Browse Topic: Bodies and Structures
This SAE Information Report contains definitions for hydrogen fuel cell powered vehicle terminology. It is intended that this document be a resource for those writing other hydrogen fuel cell vehicle documents, specifically, Standards or Recommended Practices.
Traditional safe-life methodologies for rotorcraft structural components rely on deterministic safety factors to account for uncertainty in loads, material properties, and operational usage. While effective for ensuring safety, these approaches lead to early retirement lives and reduced aircraft availability. This paper presents an updated digital twin-based probabilistic framework for rotorcraft component fatigue life assessment that integrates a probabilistic stress–life (S-N) material model, machine learning-based load estimation from flight data, and Monte Carlo uncertainty propagation. The approach is demonstrated for a critical location on the CH-146 Griffon main rotor yoke. Compared with earlier work, the present study advances the framework through independent validation of the load-estimation model and application to available in-service flight data from multiple mission categories. A probabilistic sensitivity analysis is used to examine the separate and combined effects of material variability and load-estimation uncertainty on fatigue life, cumulative probability of failure, and hazard rate. For the CH-146 demonstration case, the results indicate that the material fatigue strength uncertainty has a major impact on the lower tail of the life distribution and the corresponding reliability-based life, whereas load-estimation accuracy uncertainty has a secondary influence on risk metrics. The application of the digital twin framework to operational, search and rescue, and training mission data further shows that mission-specific usage variability plays an important role in the evolution of fatigue damage accumulation and structural risk. Overall, the proposed framework provides a more informative basis for risk-based rotorcraft life assessment by explicitly quantifying uncertainty and incorporating aircraft-specific operational data. The study is intended as a step toward validation of the framework rather than a completed operational deployment.
This organizational process survey provides insight into the technical aspects of approved airworthy aircraft modifications applied in government organization vertical lift flight test. The publication reviews processes applied by the National Research Council of Canada's Flight Research Laboratory (NRC-FRL) and its Airworthiness Control System to enable research flight testing. Dominated by the need for integrating experimental payloads, the NRC-FRL embeds a Design and Fabrication Service organization for modification of internal and external client projects and flight test aircraft. In context of experimental flight testing, this work reviews technical information on process, facilities, and methodology for airworthy integration of flight test payloads. Information is used to synthesize recommendations in experimental vertical lift flight testing that satisfy both formal (regulated compliance) and informal (compliance intent) airworthiness requirements.
This digital standard is a requirements extract of AS4159 Specification For An Automated Interchange Of Standards Data. This file contains a general requirements extraction as well as files that are optimized for use with Doors Classic, Siemens Polarian, and PTC.
This digital standard is a requirements extract of AS6500A Manufacturing Management Program. This file contains a general requirements extraction as well as files that are optimized for use with Doors Classic, Siemens Polarian, and PTC.
This digital standard is a requirements extract of AS861C Minimum General Standards for Oxygen Systems. This file contains a general requirements extraction as well as files that are optimized for use with Doors Classic, Siemens Polarian, and PTC.
This digital standard is a requirements extract of AS5127D Aerospace Standard Test Methods for Aerospace Sealants Methods for Preparing Aerospace Sealant Test Specimens. This file contains a general requirements extraction as well as files that are optimized for use with Doors Classic, Siemens Polarian, and PTC.
This digital standard is a requirements extract of AS13001A Delegated Product Release Verification Training Requirements. This file contains a general requirements extraction as well as files that are optimized for use with Doors Classic, Siemens Polarian, and PTC.
This digital standard is a requirements extract of AS50881H Wiring Aerospace Vehicle. This file contains a general requirements extraction as well as files that are optimized for use with Doors Classic, Siemens Polarian, and PTC. 
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.
