Browse Topic: Sensors and actuators
Rotor-rotor and rotor-boundary aerodynamic interactions of a quadrotor system without a fuselage in ground effect and ceiling effect for varying rotor-boundary distances and hub spacings were investigated. A GPU-accelerated Lattice-Boltzmann Method (LBM) was coupled to new unsteady actuator disk method (ADM) and actuator slice method (ASM) based rotor models for this purpose. Validation was conducted against experiments for both performance and particle image velocimetry flow field data. The trends in thrust and power were accurately predicted by both actuator methods, with high computational efficiency. Interactional flow physics were resolved, causing the consistent performance benefits very close to the ground, the performance penalties caused by the fountain flow effect between rotors occurring over a limited range of ground distances, and the persistent performance augmentation in ceiling effect. The ASM rotor model, with its individual blade representation, was found to predict rotor-rotor interactions more accurately than the ADM. However, it generally overpredicted aerodynamic loads, which was attributed to the Gaussian regularization applied along the blades. The results indicate that the GPU-accelerated LBM coupled with the introduced actuator methods is capable of computationally efficient prediction of multirotor aerodynamic performance and flow fields in ground and ceiling effects.
This paper presents a robust and adaptable control system for tilt-wing aircraft, developed by Dufour Aerospace. The transitional tilt-wing aircraft, Aero2, combines the vertical takeoff/landing capabilities of helicopters with the high-speed range of fixed-wing aircraft. Addressing the inherent control complexities required to maintain control and stability, the developed system employs established control techniques, utilizing linearization at trim points and gain scheduling based on wing tilt. The architecture comprises a Control Allocation module for optimal actuator management, a Control Augmentation System utilizing an LQRI controller enhanced with a feedforward component for precise attitude tracking, and a Unified Velocity Controller for seamless transitions between ground speed tracking in hover and airspeed tracking in cruise. Special challenges unique to transitioning aircraft to ensure control in all axes, including in windy conditions are addressed with operational strategies and control system modifications, validated through flight testing of a subscale and Aero2 aircraft. The resulting control system demonstrates effective velocity and attitude tracking across the flight envelope, enabling reliable piloted flight and automated missions.
A cooperative flight test campaign between the US Army and NASA was performed. This test sought to characterize the acoustic emissions of a fully instrumented MD530F helicopter using a snapshot array and a phased array of microphones. The snapshot array of microphones aimed to provide even coverage across the surface of a hemisphere, providing an acoustic emission hemisphere in a single 'snapshot' of time. The phased array of microphones was designed to provide enough resolution to determine noise sources from each individual blade as well as perform source separation from main rotor and tail rotor emissions. Test conditions for the characterization effort were chosen using a traditional one-factor-at-a-time approach as well as three design of experiment approaches. Characterization conditions included constant speed level flight, descent, and ascent conditions. Transient maneuver conditions were also captured over the snapshot array. The vehicle instrumentation included measurements of pilot controls, optical sensors to measure blade azimuth locations, pitch link loads, along with strain gauges to measure structural loads, blades and fuselage. This report will provide an overview of the test, document the data acquired, and provide some initial results.
As part of a US-France Project Agreement, the US Army and ONERA are investigating mid-fidelity computational approaches for rotorcraft aerodynamics. The approaches from both groups use immersed boundaries in the place of boundary-layer-resolved meshes and actuator lines in the place of rotor blades. Results are compared between the US Army and ONERA to assess strengths and limitations of the mid-fidelity algorithms. An isolated rotor case is first used to validate and compare actuator line wake structure against a high-fidelity result. Second, a static coaxial hub is used to compare immersed boundary algorithms. In the final application, immersed boundary methods and actuator lines are used together for the Dauphin 365N configuration in forward flight.
Modern aircraft have an established need for a high-performance, open standards solution to interconnect increasing number of digital components including sensors, actuators, controllers, processors, displays and data concentrators. The aircraft can be envisioned as a distributed system requiring highly available, reliable, and deterministic communication network - often termed as digital backbone - for safe operation. This paper introduces a new zonal architecture for aerospace onboard networks using Time-Sensitive Networking (TSN). TSN is an open standard based deterministic Ethernet solution for mission and safety critical networks in aerospace industry that truly meets the Modular Open Standards Approach (MOSA) requirements. This paper also presents a reference implementation of the proposed digital backbone architecture using commercial-off-the-shelf hardware from multiple vendors. Experimental data from laboratory evaluation shows stability, performance, and reliability that meets or exceeds the needs of aerospace use cases. The proposed next generation digital backbone provides significant size, weight, and power savings as well as enables hardware and software modularity using open standards. A specific use case of such a digital backbone is the US Army's Future Vertical Lift (FVL) program, but the proposed architecture is generally applicable to all aircraft networks.
