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The Dragonfly mission has the scope of studying the surface of Titan, Saturn's largest moon, using a rotorcraft lander. A mission-critical event in the timeline involves Dragonfly's initial flight as it descends into Titan's atmosphere. The preparation for this initial flight involves releasing the heat shield and utilizing the lander's rotors to damp motion while still attached to the parachute-supported aeroshell to ensure a smooth release. The aerodynamics of this event are studied in Titan conditions using high-fidelity computational fluid dynamics (CFD). The CFD model is benchmarked using available experimental measurements relevant to subcomponents of the event and include: (1) an open backshell, (2) bluff bodies similar to the Dragonfly fuselage, and (3) a rotor-body interaction from a helicopter rotor. The CFD results correlate well with measurements to provide confidence in the present studies focused on Titan. The novel studies presented in this work focus on the rotor-based control of the lander–backshell combination and its ability to mitigate spinning about the main parachute support line. This effort finds that the relatively large fuselage combined with near-fuselage rotors (driven by aeroshell space limitations) drives an unexpected aerodynamic interaction between the rotor and fuselage. Specifically, we observe low-pressure zones on the fuselage adjacent to the rotor that create a "suction" force. This first-order force creates an unexpected aerodynamic interaction that demands careful attention. In this paper, this force is documented and solutions are proposed to mitigate undesirable control character. • Aerodynamics of the Dragonfly lander interacting with a parachute-supported backshell. • Efforts supported using high-fidelity numerical models backed with a series of experimental benchmarks. • Elucidation of rotors inducing a strong force on the lander that need special consideration. • Proposed solutions highlighted by careful consideration of rotor orientation. [ABSTRACT FROM AUTHOR]

Traditional air mobility capability development consists of defining a set of requirements and subsequently designing an aircraft that satisfies these requirements. Once designed, the overall structure or topology of the aircraft is fixed. That is, the aircraft has certain fixed physical dimensions, number and sizes of rotors, engines, etc. The topology cannot be changed to accomplish missions outside of its original design envelope, and if designed for a robust mission spectrum, the aircraft may be inefficient during routine missions that do not require peak performance. The work reported here takes a different approach where a core set of components (fuselage, rotor, power plant, etc) defines the fundamental modules of the air mobility system. The modules were developed using mature and readily available technologies and designed to be quickly connected to other modules to assemble different aircraft configurations with widely varying performance characteristics. The aircraft can be reconfigured between missions to best satisfy mission requirements. Thus, the envelope of achievable performance characteristics is significantly expanded compared to a single aircraft. This concept is explored for autonomous air cargo delivery missions targeted for nominal payloads and ranges around 500 pounds and 500 nautical miles with maximum payloads up to 3500 pounds across shorter distances. Results indicate that for mission spectrums with large differences between nominal and maximum payloads and skewed towards smaller payloads, the operational cost per lb-km is significantly reduced using a modular and reconfigurable air mobility system since cargo delivery requirements can be better matched to an air vehicle system. [ABSTRACT FROM AUTHOR]

Small unmanned aerial rotorcrafts (SUARs) have garnered widespread attention in recent years due to their applications in various fields such as aerial photography, surveillance, search and rescue, and agriculture, with particular emphasis on the control issues associated with these rotorcrafts. Central to the discourse surrounding SUARs is the nuanced investigation of control methodologies pertinent to their operational dynamics. This paper contributes to this discourse by presenting a hierarchical-based self-triggered controller design approach for SUARs, aiming to reduce communication resource usage while ensuring system stability. The proposed hierarchical framework intends to optimize the self-triggered control structure wherein the upper layer serves as the trigger layer for triggering interval calculation and the lower layer as the control layer for determining control inputs. Within this hierarchical framework, a cyclic iterative solution method is introduced. After setting an initial inter-execution interval at the upper layer, the lower layer controller optimizes the specified performance index for the given triggering interval time domain to calculate the optimal control input, which is then fed back to the upper layer trigger for computing the next maximum triggering interval while ensuring system stability. Simulation results on the SUARs model validate the effectiveness and superiority of this method. [ABSTRACT FROM AUTHOR]

