Browse Topic: Propellers and rotors
This SAE Aerospace Recommended Practice (ARP) recommends a methodology to be used for the design, analysis and test evaluation of modern helicopter gas turbine propulsion system stability and transient response characteristics. This methodology utilizes the computational power of modern digital computers to more thoroughly analyze, simulate and bench-test the helicopter engine/rotor system speed control loop over the flight envelope. This up-front work results in significantly less effort expended during flight test and delivers a more effective system into service. The methodology presented herein is recommended for modern digital electronic propulsion control systems and also for traditional analog and hydromechanical systems.
This SAE Aerospace Information Report (AIR) outlines a recommended procedure for evaluation of the vibration environment to which the gas turbine engine powerplant is subjected in the helicopter installation. This analysis of engine vibration is normally demonstrated on a one-time basis upon initial certification, or after a major modification, of an engine/helicopter configuration. This AIR deals with linear vibration as measured on the basic case structure of the engine and not, for example, torsional vibration in drive shafting or vibration of a component within the engine such as a compressor or turbine airfoil. In summary, this AIR discusses the engine manufacturer’s "Installation Test Code" aspects of engine vibration and proposes an appropriate measurement method.
This document provides recommendations involving BEV battery data retention and battery design that enhance the potential for BEV battery reuse and serviceability and that can improve recyclability. These recommendations have been developed by a group of professionals skilled in the secondary-use of batteries and in the research, development, and manufacture of BEV batteries and battery systems.
Rotorcraft airfoils often feature a tab which aides in the manufacturing of composite rotor blades, but also has aerodynamic merits. This study performs a comprehensive analysis of the impact of this tab on the 2D airfoil performance, structural adjustments and 3D rotor performance. The aerodynamics are evaluated using CFD, with CFD/CSD coupled results for the rotor performance. The structural data is adjusted using an FEM based in-house process. The HART II model rotor has been taken as a baseline and modified according to the tab variation studies. These included the comparison of a sharp trailing edge versus a tabbed airfoil, various tab thicknesses, lengths, and angles. The studies showed a variation of peak Figure of Merit between 66% to 68% and peak rotor L/D from 4.2 to 4.6 The careful design of the airfoil tab is therefore advised, but similarly the structural design of rotor blades.
The TiltRotor Aeroelastic Stability Testbed (TRAST) was developed to experimentally investigate whirl-flutter stability of tiltrotor aircraft. Previous wind-tunnel testing focused on configurations representative of current generation tiltrotors utilizing gimballed rotor hubs. The TRAST platform was also designed to support a hingeless rotor system to investigate whirl-flutter mechanisms representative of stiff proprotor configurations. This paper presents analytical whirl-flutter predictions for a hingeless rotor configuration of the TRAST model. Structural mode shapes derived from a NASTRAN finite-element model are combined with comprehensive aeroelastic analyses in CAMRAD II and RCAS. The results show that the dominant whirl-flutter mechanism differs from the gimballed configuration, with instability occurring through the wing in-plane mode rather than the wing vertical bending mode. Parametric studies examining rotor speed, pitch-spring stiffness, rotor flexibility, and diaphragm spring stiffness are conducted to evaluate the sensitivity of the predicted stability boundary. Results indicate that the hingeless configuration is significantly more stable than the equivalent gimballed configuration and exhibits different trends with rotor speed and structural stiffness. These predictions help identify configurations of interest for future wind-tunnel testing and provide insight into whirl-flutter mechanisms for hingeless tiltrotor systems.
This paper presents a study of gunshot acoustic signal detectability in the near field of propeller noise, with a focus on the isolation of external gunshot signatures masked by propeller-induced noise. Controlled measurements were conducted in a Recirculation Delayed Anechoic Chamber (RDAC), where acoustic data were collected across varying rotor speeds, source locations, and propagation distances. Propeller noise characteristics were verified using UCD-QuietFly. The recorded signals were analyzed for the acoustic pressure, sound pressure level, and overall sound pressure level directivity to quantify masking effects. Results show that RPM is the dominant factor governing signal detectability. At 3000 RPM, the gunshot signal remains clearly identifiable within the low frequency range of 200–2000 Hz. At 4000 RPM, the signal becomes partially masked, while at 5000 RPM, propeller noise fully dominates and the gunshot signal becomes undetectable. Detectability is further reduced with increasing propagation distance. In-plane microphone locations provide improved detectability. A machine learning-based spectral separation framework was developed to suppress propeller noise and enhance the visibility of impulsive gunshot signatures in multichannel spectrograms. Experimental results show that learning-based denoising is effective at lower RPMs where the signal-to-noise ratio remains favorable, but performance degrades as broadband masking intensifies at higher rotor speeds.