This paper describes a mathematical framework for determining the optimal sensor set location for adequately capturing the sound generated by rotors. The approach leverages the gappy-POD method proposed by Everson and Sirovich [J. Opt. Soc. Am., Vol. 12, 1995, pp. 1657-1664], which first identifies the various mode constituents that make up the first few rotor blade-pass frequency harmonics of the sound-field. The algorithm is developed using a covariance matrix for the POD problem comprising auto- and cross-spectral densities of spatially and temporally resolved sound waves captured by an array of microphones oriented parallel to the axis of a laboratory-scale hovering rotor. Three different forms of the technique are developed and compared. These comprise a homogeneous form and two heterogeneous forms; the heterogeneous forms are referred to as XX-topos and XX-chronos and depends on which term in the error minimization equation is assigned the gappy sensor set. A greedy algorithm is then employed to determine the optimal location of the limited sensor set. The findings are analyzed for different combinations of POD modes and blade-pass frequency harmonics of the sound generated by the hovering rotor.
Wear debris monitoring and analysis is a common practice for the condition assessment of engine and transmission health. Oil debris monitoring (ODM) and electronic chip detectors (ECD) are two common methods deployed for continuous monitoring of oil wetted component health in-flight. This study evaluates the diagnostic performance of the two sensing technologies within controlled rolling element bearing (REB) fault experiments. Progressive visual inspection of the REB spall progression through failure provided a ground truth against which both systems could be compared. Quantifiable metrics of reliability, diagnostic accuracy, provided maintenance interval were defined to create a framework for condition-based maintenance (CBM) program decision making. In summary, it was found that the ODM sensor system provided earlier fault notice, but more so, vastly outperformed the ECD in reliability and avoidance of false positives.
The Human Readiness Level (HRL) scale was applied to a high criticality, disruptive, prototype laser-based aviation sensor to evaluate its readiness for human use. Applying HRL to the laser-based aviation sensor accelerated risk identification and provided enough lead time to influence design. Specifically, a protective sensor cover prototype was implemented to address key safety issues. This success demonstrated that the HRL scale is invaluable and should be applied to other technologies.
This study uses a mid-fidelity, aeromechanics coupled framework using the Lattice-Boltzmann Method (LBM) fluid solver to investigate an experimental coaxial rotor system. Co-rotating and counter-rotating rotor operation scenarios in hover are studied. The rotors are represented as an actuator line in the LBM fluid simulation. Simulation results are compared to experimental data, mid-fidelity and CFD simulation results available in literature. Results indicate that the framework can accurately predict thrust and thrust variations for both upper and lower rotors. Power prediction has deficiencies compared to experimental and CFD results, but is in line with mid-fidelity simulation results in literature. Flow field results are also compared qualitatively with CFD results. Results are sensitive to the actuator line representation of the blade, the inflow sampling, and tip corrections.
ABSTRACT This paper introduces the Shape Adaptive Blades for Rotorcraft Efficiency (SABRE) Horizon 2020 research program and presents initial comprehensive analysis results on the efficacy of adapting blade shapes as a means of reducing rotorcraft power requirements and emissions. The aims of the research program are introduced, followed by discussion of the six different morphing concepts that will be explored. The morphing mechanisms are based on active camber, chord extension, twist, and active tendon morphing technologies. SABRE will explore the use of these concepts individually and in combination, for rotor quasi-steady configuration-type morphing and up to 2/rev actuation of some of the mechanisms, with the objective being to find the best balance between emissions reductions versus complexity and added weight. Initial investigations into the potential power reductions compared to the baseline full-scale BO-105 main rotor achievable with the morphing concepts were performed using Blade Element Momentum Theory and a comprehensive analysis model that was developed using CAMRAD II. The analytical model was validated by full-scale rotor wind tunnel measurements. A combination of active twist and active chord extension achieved up to 11% performance gain in hover. Active camber morphing performance was very sensitive to the combination of deflection, morphing section radial length and placement on the blade, as well as the actuation phasing and blade loading coefficient. The active camber morphing showed power reductions of up to 5.5% in hover and 5% at an advance ratio of 0.313 with a 2/rev actuation, while the active tendon concept showed the capability to change the dynamic response of the rotor blade.