An experimental investigation was conducted to characterize the effects of partial-ground on the aerodynamics of a hovering rotor. A model-scale rotor was tested at a range of heights above ground and under partial-ground coverage, and rotor hub forces and moments were measured using a six-axis force/torque transducer during constant-power operation. The measurements were used to develop a semi-empirical thrust ratio model that accurately captures trends from out-of-ground effect to full-ground effect conditions. This model predicts realistic thrust behavior at low ground-coverage conditions, exhibiting high adjusted R2 and minimal root mean square error. Time-resolved particle image velocimetry was conducted for selected cases to examine induced flow features and to qualitatively assess changes in the downwash and edge-driven crossflow associated with partial-ground interactions. A geometric rotor-ground interaction area based on a circular-segment formulation was correlated to the thrust coefficient over the interacting region. Results show that thrust increases with ground coverage and decreases with increasing out-of-ground height, whereas moment coefficients increase with decreasing height. The increases in thrust and moment can be attributed to a reduction in the induced velocity above ground, as observed in the measured flow field, thereby increasing the effective ground-induced pressure beneath the rotor.
Rotor-rotor and rotor-boundary aerodynamic interactions of a quadrotor system without a fuselage in ground effect and ceiling effect for varying rotor-boundary distances and hub spacings were investigated. A GPU-accelerated Lattice-Boltzmann Method (LBM) was coupled to new unsteady actuator disk method (ADM) and actuator slice method (ASM) based rotor models for this purpose. Validation was conducted against experiments for both performance and particle image velocimetry flow field data. The trends in thrust and power were accurately predicted by both actuator methods, with high computational efficiency. Interactional flow physics were resolved, causing the consistent performance benefits very close to the ground, the performance penalties caused by the fountain flow effect between rotors occurring over a limited range of ground distances, and the persistent performance augmentation in ceiling effect. The ASM rotor model, with its individual blade representation, was found to predict rotor-rotor interactions more accurately than the ADM. However, it generally overpredicted aerodynamic loads, which was attributed to the Gaussian regularization applied along the blades. The results indicate that the GPU-accelerated LBM coupled with the introduced actuator methods is capable of computationally efficient prediction of multirotor aerodynamic performance and flow fields in ground and ceiling effects.
Previous researchers developed equations to model the induced flow on a 2D airfoil in the finite-state as opposed to the closed-form. Those models, however, were limited in that they could not handle an oscillating free stream that became negative. Recently, a new model was developed to include a single factor to carry the effects of the free stream changing signs. In developing this model, a Floquet instability was discovered at the instant when the flow changes direction. The effect of the instability grows with increasing number of oscillations of the sign of the free stream. The effects can be limited depending on the parameters of the flow. In this paper, the previous 2D model is amended to include a term that considers the effects of the induced flow from all previous vorticity segments that have been generated from each oscillation of the flow. This paper details the beginnings of the testing on the stability limits of the theory, based on changing the parameters of the free stream, airfoil, and timing. It is the intention of this research to further investigate the limits of this model in reversing flow in hopes of using the lessons learned to extend 3D finite-state models that are currently incapable of handling cases where the sign of the free-stream velocity changes as a result of the rotor reentering its own wake such as when a helicopter or quadcopter quickly descends after ascending to avoid an obstacle.
This paper discusses the design of a 2000-lb manned eVTOL aircraft propelled by a novel cycloidal rotor propulsion system. To systematically evaluate the performance of the proposed configuration, a coupled trim model was developed to quantitatively evaluate the performance of the configuration across a range of forward flight speeds. The trim framework integrates an efficient physics-guided neural-network-based aerodynamic model for cycloidal rotor performance with a vehicle-level dynamic response model. This framework is used to conduct a systematic parametric study to identify key cycloidal rotor and airframe design parameters. The selected configuration is verified using high-fidelity CFD simulations, and a detailed structural design, powertrain design, and CAD model of the aircraft is developed. In addition to CFD validation, the proposed cycloidal rotor underwent structural optimization to confirm the validity of such a concept at this scale. The results demonstrate that the cycloidal rotors provide a viable propulsion alternative for eVTOL aircraft with strong potential to overcome limitations of existing configurations.