ABSTRACT The desire to use additively manufactured (AM) parts to ensure the availability of military aircraft, and to build limited‐life unmanned aerial vehicles (drones), coupled with the United States Air Force (USAF) approach to the airworthiness certification of AM parts has focused attention on durability analysis/assessment, and hence on the growth of small cracks in AM parts. Previous studies have shown that laser powder fusion built (LPBF) Scalmalloy® has: i) A yield stress and an ultimate strength that are greater than that of AA2024‐T3 and comparable to that of AA7075‐T6; ii) A resistance to crack growth that is better than that of AA7075‐T6 and comparable to that of AA2024‐T3. However, since the ability to predict the durability of a part is essential for its airworthiness certification, the present paper illustrates how to perform a linear elastic fracture mechanics (LEFM)‐based durability assessment of Boeing Space, Intelligence and Weapon System (BSIWS) LPBF Scalmalloy®. The durability study includes specimens with both machined surfaces and surfaces left in the as‐built condition. As a result, it would appear that BISWS AM LPBF Scalmalloy® is an ideal candidate for building limited‐life AM replacement parts for fixed and rotary wing aircraft and drones. [ABSTRACT FROM AUTHOR]

Introduction The association between hypothermia, coagulopathy, and acidosis in trauma is well described. Hypothermia mitigation starts in the prehospital setting; however, it is often a secondary focus after other life-saving interventions. The deployed environment further compounds the problem due to prolonged evacuation times in rotary wing aircraft, resource limitations, and competing priorities. This analysis evaluates hypothermia in combat casualties and the relationship to resuscitation strategy with blood products. Methods Using the data from the Department of Defense Joint Trauma Registry from 2003 to 2021, a retrospective analysis was conducted on adult trauma patients. Inclusion criteria was arrival at the first military treatment facility (MTF) hypothermic (<95ºF). Study variables included: mortality, year, demographics, battle vs non-battle injury, mechanism, theater of operation, vitals, and labs. Subgroup analysis was performed on severely injured (15 < ISS < 75) hypothermic trauma patients resuscitated with whole blood (WB) vs only component therapy. Results Of the 69,364 patients included, 908 (1.3%) arrived hypothermic; the vast majority of whom (N  = 847, 93.3%) arrived mildly hypothermic (90–94.9°F). Overall mortality rate was 14.8%. Rates of hypothermia varied by year from 0.7% in 2003 to 3.9% in 2014 (P <0.005). On subgroup analysis, mortality rates were similar between patients resuscitated with WB vs only component therapy; though base deficit values were higher in the WB cohort (−10 vs −6, P  < 0.001). Conclusion Despite nearly 20 years of combat operations, hypothermia continues to be a challenge in military trauma and is associated with a high mortality rate. Mortality was similar between hypothermic trauma patients resuscitated with WB vs component therapy, despite greater physiologic derangements on arrival in patients who received WB. As the military has the potential to conduct missions in environments where the risk of hypothermia is high, further research into hypothermia mitigation techniques and resuscitation strategies in the deployed setting is warranted. [ABSTRACT FROM AUTHOR]

Aiming to address the significant external noise produced by helicopter rotors, an integrated process based on acoustic modes has been developed for predicting the global noise footprint. The proposed method has the advantages of global prediction, analysis, and low computational costs. The key steps involve constructing an acoustic modal database and efficient prediction of noise footprint. First, the global noise model is utilized to map the time-domain sound pressure to amplitudes and coefficients in the acoustic modal domain. Acoustic mode coefficients representing various flight states are systematically cataloged in the database. Subsequently, the coefficients are extracted based on the characteristic parameters of trajectory elements, allowing for the evaluation of noise radiation in the acoustic modal domain. To investigate the impact of descent angle and advance ratio on the noise footprint, a simulation study is conducted in the forward and approach flight with an AS350 helicopter. Results indicated that, for a region with 1800 observers, the computational time of the developed approach requires only one-fifth of the time compared to the traditional Rotorcraft Noise Model method. This remarkable reduction in computation time, along with the global predictive capability, supports the practicality of efficient noise footprint assessment in both military and civilian contexts. [ABSTRACT FROM AUTHOR]