ABSTRACT A flight simulation model for the UH-60 Black Hawk based on Sikorsky's GenHel model is modified to simulate a locked failure of a main rotor swashplate servo actuator and is compensated by using the stabilator as a redundant control effector. Steady state trim analysis is performed to demonstrate feasibility of trimmed flight in various conditions with different locked servo actuator positions for the forward, aft, and lateral actuators. A model-following, linear dynamic inversion controller is implemented and modified to account for locked actuator position. Post-failure, the control mixing and feed-forward control coupling terms are reconfigured to partially reallocate the control authority in the longitudinal axis from the main rotor longitudinal cyclic to a symmetric deflection of the stabilator. This is done by manipulation of only the control allocation relating pilot stick inputs to servo actuator positions, the feedback control gains and mechanical rigging between servo actuators and rotor pitch controls remain identical to the baseline controller. Flight simulation results demonstrate the ability of this reconfiguration to compensate for locked failure of the forward main rotor swashplate servo actuator, as well as the ability of the aircraft to decelerate from cruise at 120 knots to 50 knots which is less than the published safe rolling landing speed of 60 knots. A similar range of locked positions of the forward and aft actuators is demonstrated to be feasible for aircraft recovery using control of the stabilator. Feasibility of aircraft recovery for locked positions of the lateral servo actuator is also considered.
This study presents a statistical approach for detecting and estimating damage to multicopter propellers through a comprehensive probabilistic model. The methodology is derived from model-based analysis and applied within the time series statistical techniques. This research accounts for uncertainties in the estimation process and offers confidence intervals for assessing the extent of damage to the propellers. The framework employs functionally pooled (FP) models characterized by parameters that depend on damage sizes, proper statistical estimation, and decision-making schemes. The validation and assessment are assessed via a hexacopter flying in circles with a constant velocity and altitude under turbulence. The damage size ranges from healthy to 10 mm. The method achieves fast damage detection and precise magnitude estimation based on a segment of a single measured signal obtained from aircraft sensors during flight.
This study models flow around isolated and side-by-side three-bladed propellers in (IGE) and out of ground effect (OGE) using actuator-based techniques of varying fidelity. Actuator techniques model propellers using momentum sources distributed over the disk in actuator disk method (ADM) or distributed over moving lines in actuator line method (ALM) to reduce computational cost compared to blade-resolved DDES simulations. The lowest fidelity ADM method is observed to reasonably predict thrust with the use of a tip loss model to control runaway thrust at the tip while not resolving flow features such as blade-bound vortices and helical tip vortices at a fraction of the cost of BR-DDES (1/100). The coarser ALM model resolves these features but still requires a tip loss model to control runaway thrust at 1/10th the cost of BR-DDES. Finally, the finer ALM model used in this study accurately captures blade-related features and further predicts the tip loss trend from first principles at 1/3rd the cost of BRDDES. Demonstrating the efficacy of these techniques for a commonly encountered flow scenario - side-by-side rotors at 2.5R hub separation are simulated where turbulent fountaining flow is observed between the rotors which eliminates the thrust increase normally seen IGE, a feature captured by both ADM and ALM techniques. However, only ALM captures the impulsive 3-per-rev thrust loading seen in BR-DDES.
Full flight regime trim strategies are examined for a Lift+Cruise eVTOL aircraft. Control laws are designed for hover, transition, and cruise conditions to satisfy standard flying-qualities requirements based on the characteristic behavior of the vehicle (rotorcraft versus fixed wing) while ensuring realistic motor limits (peak and continuous) are satisfied. CONDUIT® is used to optimize control laws to minimize actuator activity while meeting flying-qualities constraints. Variable-RPM control is shown to be sufficient to satisfy Level 1 flying-qualities requirements in hover and low-speed flight where control surfaces have inadequate control authority. Time domain simulations are presented to verify controller performance and ensure actuator limits are not violated while following step commands. The aircraft is able to follow commands well in all axes and flight regimes. Transition through the full flight regime (hover to cruise) is simulated using a stitched model.
