Browse Topic: Analysis methodologies
The current work presents a methodology to estimate the mission and performance capabilities of a generic rotorcraft configuration, to satisfy the need of evaluating the integration of a full electric powertrain in the aircraft design. To include all the design steps, two different approaches are proposed. For the preliminary phase, the "Analytic Method" is considered, which exploits a purely resistive model. Conversely, a method based on look-up tables called "Table Method" is intended to be used in more advanced phase, when the battery pack is defined. Both approaches are tested by evaluating a reference mission and a hover chart. Finally, a verification of the presented methodology is carried out by comparing the mission results with a commercial software, specialized in the evaluation of the cell discharge when a given power spectrum is provided.
Emerging technologies in the field of electrified propulsion systems offer a promising solution to reduce the dependence on fossil fuels and improve efficiency. However, the design of high-power density electric machines introduces new challenges, including limited passive cooling potential and the issue of the weight of electric motors. To address these challenges, this paper considers analysis and design methods for high torque-to-weight ratio axial flux motors. A magnetic equivalent circuit model coupled with a lumped parameter thermal network is developed for design space exploration and optimization. This inexpensive analytical model predicts the performance of a single-stator dual-rotor axial flux motor based on geometry, loading condition, and slot and pole pair combination. To enable comparisons against real-world data, the optimization study was demonstrated using the hover mission requirements from the Research Aircraft for eVTOL Enabling techNologies (RAVEN) vehicle to minimize the mass of the motor. In tandem with the analytical model, a higher-fidelity finite element model was also developed, and good agreement between predicted power and efficiency was demonstrated across a range of axial flux motor designs. The lightest weight design that satisfied the hover mission requirements was the 12 pole pair 27 slot (12PP 27S) configuration with a fixed weight of 9.28 kg. The analytic model undersized the output power of the electric motor by approximately 9% across a range of slot and pole pair combinations.
NASA's successful demonstration of powered flight on Mars through the Ingenuity Helicopter, as part of the Mars 2020 Perseverance rover mission, has led to the development of next generation Martian rotorcraft. The future of Martian rotorcraft has evolved to include high payload-carrying vehicles to possibly contribute to planetary science missions, which will require improved flight dynamics and rotor aerodynamic performance to fly at nominally high forward flight speeds and at higher flight altitudes. To ensure the feasibility and viability of successful mission performance, it is also critical to mature the structural design for advanced Martian rotorcraft to bridge the gap between the best practices of the spacecraft and aircraft communities. This paper focuses on the structural analysis of a Mars Science Helicopter (MSH) blade using finite element methods. Multiple loading conditions including launch and operational flight were applied to investigate the blade’s structural integrity. The blade’s modal natural frequencies were also analyzed to investigate the blade's dynamic behavior.
The current effort presents novel investigations of rotor-wake–surface interactions for the Dragonfly lander, NASA's rotorcraft lander to explore Titan. The numerical framework couples unsteady RANS with blade-element and virtual disk rotor models and a coupled Lagrangian particle tracking method to examine rotor–ground interactions and brownout. Simulations span a range of complexity, from isolated rotor benchmarks and rotor pairs to full eight-rotor configurations without a fuselage and the eight-rotor configuration with a simplified Dragonfly fuselage. To quantify model fidelity and near-ground shear, blade-resolved simulations of the isolated rotor are performed using Spalart–Allmaras and Reynolds Stress turbulence models with vorticity confinement, demonstrating that virtual blade models under-predict tip-vortex strength and local inflow distortion but reproduce wall shear reasonably well, whereas blade-resolved RSM solutions yield higher peak shear levels relevant to brownout prediction. These findings improve understanding of planetary rotorcraft aeromechanics and sediment transport in ground-effect while supporting ongoing efforts to assess environmental risks for Dragonfly operations and inform multi-rotor VTOL design for terrestrial applications.
