Technical Papers - SAE Mobilus
SAE Technical Papers are written and peer-reviewed by experts in the automotive, aerospace, and commercial vehicle industries and provide the latest advances in technical research and applied technical engineering information.
ABSTRACT Over time, the National Institute of Standards and Technology (NIST) has refined the 4Dimension / Real-time Control System (4D/RCS) architecture for use in Unmanned Ground Vehicles (UGVs). This architecture, when applied to a fully autonomous vehicle designed for missions in urban environments, can greatly assist in the process of saving time and lives by creating a more intelligent vehicle that acts in a safer and more efficient manner. Southwest Research Institute (SwRI®) has undertaken the Southwest Safe Transport Initiative (SSTI) aimed at investigating the development and commercialization of vehicle autonomy as well as vehicle-based telemetry systems to improve active safety systems and autonomy. This paper will discuss the implementation of the 4D/RCS architecture to the SSTI autonomous vehicle, a 2006 Ford Explorer.
Operational shifts in natural frequency constitute a critical phenomenon for structures exposed to multiple excitation sources over a broad frequency range. To investigate this behavior on the helicopter windshield, finite element models were correlated with ground modal test results and assessed using in flight measurement data. In rotorcraft, the large number of operational excitation sources makes it particularly difficult to distinguish structural responses associated with the inherent dynamic characteristics of the windshield from those induced by the aircraft's periodic excitations. To address this challenge, time synchronous averaging was employed to remove the dominant main rotor frequency components from the measured flight data. The residual response was then analyzed using joint time frequency analysis techniques, revealing that the windshield natural frequencies shift with increasing flight speed and the associated aerodynamic pressure variation. For further correlation of the finite element models, in flight impact hammer tests were performed. The resulting responses were processed through operational modal analysis methods to extract the exact natural frequencies more accurately under operational conditions. The validated methodology was subsequently applied to windshields of different thicknesses, demonstrating that windshield thickness has a pronounced influence on the observed frequency shift. These findings indicate that the problem should be treated as a coupled dynamic, static, and aerodynamic interaction. An illustrative mitigation study was also conducted to examine possible interventions for this phenomenon, and the influence of structural modification on the variable dynamic characteristics of the structure was evaluated.
Enterprises that develop complex products or systems often struggle to reuse technology efficiently across their portfolios. This challenge is especially prevalent in aerospace, transportation, energy, and defense industries, where preserving freedom of action is critical. In this context, freedom of action is defined as the ability to avoid vendor lock imposed by integrators or third parties, while enabling competition within clearly defined functional boundaries that establish effective market segments for system components. This paper presents eight best practices for Enterprise Reference Architecture (ERA) development to address this challenge and applies them to aviation functionality spanning both vertical lift and fixed wing platforms. Because complex systems can be modularized in many ways, a consistent set of guiding rules is required to produce an organized set of modules that are reusable across an enterprise portfolio. The best practices presented in this paper are intended to fulfill that role.
This study investigates the aerodynamics and performance of aerial screw rotors in axial climb and off-axis flight for the first time. Additionally, this work highlights comparisons between a hover-efficient aerial screw and a two bladed conventional rotor in various flight states. Using high-fidelity Computational Fluid Dynamics (CFD) analysis, the research identifies the formation of the "da Vinci vortex"—a shape-conforming helical structure to be crucial to thrust generation and performance. Investigations of linear and bilinear variations of the screw pitch and taper reveal nuanced effects on loads on the screw surface and climb efficiency. The effect of the pitch rate is found to dominate over the effect of the taper rate in axial and off-axis climbs. Efficiency in axial flight states is heavily dependent on the pitch rate of the aerial screw. Investigation of all forces on the aerial screw surfaces identifies significant phase differencing in loading between linear and bilinear varying designs, suggesting potential for canceling oscillations in one bladed aerial screws. This study builds on the groundwork for the application of aerial screws in next-generation Vertical Take-Off and Landing (VTOL) vehicles by enhancing the understanding of the unique rotor’s aeromechanics.
Electric vertical takeoff and landing (eVTOL) aircraft have rapidly emerged as a potential promising solution for sustainable and scalable urban air mobility. These vehicles are at a crucial stage of development, with sub-scale and full-scale flight test campaigns already underway. In forward flight, critical structural components of these variable RPM rotary wing aircraft can experience strong time varying aerodynamic loads. These components are lightweight, stiff, and often have very little intrinsic damping. Thus, from a life-cycle perspective, vibratory strains are of significant concern. This work investigates the use of Impact Dampers (IDs) in attenuating the bending and torsional response of a characteristic eVTOL boom. A computational reduced order model of an ID is developed using the Hunt-Crossley nonlinear contact model and linear beam finite elements. Predicted damper performance is compared with prior experimental measurements. Parametric studies are performed by varying key damper design variables, and predicted trends are shown to be consistent with experimentally observed behavior. The IDs attenuated the system structural response at resonance by 60.7% in bending and 52.1% in torsion, for a 5% mass penalty. Building on works presented at the 80th and 81st Vertical Flight Society Annual Forums, this publication is a next step towards establishing IDs as a viable multi-modal passive vibration control device in variable RPM multirotor vehicles.
Given the necessity of performing System Certification according to SAE ARP4754, accepted as guideline by aeronautics certification authorities for development of aircrafts and complex systems, the need to define a robust and adaptable system requirements Validation and Verification (V&V) process has become a priority. SAE ARP4754 compliant processes shall be applied for certification of new complex systems, as well as to existing ones. Defining suitable and compliant processes for projects that were already in an advanced development stage when compliance to ARP4754 became mandatory is even more challenging with respect to the application to new projects, as the need of rearranging existing certification documentation naturally arises. This paper illustrates a process compliant with ARP4754 guidelines to achieve the System level requirement V&V. The presented process – based on the Function-Based Systems Engineering (FuSE) – has been applied to the civil certification of the Fly-By-Wire Flight Control System (FCS) of the AW609 tiltrotor by Leonardo Helicopters and has been reviewed by the Federal Aviation Administration (FAA).
