Unmanned helicopter flight control actuator specification through mission profile analysis

Aircraft maintenance is one of the most important cost items faced by the operators of air fleets and is a major contributor to the aircraft life cycle cost. An aircraft fly-by-wire flight control system has a total of primary flight control actuators ranging from 10 to 20 depending on the aircraft type, with a failure rate of 1/1000 flight-hours; therefore, a health monitoring system for primary flight control actuators, able to recognize an actuator degradation in its early stage could greatly contribute to optimize the maintenance operations, reduce the airplane downtime and prevent missions interruptions.This note presents the initial part of an ongoing research project aimed at developing a prognostic and health management system for fly-by-wire primary flight control actuators. A key feature of the project is to develop a PHM system for these actuators suitable for the flight control actuators of legacy airplanes, which are poised to operate for still a long time, and not only for those of new aircraft. The primary flight control actuators of fly-by-wire flight control systems of existing aircraft are electrohydraulic servoactuators with a typical configuration and complement of transducers, and there is no practical possibility of introducing additional sensors. For this reason, the research activity was directed towards the study of algorithms able to identify faults only by using the already available information of the servoactuators state variables.The implemented algorithms are a combination of mathematical and neural network based ones, and the identification of degradations was performed by the analysis of the response of the servoactuators to a sequence of selected stimuli provided in preflight or postflight. The servovalve current and the feedback position are processed by dedicated algorithms in order to obtain significant indicators of the servocatuator health condition. The values of the indicators obtained during the sequence of stimuli are analyzed in combination with those obtained in the past.This is performed by the neural network part of the algorithm which allows a reliable identification of presence and of type of a degradation.The results obtained from the initial part of the research activity are interesting and encouraging. Individual degradations of the servoactuator parameters have so far been addressed and the algorithms for identifying them have been developed. All that makes up the foundations of the future research activity which will be focused on analyzing the effects of simultaneous multiple degradations and to the estimation of the remaining useful life.

This paper introduces a real field mission profile oriented design tool for the new generation of grid connected PV-inverters applications based on SiC-devices. The proposed design tool consists of a grid connected PV-inverter model, an ElectroThermal model, a converter safe operating area (SOA) model, a mission profile model and an the evaluation block. The PV-system model involves a three level bipolar switch neutral point clamped (3L-BS NPC) inverter connected to the three phase grid through a LCL-filter. Moreover, the SOA model calculates the required converter heatsink thermal impedance Zth_H in order to perform in a safe mode for the whole operating range. Furthermore, the proposed design tool considers the mission profile (the measured solar irradiance and ambient temperature) from the real field where the converter will operate. Thus, a realistic loading of the converter devices is achieved. To consider one year real field measurements of the mission profile, an accurate long term simulation model is developed. The model predicts the junction and case temperature of the converter devices, for three different case scenarios. In the first case, a one year mission profile is used into the model with a sampling rate of five minutes. For the second and third case, a more detailed analysis is performed for one week mission profile (in the wintersummer time) with a sampling rate of 25 s. The simulation results shows the thermal loading distribution among the converter devices (MOSFET, IGBT+FD) in terms of junction (average, peak, ΔT) and case (average) temperature for all three simulation cases. Finally, the evaluation block is used to analyze the results in order to perform a thermal loading-based classification of the converter devices.

Model-based design and experimental verification of a monitoring concept for an active-active electromechanical aileron actuation system

Electromechanical actuators (EMAs) are considered a promising energy-efficient technology for actuating the flight controls of future aircraft. When it comes to aerospace systems, the degradation of EMAs should be checked during regular maintenance events or through condition-based maintenance. Ball bearings have a significant failure rate for flight control EMAs and are usually monitored by vibration noise. A challenge for detecting bearing faults, using state-of-the-art industrial methods, is the presence of a ball-screw mechanism that produces nominal vibration noise similar to that of faulty bearings. No prior research has investigated this problem. This paper explores vibration noise generated from a set of healthy and faulty bearings included in a typical ball-screw EMA. In addition, a method is introduced for evaluating fault diagnosis performance for different time and frequency vibration features. The technique has been validated on an EMA actuator at the German Aerospace Center (DLR).

In this chapter, preliminary sizing and performance calculations of unmanned air vehicles (UAV), with relation to its propulsion system, are explained. Starting with the mission profile of the air vehicle, which is one of the important design drivers, the UAV is sized for the requirements. The requirements in design consist of a set of target values (e.g., payload amount to be carried, range, endurance time, cost, etc.). Additionally, the parameters within mission profile such as cruising altitude and speed of aircraft affect engine type, power level required, fuel quantity, and therefore general dimensions and the gross weight of the aircraft. It is often an iterative process to size the air vehicle and engine together. UAV is designed in a loop of calculations in which sizing, flight performance, and engine performance are done for each phase of the mission or flight profile to satisfy the overall design mission and requirements.