RPM-controlled hexacopters offer mechanical simplicity and inherent redundancy, but are unable to re-trim under all failure cases in forward flight. This paper investigates the use of reverse-enabled rotors as a means of expanding the attainable trim envelope and improving fault tolerance in RPM-controlled hexacopters. Isolated rotor experiments are conducted to characterize thrust and torque behavior under forward and reverse rotation, providing validation data for aerodynamic modeling. A blade-element-based model implemented in the Rensselaer Multicopter Analysis Code (RMAC) is then used to perform comprehensive trim analyses for a 1200-lb-class hexacopter in hover and in cruise at the best-range speed of 65 kts. Post-failure trim solutions are evaluated for four configurations, including edge-first and vertex-first orientations with different rotor spin directions. Results show that enabling reverse rotation allows trim recovery for all single-rotor failure cases in cruise, including aft-rotor failures that are not trimmable with conventional RPM-controlled rotors. A systematic comparison of peak rotor torque, peak rotor power, and total aircraft power reveals that failure severity is governed primarily by yaw moment deficits arising from the combined loss of hub torque and aerodynamic drag. Among the configurations examined, the edge-first configuration with a counter-clockwise spinning rotor 1 exhibits the lowest rotor torque and rotor power requirements, post-failure.
This study presents a comprehensive analysis of single-rotor failure tolerance for a classical octocopter configuration, examining both hover and forward flight at the best range speed. Using a state-of-the-art eVTOL comprehensive analysis to retrim the octocopter post-failure, the redistribution of rotor thrust, torque, and power following individual rotor failures was quantified, along with resulting aircraft-level power penalties. In hover, orthogonal rotors to the failed rotor provide primary lift compensation, the opposing rotor operates mostly unchanged, and the four opposite spinning rotors primarily provide pitch/roll moment compensation. This results in a total aircraft level power increase of approximately 10.4%, roughly half that of comparable hexacopters. In forward flight, at best range cruise speed, load redistributions were again calculated for various individual rotor failures. In the worst case, a maximum individual rotor torque increase of 62% and power increase of 108% was observed, while total aircraft power requirements increased between 7-12%. These results demonstrate the fault-tolerant capabilities of octocopters and provide practical guidance for propulsion system sizing, energy management, and failure-case assessment on classical octocopters.
Propeller driven rotors utilize propellers on the main rotor blade to spin the rotor. Past research efforts have highlighted dynamic issues that arise from the rotor-propeller Coriolis interaction. For this paper, a comprehensive multi-body analysis methodology, called Elastic Rotorcraft Analysis (ERA), was applied to various propeller driven rotor datasets. The focus of the modeling effort was on propeller driven rotor twirl phenomenon, which arises from rotor-propeller inertial couplings interacting with rotor blade modes. After describing the phenomenon, the paper is split into two parts: validations and predictions. In Part I of the paper, the ERA propeller driven rotor model was validated using three datasets: (i) a propeller flapping vacuum chamber experiment, (ii) a propeller/rotor loads vacuum chamber experiment, and (iii) a propeller driven rotor hover experiment. The ERA model showed good agreement with the data, and captured the important rotor-propeller Coriolis interaction. In Part II of the paper, predictions for several propeller driven rotor configurations were generated and analyzed. Loads were computed for an isolated propeller and are compared to propeller loads during propeller driven rotor operation. The analysis showed that operating the propeller on the rotor blade introduces significant inertial loads on the propeller. Finally, propeller placement along the main rotor blade span was investigated. The results of the present study agree with earlier research, which showed placing the propeller at the midspan location reduced the electrical power coefficient by nearly half compared to a tip mounted propeller.