This paper introduces a novel concept for an AI-powered system designed to manage vertiport takeoffs and landings by proactively addressing the safety-critical issues of downwash and outwash. The proposed system utilizes a stream of live feedback from on-site sensors, combined with a robust predictive modeling engine, to generate optimal, aircraft-specific approach and landing trajectories in real-time. By leveraging a comprehensive database of pre-computed downwash/outwash scenarios for a multitude of UAM aircraft configurations, the AI can accurately predict the unique outwash operational footprint for each individual landing operation, based on the approaching aircraft and under the prevailing conditions. This powerful predictive capability allows the system to calculate and assign an optimal approach vector that actively minimizes risk by directing hazardous airflows away from personnel, active walkways, and other sensitive ground assets. This represents a paradigm shift from static, reactive safety measures to a proactive, intelligent, and performance-based operational model; thereby paving the way for the safe, efficient, and large-scale implementation of UAM aircraft and air taxi operations.
T-tail architectures show potential for enhancing vertical tail-efficiency and lowering fuselage download and hub load cycles during low-speed transition. However, a horizontal stabilizer is principally susceptible to rotor wake impingement during cruise flight, which, in unfavorable conditions, could induce dynamic loads along with associated vibrations and structural fatigue. Predicting this phenomenon is challenging due to the complex aerodynamics and sensitive structural dynamics involved. This paper demonstrates the capabilities of a mid-fidelity simulation methodology for predicting empennage structural loads and vibrations. The approach utilizes mid-fidelity interactional aerodynamics modeling, building upon previously published Vortex-Lattice Model (VLM) results and extending them to include a Viscous Vortex Particle Wake (VVPM) analysis, coupled with a modal structural dynamics model of the fuselage. The study extends the simulation model's validation against experimental data across various flight states and sensors, incorporating a sensitivity analysis of the aerodynamic modeling. Additionally, the work presents flight state sensitivities, as well as design sensitivity studies examining the influence of main rotor blade number and T-tail planform. The results indicate that the mid-fidelity tool chain is a valuable industrial asset supporting the aeroelastic and aeromechanical design of airframe tailplanes affected by interactional effects. It allows efficient analyses with adequate numerical accuracy over a large range of operating conditions on the one hand and covering a variety of architectural choices on the other hand in view of vibratory loads and tailplane vibrations. The sensitivity study demonstrates the advantages of a high number of main rotor blades and a swept T-tail planform design for reducing vibratory loads, considering both, aerodynamic excitation and structural response.
This paper presents the integration and use of state-space free-vortex wake models within closed-loop rotorcraft flight dynamics simulations. The free-vortex wake models are formulated in state-variable form, such that they constitute a system of nonlinear, time-varying ordinary differential equations in first-order form that augment the baseline rigid-body and rotor dynamics. The wake models considered include a tip-vortex-only formulation, as well as a formulation combining a vortex-lattice near wake with a tip-vortex representation of the far wake. Following trimming, linearization, and model-order reduction of the flight dynamics at discrete increments in flight speed, Dynamic Inversion (DI) flight control laws are synthesized to enable automatic transition from hover to cruise flight. Two-way-coupled losed-loop simulations are then performed for a generic utility helicopter representative of an H-60 in transition from hover to forward flight using three inflow models: (i) Pitt-Peters, (ii) tip-vortex-only wake, and (iii) vortex-lattice near wake with tip-vortex far wake. These simulations are compared both for validation purposes and to assess whether the different wake representations lead to significant differences in the predicted closed-loop response.
A velocity potential-based finite state model (VPBFSM) has been developed to analyze an isolated rotor in ground effect. The model represents the ground using mass source distributions and imposes the non-penetration of flow boundary condition at the ground. In this paper, VPBFSM predictions of the inflow distribution are compared with experimental results for full and inclined ground effect cases using a model-scale rotor. The VPBFSM shows good agreement with the experimental results and captures the expected trend of decreasing inflow as the rotor approaches the ground, with a larger reduction on the side closest to the ground. Differences in magnitude are observed, but remain acceptable and are attributed to reduced-order modeling assumptions in the VPBFSM and uncertainty in the experimentally derived inflow measurements.
An aspect of the ship-helicopter dynamic interface (DI) is the highly unsteady flow environment generated by ship-rotor aerodynamic interactions, which challenges safe launch and recovery operations. To investigate these interactions without the constraints of conventional rotor scaling, a novel airflow-and-blade-frequency (ABF) system was developed, decoupling rotor thrust from blade-passing frequency and enabling independent control of disk loading and periodic excitation. Mean-flow superposition and spectral analyses were used to assess the validity of linear-superposition approaches for DI modeling. While superposition reproduced portions of the interacting mean flow, it failed to capture key features such as superstructure sheltering. Spectral results showed that momentum injection and blade-passing frequency modified the interacting flow through distinct mechanisms. Across all operating conditions, the interacting flow exhibited elevated turbulent kinetic energy at pilot-relevant frequencies over a broader spatial extent than either the isolated airwake or the superposed field, indicating that nonlinear aerodynamic interactions generated flow features that super-positional models did not capture. The persistence of these trends across different ABF operating parameters suggested that correction-based approaches may approximate rotor-feedback effects without requiring fully resolved aerodynamic interactions between the ship and rotor (air)wakes.
This paper develops and tests a feature-based autonomous landing system for vertical lift aircraft on stochastically moving ship decks, under degraded visual conditions. The system is tested with a custom-built quadrotor on a six-degree-of-freedom 1.5-ton Stewart platform reproducing stochastic motions up to Sea State typical of DDG-51-class ships. Experiments began with nominal conditions, followed by a stepwise degradation of deck features through occlusion, low illumination, water distortion, and glare. The vision algorithm tracked the platform and achieved landing across all scenarios, with tracking errors of up to 14% of the vehicle footprint, and up to 2.2° of pitch and roll. Overall, it demonstrated the ability to land in a GPS/Lidar-denied, difficult environment with on-board vision alone, achieving deterministic, repeatable results.
A single pilot, full-scale, proton exchange membrane fuel cell powered helicopter is flight tested with 700 bar compressed gaseous hydrogen as fuel. Models are developed for the fuel cell, hydrogen and the helicopter and validated with flight test data. The data covers powerplant architecture, stack electrical characteristics, hydrogen flow, detailed component weights, radiator drag, and full aircraft power measured in hover and forward fight. The validated models are then used to conceptually explore the conversion of a larger, more capable, turbine engine Robinson R66-like airframe with liquid hydrogen supplied fuel cell. Predictions indicate that payloads of 300−600 lb can be carried over a range of 200 nautical miles with current fuel cell technology if hydrogen storage weight fractions of 0.2−0.3 can be achieved and the tank and baggage compartment both are used for fuel. The key conclusion is that hydrogen fuel cell helicopters are feasible and the test data and validations presented here open the door for advancing its performance in the future.