Dimensional reduction of data can be accomplished through various methods and has applications critical to machine learning and surrogate modeling. Within the rotorcraft community, leveraging these techniques allows for improved rotor parameterization and performance prediction. Machine learning models generally perform faster and better with lower input dimensions, so long as all necessary information is retained, making appropriate dimension reduction paramount. Data can also be arranged in a one-dimensional (concatenated/stacked) or two-dimensional arrays to take advantage of function correlations, and this arrangement may allow for greater reduction at lower reconstruction costs. Principal Component Analysis with a stacked input shape proves to be the most effective reduction method considered, with reconstruction accuracy being validated though a suite of mid-fidelity aerodynamic simulations. A blade geometry defined using 204 original parameters can be fully described using just 10 component parameters with the reconstructed blade maintaining performance figures within 1% of the original blade.
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.
Large-Eddy Simulations of a boundary layer over a rotor blade are performed with and without inclusion of the rotational sources in the code. The numerical setup matches the one of a wind tunnel test available in the literature, and the numerical results are compared to each other and to the experiment. The mean boundary layers obtained from the simulations are studied by means of the linear stability analysis techniques with the aim of reproducing the transition location. It is shown that the non-rotating scenario, even when performed for the matching Mach and Reynolds numbers, predicts the transition location that is much farther downstream than the one seen in the experiment. However, inclusion of the rotational sources in the code moves the transition forward to the location that better agrees with the experiment. This shows that non-inertial forces associated with rotation play a crucial role in the transition in the considered setup. The character of the transition is different in two scenarios, being dominated by the Tollmien-Schlichting waves in the non-rotating case and by the travelling crossflow vortices in the presence of a laminar separation bubble in the rotating case. The linear stability analysis can reproduce the transition location in both cases, although the rotating scenario demonstrates noticeably lower amplification factors.
This paper presents the design, development, and subscale flight testing of an optionally-autonomous lift-plus-cruise (LPC) eVTOL aircraft for emergency response missions that bridges the gap between existing aerial capabilities and the needs of first responders. A 4+1 LPC configuration consisting of four vertical lift propellers and a single pusher propeller was selected to balance hover performance and cruise efficiency. The vehicle is sized around a 600 lbs gross takeoff weight with a 125 lbs payload capacity. VTOL and Pusher propeller blades were optimized using parametric studies, resulting in a high Figure of Merit and propulsive efficiency. Trim analysis demonstrates efficient hover to cruise transition, lift-to-drag ratios of 10-11 between 70-90 knots, and propulsive efficiency exceeding 0.9 at the cruise speed of 100 knots. The subscale configuration utilized a simulation framework for trim and optimization of flight control laws, which were subsequently implemented on a 1/3-scale subscale demonstrator. Subscale flight tests showed stable hover and robust trajectory tracking under wind disturbances throughout the flight envelope, which demonstrates the feasibility of the proposed LPC architecture.
Forward flight rotorcraft analyses typically require time-marching aeroelastic trim of coupled rotor-airframe models, which is expensive for repeated evaluations. This paper presents a non-intrusive model-order reduction framework based on Dynamic Mode Decomposition with control (DMDc) identified from snapshot data. A POD projection reduces the state dimension; the DMDc operators are identified in the reduced coordinates and used for fast time-marching. Two sequential maps are constructed: DMDc-A reconstructs aeroelastic sectional airloads from low-cost rigid-blade airloads, and DMDc-S predicts coupled deformation, including blade and airframe degrees of freedom (DOFs), from the reconstructed airloads. The method is demonstrated for the XV-15 airplane mode configuration using a stick airframe model and a coupled rotor-airframe solver. Over 160-400 knots, it is found that the surrogate reproduces blade airloads and structural deformation of blade and airframe.