As per Committee/Henry E. Harschburger recommendations
This study examines the ability of a large (1200 lb gross weight) hexacopter with collective pitch controlled rotors to tolerate single motor failure. The hexacopter is considered in various orientations, and the vehicle is trimmed with one motor inoperative (OMI). Unlike RPM-controlled hexacopters, which were trimmable but uncontrollable in hover, and were untrimmable in cruise with an aft-rotor failure; with pitch-control the hexacopter is controllable in hover as well as trimmable for failure of any rotor in cruise (including an aft rotor failure). The study examines how pitch controls, and thrust are redistributed amongst the operational rotors, post-failure, for the different hexacopter orientations. For each case, the maximum thrust and torque increases on any individual rotor, and the total power increase, post-failure is examined. It is found that the hardest to trim cases are those where the hub torque and the hub drag induced yaw moment of the failed rotor add, and fault compensation for these cases usually comes at a high cost in terms of torque penalty and power requirement (necessitating use of more powerful motors and batteries). The results from the study indicated that operating the pitch-controlled hexacopter in an edge-first configuration with a clockwise spinning rotor 1 is the best choice from a fault tolerance perspective.
Electric aviation is advancing rapidly, with aircraft from manufacturers like Joby and Archer well on their way to certification, aircraft electrification will continue and begin to apply to larger aircraft. To support larger electrified rotorcraft, rotors will need to grow if disc-loading and hover efficiency are to be maintained. A consequence of this is the need to reduce rotor speed to maintain an acceptable acoustic signature, especially for operation in urban environments. Most current applications utilize radial flux motors, sometimes with a reduction gearbox. Gearboxes can improve overall propulsion system power density by enabling higher motor speeds but are generally not preferred as they introduce additional potential failure modes and maintenance schedules. In this paper a holistic approach is used to understand the trade-offs between rotor and motor and their consequences on propulsion system power density.
The paper describes a method for optimal design of a helicopter tail shaft that considers rotordynamic effects from long shaft assembly. The tail shaft transmits power from the main gearbox (MGB) to the tail rotor of the helicopter and operates at high speeds that may exceed 6000 rpm. While higher speeds allow for weight reduction, they also pose risks associated with supercritical operation, necessitating careful design optimization. The objective of the optimization is to maximize the first three transverse natural frequencies with the constraint of the safety parameter (avoidance of the resonance/critical zone) while minimizing the weight of the system. A Non-Dominated Sorting Genetic Algorithm (NSGA-II) is used to obtain the solution to this multiobjective optimization problem, which involves shaft design variables such as length, outer diameter, and wall thickness. In addition, the optimization framework also incorporates system related design variables, including the stiffness of tail shaft bearing supports, the location of bearings, and coupling characteristics, to comprehensively evaluate and enhance the system performance. A rotordynamics model utilizing Timoshenko beam finite elements is deployed to predict natural frequencies.
The vertical flight industry is on its way to a transformative era, with autonomous technologies set to alter aerial vehicle operations. While it seems certain that fully autonomous helicopters will eventually be deployed for a variety of missions, some high-stakes situations—like medical evacuations (MEDEVAC)—will for the foreseeable future demand human participation in the form of Emergency Medical Care-giving Crew. This study describes the testbed built to run and investigate hypothetical future situations in which a helicopter is autonomously piloted while a human medic with no aviation training, subjected to aviation and medical emergencies, manages patient care onboard. A total of 22 participants, with emergency medical technician certification, nursing or a medical board certification, were invited to run and evaluate the use of AI pilot (AP) in different scenarios of medical evacuation under the following emergencies: medical, empty fuel tank, pressure sensor miscalibration, and engine failure. A comprehensive evaluation of both objective and subjective performance metrics revealed that novice medical professionals could effectively execute medical evacuation operations in conjunction with an AI pilot, even during unforeseen circumstances. The analysis of response times unveiled distinct perspectives on how medics perceive and manage various emergency situations when an AP functions as a collaborative and effective team member.
Gearbox casing cracks in helicopters would be critical impacting the aircraft's reliability and operation safety directly. The Defense Science and Technology Group (DSTG) HUMS2025 gearbox casing failure data set was the unexpected result of a test stand operation. The gearbox undergoes high cycle (> 400 acquisitions) under high torque (100% and 125% nominal torque) conditions. We hypothesized that the any cracking would be due to the planet/ring gear interaction. A condition indicator (CI) would be sensitive to a crack feature and this would be sensitive to change in gearbox torque. This paper explores the development of both a cyclo-stationary based CI (frequency-domain) and a time synchronous average CI (time-domain). The trend shows that proposed methods can help to detect localized defects in gearbox casing at an early stage and trend as the crack propagates before catastrophic failure occurs.