The present work develops a computational framework for simulating the two-way coupled ship-helicopter dynamic interface using large-eddy simulation. The Simple Frigate Shape 2 geometry is modeled using the immersed boundary method, and baseline simulations under both uniform inflow and neutral atmospheric boundary layer (ABL) conditions are validated against wind tunnel measurements for two wind-over-deck angles. Rotor modeling techniques, including the actuator line model (ALM) and actuator disk model (ADM), are verified and validated across several configurations: the Knight and Hefner rotor, ONERA HAD-1 propeller, and NASA Dragonfly Phase B* coaxial rotor. The lower-fidelity ADM captures wake characteristics consistent with the ALM with up to a 9× speedup. The ADM maintains strong agreement with experimental and numerical results for integrated performance metrics and is suitable for two-way coupled simulations. The developed framework is applied to a rotor-obstacle configuration based on a GARTEUR 22 experimental survey and subsequently to a fully-coupled ship-rotor-ABL system.
This experimental study showcases the aeroacoustic sources measured on a NACA0012 airfoil subjected to dynamic stall due to sinusoidal plunging motions. The flow fields are measured on the upper surface of the airfoil using time-resolved particle image velocimetry (PIV), and the broadband surface pressure fluctuations were measured using a flush-mounted microphone probe and the reconstructed pressure field from PIV. Boundary layer separation occurs as the plunging airfoil approaches the maximum plunging velocity. A dynamic stall vortex (DSV) forms on the upper surface near the leading edge. Pressure distribution over the upper surface evolves in response to the movement of the DSV, with the lowest surface pressure observed at the DSV location. Full boundary layer separation results in a temporary reversal of the adverse pressure gradient, and the lowest pressure during moments of full detachment is at the trailing edge. The overall magnitude of the power spectral density (PSD) of the surface pressure fluctuations increases during the stages near the maximum plunging speed, with greater increases observed for the downstroke phase where the DSV is proceeding over the surface. The low-frequency tonal peaks observed at both the DSV location during downstroke and the maximum velocity during upstroke. However, high-frequency broadband fluctuations were measured from the DSV passage during downstroke. These variations of the surface pressure and its broadband components indicate significant unsteady loading and broadband noise sources from a plunging wing.
After systematic testing of scaled propeller-driven rotor models in hover, and successful correlation of the test data with prediction methodology, a scaling study was conducted for a range of aircraft gross weights from 100 lbs to 20,000 lbs. From previous studies it became evident that propeller sizing and spanwise location on the main rotor blade were key for overall good main rotor performance. The impact of propeller design and placement on propeller-driven rotor hover power was estimated from the isolated propeller performance and propeller-driven rotor configuration. Propeller-driven rotor hover power was compared to a conventional shaft-driven isolated main rotor and a shaft-driven single-main rotor helicopter. The scaling study showed that, for a well designed propeller-driven rotor, the hover power was comparable between the propeller-driven and shaft-driven isolated rotor/single main rotor designs. Finally, a disk loading sensitivity analysis was performed which found that the propeller-driven rotor should have a low disk loading to prevent excessive rotor and propeller tipspeeds.
A wind tunnel investigation to assess the impact of rotor-fuselage spacing on the development of the Vortex Ring State and flow topology is presented. Particle Image Velocimetry was utilised to investigate flow mechanisms across a range of rotor-fuselage spacings and descent ratios, which were compared to that of an isolated rotor configuration. Mean flow data was used to identify coherent flow structures, whilst flow unsteadiness was investigated through statistical analysis of the velocity fluctuations. It was found at cases of Vortex Ring State onset, the presence of the fuselage delays the development of the Vortex Ring State for all rotor-fuselage separation distances tested. Furthermore, certain cases of rotor-fuselage spacings display a rotor-fuselage aerodynamic interaction that results in an increased effective descent ratio.
When surveying the current landscape of Deterministic Ethernet avionics solutions in the aerospace industry, the three main technologies in the market are ARINC 664 part 7 rate-constrained Ethernet (commonly known by its trademark name "AFDX®"), TTEthernet (which combines ARINC 664 part 7 with Best-Effort Ethernet, while adding a new class of synchronous determinism defined in SAE AS6802 [Time-Triggered Ethernet]), and IEEE 802.1 Time-Sensitive Networking (TSN). No single deterministic Ethernet technology optimally satisfies certification, MOSA, and lifecycle goals across all avionics domains. Instead, successful digital backbones require intentional partitioning of responsibilities across technologies. This paper will seek to identify a number of those considerations and provide guidance on which technologies offer the best fit. After first opening with an explanation of the market forces driving the trends towards these technologies, this paper will delve into a short outline of each of these Ethernet standards. Following that, this paper will compare a variety of characteristics of each of the above Ethernet technologies pertaining to certifiability, MOSA conformance for DoD use cases, and supply chain considerations. Finally, the paper will synthesize those comparisons into a number of recommendations regarding the most appropriate uses for each Ethernet technology.
Helicopter air tours operate in one of the most challenging and least-controlled environments of commercial aviation, yet the safety outcomes of these operations remain inconsistent across regulatory frameworks. This study examined 55 helicopter air tour accidents in the United States from 2014 to 2024 using data from the NTSB Case Analysis and Reporting Online database. Defining event narratives, contributing factors narratives, and probable cause were coded to identify causal relationships between accidents and identify safety trends between 14 CFR Part 91 operations and Part 135 operations. CFR Part 91 operations exhibited accident rates approximately three times higher than Part CFR 135, averaging 3.94 per 100,000 flight hours compared to 1.23. Maintenance/mechanical was the most common initiating cause for accidents under CFR Part 91, accounting for 52% of cases compared to 37% under Part 135. Pilot/related cases were more prevalent under CFR Part 135, accounting for 53% of accidents. The two regulatory frameworks operated substantially different fleets, with CFR Part 91 relying on reciprocating-engine helicopters (76%) and Part 135 on turbine-powered aircraft (81%). Engine and powerplant/related events accounted for 27% of all defining events, and nearly half of all events involved a technical or mechanical initiator.
