satisfy this need by the construction of a test bench that allows for the extraction of the unknown parameters instilling confidence in the modelling process. The established relationships are then used as inputs into a developed six degree of freedom Euler based mathematical model. A mission profile was constructed with distinct phases of which the mathematical model was used to simulate. Each phase in the mission profile excited different modes of the quadrotor dynamics creating an ideal simulation environment in which changes can be implemented and studied. [ABSTRACT FROM AUTHOR]

AeroVelo initiated the Atlas Human-Powered Helicopter Project in August 2011 to win the American Helicopter Society Sikorsky Prize, which despite prior attempts had remained unclaimed for over 30 years. The Sikorsky Prize required a human-powered helicopter to sustain flight for 60 s, momentarily reach a height of 3 m, and maintain position within a 10-by-10-m area. A configuration study was undertaken using low-fidelity aerodynamic analysis and estimated mass figures. An aerostructural optimization framework was developed for rotor design, including a novel vortex-ring aerodynamic model with included ground effect prediction, finite element analysis including integrated composite failure analysis, and a detailed weight estimation scheme. The airframe comprised a wire-braced truss structure, and innovative designs were developed for many of the aircraft's lightweight-focused subsystems. After initial flight testing in August 2012, experimental optimization and performance improvement led to a second testing program beginning in January 2013. Testing in 2013 led to a reduction in required power, and improved understanding of the structural dynamics and control strategy. The project culminated with the successful American Helicopter Society Sikorsky Prize flight on 13 June 2013. [ABSTRACT FROM AUTHOR]

The ability of a helicopter to carry externally slung loads makes it very versatile for many civil and military operations. However, the piloted handling qualities of the helicopter are degraded by the presence of the slung load. A control system is developed that uses measurements of the slung load motions as well as conventional fuselage feedback to improve the handling qualities for hover/low-speed operations. Prior research has shown a fundamental trade-off between load damping and piloted handling qualities for a feedback control system with cable angle/rate feedback. A new task-tailored approach proposed and implemented herein uses a method of switching between a load damping mode and a piloted handling qualities mode. These modes provide appropriate load feedback depending on the piloting task and flight regime. This provides improved handling qualities for maneuvering flight and for improved precision load control at hover. A new mission task element for precision load placement is developed (for possible inclusion into ADS-33E-PRF) to test the ability of the cable feedback system to improve load placement task performance. The improvements provided by this control system are demonstrated in a piloted flight test on the JUH-60A RASCAL fly-by-wire helicopter. The average load set-down time was reduced by a factor of two for the 1000-lb load on a 56-ft sling. [ABSTRACT FROM AUTHOR]

This paper describes a framework for performance evaluation of autonomous guidance systems. The elements of the framework consist of a set of spatial geometries, flight tasks, performance metrics, a flightdynamic model, and baseline solutions. The spatial benchmarks consist of six tasks in simple geometrical environments and 10 tasks in more complex urban environments based on a real digital terrain elevation map. The framework also includes a set of performance metrics used to compare trajectories. The performance baselines used in the proposed framework are near-optimal solutions computed using one of two trajectory optimization methods: numerical optimization based on nonlinear programming for the simple geometric environments and a motion primitive automaton for problems involving the urban environments. The paper concludes with a demonstration of the benchmarking framework using the Obstacle Field Navigation system developed by the Army Aeroflightdynamics Directorate. [ABSTRACT FROM AUTHOR]

This paper presents a mathematical model for the simulation of rotor blade flexibility in real-time helicopter flight dynamics applications that also employs sufficient modeling fidelity for prediction of structural blade loads. A matrix/vector-based formulation is developed for the treatment of elastic blade kinematics in the time domain. A novel, second-order-accurate, finite-difference scheme is employed for the approximation of the blade motion derivatives. The proposed method is coupled with a finite-state induced-flow model, a dynamic wake distortion model, and an unsteady blade element aerodynamics model. The integrated approach is deployed to investigate trim controls, stability and control derivatives, nonlinear control response characteristics, and structural blade loads for a hingeless rotor helicopter. It is shown that the developed methodology exhibits modeling accuracy comparable to that of non-real-time comprehensive rotorcraft codes. The proposed method is suitable for real-time flight simulation, with sufficient fidelity for simultaneous prediction of oscillatory blade loads. [ABSTRACT FROM AUTHOR]

• Anomaly detection in condition monitoring using healthy vibrations for training. • Design of data-driven model architectures that integrate a physical/statistical understanding of different faults. • LSTM/SVM capable of detecting new deterministic components introduced by gear faults. • A two-step LSTM regression + one-class SVM for detecting new random components caused by bearing faults. • SVM capability in fault detection when vibrations collected using different sensors. Fault detection is a critical step for machine condition monitoring and maintenance. With advances in machine learning technologies, automated faulty condition identification can be achieved by training an artificial intelligent (AI) model if sufficient data is available. In most practical applications, it is unlikely that enough data from faulty cases are available for supervised training of AI models for anomaly detection. This work is based on training data that are exclusively constituted of healthy signals (i.e., semi-supervised) and aims to develop an automated algorithm capable of identifying any abnormal mechanical behaviour captured by vibration measurements. A deep learning method with long short-term memory (LSTM) architectures combined with a one-class support vector machine (SVM) is used to separate abnormal data from normal vibration signals collected during an endurance test of a reduction gearbox and helicopter test flight data measured by multiple sensors situated at different locations of the aircraft. For the gearbox dataset, a detailed understanding of the physical mechanisms by which different types of faults (gear wear and bearing faults) affect the vibration signal led to the design of two anomaly detection architectures: i) an LSTM regression + one-class SVM for detecting new deterministic components introduced by gear faults; and ii) a two-step LSTM regression + one-class SVM for detecting new random components caused by bearing failures. For the helicopter dataset, which does not contain consecutive time-series, we show that the LSTM regression is not advantageous, and a better performance can be achieved by a simpler one-class SVM outlier detection based on statistical features. This work contributes to the field of machine condition monitoring by introducing a novel two-step LSTM configuration for removing deterministic components associated with dominant gear signals in the first step and removing the 'residual deterministic' components associated with varying gear signals in the second step. [ABSTRACT FROM AUTHOR]