This paper presents the investigation of experimental data belonging to main rotor loads during Never-Exceed-Speed demonstration of T625 Gökbey helicopter. Load data from the critical flight conditions in the VVNNNN envelope including cold-weather testing are collected. Maximum advancing tip Mach number demonstration, power-on and power-off flight conditions are investigated in terms of pitch link loads and blade loads. Blade loads including flapwise and chordwise bending moments, torsional moments and pitch link loads are examined to assess any divergence due to compressibility effects and the onset of stall. Load trends that are correlated with the tip Mach number are isolated from the effect of increasing dynamic pressure. Compressibility effects are observed to be the most dominant factor on the blade torsional moment and pitch link loads in advancing blade. The retreating blade stall phenomenon is apparent cases with a high advance ratio and mainly leads to dynamic stall cycles on the retreating blade, resulting in torsional moments and pitch link loads. Experimental results are compared with blade-resolved Unsteady Reynolds-Averaged Navier-Stokes simulations.
An experimental investigation was conducted to explore the loads, acoustics, and tip vortex trajectories of coaxial counter-rotating (CCR) rotor with unequal upper and lower radii. The upper and lower rotor radii were tested both at the nominal radius of 1.108 m, and also with a lower rotor radius of 90% nominal radius, for a constant rotor speed of 1180 RPM and a constant inter-rotor spacing of z/R = 0.108. Rotors were torque balanced and tested for a range of upper rotor collective pitch from -2◦ to 10◦ . The power required for both CCR systems was within 0.9% for most trim conditions, and equal thrust was produced at upper rotor collectives of 6◦ and 8◦ (within 1.0%). At low loading conditions the unequal radii configuration produced more thrust for the same power due to a reduction in profile drag. The overall sound pressure level (OASPL) was lower for the CCR rotor with shortened lower rotor blades at all angles of elevation. Larger reductions in A-weighted OASPL(A) were observed, due to a larger contribution of broadband noise to the total OASPL(A).
A technique for rapidly designing roughness tolerant low drag airfoils has been developed. Airfoils of varying thickness to chord ratio, ranging from 10% to 22% have been designed. A target pressure distribution is specified by the designer for a notional lift coefficient, Reynolds number, and Mach number. The specified pressure distribution is first analyzed using classical integral boundary layer analyses and empirical transition criteria for smooth and rough airfoils to ensure laminar flow over much of the airfoil under design conditions. The resulting airfoil is subsequently analyzed under natural transition, and forced transition caused by the tripping of the boundary layer due to roughness near the leading edge. It is found that the present approach performs well for a broad range of lift coefficients. An in-house propeller design and analysis tool has been used to examine the impact of the low drag airfoil on the pusher propeller performance designed for a fixed wing UAV drone configuration.
This paper presents the design and validation of the Rotor Optimization for the Advancement of Mars eXploration (ROAMX) Hover Test Stand, a vacuum compatible stand developed to measure rotor performance under Mars aerodynamic conditions. The stand integrates a high-speed water-cooled motor, a variable pitch hub, and a structural safety system capable of withstanding the high loads resulting from rotor blade loss while enabling continued experimental operations. The stand also maintains the measurement fidelity required for thrust and torque characterization at low Reynolds number. Aerodynamic blockage was limited to less than 20% through geometric constraints on the stand architecture. Calibration and sensor procedures ensured correct load transfer through the intended structural load path and also verified sensor accuracy. Test Entry 1 demonstrated successful test stand performance and testing of experimental blades. The stand spun blades to 4,000 Revolutions per Minute (RPM) for Test Entry 1, but the motor is capable of continuous operation at 5,900 RPM. The ROAMX Hover Test Stand provides a reusable experimental capability that enables high-fidelity evaluation of Mars rotor designs.