The paper deals with the status of development and qualification/certification of electromechanical actuation for Helicopters and VTOL applications with the focus on aspects relevant to the Fault-Tolerance. In particular a linear Electromechanical Actuator (EMA) architecture is presented, derived from a fault tolerant ballscrew-based differential (speed-summing arrangement) actuation system patented by UMBRAGROUP S.p.A. The focus is on safety-critical and high reliability/availability requirements for electromechanical actuation certification. The main characteristic is the use of two independent mechanical actuation channels in the same envelope driven by independent Motor Control Electronics (MCEs). At the state of the art, the presented fault-tolerant architecture is under development in flight-critical swashplate application for eVTOL platform and under feasibility study in flight-critical swashplate application for CS27 platform.
The Adaptive Digital Automated Pilotage Technology (ADAPTTM) flight control software package aims to take advantage of redundant controls to improve safety, survivability, and performance for advanced rotorcraft. Vehicle Maneuver Optimization (VMO) is one component of the ADAPTTM architecture intended to increase maneuverability. VMO uses feedforward actuation within the control null space of over-actuated aircraft to minimize power required during quasi-steady maneuvers. In this study, the system is applied to a generic tiltrotor aircraft and evaluated in piloted simulations at the Penn State Rotorcraft Simulator. In this application, VMO uses flap deployment and nacelle tilt to reduce power required in turn maneuvers. Piloted simulation results show that the system effectively reduces power required during Break Turn and Maximum Performance Turn Mission Task Elements (MTE), while handling qualities are equivalent to the baseline controller without VMO. The system was also tested for a terrain flight mission scenario. Pilot comments indicated better handling with VMO in the aggressive maneuvering phases of the flight.
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This paper reports on the integration, test and evaluation of a Degraded Visual Environment (DVE) system installed on an Airbus H145 (BK117 D-2) civil certified helicopter. The DVE system consists of a LiDAR sensor, an EVS camera and a head-tracked helmet mounted display system (HMD) integrated into the onboard HELIONIX® digital avionics suite. The DVE system combines sensor enhanced and synthetic elements of the external scene and provides an accurate representation of the real world for visual reference and safe manoeuvring in DVE. All systems were prototypically integrated into the H145 demonstrator in a serial-like manner, allowing for a potential serialization of the system. Extensive flight trials were conducted focusing on military as well as on civil HEMS missions and were used to verify the intended function and evaluate installed DVE system performance. The activities described herein are partially performed in the frame of a research project supported by the German Federal Office of Bundeswehr Equipment, Information Technology and In-Service Support (BAAINBw). Between November 2018 and March 2019 the system was successfully deployed in both ground and flight tests.
Computations were performed to assess the effect of fluidically-oscillating jets on a ROBIN-mod7 helicopter fuselage. The simulations utilize previously experimentally validated methodologies that rely on a new boundary condition formulation at the actuator throats, based on phase-averaged flow variables, which obviates the need to resolve the internal cavities simultaneously with the outer flow. Predictions of the base flow past the helicopter fuselage were validated against experimental and computational data available in the literature. The fluidic oscillator characteristics were then evaluated at different scales and pressure ratios, and invariant quantities were identified. In the flow control evaluation, flow separation was significantly reduced and, in some cases, suppressed. However, drag reduction was not obtained, indicating the sensitivity of the actuation location and operating conditions to the vehicle design and flight orientation.
In the course of operation, rotorcraft structural components will encounter loading cycles that are often associated with excursions in the main rotor and tail rotor torque. Rotorcraft designs anticipate these load excursions via a design usage spectrum that factors a certain number of such load conditions per flight hour. In general, the usage spectrum assumes worst-case usage regardless of how the helicopter is actually flown. Thus, maintenance intervals are the same regardless of how the helicopter operator uses the aircraft. If a means exists to detect the actual occurrence of these load cycles (per flight hour) and the measured occurrence is less frequent than the assumption in the design usage spectrum, a usage credit could apply. Such a usage credit would reduce maintenance cost by allowing longer intervals between inspections and longer intervals between replacement lives for certain life-limited components. Such a methodology is being sought for the Bell 525, which has a tachometer-based tail rotor torque measurement system. Main rotor torque is also available and is calculated from known engine torque, tail rotor torque, and losses. The torque measurement system has almost two thousand hours of flight test results that show it to be highly accurate. A methodology to use the tail rotor torque measurement and main rotor torque calculation to safely substantiate future usage credits and achieve reduced maintenance cost is put forward. Finally, an application for this sensor technology to detect drive system faults is introduced.