Purpose: With the advancements in electric vertical take-off and landing (eVTOL) aircraft technology such as batteries, mechanisms, motors, configurations and so on, designers and engineers are encouraged to create unique and unconventional configurations of eVTOL aircraft to provide better capabilities and higher efficiencies to compete in the market. The box fan-in-split-wing tiltrotor eVTOL aircraft is an innovative design that aims to address the aerodynamic inefficiencies such as propeller effects in cruise and engine mounts drag that existed in traditional eVTOL aircraft designs such as vectored thrust, rotorcraft, lift + cruise and multi-copter configurations. This paper aims to propose a multi-disciplinary design process to conceptually design the box fan-in-split-wing Tiltrotor eVTOL aircraft. Design/methodology/approach: An unconventional methodology was used to design the UAM aircraft, and the following parameters are considered: capable of vertical take-off and landing, highly aerodynamic with a high lift-to-drag ratio, low Cd0 modern and appealing, rechargeable or battery swappable and feature to minimise or negate propeller drag. A heavy emphasis on improving performance and weight based on aerodynamics was enforced during the conceptual design phase. MAPLA and XFOIL were used to identify the aerodynamic properties of the aircraft. Findings: Upon determining the key parameters and the mission requirements and objectives, a list of possible VTOL configurations was derived from theoretical and existing designs. The fan in the wing/split wing was selected, as it could stow the propellers. A tiltrotor configuration was selected because of its ability to reduce the total number of lift props/motors, reducing powerplant weight and improving aerodynamic efficiency. For the propulsion configuration, a battery–motor configuration with a hexa-rotor layout was chosen because of its ability to complement the planform of the aircraft, providing redundant motors in case of failure and because of its reliability, efficiency and lack of emissions. Coupled with the fan-in-wing / split wing concept, the box wing seamlessly combines all chosen configurations. Originality/value: The box fan-in-split-wing Tiltrotor eVTOL aircraft aims to address the aerodynamic inefficiencies of earlier designs such as propeller effects in cruise and engine mounts drag. The potential benefits of this aircraft, such as increased range, endurance and payload capacity, make it an exciting prospect in the field of Urban Air Mobility. [ABSTRACT FROM AUTHOR]

Multirotor vehicles typically use either complex mechanisms to vary blade pitch angle or speed control to vary thrust and torque to attain stability and maneuverability. Alternatively, thrust and torque can be varied by morphing the blades without the use of complex mechanisms and without varying the rotor speed. This method could be faster to change thrust, more efficient, and it may require fewer mechanical components. This paper presents an experimentally validated aerodynamic model with an empirical static-aeroelastic correction to predict the thrust and torque coefficient of a mechanism-free rotor with Macro-Fiber Composite actuators. A variable-camber piezocomposite rotor prototype is experimentally tested on a static-thrust setup. The experiment is used to identify the so-called unknown coefficients of the model (e.g., related to the shape of the blade induced by the piezoelectric device) and to identify a correction to account for flow-induced deformations. Using the identified model, several parametric analyses are conducted to predict thrust and torque coefficients as a function of rotor speed and the electrical excitation of the piezocomposite actuators. In addition, the model is used for design optimization, leading to several optimal designs based on a limited parameter space. Three different objective functions are used: maximizing thrust, minimizing torque, and maximizing thrust-to-torque ratio. The operational aerodynamic characteristics of these designs are also presented. [ABSTRACT FROM AUTHOR]

The coaxial compound helicopter has been offered as an idea for a high-performance rotorcraft that possesses high-speed qualities that conventional helicopters do not. This work seeks to create a flight dynamics model of a coaxial compound helicopter and investigate its trim, stability, and controllability. Regarding this, this study builds aerodynamic models for each component of a coaxial compound helicopter (coaxial rotor, fuselage, horizontal stabilizer, vertical stabilizer, and propeller). For coaxial rotor configuration, an equivalent flapping model, a reflected model, an inflow model taking the upper and lower rotor's impact into account, and a phased cyclic pitch model are done. A nonlinear coaxial compound helicopter mathematical model is constructed, and the trim across the entire flight envelope is carried out. The nonlinear model is linearized in order to evaluate stability and control derivatives. The trim results exhibit a high degree of concordance with the reference, thereby validating the flight dynamics model. [ABSTRACT FROM AUTHOR]

HeliCAT-DARTS is a high-fidelity rotorcraft dynamics simulator developed for the design and development of rotorcraft for planetary exploration. While initially developed for the life cycle use of the Ingenuity Mars Helicopter mission, the simulator now supports a broad range of rotorcraft configurations and applications. HeliCAT provides a GNC testbed and aerial mobility analysis platform for rotorcraft design, closed-loop flight software development, verification and validation (V&V), and mission operations. This article discusses the design and use of the HeliCAT simulator and results from technology demonstrations and missions. [ABSTRACT FROM AUTHOR]