This paper deals with the uncertainty estimation of identified frequency and damping trends of whirl flutter modes, obtained by applying system identification methods on experimental data. In particular, two different identification approaches are considered, namely the free-decay analysis by using Matrix Pencil algorithm and the Data-Driven Stochastic Subspace Identification method (SSI), applied to system response to stochastic input. The two approaches lead to as many uncertainty estimation methodologies, both leveraging the bootstrapping statistical process. A full validation procedure is then set up to assess the accuracy of such methods in correctly quantifying the uncertainty of the estimated statistics. To do so, a wing-rotor state-space linear numerical model is used to simulate system response to both dwell and stochastic inputs. The state space numerical system aims to replicate the ATTILA wing-rotor wind-tunnel model, which falls in the framework of Clean Sky 2 European program to investigate the possible occurrence of whirl-flutter instability in tiltrotor configurations. Hence, one of the proposed methodologies is applied to ATTILA experimental data for both modal identification and uncertainty estimation, and the processed flutter trends are reported in a few conditions of particular interest due to the progressive approaching of whirl-flutter condition.
A robust velocity stability augmentation system was developed for the CoAX 600/2D coaxial-rotor helicopter to enable safe testing of a fly-by-wire system on an optionally piloted variant of the aircraft, developed by Piasecki Aircraft Corporation. The control law design and subsequent stability analysis were based on a validated nonlinear model of the CoAX 600 rotorcraft. A subset of helicopter handling qualities were evaluated through both analytical methods and piloted simulations, conducted with and without the stability augmentation system. Additionally, flight test data contributed to the analysis, albeit to a limited extent.
Electric vertical take-off and landing vehicles are proposed as a viable solution for urban air mobility due to their potential for reducing carbon emissions, noise, and operational costs. However, the shift towards electrified aircraft introduces new thermal management issues due to the excess heat generated by electric motors and power electronics. This heat is challenging to dissipate during the mission, resulting in transient motor temperatures, especially during high-power mission segments. In addition, electrified aircraft also encounter design challenges associated with the fixed weight of electric motors and batteries. To address these challenges, this work presents a multifidelity framework for performing shape optimization of an electric motor subject to performance, geometric, and thermal transient constraints. A preliminary sizing of the electric motor is performed using a low fidelity Fourier series model. Next, the sizing is refined by utilizing a coupled electromagnetic-thermal finite element model of the the motor physics. To demonstrate the framework, a numerical optimization is performed for a hover hold mission with a relevant scale eVTOL vehicle, reducing the structural mass of the motor by 2.7kg while ensuring power and thermal constraints are satisfied.
Finite element analysis of gearbox and transmission assemblies is usually a process that takes several weeks of meshing and geometry cleanup, running the simulation, checking the results, fixing the inevitable mistakes, then rerunning and checking again. This process doesn’t even account for the addition of design changes which make the whole process start over. This paper is meant to explore some of the new meshless simulation tools and see how they can be leveraged to more quickly produce results, and allow for better design changes that are driven by the simulation needs. We will show how leveraging meshless analysis leads to better designs for complex assemblies and parts.
Carbon/epoxy stiffened panels are being increasingly used in transport rotorcraft. The reduced mass density and high stiffness of carbon/epoxy composites can lead to higher levels of vibration relative to comparable metallic structures, which themselves can have vibrations and interior noise high enough to damage the hearing of crew and passengers. The current investigation explores a method to reduce the vibration of carbon/epoxy stiffened panels by introducing thickness tapers known as acoustic black holes (ABHs). The ABH feature is integrated into either the stiffeners or plate of a representative stiffened panel configuration. A finite element (FE) parametric study was used to guide designs that reduce the vibration of the panel without compromising the compressive buckling capability or mass of the panel. FE studies showed that a 30 ply to 12 ply thickness taper longitudinally oriented in the blade stiffener can reduce vibrations and increase compressive buckling capability. Carbon/epoxy panels were manufactured using a low-cost out-of-autoclave material with simple molding. Experimental testing concluded that integrating the ABH into the stiffeners longitudinally helped to reduce the broadband vibration by 5 dB and increase the buckling load (+4.3%) and collapse load (+16.5%) without increasing the mass greatly compared to a traditional baseline design.