The objective of NASA's 4th New Frontiers Mission, Dragonfly, is to explore the surface chemistry and habitability of Saturn's largest moon, Titan. With its thick nitrogen atmosphere, liquid methane cycle, and rich, organic surface materials, Titan holds clues to prebiotic chemistry to answer fundamental scientific questions about the building blocks of life. The combination of high fluid density (4.4x) and low gravity (1/7th) compared to Earth makes exploration of this cryogenic ocean world in the outer solar system feasible by means of a relocatable lander - this is Dragonfly, a multi-rotor vehicle designed for the unique atmospheric conditions and environment at Titan. Dragonfly enables flight in a quad-rotor configuration with two counter-rotating, canted rotors mounted on each of four sting arms. All eight rotors are three-bladed, stiff metal rotors that are controlled by variable-speed electric motors. The objective of this paper is to tell the story of Dragonfly's rotor blade design and optimization, starting with the conceptual design based on flight requirements for Titan, preliminary design iterations of the rotor blades, and detailed design and optimization of the final configuration. Details are given with respect to design constraints driven by the cryogenic Titan environment, resulting from scientific instruments located on Dragonfly, and the overall mission flight profile. Design tools ranged from momentum theory to free-wake methods, hybrid computational fluid dynamics (CFD), and blade-resolved CFD analyses compared to wind tunnel measurements of rotor and lander combinations.
Ground resonance, a self-excited instability typical of helicopters, that can lead to catastrophic failure. Its analysis becomes particularly complex in non-symmetric rotor configurations, such as those with an inoperative damper, which give rise to Linear Time-Periodic (LTP) systems. The stability investigation of the equations of motion can be handled through the Floquet method of characteristic exponents, or with the more recent linear time-invariant (LTI) harmonic decomposition (HD) method. The two methods are presented and their relationship is explored through Hill's infinite matrix and its modal solutions. Numerical results are given for rotors with different blade counts and two damper arrangements—blade-to-hub (BH) and interblade (IB)—under single and multiple damper failure conditions. A minimum-harmonics rule is derived that links the required number of harmonic terms directly to the multi-blade coordinate (MBC) structure of the rotor. Finally, the LTI state-space system produced by the HD is coupled with a describing-function framework to extend the stability analysis to rotors equipped with nonlinear dampers, enabling efficient limit-cycle prediction without recourse to time integration.
This paper presents two distinct Lagrangian models developed for efficient rotorcraft inflow prediction: an Extended Vortex Ring Wake (EVRW) model and a Viscous Vortex Particle Method Actuator Disk (VVPM-AD). The EVRW model represents wake vorticity as circular rings with a harmonic distribution of bound circulation, thereby avoiding explicit coordinate-by-coordinate calculations. In contrast, the VVPM-AD discretizes shed vorticity into an independent Lagrangian particle cloud. This formulation naturally captures wake roll-up and viscous effects via a meshless Large Eddy Simulation solver, eliminating the need for empirical core-radius formulations. Validated against experimental data, these separate modeling tracks provide a scalable toolkit that balances computational speed with simulation fidelity.
Fuel cell systems have achieved a significant level of technological maturity in ground-based mobility over the past two decades. In particular, commercially available fuel cell propulsion systems are now in serial production for passenger cars and city buses, and are already in regular commercial operation. In the segment of heavy-duty vehicles - such as trucks and other long-haul applications - small-series production and technology demonstrators are currently available and are on the verge of entering the mainstream market. These developments have resulted in well-proven system architectures, sophisticated balance-of-plant components, and established supply chains. In contrast, the utilization of fuel cell propulsion in aviation is still at a very early stage. At present, only a handful of individual prototypes and technology demonstrators - mostly for small aircraft - exist, while serial production remains far in the future. Particularly in the field of lightweight, small, electrical vertical take-off and landing (eVTOL) aircraft there is a unique opportunity to leverage the proven fuel cell systems developed for ground vehicles, adapt them, and further develop them to meet aviation-specific requirements. Such an approach can shorten development timelines and reduce technical risks. Transferring existing fuel cell technologies into aviation, however, is far from a straightforward process. One decisive difference lies in the required specific power density. Aircraft - especially eVTOL - demand significantly higher power densities than those delivered by current commercial fuel cell systems from the automotive sector. This requires a direct adjustment of the stack design and system architecture. Likewise, thermal management poses particular challenges. Whereas piston engines and gas turbines discharge a large portion of their waste heat via exhaust gases, fuel cells must remove all waste heat directly through their cooling systems. This requires efficient radiators capable of transferring heat from the coolant to ambient air. Larger radiator surfaces, however, increase both total aircraft mass and aerodynamic drag, making compact radiator designs essential for aviation applications.
This study investigates the acoustic performance of a single rotor representative of those seen on multi-passenger UAM-sized vehicles, focusing on the effects of blade count, disk loading, solidity, and tip Mach number in both hover and propeller operating conditions. Using PSU-WOPWOP and ANOPP2, unweighted and A-weighted overall sound pressure levels (OASPL) are computed in-plane for 2- and 5-bladed rotors across a range of design parameters and operating conditions. Unweighted results show that reducing blade count significantly increases total noise levels (14.1 dB on average) and reduces sensitivity to design parameters. In contrast, A-weighted results demonstrate that broadband noise dominates perceived acoustic performance and shows a decreased sensitivity to blade count (1.9 dBA average difference). Minimum noise levels occur at tip Mach numbers ranging from 0.35-0.45 for unweighted results and 0.4-0.5 for A-weighted results, and are primarily governed by broadband noise sensitivity to disk loading and solidity. The rotor in propeller mode, with axial flow and reduced disk loading, showed less sensitivity to variation in disk loading and solidity than the rotor in hover, indicating weaker acoustic dependence in cruise conditions.
After four decades of research and 3.5 year prototype testing campaign, Penn State's pericyclic transmission technology demonstrator, dubbed the 'Pericycler', has achieved its operating speed of 5,000 RPM at 17 HP. The characterization of this system by experimental efficiency and vibration represents a major milestone in pericyclic gear technology. A post-test inspection procedure was performed to analyze component wear and validate hypotheses on mesh behavior. This work concludes with structural, tribological, and instrumentation modifications to the Pericycler for future testing.