Advanced nonlinear controllers are a desirable solution to rotorcraft flight control as they can solve the system high nonlinear dynamic behavior. However, conventional nonlinear controllers such as Nonlinear Dynamic Inversion (NDI) controller heavily rely on the availability of accurate model knowledge and this can be problematic for rotorcraft. Therefore, incremental control theory can solve the modelling errors sensitivity by relying on the information obtained from the sensors instead. The paper applied the Incremental Nonlinear Dynamic Inversion (INDI) controller to rotorcraft case. It will be demonstrated that, for rotorcraft, the incremental nonlinear controllers depend on the delays introduced in the controller by the rotor dynamics. To correct this behaviour, residualization and synchronization methods need to be applied accordingly in order to remove the effects of rotor flapping (disctilt) dynamics from the controller. These particularities of rotorcraft in dealing with advanced controllers shows that incremental nonlinear controllers can have relatively small stability robustness margin and careful controller design is needed in order to account properly for rotorcraft time delays and unmodelled dynamics.

Copyright of Telecommunication Engineering is the property of Telecommunication Engineering and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

A coaxial helicopter with a maximum take-off weight of 600 kg was converted to an unmanned aerial vehicle. A minimally invasive robotic actuator system was developed, which can be retrofitted onto the copilot seat of the rotorcraft in a short period of time to enable automatic flight. The automatic flight control robot includes electromechanical actuators, which are connected to the cockpit inceptors and control the helicopter. Most of the sensors and avionic components were integrated into the modular robotic system for faster integration into the rotorcraft. The mechanical design of the control system, the development of the robot control software, and the control system architecture are described in this paper. Furthermore, the multi-body simulation of the robotic system and the estimation of the linear low-order actuator models from hover-frame flight test data are discussed. The developed technologies in this study are not specific to a coaxial helicopter and can be applied to the conversion of any crewed flight vehicle with mechanical controls to unmanned or fly-by-wire. This agile development of a full-size flying test-bed can accelerate the testing of advanced flight control laws, as well as advanced air mobility-related functions. [ABSTRACT FROM AUTHOR]

This article presents the development and implementation of active noise abatement flight control laws utilizing redundant control allocation in rotorcraft. The study starts by finding a periodic equilibrium of coupled rotorcraft flight dynamics and acoustics using a modified harmonic balance trim solution method. Next, the rotorcraft nonlinear time-periodic dynamics are linearized about its periodic equilibrium and transformed into an equivalent higher-order linear time-invariant system in harmonic decomposition form. Composite aeroacoustic measures are formulated as outputs of this system. To enable faster runtime and tractable control design, a reduced 8-state model representative of the coupled rigid-body dynamics and acoustics is obtained through residualization. An explicit model following control law with pseudoinverse allocation to redundant control surfaces--an active horizontal stabilizer in this study--is synthesized using the 8-state model. The article further assesses the impact of noise alleviation control on handling qualities of conventional and compound rotorcraft in forward flight. The results demonstrate that redundant control allocation is an effective method for active reduction of unsteady rotor noise, with increasing effectiveness during aggressive maneuvers. Future-generation rotorcraft, particularly future vertical lift configurations featuring high control redundancy and aggressive maneuvering capabilities, can significantly benefit from noise abatement through redundant control allocation. Furthermore, the article highlights that while implementing noise abatement control in conventional helicopters can degrade handling qualities, the use of redundant control allocation in compound configurations offers noise reduction without compromising handling qualities. [ABSTRACT FROM AUTHOR]