Urban Air Mobility (UAM) concepts require multidisciplinary analyses across multiple modes of operation and often involve discrete architectural differences such as propulsion type, rotor configuration, and mission context. Existing optimization and workflow frameworks support continuous design variables but provide limited mechanisms for handling discrete variants, multi-modal vehicle definitions, and vehicle management for UAM vehicles. This paper presents uam4x, an open-source Python framework that addresses these challenges through a structured problem definition representation, a plugin-based execution engine, integrated version control, and a function-based branching script mechanism for constructing analysis scenarios. The framework provides integration of existing tools including Open Vehicle Sketch Pad (OpenVSP), NASA Design and Analysis of Rotorcraft (NDARC), M4 Structures Studio (M4SS), and Intelligent Cross Section Generator (IXGEN) through unified plugin interfaces. Parameter sweeps, nested analyses, and optimization via OpenMDAO are supported within the same architecture. This paper also presents demonstrations that were created to illustrate the various capabilities and integration efforts of the framework.
This paper details comprehensive analysis modeling and analysis supporting the development of the Research Aircraft for eVTOL Enabling techNologies (RAVEN). An isolated rotor model was developed in CAMRAD II, and predictions of rotor performance and rotor aeroelastic stability were generated. The rotor stability predictions are part of assessing airworthiness of the RAVEN vehicle. The performance predictions were used to calibrate the surrogate model for the NASA Design of Rotorcraft (NDARC).
A comprehensive numerical study was conducted to reduce helicopter rotor hub vibratory loads and fuselage vibrations using the Higher Harmonic Control (HHC) technique. A CAMRAD II model of a medium utility helicopter was developed for aeromechanical simulation, and a linear system model representing both hub vibratory load and fuselage vibration characteristics was identified offline. Optimal control inputs were then computed to minimize vibration responses under different weightings on hub vibratory load and fuselage vibration in the objective function. The predicted performance was verified through CAMRAD II simulations. Additionally, a closed-loop HHC system incorporating actuator amplitude limitations was investigated. A control algorithm regulated actuator amplitudes while maintaining phase consistency, dynamically adjusting control inputs after each iteration. The results demonstrate that the amplitude-limited closed-loop control limits excessive pitch link loads while maintaining effective vibration suppression, thereby maintaining structural safety and ensuring practical feasibility for real-world helicopter applications.
The Rotor Optimization for the Advancement of Mars eXploration (ROAMX) project has demonstrated that rotor designs optimized for the Mars aerodynamic regime can provide substantial improvements in aerodynamic efficiency relative to heritage designs. This paper evaluates the vehicle-level performance implications of these improvements using the NASA Design and Analysis of Rotorcraft (NDARC) tool. Performance predictions for Ingenuity-class and Mars Science Helicopter (MSH)-class rotorcraft are generated using ROAMX rotor aerodynamic inputs and are compared against a configurations using the Ingenuity rotor. Parametric studies are conducted to investigate the trade between increasing payload mass and resulting changes in vehicle range and hover time, including the effects of rotor solidity and rotor type. The results show that ROAMX rotor designs enable significant increases in payload capability and operational range across both vehicle classes. These findings demonstrate how ROAMX rotor performance gains translate directly into enhanced mission capability at the vehicle level and provide a quantitative foundation for future Mars rotorcraft designs.
The induced and profile power of a hovering rotor was evaluated using experimental and computational methods. Momentum theory principles were coupled with experimental measurements over a range of thrust conditions to characterize the induced and profile power consumption at low Reynolds number conditions ∼ 105. An empirical induced power factor, κi, was extracted to quantify the non-ideal losses. Results show that these losses increase as the Reynolds number reduces, and nearly twice the power is required at Retip = 0.27×105 than the ideal momentum theory prediction. These results were compared with high-fidelity computational fluid dynamics simulations using the partial-pressure field (PPF) force/power decomposition to extract the induced and profile power contributions of the rotor. The PPF method decomposes the static pressure field of a numerical Reynolds-averaged Navier-Stokes solution into Euler and dissipative partial pressure fields. Simulations were performed across a range of thrust conditions, from which the induced power factor and profile drag coefficient,Cd0 , were computed for each simulation from the extracted power contributors.