Accurate simulation of fluid-structure interactions (FSI) is critical for designing aircraft systems, particularly for applications involving fuel tank sloshing and large deformations. Traditional added mass methods often fail to capture the nonlinear and frequency-dependent behavior of these coupled systems. This study applies the Finite Pointset Method (FPM), a mesh-free computational fluid dynamics (CFD) technique, coupled with an explicit finite element solver, to predict complex FSI phenomena. Validation is performed using benchmark experiments, including a harmonic tank sloshing test and a guided plate ditching scenario, with results demonstrating strong agreement with measured pressures and structural responses. Additional validation on a composite fuel tank drop impact test confirms FPM's ability to model large deformations and rupture under dynamic loading. The findings highlight FPM's robustness and adaptability for aerospace FSI problems, offering a powerful alternative for virtual prototyping and certification workflows where conventional methods are insufficient.
This study investigates the stress concentration and damage tolerance of lug structures, with an application example using the horizontal tail plane lug of a light utility helicopter. Using Finite Element Analysis (FEA), stress distributions around the lug hole were simulated under varying load conditions to understand how different loading angles and magnitudes affect stress concentrations. A machine learning approach was employed to predict stress distributions based on a dataset generated from FEA simulations. Several regression models were tested and Random Forest Regression model yields the best predictive accuracy among the others. The study also incorporates a flaw tolerance analysis using the NASGRO® software to calculate the crack growth characteristics of the horizontal tail plane lug structure under service loading. The results highlight the importance of stress distribution variability and identifying the most critical point of the lug structure under service loading with changing angle and amplitude. This research provides insights into the structural integrity of lug components under dynamic loading, contributing to more reliable flaw tolerance assessments for aerospace applications using machine learning techniques.
Advanced structural analysis methods, known as progressive damage and failure analysis tools, are being developed to predict initiation and propagation of damage under repeated loading based on capturing individual and interacting damage modes. This work develops structural fatigue life prediction capability in state-of-the-art emerging progressive damage failure analysis tool CDMat developed at the University of Texas Arlington Advanced Materials and Structures Lab. While JIntegral, implemented in CDMat, appears as the most objective and rigorous approach to predict delamination growth-based fatigue life of composite structures, the key material properties of the J-Integral fatigue model have not been measured with the adequate accuracy. This work addressees a fundamental challenge of eliminating the established and routine assumptions and developed a methodology to determine the key material properties meeting the material input data requirements for the JIntegral based structural fatigue life prediction models. This work generated input data for fatigue crack growth propagation, including optimized input data parameters for the in-situ cohesive law and modeling as-manufactured specimen conditions. In addition, a methodology to account for and characterize the effects of fiber bridging in static tests is presented. It uses a simple standard unidirectional panel for testing, before a component is manufactured, fabricated from the same batch of prepreg to qualify the effects of fiber bridging. Also, this work attempted to determine a minimum conservative initial crack size to streamline the fatigue crack propagation prediction. Fatigue predictions have been demonstrated on a representative composite skin–hat stiffener sub-component section and compared with tests.