Copyright of Gazi Üniversitesi Fen Bilimleri Dergisi Part C: Tasarım ve Teknoloji is the property of Gazi University and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Rotorcrafts experience severe vibrations during operation. To ensure the safety of rotorcrafts, it is necessary to perform anomaly detection to detect small-scale structural faults in major components. To accurately detect small-scale faults before they grow to a fatal size, HR (high sampling rate) vibration data are required. However, to increase the efficiency of data storage media, only LR (low sampling rate) vibration data are generally collected during actual flight operation. Anomaly detection using only LR data can detect faults above a certain size, but may fail to detect small-scale faults. To address this problem, we propose an anomaly detection technique using the SR3 (Super-Resolution via Repeated Refinement) algorithm to upscale LR data to HR data, and then applying the LSTM-AE model. This technique is validated for two datasets (drone arm data, CWRU bearing data). First, the necessity for HR data is illustrated by showing that anomaly detection using LR data fails, and the upscaling performance of the SR3 algorithm is validated in the frequency and time domain. Finally, the anomaly detection for a structural fault diagnosis is performed for the upscaled data and the HR data using the LSTM-AE model. The quantitative evaluation of the Min–Max normalized reconstruction error distribution is performed through the MSE (Mean Square Error) value of the anomaly detection results. As a result, it is confirmed that the anomaly detection using upscaled test data can be performed as successfully as the anomaly detection using HR test data for both datasets by the proposed technique. [ABSTRACT FROM AUTHOR]

Rotorcrafts are very complex systems that require a huge systems engineering effort to design, implement and integrate. A successful aircraft design is a matter of good integration between engineering disciplines and suppliers just as much as it is about finding a good technical solution to the customers' expectations. This issue is well understood within the industry and competent authorities. Companies now face the challenge of transitioning from integrating complex systems to make an aircraft (as per ARP4754) to engineering and collating a complex aircraft system as per ARP4754A: in other words, the focus shifted to a more holistic view to the aircraft development, a game changer in all respects. Leonardo Helicopters is tailoring its internal processes to reflect this change and challenge. While the ideal process can be defined today, the transition takes time and a significant change in culture and organization needs to take place. To support the transition, the authors have developed a hybrid approach to the aircraft architecture and system requirements definition process. This new approach leverages existing expertise at a system level to facilitate the integration between systems and the subsequent migration to bridge the gap with the aircraft engineering activities required by the ARP4754A. [ABSTRACT FROM AUTHOR]

Due to their complex aerodynamics, helicopters may enter different dangerous aerodynamic conditions under certain adverse circumstances. In this paper, we examine one such phenomenon—the Vortex Ring State (VRS). We present a simulation of the formation and evolution of a vortex ring around a helicopter's main rotor. The calculations were carried out by solving Navier–Stokes equations using the Ansys CFX code. The simulations modeled a real helicopter using the rotor wing concept, assuming that only the main rotor blade's geometry was modeled. A sensitivity study assessed the impact of the calculation domain and mesh size on main rotor thrust and required moment parameters. Simulations were conducted to determine the VRS region by observing the transition of the helicopter from a level flight, with the main rotor blades held at a fixed pitch position, to a gradual increase in vertical descent. The VRS region was compared with experimental results obtained from other authors, revealing sufficient coincidences. The main characteristics of the identified region were then described. [ABSTRACT FROM AUTHOR]

Three-dimensional (3-D) solid finite element analysis (FEA) is used to model and study coaxial helicopter rotors. The 3-D FEA is coupled with a lifting line aerodynamic model with free wake to capture rotor-rotor interactions. Two open-access models are developed: one is the Metaltail (a hingeless coaxial rotorcraft), and the other is the coaxial rotor built from articulated UH-60A-like rotors. The former is the main focus of this work and is developed as an example case for the U.S. Army/DoD rotorcraft simulation software CREATETM-AV Helios, while the latter is merely for a regression test. The analysis is performed on a low-speed transition flight for which qualitative data are available for the Sikorsky S-97 Raider aircraft for comparison. Predictions of performance, airloads, vibratory loads, and 3-D stresses are discussed. In the absence of available geometry or property data of the S-97, this study is primarily meant to serve as a capability demonstration of 3-D rotor structural dynamic modeling for such rotors. [ABSTRACT FROM AUTHOR]

In this study, an Octorotor UAV capable of carrying 10 desi/2 kg cargo was designed. It is an eight-rotor rotary wing aircraft and its body consists of eight arms connected to each other by two moving mechanisms. When requested in emergency situations, the fastest intervention can be provided by aircrafts. The auxiliary battery and generator system carried by the aircraft is to provide enough energy to reach the nearest electric charging station for the stranded vehicles. In cases such as disasters, the rapid delivery of health supplies can also be provided by this method. The aircraft designed for this purpose has a rotary wing and a feature that can change its body structure according to different load and weather conditions. The shape change improves its performance compared to conventional aircraft and increases its ability to fulfil its mission. For this purpose, the design of a shape-changing aircraft was realized. [ABSTRACT FROM AUTHOR]