Achieving noise reduction in rotorcraft requires an analysis of various design parameters and flight conditions. However, high-fidelity methods are computationally expensive. To overcome this limitation, reduced order model (ROM)-based surrogate models have been applied to aerodynamics and aeroacoustics prediction. This study proposes a ROM-based surrogate model employing a variational autoencoder (VAE) to predict rotor aerodynamic loads and associated noise. Train and test datasets were generated using reformulated vortex particle method across a wide range of flight conditions. The proposed framework was applied to a single rotor, and its performance was evaluated qualitatively and quantitively in comparison with proper orthogonal decomposition (POD)-based surrogate model. The results show that VAE-based model consistently outperformed the POD model in noise prediction. These results demonstrate that the proposed framework enables accurate rotor noise prediction under various flight conditions and provides a promising approach for low-noise rotorcraft design and operational planning.
Helicopter performance relies on well-designed rotor blades. This typically aims to improve two metrics: hover efficiency (Figure-of-Merit) and cruise efficiency (Lift-to-Drag ratio). This work extends a previous multi-objective optimization framework based on the University of Maryland's Aeromechanical Rotorcraft Analysis Code (UMARC-II) and Genetic Algorithm. That framework used aeroelastic analysis to design blades with non-linear twist, chord distributions, and spanwise airfoil selections, achieving significant aerodynamic gains within structural limits. In this study, we add vibration reduction as a third design objective. Helicopter vibrations peak during low-speed transitional flight, a regime dominated by unsteady aerodynamics that drives passenger discomfort and component fatigue. We calculate vibration index at low advance ratios and at the cruise speed and include it directly in the multi-objective optimization. The result is a tri-objective process that finds Pareto-optimal blade designs offering high Figure-of-Merit, high Lift-to-Drag ratio, and low vibratory hub loads. We also replace the linear inflow model used previously with the mid-fidelity Maryland Free Wake model, which accounts for vortex roll-up and wake distortion to give more accurate performance and load predictions. This extended framework generates blade designs that improve aerodynamic efficiency while reducing vibrations in both the critical transitional flight regime and at cruise speed.
A 4.75-ft diameter hingeless hub proprotor model was wind tunnel tested up to the very high speeds of 205 knots, loosely corresponding to 480 knots full-scale, with parametric variations in blades, wing spar, and pylon center of gravity. Testing revealed that a gimballed-hub configuration that reached whirl flutter at 160 knots was completely stabilized when converted to a hingeless hub – using identical blades, span, and pylon. While the gimballed-hub model encountered whirl flutter at 160 knots, the hingeless-hub configuration remained stable throughout the entire test envelope up to 205 knots. The key conclusions are that a hingeless hub can eliminate whirl flutter, and that the most stable configuration is a swept-tip blade hingeless-hub rotor with the pylon center of gravity aft of the wing spar.
This study examines the capability of medium-fidelity comprehensive analysis models to predict the acoustics for manned and unmanned rotorcraft configurations. Using the automated tool NDARC2RCAS developed at DEVCOM Army Research Laboratory, multiple configurations including a single main rotor, tilt rotor, coaxial and pusher, quadcopter, and hexacopter are evaluated at various mission segments including hover, advancing climb, and forward flight. Each configuration and condition is evaluated using a range of aerodynamic models from lower to higher fidelity, including uniform inflow, dynamic inflow, prescribed wake, free wake, and viscous vortex particle method (VVPM). These evaluations are then used with another automated tool, RCAS Acoustics, to predict noise on a Voronoi observer sphere. A comparison of the results for the single main showed good agreement between all of the aerodynamic models except VVPM. For the tilt rotor in forward flight, the higher-fidelity models produced changes in rotor loads due to the interaction with the wing. With prescribed and free wake models, this change in load is sharp and causes noise increases of up to 40 dB in front of and behind the vehicle, while the VVPM model produced a smoother change that results in a smaller, 20 dB increase in noise. The quadcopter and hexacopter show similar in-plane noise levels for all models, with alternating cancellation and amplification patterns due to rotor phasing, while out-of-plane noise is increased on the hexacopter when using the higher fidelity models.