This paper presents a method of reducing fatigue damage to critical components on the main rotor hub of a helicopter through load alleviation in the primary flight control system. The method is demonstrated using a simulation representative of the CH-53E Super Stallion. The components in question were the pushrods, rotor shaft, and the stationary scissor links. The component loads and rate of fatigue damage were correlated with aircraft states which are in turn regulated and tracked by a fly-by-wire control system. Using dynamic inversion control laws with command filter shaping, the accumulated damage to each component is shown to decrease during aggressive maneuvers while causing only a small degradation to the handling qualities.

Magnetostrictively actuated control flaps for vibration reduction in helicopter rotors—design considerations for implementation

Future space power requirements will vary from the subkilowatt range for deep space probes, to the hundreds of kilowatts range for a lunar base, to the multimegawatt range for interplanetary propulsion systems. Closed Brayton cycle (CBC) power conversion has the flexibility to be used in all these power ranges and with a variety of heat source options such as isotope, solar, and nuclear. Each of these types of heat sources has different characteristics that make it more appropriate for particular mission profiles and power output ranges. Heat source characteristics can also be major design drivers in the closed Brayton cycle design optimization process. This paper explores heat source selection and the resulting CBC system designs, and discusses optimization methods as a function of the main design drivers. Such power system requirements as power level, man-rated radiation shielding, fuel costs, eclipse/darkness duration, system mass, radiator area, reliability/mission duration, and insolation level are evaluated through several CBC parametric case studies. These cases include: (1) a 500 We power system for deep space probes; (2) a 50 kWe solar dynamic system for earth orbit and other applications; (3) a 100 kWe man-rated lunar/Mars stationary /rover power system; (4) a 200 to 825 kWe power system for the lunar outpost; and (5) 3300 kWe modules for interplanetary propulsions.

In the last years, the layout of servomechanisms used in the aeronautical field to actuate the flight controls has changed radically and, nowadays electromechanical actuators (EMAs) are increasingly replacing the older hydraulic powered actuator types. The definition of special monitoring procedures, based on the analysis of the system response and aiming to evaluate the evolution of faults, represents an important task of the modern system engineering taking into account that onboard actuators are typically safety critical items. The present paper proposes a new prognostic procedure centered on the characterization of the state of health of an EMA used in aircraft primary flight controls. This approach, based on the innovative use of a model-based fault detection and identification method (FDI), identifies the actuator actual state of wear of the actuator analyzing proper system operational parameters, able to put in evidence the corresponding degradation path, by means of a numerical algorithm based on spectral analysis techniques. The proposed FDI algorithm has been tested in case of EMA affected by two progressive failures (rotor static eccentricity and stator phase turn-to-turn short-circuit), showing an adequate robustness and a suitable ability to early identify EMA malfunctions with low risk of false alarms or missed failures.

Electro-Mechanical Actuator (EMA) is a key component in the flight control actuation of the more electric aircraft. The degradation of EMA has a significant influence on its performance. Hence, the degradation estimation of EMA needs to be studied, which is an important part of the health monitoring of EMA. However, the multiple failure modes of EMA bring much difficulty for implementing degradation estimation. To solve this problem, a fusion degradation estimation method based on Continuous Hidden Markov Model (CHMM) and Generalized Linear Model (GLM) (i.e., Multiple CHMM-GLM) is proposed for EMA with multiple failure modes. Firstly, Health Indicator (HI) extraction models are formulated based on CHMM for multiple failure modes to extract HIs, which can be used to indicate the degradation degree of EMA. Secondly, the fusion degradation model for EMA with multiple failure modes is formulated based on the extracted HIs and GLM, in which the nonlinearity of a single HI and the coupling among multiple HIs are both considered. Finally, the output of the fusion degradation model is regarded as a HI to indicate the degradation degree of EMA. This study provides a new way for the degradation estimation of EMA with multiple failure modes, which can contribute to the Prognostics and Health Management (PHM) development of EMA. To evaluate the effectiveness of the proposed method, experiments are conducted utilizing EMA simulation and practical data. Experimental results show that the proposed method has a good performance in the degradation estimation for EMA with multiple failure modes.