Mercer Engineering Research Center (MERC) is supporting Naval Air Systems Command (NAVAIR) in the determination of external airframe loading requirements and test rig design support for an MH-60 full scale fatigue test demonstrator project being conducted in collaboration with the Australian Defence Science and Technology Group (DST Group). The analyses included determination of loads for quasistatic and vibratory flight conditions, sensitivity of the structural response to the loads, displacements at actuators across the MH-60R usage spectrum, and feasibility of driving aircraft vibrations at frequencies lower than those measured in flight - specifically, obtaining vibration levels measured at 17.2 Hz by imposing forces at only 2.15 Hz. The studies also addressed the minimum number and locations of actuators required for static and vibratory loading.
An elastic blade trim model of a coaxial-pusher helicopter with aerodynamic interference between the rotors is described and validated against existing experimental data for coaxial rotor systems and helicopters. With the trim model in place, parametric sweeps of trim controls are performed to examine different allowable control settings in terms of the swashplate actuator positions on a generalized swashplate geometry at 3 different flight speeds representing a low speed, moderate speed, and high speed flight condition. The effective allowable ranges of locked-in-place positions are established for the 3 actuators on each swashplate, and explanations for the relative ranges are discussed. In low speed, differential moment variation between the rotors allows for actuator settings accounting for approximately 30% of the total range. In moderate and high speed these ranges change due to the moment balance between the rotor and aerosurfaces of the vehicle, with the aft actuator on each rotor trimmable over the entire allowable range, whereas the forward and lateral actuator on the two rotors have allowable ranges accounting for 40-60% and 20-25% of the total range, respectively.
Current rotorcraft gas turbine engines typically use titanium alloys and steel for compressor section and single-crystal nickel superalloys for the hot-section turbine stator vanes and rotor blades. However, these material selections are rapidly changing due to increased requirements on power-density and efficiency. Future Army gas turbine engines will be using ceramic matrix composites for many high temperature engine components due to their low density and improved durability in high temperature environment. The gas turbine industry is also actively developing adaptive concept technologies for production and assembly of modular gas turbine engine components with integrated sensing. In order to actively monitor engine components for extended seamless operation and improved reliability, it is essential to have intelligent embedded sensing to monitor the health of critical components in engines. Under this U.S. Army Foreign Technology Assessment Support (FTAS) program funded research project, surface bonded and embedded sensor patches from a U.K.-based company, Epsilon Optics Ltd., were experimentally evaluated to measure temperature responses on typical propulsion component material coupons. The temperature responses from this foreign technology sensor were assessed using a thermomechanical fatigue tester with a built-in furnace to conduct thermal cycling durability experiments. The experimental results obtained from the durability performance of this fiber-optic based embedded sensor are reported in this paper. This sensor technology upon maturation to higher TRL (Technology Readiness Level) levels can greatly reduce the lifecycle cost of future Army gas turbine engines.
The National Research Council of Canada and Université de Sherbrooke performed flight testing of an Actively Stabilized Slung Load on the NRC Bell 206 Research Aircraft. Hover, Attitude Capture, NRC designed Lateral Precision Hover, and Frequency Sweep mission tasks were performed for bare airframe and slung load aircraft configurations. The load mass ratio was 0.12 while the slung load pendulum mode was 1.3 rad/sec at a damping ratio of 0.2 for the 40-pound per active tether saturation load system setting. Time domain response indicated that the load remained controllable with damped and underdamped behaviors. Frequency domain analyses confirmed pilot comments indicating HQR 4 handling qualities ratings for bare airframe and stable slung load behavior. This rating degraded to HQR 5 for task execution with slung load oscillation. Pilot workload was due to lateral cycle input requirements of 2 to 3 inch amplitudes at 1 to 2 Hz frequency. Operationally, the coincidence of pilot inputs with active tether induced airframe and short period modes led to high compensation requirements for lateral axis tasks under study. The complexity of active tether management, an actuator failure, and telemetry faults represented system deficiencies. Comparable bare airframe and slung load configuration task execution results indicate the magneto-rheological actuation system’s potential to improve slung load mission task performance.
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