A high-fidelity computational study investigates the aerodynamic behavior, flight response, and control effectiveness of a multirotor electric Vertical Take-Off and Landing (eVTOL) configuration. The investigation is organized into two parts. Part I employs an unsteady computational fluid dynamics (CFD) framework coupled with a six-degree-of-freedom (6-DoF) rigid-body dynamics module. Simulations for isolated coaxial rotors and a complete eVTOL isolate rotor aerodynamics and rotor–airframe interactions under constrained kinematics, quantifying lift capability, fuselage download, and a residual nose-up pitching moment arising from fore-aft rotor lift imbalance. Fully coupled 6-DoF free-flight simulations capture the transient vehicle response to a motor failure and recovery sequence during hover. Part II assesses flight control response through a cascade Proportional-Derivative (PD) controller implemented in MATLAB/Simulink across two maneuver cases: hover stabilization and climb rate tracking, which are parameterized using aerodynamic data extracted from the isolated rotor CFD simulation. This decoupled approach enables systematic gain tuning and controller assessment without the computational overhead of fully coupled closed-loop CFD simulations. The results confirm that the CFD–6-DOF framework effectively resolves tightly coupled aerodynamic-dynamic interactions inherent to distributed electric propulsion configurations, and that the cascade PD architecture provides initial control authority assessment across the primary flight axes. These findings establish a foundation for trim strategy development, advanced control law design, and future integration toward robust flight control for urban air mobility operations.
For many years there has been a keen focus on pilot workload and its associated assessment methods, and it remains a highly relevant aspect of flight testing. This paper provides a synopsis of a novel workload rating scale and index, the Comeau-Duggan Pilot Workload Index, designed to bridge the gaps in existing subjective workload metrics – such as casual factor identification – that are present in the most widely used rating scales in flight tests. The conceptualization and development of this index represent a multi-year, dual-national research effort that builds on the foundational concepts and core principles underlying widely accepted workload rating scales used in Human Factors and Handling Qualities engineering. The Pilot Workload Index provides a structured and rigorous methodology for ascertaining and distinguishing factors that have contributed to the pilot workload of a given flying task, evaluating their impact using a systematic suffix-flowchart framework. The Pilot Workload Index was developed and assessed in piloted flight simulation trials conducted by University of Liverpool and in real-world flight trials conducted by the National Research Centre Canada for Ship-Helicopter Operating Limit and Roll-Step Mission Task Element assessments. Other scales were used during these trials (Bedford Workload Rating scale, NASA Task Load Index and Deck Interface Pilot Effort Scale, as appropriate) to compare results with the new rating scale. This paper presents a subset of the detailed Pilot Workload Index data recorded during flight simulation trials. It also explores the insights that can be derived from the data, highlighting the benefits of the Pilot Workload Index in workload assessments in flight tests.
Urban Air Mobility (UAM) represents a paradigm shift in metropolitan transportation, introducing electric vertical takeoff and landing (eVTOL) aircraft into dense urban ecosystems. This transformation is driven by advances in electrification, digital infrastructure, and integrated airspace management. According to the U.S. Department of Transportation's Advanced Air Mobility National Strategy 2025, UAM is expected to become a cornerstone of multimodal urban transport, with commercial operations projected in multiple U.S. cities before 2030 [1].
The bird strike performance of rotorcraft components must be demonstrated to the airworthiness authority in accordance with the certification requirements of CS 29.631. This necessitates continuous efforts to design and validate birdstrike-resistant structures through a combination of experiments and simulations. In this study, an integrated experimental and numerical investigation is conducted to evaluate the structural response and failure characteristics of the main rotor pitch link subjected to bird impact. In the experimental program, high-speed imaging and strain measurements were used to capture the transient deformation and impact force history. In parallel, a highly nonlinear finite element model was developed using the LS-DYNA solver. The numerical model was validated against experimental results. Results demonstrate that localized plastic deformation and stress concentrations occur near the impact region, consistent with damage patterns observed in real-world incidents. This research provides a validated methodology supporting CS 29.631 bird strike certification for rotorcraft flight-control components.
Autorotation is an emergency flight maneuver in which a helicopter descends safely without engine power by using rotor energy. This paper investigates the use of reinforcement learning (RL) for autorotation trajectory generation and systematically evaluates it against optimal control problem (OCP) solutions. A one-degree-of-freedom powered descent problem is first solved as a surrogate to identify robust hyperparameter settings. The surrogate case results demonstrate that the RL policy closely matches the OCP solution in terms of landing time, confirming its effectiveness. The autorotation problem is then solved under both frameworks, and the resulting Height-Velocity diagrams are compared, with crash behavior in the deadman zone analyzed for each. The RL framework is shown to produce autorotation trajectories comparable to OCP, establishing it as a viable real-time alternative. Warm-starting the OCP with RL-derived solutions improves convergence compared to conventional initialization. Finally, the RL policy's versatility is discussed with an example of varying initial helicopter weight, capturing different fuel states at engine failure.
This study evaluates the predictive accuracy and computational efficiency of a mid-fidelity Lattice-Boltzmann Method (LBM) framework in simulating the complex aerodynamic interactions of a tilting proprotor–wing configuration. The analysis focuses on the tiltrotor conversion maneuver, investigating a range of proprotor tilt angles from forward towards edgewise and vertical flight. To resolve the interactional flow physics, the LBM framework was integrated with two distinct proprotor modeling approaches, an Actuator Line Method (ALM) and an unsteady Actuator Disk Method (ADM), and two wall model boundary conditions, the explicit power-law and Reichardt’s log-law. The computational models were compared with experimental wind tunnel measurements and high-fidelity computational fluid dynamics (CFD) simulations. The ALM significantly outperformed the ADM in capturing discrete tip vortices and wake turbulence, which were critical for resolving the complex flow fields and wing surface pressure distributions. Reichardt’s log-law wall model demonstrated greater physical accuracy over the power-law model by correctly preserving proprotor wake alignment and improving surface pressure predictions. Modifications to the wall models and co-locating both the proprotor and the wing within the finest mesh refinement level enabled the LBM to accurately capture lift coefficient trends and magnitudes across the conversion maneuver. While limitations were identified in resolving leading edge flow attachment at high tilt angles and resolution-sensitive wing drag, this study demonstrated that the LBM-ALM approach with log-law wall model is a highly effective tool for rapid and accurate aerodynamic analysis of tilting proprotor-wing systems.