Helicopter rotor blades with several different parts, multiple load paths and/or springs and dampers can be modeled as a multibody system, into which finite element descriptions of flexible bodies can be integrated. When doing so, model order reductions can be necessary for robustness and/or performance reasons. A known drawback of such reductions is that the isolated modes of the particular bodies may not adequately describe their actual deformations in the multibody system. To alleviate this problem, the paper proposes a Craig-Bampton reduction for the flexible bodies. Compared to a standard modal reduction, the additional consideration of static interface modes in the Craig-Bampton approach significantly improves the prediction of eigenfrequencies and mode shapes, as demonstrated for a segmented steel beam with a single load path. Using the same approach, a bearingless rotor blade with multiple load paths is modeled by two beam segments. The model is assessed by code-to-code comparison with a CAMRAD II reference model. The consideration of the tension interface degree of freedom of the inner segment (flexbeam) allows for the accurate prediction of eigenfrequencies at varying rotor speeds, including the effect of centrifugal stiffening – which is not appropriately covered with a standard modal reduction. With the Craig-Bampton method, the mode shapes are plausible and in overall good agreement with the reference model. Despite these benefits, deficiencies are identified that still need to be improved. Discrepancies in the blade’s torsion eigenfrequency prediction are observed, as well as anomalies in the mode shapes that are potentially related to the flexbeam, that is – despite the consideration of interface modes – too kinematically restricted.
This paper presents an experimental and analytical investigation of whirl-flutter stability in tiltrotor aircraft, focusing on the influence of pitch-flap coupling on stability boundaries. Wind-tunnel tests were conducted using the TiltRotor Aeroelastic Stability Testbed (TRAST), a semi-span model designed for test-analysis correlation. This study examines variations in pitch-flap coupling and compares measured frequency and damping trends with predictions from RCAS and CAMRAD II. Results indicate that less pitch-flap coupling increases stability, with both analytical models capturing general trends. The analysis accurately predicts the wing inplane mode stability, but larger deviations are observed in the vertical bending mode, suggesting missing physical effects in the modeling approach. Differences in damping trends at higher speeds indicate that improvements in modeling may be necessary to refine stability predictions. These results provide valuable insights into the capabilities and limitations of current whirl-flutter analysis methods and inform future refinements in tiltrotor aeroelastic modeling.
This study investigates Reynolds number effects on rotor wake vortex development using a hyperbaric rotor facility capable of pressurizing air up to 100 bar. Background-oriented schlieren (BOS) and hot-wire anemometry (HWA) were applied to characterize vortex trajectories, core growth, and circumferential velocity distribution. BOS measurements revealed consistent blade-to-blade trajectory deviations and vortex pairing across all operating conditions, despite that the investigated three-bladed rotor was milled from a single piece of aluminum, ensuring precise manufacturing and a highly symmetric geometry. A statistical scheme was developed to analyze the radial structure of fluctuating tip vortices, which traverse the pointwise fiber-film sensor in a fixed position. With increasing vortex Reynolds number, the tip vortices are more compact with a reduction in core growth. The circulation in the vortices grows with the vortex radial coordinate, and converges at a radial position basically independent of the vortex core size. Observed asymmetries in young vortices at low Reynolds numbers indicate enhanced roll-up dynamics. The results demonstrate the facility’s ability to isolate Reynolds number effects in rotor wake dynamics.
A typical helicopter drive system consists of a multi-stage gearbox with highly loaded dynamic components such as gears, shafts, and bearings, crucial for safe flight and landing. Planetary reduction stages are commonly used in the final reduction stage of rotorcraft main gearboxes due to their ability to handle high torques at high gear ratios within a compact envelope. The planet gear, a critical component in this arrangement, is subjected to significant loads on both flanks of its teeth and must meet stringent weight and assembly requirements, leading to a thin rim design with integrated bearing races. This design makes the planet gear susceptible to relevant reduction of its fatigue life. This paper explores analysis methods to evaluate the damage resistance of the planetary stage assembly, focusing on the planet gear. The study aims to assess the "growth" or "no growth" condition of the planet gear against defined flaw defects. An iterative calculation loop determines the critical length and position of a crack that may lead to full crack propagation and, in worst cases, to system jamming. Initial crack propagation simulations use NASGRO software, with stress fields derived from a non-linear FEM of the planet gear availing of detailed Transmission3D model for teeth meshing forces evaluation. Further additional analysis can involve a dedicated FE model of the crack, iteratively updating its geometry. The impact of crack propagation on the remaining components of the planetary stage assembly is also addressed, considering the unbalanced load conditions caused by stiffness loss in the planet gear. The dissertation object of this paper aims to outline a comprehensive, effective and efficient procedure to determine the maximum allowable defect size for "no growth" condition and the operational hours until failure, providing a robust approach to support the strength substantiation of the involved components.