This study presents a comprehensive numerical framework for auralizing multirotor noise during complex maneuvers, incorporating rotor tonal and broadband noise prediction, synthesis, and psychoacoustic analysis. Validation studies for various flight scenarios, including single rotors and quadrotors, confirm the capability of this framework in the prediction-based psychoacoustic analysis of multirotor configurations. The primary focus of this research is to assess the flyover and takeoff noise of multirotors under diverse operating conditions. When tracking the mission profile under gusty wind conditions, the multirotor noise is characterized by frequency and amplitude modulations induced by rotational speed control for each rotor. A high-resolution time-frequency analysis is conducted for the tonal noise to highlight these modulation characteristics. Additionally, the prediction-based psychoacoustic analysis is performed for cross- and plus-type quadrotor configurations to clarify the effects of gusty wind profiles and flight control. The results demonstrate that variations in the rotational speed of each rotor, influenced by mission profiles and gusty wind conditions, contribute to the distinct acoustic characteristics of multirotor configurations in both physical and psychological aspects. The synthesized noise time signal and modulation characteristics provide valuable insights into the impact of operational environments on psychoacoustic metrics and annoyance.

Neptec has developed a family of obscurant-penetrating 3D laser scanners called OPAL 2.0 that are being adapted for rotorcraft platforms. Neptec and Boeing have been working on an integrated system utilizing the OPAL LiDAR to support operations in degraded visual environments. OPAL scanners incorporate Neptec’s patented obscurantpenetrating LiDAR technology which was extensively tested in controlled dust environments and helicopters for brownout mitigation. The OPAL uses a scanning mechanism based on the Risley prism pair. Data acquisition rates can go as high as 200kHz for ranges within 200m and 25kHz for ranges exceeding 200m. The scan patterns are created by the rotation of two prisms under independent motor control. The geometry and material properties of the prisms will define the conical field-of-view of the sensor, which can be set up to 120 degrees. Through detailed simulations and analysis of mission profiles, the system can be tailored for applications to rotorcrafts. Examples of scan patterns and control schemes based on these simulations will be provided along with data density predictions versus acquisition time for applicable DVE scenarios. Preliminary 3D data acquired in clear and obscurant conditions will be presented.

Helicopters have many good flight characteristics. Recently, helicopters are used for various situations. However, there are few studies about cooperative control between different scale helicopters. In this paper, an approach to cooperative control between different scale helicopters is proposed to expand the field of application of using helicopters. Leader-following is applied for the control scheme of formation flight to multiple autonomous helicopters. Two helicopters model is discussed as a first step for formation flight of multiple helicopters in this paper. The cooperative control system is based on the weighted pseudo-inverse matrix and null-space. This proposed system is scalable. Therefore the control system can be expanded easily. Additionally, the control input generator is designed with cascade controller. Four control inputs should be generated suitably to control 6-DOF of the helicopter because helicopters are the underactuated robots and there are nonholonomic constrains. The workspace observer are adopted in order to compensate disturbance, parameter uncertainty, interaction of helicopters, influence of the wind, and so on.

The application of a new class of actuators is considered. The actuators under development combine a high energy density smart material, specifically a piezoelectric material, with internal servohydraulic components. Large displacement outputs are produced, while the high force capacity of the stiff smart material is retained, for a net high-energy output. The actuator is considered “power-by-wire” because only electrical power is provided from the vehicle or system controller. A primary motivating application is in unmanned combat air vehicles (UCAVs). The particular actuation needs of these vehicles, in flight control and other utility functions, are described and distilled to a set of relevant device requirements. Other potential applications, such as flight motion simulation, are also highlighted. The new actuation architecture offers specific advantages over centralized hydraulic systems and has capabilities not present in electromechanical actuators (EMAs). The main advantage over centralized hydraulic systems is the elimination of the need for hydraulic lines. Compared to motor-driven ball screw type EMAs, the new actuators offer higher frequency response, and a larger peak-to-average output. A laboratory test facility designed to represent the loading experienced by a UCAV control surface is described. Key steps necessary to flight qualify the actuator are introduced.

The wing on the NASA F-111 transonic aircraft technology airplane was modified to provide flexible leading and trailing edge flaps. This wing is known as the mission adaptive wing (MAW) because aerodynamic efficiency can be maintained at all speeds. Unlike a conventional wing, the MAW has no spoilers, external flap hinges, or fairings to break the smooth contour. The leading edge flaps and three-segment trailing edge flaps are controlled by a redundant fly-by-wire control system that features a dual digital primary system architecture providing roll and symmetric commands to the MAW control surfaces. A segregated analog backup system is provided in the event of a primary system failure. This paper discusses the design, development, testing, qualification, and flight test experience of the MAW primary and backup flight control systems.