Newly designed eVTOL aircraft utilize propellers that operate with a large range of propeller rotation rates. Traditional nomenclature uses nondimensionalization based on the blade tip speed, and input reduction based on a similarity assumption under constant advance ratios. In this study, we explore the validity of this similarity assumption in the context of hover and descent scenarios for a variable pitch eVTOL propeller with rotation rates ranging from 54%-100% of the maximum value. In hover, the relative Reynolds number and Mach number effects are found to be relatively minor. As the axial descent ratio increases, prior to the onset of vortex ring state, the similarity assumption breaks down, and the mean thrust coefficient varies up to ±10% under different rotation rates. A similar breakdown is observed for descent conditions with higher edgewise flow. A detailed exploration shows that the effect is primarily due to relative Mach number effects, which alters the tip vortex wake interaction at the disk, and to a lesser extent changes the loading along the blade. Since the breakdown in the similarity assumption is isolated to a small region of the propeller operating envelope, a surrogate model is developed to preserve the reduced input dimensionality while accounting for the relative tip speed effects in the affected region of the envelope. This is achieved by first generating the surrogate over the full domain, assuming validity of the similarity assumption. A separate high-RPM surrogate correction layer is then built to capture the worst-case tip speed effects in the descent region of the envelope.
This paper presents the development flight test campaign of autopilot Upper Modes for T-625 Gökbey helicopter. The primary objective of the test campaign is to evaluate the newly developed Upper Modes in the frequency and time domain across the operational flight envelope. For quantification of performance and stability, various metrics are selected from the literature. Flight tests are designed to extract the metrics from time domain data and tests are conducted. Initial flight tests revealed discrepancies between theoretical design models and actual aircraft dynamics, requiring iterative control law gain optimizations. Furthermore, combined mode engagements required targeted simultaneous tuning of different modes to maintain stability margins in combined engagement. By integrating quantitative data analysis with qualitative pilot feedback, engagement logic and control parameters were successfully refined.
This paper experimentally investigates the effect of positioning of individual blades of a two bladed propeller around the rotor hub on overall noise generated by it. An experimental setup was created to measure noise and performance in an anechoic chamber to carry out parametric study in which axial and azimuthal separations between the two blades were introduced through a custom built rotor hub and balancing weight. The propeller noise that is dominated by tonal components associated with blade passage frequency appears to be influenced by azimuthal separation between individual blades and the broadband components generated by turbulent blade-wake interactions is primarily affected by the axial separation between individual blades. From the present study, it is identified that the rotor configurations with 60° azimuthal and 6 mm (3.2% of rotor radius) axial separation resulted in up to 4.4 dB reduction in Overall Sound Pressure Level (OASPL) and 63.9% reduction in acoustic energy, while the 120° configuration with 6 mm (3.2% of radius) axial separation showed up to 3.8 dB reduction in OASPL and 58.7% reduction in acoustic energy through redistribution in acoustic energy across wider frequency range. Flow-field measurements using Particle Image Velocimetry (PIV) revealed that these improvements are associated with merging of tip vortices, thereby reducing blade-wake interactions, demonstrating that controlled blade spacing through hub modification is an effective passive strategy for reducing UAV propeller noise without affecting the performance significantly.
Evaluating rotor component clearances is a multidisciplinary process aimed at ensuring that no contact occurs between rotor parts during a rotorcraft's operational life. It begins with calculating relative distances between components across all possible displacements and deformations combinations using a rotor kinematic model, and ends with clearance verification through flight data analysis and simulation. This task requires coupling detailed rotor aeroelasticity with flight mechanics to predict deformation under load, which is computationally expensive and unsuitable for real-time use. This work proposes a machine learning–based alternative: a neural network to estimate rotor clearances from flight mechanics inputs, with a specific application demonstrated in a simulated tiltrotor emergency maneuver with a pilot in the loop. The trained model successfully captures nonlinear relationships between maneuver parameters and rotor structural response, providing accurate predictions with reduced computational cost. The proposed framework can support both preliminary and detailed rotorcraft design and offers potential for real-time load prediction in applications where conventional aeroelastic tools are computationally expensive.
This paper presents the design, development, and successful demonstration of the first-ever ballistically tube-launched tailsitter unmanned aerial system. The vehicle expands upon the capabilities of existing tube-launched systems by simultaneously integrating the hovering capability of a rotary-wing aircraft with the efficiency and speed of a fixed-wing aircraft. To achieve this, the platform's design incorporates a novel coaxial thrust-vectoring propeller system for control in vertical flight and a unique foldable wing design for ultra-compact storage in the launch tube. The aeromechanics of the foldable wings during deployment are studied through a combination of wind tunnel experiments and flight dynamics model simulations, and the results are used to formulate a methodology for executing the ballistic launch. Simulations are also performed to characterize the robustness of the system against asynchronous deployment of the left and right wings. Experimental data, collected from flying a prototype in vertical, horizontal, and transitioning flight, demonstrate the aircraft’s flying performance. The study ultimately culminates with a demonstration of the prototype being rapidly launched from a tube at 25 m/s (56 mph) and autonomously unfolding, stabilizing, and transitioning into self-powered cruising flight.
This study investigates the post-failure flight dynamics of a 1200 lb classical octocopter under single motor inoperative condition using nonlinear time-domain simulations with a baseline feedback controller. A physics based propulsion sizing strategy is developed using IEC duty cycle definitions where continuous requirements are derived from nominal hover with margin and short time capability is used to accommodate elevated post failure loads. The selected motor satisfies both regimes and enables transient overdrive without excessive weight penalty. Simulation results in hover and forward flight at the best range speed showing that the vehicle can recover from any single motor failure and retrim using inherent redundancy without fault identification. However, recovery involves significant transient attitude excursions and altitude loss, and requires substantial increases in motor power, with multiple motors exceeding S1 power limits. Post-failure maneuver simulations indicate retained controllability with some degradation and increased coupling. These simulations demonstrate that the proposed motor sizing enables necessary operation post-failure while avoiding unnecessary oversizing.
This paper demonstrates the sizing and optimization of a hybrid-electric multi-tilt rotor configuration of both conventional and vertical takeoff and landing capabilities. The study uses Parametric Energy-Based Aircraft Configuration Evaluator to design and optimize the aircraft. To explore the design space comprising both discrete and continuous design variables, a genetic algorithm is used for optimization. The design variables are not limited to conventional aero-propulsive parameters such as wing loading, aspect ratio, and disk loading. Battery-related parameters such as the maximum permissible depth of discharge, maximum permissible discharge rate, and the number of parallel strings in a battery pack are also considered in this work to study their impact on aircraft gross weight and fuel consumption. The Non-dominated Sorting Genetic Algorithm-II (NSGA-II) optimization framework is used to solve the multi-objective optimization problem, with objectives to minimize the maximum take-off mass and fuel weight. The sensitivity studies showed that higher wing-loading and lower-aspect-ratio designs resulted in lower gross weight. A higher permissible depth of discharge led to lower fuel consumption, despite a slight increase in gross weight. But increasing the number of parallel strings in a battery pack increased gross weight with negligible change in fuel consumption.