Bench-level tribological experiments were utilized to evaluate material, coating, and lubricant formulation effects on the loss-of-lubricant survivability of tapered roller end and cone rib contacts. Cone rib and roller end contacts were simulated using a single rotating roller and rotating flat disk. The applied load and rotational speeds of the roller and disk were controlled to simulate representative rotorcraft gearbox bearing operating conditions. The contacts were lubricated for an initial period before the lubricant supply was shut off, and the supply tube was then removed. Tests continued to run, without additional oil, until the measured friction force reached a predetermined cutoff value. Weibull-based statistical analysis was used to compare the loss-of-lubrication runtimes.
Composite materials have long been used in rotorcraft structural applications due to their high strength-to-weight ratio. However, they can also be prone to unique types of defects such as resin pooling, porosity, and delamination which can impact their structural performance. Resin pooling can occur when excess resin accumulates in certain areas of the composite laminate. This paper presents a method of finite element analysis to quantify the effect of the resin lens on the structural strength of a specific laminate. This method was used to successfully validate the structural integrity of several composite torque tubes with resin-rich indications, which were initially rejected at inspection.
Dufour Aerospace designs and manufactures an automated tilt-wing aircraft for critical cargo delivery missions. Emphasizing operational efficiency, the platform integrates path generation and tracking techniques tailored for the unique dynamics of tilt-wing flight and builds upon the existing lower level control. While there exist a myriad of methods for high-level aircraft automation ranging from PID to MPC, they often require a trade-off between complexity and the capability to handle non-linear dynamics of the system they are controlling. Hence, a lightweight, deterministic geometric path generation approach using clothoid-based transitions between three waypoints and a robust SO(3)- based path tracking controller adapted for tilt-wing dynamics are presented. Additionally, a high-level automation framework is introduced that includes failure mode handling for GNSS loss and communication breakdowns. This system ensures mission continuity and operational safety while supporting flexible mission planning. The methods are validated through extensive flight testing on both small and large-scale aircraft. The latter prove the scalability, safety, and reliability of the presented solution for tilt-wing aircraft automation and enhance the aircraft's capabilities in real-world emergency response and complex operational scenarios.
This paper investigates an output-based approach for tiltrotor whirl flutter bifurcation analysis. The approach uses free decay output data for a quantity of interest at various forward speeds to estimate the system's recovery rate to equilibrium while capturing its variation with amplitude. The recovery rate is then extrapolated to predict the bifurcation diagram, which gives the limit-cycle oscillation amplitude for the quantity of interest as a function of the forward speed. The approach is demonstrated using output data from transient simulations of a notional tiltrotor model with polynomial structural nonlinearities. The approach accurately predicts the tiltrotor whirl flutter speed and limitcycle oscillation amplitudes while only requiring two free decays. This approach can facilitate whirl flutter bifurcation analyses of tiltrotor systems exhibiting nonlinear dynamics.
As per certification requirements, for a large rotorcraft that does not meet the Category A requirements, the Height-Velocity (HV) avoid region must be determined in total power failure condition. The development of a digital twin representative of the real rotorcraft behaviour allows to reduce flight testing hours and to increase flight tests safety, especially in such critical conditions, thus decreasing risks and costs. In this work, an extensive simulation activity has been carried out to generate HV charts for a medium twin-engine helicopter in case of loss of both engines. An in-house software that emulates pilot logics has been exploited, coupled with a Flightlab model representative of the rotorcraft and validated against flight data. Manoeuvres performed after a dual engine failure were simulated starting from an all engine operative hover out of ground effect (HOGE) and in ground effect (HIGE) or level flight condition until landing, in a grid of heights and velocities and for different weights and altitudes combinations. Sensitivity analyses on the parameters that affect the manoeuvres the most have been performed and the effects of the assumptions and requirements at touchdown on the HV avoid region investigated.
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