This paper proposes a first iteration towards a framework for enhancing the trustworthiness of machine learning in the health and usage monitoring of in-service helicopters. This bottom-up approach is based on our experience operating machine learning models for monitoring Airbus Helicopters' customer fleets. Key factors for improving trustworthy machine learning have been identified for both the development and execution phases, with specific methods defined for each enabler. These methods have been implemented in two use-cases involving machine learning models for regression tasks: monitoring the helicopter's main gearbox lubrication system, deployed in the FlyScan predictive maintenance service, and tracking the usage of the main rotor lead-lag damper loads. The results from both use cases show that confidence in machine learning model predictions can be effectively improved.
With performance advances proposed for the Future Vertical Lift suite of aircraft and advancements in the electronic battlefield, it is imperative that advanced materials and concepts be included in the vehicle designs to meet the aggressive weight reduction objectives, structural requirements, and operational environment capabilities. Integrating electromagnetic (EM) shielding during the design process offers an opportunity to make progress towards the performance goals. To this end, efforts must be made to minimize the impact of this shielding to platform weight and structural performance. This article presents work to develop a hybrid multifunctional composite material technology that incorporates copper mesh into a carbon fiber and thermoplastic matrix structural composite material to achieve required levels of EM shielding and high levels of structural efficiency while reducing the overall weight of the system. This article focuses on the design of a representative helicopter tailcone as means of illustrating expected improvements in weight and manufacturing cost. Baselines that include structure fabricated from aluminum and thermoset composite are established and predicted improvements quantified.
With advanced air mobility (AAM) vehicles becoming an increasingly popular topic in aviation, the Eagle Flight Research Center (EFRC) at Embry-Riddle Aeronautical University continues to investigate control strategies that enhance aircraft resilience to total power unit failures. Utilizing a distributed electric propulsion (DEP) quad-heli test bed, the EFRC has explored a variety of control laws and hardware configurations to evaluate their effectiveness under failure conditions, including sustained flight with a completely inoperative rotor. The aircraft utilizes a fractional-order PID (FOPID) controller that has recently been developed and shown to outperform conventional PID controller used previously in both nominal and failure scenarios. The use of a FOPID controller offers improved stability and tracking performance. Another development is the implementation of a split-rotation rotor configuration—where the left-side rotors rotate clockwise and the right-side rotors rotate counterclockwise—which, when combined with a control law leveraging cyclic rotor inputs for yaw control, provides a robust alternative to conventional quadrotor designs. This configuration leverages the redundancy offered by equipping each of the aircraft's four rotors with full helicopter controls, enhancing the potential for sustained flight in the event of two rotor failures while preserving the maneuverability and controllability characteristic of the conventional configuration. Additionally, to support expanded testing and operational capability, the aircraft has been outfitted with a custom-designed landing gear system, designed and fabricated by the EFRC team.
Future military missions for Agile Combat Employment (ACE) and next generation Special Operations Forces need an aircraft with effective hover and the ability to operate in transonic cruise. Hover requires significant power that can only be mitigated by larger diameter rotors, but large diameter rotors become a detriment to achieving transonic flight. The stop-fold rotor configuration can “make the rotor disappear” in cruise and stands out as the most viable option for meeting these next-generation air vehicle requirements. This paper discusses the progress Bell has made in developing enabling technologies for a practical and scalable high-speed VTOL (HSVTOL) based on the stop-fold configuration. To this end, a unique Track-Guided Test Vehicle (TGTV) was developed at Bell and tested at the 10-mile High Speed Test Track at Holloman Air Force Base. The test vehicle integrates all subsystems required to demonstrate the key technologies in a representative environment, including multi-mode propulsion, folding proprotors, and fly-by-wire transition controls to automatically manage each step of the prop-to-jet (and reverse) process. The TGTV demonstration validated the stop-fold technical approach and generated critical data to substantiate engineering models. Test results provide confidence that each step in the stop-fold transition process is thoroughly understood and that engineering tools can be used confidently for future aircraft design efforts. This paper documents the first known powered prop-to-jet and jet-to-prop transition and represents a significant milestone in vertical flight technology development.