Vertical Take-Off and Landing (VTOL) aircraft represent one of aviation's most complex design challenges, balancing lift, thrust, stability, and control within an inherently unsteady aerodynamic environment. Since the 1940s, computational methods used to design VTOL systems have undergone a profound transformation, progressing from hand-drawn airflow approximations and wind-tunnel testing to today's high-fidelity digital twins, computational fluid dynamics (CFD), and AI-assisted optimization. The evolution of these methods mirrors the broader technological shift from empirical design toward simulation-driven innovation. The greatest transformation in VTOL design of the past 80 years is the shift from material and mechanical innovation to computational and cognitive design. Modern aircraft are as much products of computation and data as of metal and composites. As electric propulsion, autonomy, and digital twin technology converge, the next generation of designs, particularly configurations inspired by power systems such as hybrid-electric, hydrogen, battery only, will extend this century-long trajectory into a new paradigm: sustainable, intelligent, and continuously self-optimizing VTOL flight.
This paper explores the potential of three different hybridization solutions for a medium-sized rotorcraft: an electric tail rotor, an "eco-mode", and a "boost-mode". The solutions were evaluated as a retrofit to a generalized medium lift rotorcraft and the impact on performance across five mission types, representative of the typical use cases for a military rotorcraft, was assessed. Two separate rotorcraft performance modelling tools were used to carry out the assessment, allowing for the results to be cross-examined. The models predicted performance gains for the eco-mode configuration when utilizing the single engine cruise capability for low-speed applications. Likewise, the models predicted improved performance for the boost-mode configuration when operating at hot and high (6,000 ft, 95°F) conditions due to the increased power provided by the battery system. However, all three solutions suffered from increased platform empty weight which negatively impacted performance at certain flight states.
Advanced air mobility (AAM) seeks to develop a large-scale transportation system to revolutionize how people live and work, with electric vertical take-off and landing (eVTOL) aircraft serving a central role due to reduced emissions and noise impact. An important aspect for eVTOL aircraft certification is safe urban operations, which require understanding of the response due to aerodynamic disturbances. Experimental data are required to support eVTOL aircraft development with respect to flight dynamics and controllability, as well as design specification development. While flight testing of the full-sized air vehicle will be necessary as part of the certification process, subscale testing offers many advantages with respect to cost and flexibility, in addition to examining operational conditions that one would be reluctant to test in flight at full scale such as emergency conditions. These advantages only may be seen if the underlying scaling principles of flight dynamics / control, aerodynamic interactions, and propulsion-airframe integration are understood. This paper describes initial work towards development of a general subscale testing methodology for eVTOL aircraft flight dynamics and disturbance response characteristics including limited degree of freedom (DOF) and free flight testing. An overview of the initial development work is provided, including discussion of scaling relationships, subscale air vehicle model development, and testing activities focusing on flying qualities and stability / control characteristics.
This paper presents the development, optimization, and flight test validation of a Trajectory Control System (TCS)-based flight control system for a tiltwing unmanned aerial vehicle. The TCS is a configuration-independent middle-loop longitudinal controller for vertical takeoff and landing aircraft and is integrated here with explicit model following inner-loop controllers, inverse propulsor models, and a tiltwing-specific control allocation scheme. The resulting flight control system provides coordinated control across vertical flight mode, hybrid flight mode, transition flight mode, and forward flight mode while relying on a concise feedback set and requiring only airspeed from the air data system. The control laws are obtained using a formal constrained optimization framework and transferred directly from simulation to flight without additional on-site retuning. Flight test results from piloted, semi-autonomous, and fully autonomous operations demonstrate stable and predictable behavior throughout the flight envelope, including tight hover performance, simultaneous climb rate and speed tracking in hybrid flight, and successful departure and arrival transitions at multiple speeds. Selected simulation-versus-flight comparisons further show that the nonlinear model captures the dominant trends in the measured response while also identifying specific aerodynamic and transition regime effects that warrant further refinement. Overall, the results demonstrate that the TCS + EMF architecture provides a practical and effective control solution for tiltwing VTOL aircraft.
Within the next years, it is expected that the capabilities that are demanded to the rotorcraft fleet would be enhanced with respect to the current ones. Very long range, speed above typical rotorcraft performance, hot and high HOGE capability and high payload capacity are foreseen, together with limitation on aircraft take-off weight (TOW): among these sizing cardinal requirements, speed characteristics and long-range operations drive the sizing towards innovative solution, to overcome the physical limitation of a conventional rotorcraft. The work starts with a performance-based comparison of different fast rotorcraft architectures, comparing it with respect to the conventional helicopter, used as benchmark. Once first investigation loop is completed with a preliminary sizing analysis, a detailed one is focused on tiltrotor architecture, showing the impact of hover and high-speed capability on lifting and powerplant systems, as well as the impact of sizing criteria on the overall performance. In such second step, a matrix scenario is proposed, where both requirements and sizing criteria are evaluated to show the peculiarity on tiltrotor solution. In conclusion, considerations on balanced criteria for tiltrotor sizing are reported, with focus on sizing trade-off.
Bench-level boundary-lubricated fretting experiments were conducted to compare the relative wear of all-steel and hybrid material pairs. Roller-on-raceway contacts were simulated using both AISI M50 steel and Si3N4 cylindrical rollers on flat AISI M50 steel disks. The rollers were 9 mm long with a 9 mm diameter. Tests were conducted with constant amplitude, oscillation frequency, and load. All tests were boundary-lubricated with 0.1 ml of DOD-PRF-85734, MIL-PRF-32538, MILPRF-23699, or unclassified ISO VG 68 aviation gear oil. Wear volume was calculated from 3D measurements on the roller and disk samples after each test. Wear tracks were inspected with light and scanning electron microscopy. It was concluded that hybrid pairs exhibited less wear than all-steel pairs when boundary-lubricated with three of the four aviation gear oils. Both hybrid and all-steel pairs exhibited similar wear when boundary-lubricated with MIL-PRF-23699 oil.