The paper presents the research carried out to handle the air resonance mode that occurs on the DLR EC135 ACT/FHS (Active Control Technology/Flying Helicopter Simulator) research helicopter. The air resonance mode on the ACT/FHS is excited in particular when roll rate feedback loops in the flight control system are closed. A flight-test campaign was conducted to gather data of the air resonance mode for different roll rate feedback gains. A model, determined by system identification, is tuned to optimally match the model response with the flight-test data for different roll rate feedback gains. Based on this model, an air resonance controller was subsequently developed and tested that effectively suppresses the excitation of the air resonance mode resulting from roll rate feedback. The final air resonance controller concept was extensively flight-tested to study the influence on the flight control performance and handling qualities. [ABSTRACT FROM AUTHOR]

This paper presents a new active in-flight rotorblade tracking control system for helicopters equipped with trailingedge flap actuators on the rotor blades. The objective of the in-flight rotorblade tracking controller is to minimize vibrations in the helicopter cabin and vibratory loads at the rotor hub induced by static blade dissimilarities. The control design is based on a steady-state transfer matrix from flap inputs to vibration outputs, which is first identified experimentally and then symmetrized. For this design model, an analytically derived Moore-Penrose pseudoinverse, which plays a key role in overcoming restrictions of existing approaches, such as the dependency on full-rank transfer matrices, is used. The system is verified through flight tests carried out with a full-scale experimental helicopter. The results demonstrate that the presented in-flight tracking system effectively minimizes the tracking error induced vibrations. [ABSTRACT FROM AUTHOR]

Background-oriented schlieren (BOS)methods suited for large-scale and in-flight testing are presented with special emphasis on the detection and tracing of blade tip vortices in situ. Retroreffective recording and photogrammetric epipolar analysis for the computation of the vortices' spatial coordinates in the wind tunnel are described. Feasibility and fidelity of reference-free BOS in conjunction with natural formation backgrounds and related evaluation methods are discussed, additionally, illustrating their simplicity and robustness. Results of successful image acquisition from a chaser aircraft are presented allowing vortex wakes to be identified at a wide range of flight attitudes, including complex maneuvers. [ABSTRACT FROM AUTHOR]

A methodology to predict the airflow hazard of helicopter flight in icing conditions is developed. By incorporating the existent ice accretion codes into an established basic helicopter flight dynamic model and considering airflow disturbance that mainly covers downdraft, head/tail wind, and left/right wind, the hazardous effects on trims, stability, and controllability of UH-60A single rotor helicopter in icing/ice-free conditions and within/without different types of wind field are investigated. The stability and controllability of helicopter that encounters airflow disturbance from wind velocity of 0 to 2.5–5.0 m/s for forward flight are examined. The indications of all the work are summarized at the end of this article. Furthermore, this method can be used to helicopter inflight safety prediction or airflow hazard avoidance analysis in icing conditions. It can also be laid as the foundation of the further research about the more complex airflow hazard prediction in icing conditions for helicopter flight safety. [ABSTRACT FROM AUTHOR]

Highlights: [•] An application was developed to support pro-active routing of police helicopters. [•] The in-time arrival of police helicopters can be improved with a factor nine. [•] Ten flights can be scheduled in approximately five to six minutes. [•] All historic data can be used, regardless of location, date, and time. [•] The application has been validated and is currently used by the police. [Copyright &y& Elsevier]

In this paper, a proportional-integral sliding mode control (PISMC) method is presented to achieve the attitude and position control of an unmanned helicopter system for autonomous hovering. The PISMC includes the advantages of both conventional sliding mode control (SMC) and integral sliding mode control (ISMC). The proposed PISMC algorithm is capable of stabilized attitude and position control for unmanned helicopter autonomous hovering, and it offers a performance improvement over the conventional SMC and ISMC algorithms as well. The system performance is quantified and an experimental unmanned helicopter system is developed to access the performance of the PISMC. Both simulations and actual flight tests are conducted to verify the PISMC attitude and position control. The simulation results show that the PISMC tracking error is smaller and that PISMC converges faster than SMC does. The flight test results prove that the proposed PISMC is effective in the attitude and position control of unmanned helicopter autonomous hovering. [ABSTRACT FROM AUTHOR]

In this paper, a mathematical model of a helicopter NO TAil Rotor (NOTAR) antitorque system is developed for real-time flight simulations. The model consists of the circulation control tail boom, direct jet thruster, and the vertical stabilizers. The airflow inside the tail boom is modeled by dividing the flow into aerodynamic control volumes. The model features a blade-element-type approach for modeling the mass flow through the axial fan blades as well as aerodynamic mass and momentum conservation calculations in each control volume. The side force produced by the direct jet thruster and circulation control tail boom is calculated at different flight conditions. A simple rectangular plan form with symmetrical airfoil shape is used to model the vertical tail. The developed NOTAR antitorque model is integrated with a nonlinear helicopter model. Flight tests are performed on the MD600N, and the data obtained are used for verification. Results show a reasonable match between the model and flight-test data in the antitorque response in hover and forward flight to pilot pedal input. [ABSTRACT FROM AUTHOR]