The capability for solving compressible fluid flows in the rotating frame of reference is added to an existed open-source CFD solver, namely, HiSA solver. The new implementation is explained and validated using the experimental data of the Sikorsky S-76 rotor. A comparison is presented between the moving mesh results obtained from the original HiSA code and the single rotating frame results achieved through the new implementation. The comparison includes an analysis of torque and thrust values, as well as computational costs. The results imply that, for evaluating the performance of an isolated rotor or for shape optimization purposes at the transonic regime, the single rotating frame method, like the one introduced in the current work, can provide accurate results within an acceptable computational budget. Furthermore, the results show that, at least 25 revolutions are required for the transient analysis to reach an acceptable steady-state converged solution like the one obtained by the single rotating frame method. [ABSTRACT FROM AUTHOR]

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

In rotorcraft, the blade-vortex interaction (BVI) has been recognized as one of the main sources of noise and vibrations, affecting its aerodynamics performance and stability. Active and passive flow control techniques, for the reduction of BVI, have been investigated. The passive flow control techniques have proved to be inefficient due to the complex aerodynamics of rotorcraft. Thus, active flow control techniques has become of particular interest. In the present research we propose a novel approach, for the reduction of rotorcraft's noise and vibrations due to BVI. The main idea of this approach is to inject air at the leading-edge of the blade to alter the characteristics (strength and core size) of the incoming vortex and thus, minimize the noise and vibrations associated with BVI. Computational studies are carried out using the large-eddy simulation (LES) with the dynamic Smagorisky sub-grid scale model. The computations are carried out for a flow of Reynolds number R e = 1.3 × 10 6 , based on the chord of NACA0012 airfoil and free-stream velocity. The studies show that injecting air at the leading edge of the blade, the influence of blade-vortex interaction on the aerodynamic coefficients is significantly reduced. Higher reduction of BVI effect, on the aerodynamic coefficients, was achieved with the increase of the speed of the synthetic jet. The reduction of the BVI, also, leads to the reduction of the vibrations. [ABSTRACT FROM AUTHOR]

Comprehensive rotorcraft simulation codes are the workhorses for designing and simulating helicopters and their rotors under steady and unsteady operating conditions. These codes are also used to predict helicopters' limits as they approach rotor stall conditions. This paper focuses on the prediction of maximum rotor thrust when hovering (due to stall limits) and the thrust and power characteristics when the collective control angle is further increased. The aerodynamic factors that may significantly affect the results are as follows: steady vs. unsteady aerodynamics, steady vs. dynamic stall, blade tip losses, curvature flow, yaw angle, inflow model, and blade-vortex interaction. The inflow model and tip losses are found to be the most important factors. For real-world applications vortex-based inflow models are considered the best choice, as they reflect the blade circulation distribution within the inflow distribution. Because the focus is on the impact of aerodynamic modeling on rotor stall, the blade design and its flexibility are intentionally not considered. [ABSTRACT FROM AUTHOR]

This article demonstrates the linearization of the coupled rotorcraft flight dynamics and aeroacoustics to provide realtime acoustic predictions in generalized maneuvering flight. To demonstrate the methodology, the study makes use of a nonlinear simulation model of a generic utility helicopter (PSUHeloSim) that is coupled with an aeroacoustic solver based on a marching cubes algorithm. A periodic equilibrium of the coupled rotorcraft flight dynamics and acoustics is first found at a desired flight condition using a modified harmonic balance trim solution method. Next, the nonlinear time-periodic dynamics are linearized about that periodic equilibrium and transformed into an equivalent higher order linear time-invariant system in harmonic decomposition form. Composite aeroacoustic measures are included as an output of this system. To speed up runtime and make control design tractable, the order of these harmonic decomposition models is reduced via residualization to an 8-state model where the states are representative of the rigid-body dynamics of the aircraft. This 8-state model is shown to provide accurate acoustic response predictions for small-amplitude pilot inputs and to abate runtime by a factor of approximately 104, thus enabling acoustic predictions in generalized maneuvering flight that are significantly faster than real time. The 8-state model is subsequently used to demonstrate the use of linear system tools for the dynamic analysis of the coupled rotorcraft flight dynamics and acoustics. [ABSTRACT FROM AUTHOR]