A key tenet of the Modular Open Systems Approach (MOSA) is to modularize systems onboard the aircraft by standardizing interfaces and interoperability between various aircraft systems and subsystems. This is intended to decrease the time and cost required to install upgrades or replacements while minimizing the amount of regression testing required. This approach enables the evolution of capabilities over a system's lifecycle, allowing program teams to build, upgrade, extend and support systems at a lower cost and effort, ideally without extensive systems testing. This flexibility and optionality can be boiled down to the principle of "Freedom of Action", defined as “The ability to accommodate the desire of a system user to update and change the system to adapt to changing conditions, independently from a single vendor to encourage competition and lower price." Given that requirement, Open Systems Verification Demonstration (OSVD) is a methodology designed to provide evidence of a program’s MOSA conformance. It is intended to prove the degree of "Freedom of Action" that a particular system provides to its users. An OSVD event is intended to demonstrate alignment to these goals by showing the solution components meet required key interfaces and open standards, that the technical data documenting these interfaces is sufficient, and that the vendors provide the appropriate data rights. Prior to the U.S. Army's decision to cancel the Future Attack Reconnaissance Aircraft (FARA) program, it provided funding to Lockheed Martin for a total of 4 OSVD events specifically for the FARA aircraft. The purpose of these events was to evaluate FARA's planned Digital Backbone solution for modularity and openness. Following the cancellation of FARA, the 2 remaining OSVD events were redirected to focus on the UH-60 Black Hawk. This paper will summarize the recently concluded OSVD events for FARA / UH-60 as well as discuss the results and lessons learned. While OSVD's objectives cover a wide range of interoperability and integration goals, this paper will specifically focus on the network integration of the Digital Backbone.
Rotorcraft pilots operating in degraded visual environments encounter significant challenges during hover flight, where the absence of critical visual cues increases the risk of spatial disorientation. At low altitudes and in obstacle-rich environments, even minor losses in situational awareness can have severe consequences. Understanding the visual cues that support stable hover in good visual environments, and how their absence impacts performance and cognitive workload, is essential for mitigating these risks. This study examined key human factors in hover flight, focusing on the role of peripheral vision and microtextures in supporting pilot performance. It evaluated whether naturally relied-upon visual cues in good visual environment conditions can be artificially replicated to restore visual dominance in simulated degraded visual environments. Analysis included flight performance metrics, control inputs, physiological workload indicators, subjective assessments, and pilot feedback. The findings contribute to improved understanding of visual cueing and pilot adaptation in degraded conditions.
This paper presents the results of a flight test effort examining fully autonomous shipboard operations for small unmanned aerial vehicles (UAVs). Experiments were conducted at the Maneuvering and Seakeeping Basin (MASK) located at the Naval Surface Warfare Center, Carderock Division using custom-built quadrotor UAVs landing on an unmanned surface vessel (USV). These tests build upon previous ship landing algorithm testing in order to expand the envelope of operations and be more representative of a real-world mission. Several new flight modes were implemented, including takeoff and pattern flying, and a finite state machine was developed to allow smooth and autonomous transition between the different flight modes. The results from testing show smoothly executed missions both in still water and in the presence of waves. However, it was found that the initial conditions for the command filters in the position controller needed to be carefully selected. Without the correct initial conditions, discontinuities in the commands were seen when switching between modes that used the command filters and modes that bypassed them. The results of this work will help bridge the gap between ship landing-specific research and real-world applications encompassing multiple flight modes.
This work extends an integrated VABS (iVABS) framework for rotor blade structural optimization by incorporating enhanced cross-sectional parameterization and manufacturing-aware design considerations. The refined model includes features such as curved spar corners, continuous wrap-around skins, conformal non-structural mass, and ply-level discretization, enabling a more realistic representation of composite blade structures within a scalable optimization framework. Multi-objective blade-level optimization studies are conducted on a UH-60A-based blade using three representative cross-sections, considering both unconstrained and strength-constrained design cases. For the unconstrained problem, the optimal design achieves close agreement with target stiffness properties while also providing improved matching of mass center and shear center locations compared to prior design. When a minimum strength ratio is enforced, the feasible design space is significantly reduced, resulting in increased deviation from the target stiffness properties. However, strength-constrained designs exhibit improved agreement in mass per unit length, indicating a positive correlation between strength requirements and mass matching. The resulting structural trends are physically consistent, with thicker spars near the blade root and thinner, more elongated spars toward the tip. To connect computational design with physical realization, a manufacture–test campaign is conducted using a representative spar section. A composite spar based on an iVABS-derived layup is fabricated using aerospace-grade prepreg materials and standard tooling procedures. Experimental measurements show good agreement with iVABS predictions, including approximately 5.3\% error in mass per unit length and 1 2\% error in center-of-gravity location. These results provide preliminary validation of the framework and demonstrate its capability to generate manufacturable designs with consistent structural predictions.
This work presents the development of an interconnected gearbox drive system for a tandem rotor unmanned aerial vehicle (UAV) designed for a power rating of 20 kW per rotor. Development of the facility for dynamic testing of the interconnecting drive system is also presented. Tandem rotor configurations offer superior payload capacity and aerodynamic efficiency but pose challenges in transmission design due to the need for synchronized power distribution between counter-rotating rotors. The proposed gearbox employs a simple two-stage reduction system combining bevel and planetary gears to achieve compactness, high torque transmission, and ease of fabrication. A dedicated test rig is concurrently being developed to evaluate gearbox performance under variable torque and speed conditions for the first stage with the mechanical interconnection. The test set-up integrates a variable-speed drive, torque and vibration sensors, and a data acquisition system to measure efficiency, losses, and thermal characteristics. While fabrication and assembly are ongoing, this work lays the foundation for experimental validation and optimization of drivetrain performance. The project contributes toward developing reliable, lightweight, and scalable transmission solutions for heavy payload multi-rotor vertical lift UAVs. The drive system dynamics, in particular, the torque transmission through the cross shaft during maneuvers will be the investigated using the current test rig configuration.
This study presents a high-fidelity aeroelastic analysis for lift-offset coaxial rotors based on a three-dimensional (3D) finite element (FE) multibody dynamic analysis. The structural model is based on an updated Lagrangian formulation to capture geometrically nonlinear behavior. The internal aerodynamic model uses lifting line theory with linear inflow model, while the external aerodynamic model employs a panel/vortex particle method to predict aerodynamic loads. The lift-offset coaxial rotor developed by Korea Aerospace Research Institute is employed to investigate the aeroelastic response and the coupling analysis is performed on hover flight condition. The results obtained from the aeromechanics analysis using uniform inflow are compared with CAMRAD II in terms of blade displacement and sectional loads. Furthermore, through high-fidelity aeroelastic analysis using panel/vortex particle method, rotor–rotor aerodynamic interactions and structural loads, and 3D stress and strain distributions are investigated.
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