This work presents an extension of previous studies on the analytical linearization of state-space vortex particle methods (VPM) by incorporating viscous effects. The developed framework couples a panel method to capture blade surface and near-wake aerodynamics with a viscous vortex particle method (VVPM) for modeling the far-wake. The resulting formulation yields a nonlinear time-periodic (NLTP) system described by ordinary differential equations (ODEs) in first-order form. To enable linear analysis, the NLTP system is linearized into a linear time-periodic (LTP) representation using two techniques: finite differencing and a novel analytical linearization approach. Harmonic decomposition is then applied to transform the LTP system into a higher-order linear time-invariant (LTI) model, enabling the use of time-invariant analysis tools. The methodology is implemented in MATLAB® and applied to a representative utility helicopter rotor blade. Validation is performed against experimental data, and the linearized models are evaluated through time and frequency domain comparisons with the nonlinear system. Results demonstrate that the linearized models accurately capture unsteady wake dynamics, particularly under low to mid-frequency excitation. Moreover, the analytical linearization method significantly reduces computational cost – achieving an efficiency improvement of O(n2), where n is the number of system states – compared to finite-difference approaches.
A joint acoustic flight test was conducted by NASA Langley Research Center and the U.S. Army Combat Capabilities Development Command Aviation & Missile Center, with the goal of investigating new methods for acoustic data collection. The impetus for the effort is the anticipated growth of Urban Air Mobility and Future Vertical Lift vehicles. Many of these vehicles are expected to have distributed propulsion systems that may result in unsteady vehicle state conditions even during steady flight. This work examines the acoustic measurements collected during purposefully unsteady maneuvers performed by an MD530F helicopter. A snapshot microphone array design was deployed for this test to capture the acoustic signature on the ground from the helicopter under maneuver conditions. An analysis of the acoustic emissions indicated the presence of blade-vortex interactions, not only during the rolls towards the advancing side of the main rotor, but also rolls towards the retreating side and during pitch-up maneuvers. The strength of the interaction noise was found to be strongly dependent on the acceleration of the vehicle.
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.
Wind tunnel tests and comprehensive rotorcraft analysis were carried out on a slowed main rotor full-wing lift and thrust-compounded helicopter with a trailing propeller to investigate the effects of rotor and wing configuration on performance, blade structural loads, and hub vibratory loads. Experiments were conducted at advance ratios up to 0.7, incorporating three full-wing configurations with symmetric and asymmetric incidence angles and three different rotor shaft tilt angles. Propulsive thrust was measured by a trailing pusher propeller with its own balance system. The wind tunnel test data was used to validate the University of Maryland Advanced Rotorcraft Code (UMARC). Results showed that the maximum lift-to-drag ratio is achieved using either of the symmetric or asymmetric full-wing lift-compound configurations with high lift offloading and aft shaft tilt. Both blade structural loads and hub vibratory loads are significantly reduced when rotor lift is offloaded to the wings. Lift compounding enables efficient high-speed flight, with a high lift-to-drag ratio of 7.5 maintained at an advance ratio of 0.7. Lift and thrust-compounded rotorcraft propulsive trim was achieved up to an advance ratio of 0.5, demonstrating that efficient high-speed flight, with peak lift-to-drag ratios exceeding 4.5, is enabled through lift offloading.
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