Advanced nonlinear controllers are a desirable solution to rotorcraft flight control as they can solve the system high nonlinear dynamic behavior. However, conventional nonlinear controllers such as Nonlinear Dynamic Inversion (NDI) controller heavily rely on the availability of accurate model knowledge and this can be problematic for rotorcraft. Therefore, incremental control theory can solve the modelling errors sensitivity by relying on the information obtained from the sensors instead. The paper applied the Incremental Nonlinear Dynamic Inversion (INDI) controller to rotorcraft case. It will be demonstrated that, for rotorcraft, the incremental nonlinear controllers depend on the delays introduced in the controller by the rotor dynamics. To correct this behaviour, residualization and synchronization methods need to be applied accordingly in order to remove the effects of rotor flapping (disctilt) dynamics from the controller. These particularities of rotorcraft in dealing with advanced controllers shows that incremental nonlinear controllers can have relatively small stability robustness margin and careful controller design is needed in order to account properly for rotorcraft time delays and unmodelled dynamics. [ABSTRACT FROM AUTHOR]

In rotary wing aircraft, composite lug structures are perforated structures that connect the rotor and blades. They play a significant structural role in rotary wing aircraft. In this study the effect of lay-up parameters on mechanical strength of a thick composite part that represents a blade root. This part is autoclave manufactured polyetherketoneketone matrix carbon fiber reinforced (CF/PEKK) composites, which have recently been researched for usage in defense and aerospace is investigated. The chosen material carbon-fiber-reinforced polyetherketoneketone composites (CF-PEKK) have an excellent mechanical, physical, thermal performance. Aerospace sector has a special interest on this material due to material's low density and versatility. In this study, static and fatigue performance of rotorcraft lug part were investigated by applying non-destructive-test (NDT) methods and the mechanical strength values were discussed according to the experimental results. It has been observed that the component parts with ∓45° layup has the highest mechanical strength. As a result of the static and fatigue tests, it is seen that satisfactorily performance in the view of the both static and dynamic loading states. [ABSTRACT FROM AUTHOR]

The control problem for the multivariable and nonlinear dynamics of unmanned rotorcrafts is treated with the use of a flatness-based control approach which is implemented in successive loops. The state-space model of 6DOF autonomous quadrotors is separated into two subsystems, which are connected between them in cascading loops. Each one of these subsystems can be viewed independently as a differentially flat system and control about it can be performed with inversion of its dynamics as in the case of input-output linearized flat systems. The state variables of the second subsystem become virtual control inputs for the first subsystem. In turn, exogenous control inputs are applied to the second subsystem. The whole control method is implemented in two successive loops and its global stability properties are also proven through Lyapunov stability analysis. The validity of the control method is further confirmed through simulation experiments showing precise tracking of 3D flight paths by the 6DOF quadrotor. [ABSTRACT FROM AUTHOR]

Tail-sitters aim to combine the advantages of fixed-wing aircraft and rotorcraft but require a robust and fast stabilization strategy to perform vertical maneuvers and transitions to and from aerodynamic flight. The research conducted in this work explores different nonlinear control solutions for the problem of stabilizing a tail-sitter when hovering. For this purpose, the first controller is an existing strategy for tail-sitter control obtained from the literature, the second is an application of Nonlinear Dynamic Inversion (NDI), and the last one is its incremental version, INDI. These controllers were implemented and tuned in a simulation in order to stabilize a model of the tail-sitter, complemented by estimation methods that allow the feedback of the necessary variables. These estimators and controllers were then implemented in a microcontroller and validated in a Hardware-in-the-Loop (HITL) scenario with simple maneuvers in vertical flight. Lastly, the developed control solutions were used to stabilize the aircraft in experimental flight while being monitored by a motion capture system. The experimental results allow the validation of the model of the X-Vert and provide a comparison of the performance of the different control solutions, where the INDI presents itself as a robust control strategy with accurate tracking capabilities and less actuator demand. [ABSTRACT FROM AUTHOR]

On the basis of a hardware-in-the-loop prototype test bench of an integrated diagnostic system for assessing the technical condition of the main functional systems and units of the rotary-wing aircraft, the authors present a promising scheme for the use of artificial neural networks and form a methodological approach to the introduction of neural network technologies to assess the lifetime of each component of the performance monitoring system determining the flight characteristics of the rotorcraft. [ABSTRACT FROM AUTHOR]