A multi-objective optimization of a rotor blade airfoil is presented using compressible unsteady Reynolds-averaged Navier-Stokes simulations directly within the optimization loop. The baseline SC1095 airfoil is optimized using NSGA-II with two objectives: pre-stall aerodynamic efficiency representing hover performance, and lift hysteresis loop area representing dynamic stall severity. The optimized airfoil exhibits increased maximum thickness with an aft-shifted crest and substantially higher camber. Static polars show improved lift-to-drag ratio at $Ma = 0.5$ and $0.6$. Hover performance is essentially unchanged relative to the baseline. In forward flight, a progressive power penalty is incurred above $\mu = 0.2$, attributed to higher profile drag at advancing blade Mach numbers. Dynamic stall simulations show an 80% reduction in peak drag and a 50% reduction in peak pitching moment excursion relative to the SC1095, demonstrating the effectiveness of the optimization for retreating blade conditions.
This paper presents a higher harmonic control (HHC) simulation by using a coupled actuator line method (ALM) and computational structural dynamics (CSD) framework. The ALM is implemented within an in-house computational fluid dynamics (CFD) solver based on a compressible large-eddy simulation (LES) approach DYMORE, a multibody dynamics software, is utilized as the CSD solver. The proposed framework is validated by the trim analysis of UH-60A rotor system under the Flight Counter 8513 flight condition, which represents a lower-speed forward flight maneuver. Using such trimmed state as a baseline, comprehensive HHC simulations of the UH-60A rotor will be executed. The orthogonality and linearity of the HHC input are evaluated, and their control authority over various rotor hub load components is investigated. Furthermore, the effect of the HHC input on the vibratory rotor blade and pitch link loads will be analyzed.
Stacked co-rotating rotors offer a mechanically simple alternative to conventional coaxial counter-rotating systems, but their aerodynamic performance is strongly dependent on both axial and azimuthal blade spacing. This study experimentally and numerically investigates the effects of rotor spacing on the performance and wake structure of model-scale stacked rotors in hover. A dedicated test platform was developed to measure thrust, power, and phase-resolved 2D-3C particle image velocimetry flow fields for two-bladed stacked rotors over axial spacings of Δz/c=0.75 to 5 and azimuthal spacings of ϕ = 0° and 90°. Relative to isolated two- and four-bladed baseline rotors, the stacked configurations exhibited measurable variations in total hub loading and induced flow structure as a function of spacing. The flow field results show that changes in axial spacing alter the relative position of the lower rotor within the convected wake of the upper rotor, producing corresponding changes in inflow, effective angle of attack, and total thrust. Azimuthal spacing further modifies these trends by shifting the phase relationship between upper- and lower-rotor blade passages. To interpret these effects, a coupled blade element momentum theory and Blade Interaction Prediction model was developed. The model captures the primary trends in measured thrust and sectional loading, demonstrating that both wake convection and chordwise blade interaction are required to predict the aerodynamic behavior of closely spaced stacked rotors.
A new approach that enables the synthesis of fully coupled system dynamics is described in this paper. The approach facilitates collaboration between solver developers by explicitly avoiding inter-code coupling and instead uses a generic interface that enables inputs to be set and flags outputs available to other solvers. The assembly of a fully coupled linearized system matrix is obtained entirely from the existence of coupling maps, and does not rely on user intervention. Tiltrotor whirl flutter in cruise conditions is systematically investigated by careful examination of results obtained through various combinations of domain synthesis and obtained from the Hermes coupling framework. Investigated solver domains include nonlinear rotor dynamics, nonlinear aerodynamics, and linear structural dynamics. Utilized software modules include RCAS; Project Chrono, a purpose-built lifting line aerodynamics solver; and FuselageSolver for linear structural dynamics. Aeroelastic predictions are synthesized and verified for a pitch/plunge airfoil and a low-speed wing. Rotor aeroelastic effects are verified by coupling rotor dynamics with rotor aerodynamics. Structure-to-structure results are synthesized by coupling both RCAS and Project Chrono rotor dynamics models to a pylon structure modeled in FuselageSolver. Finally, full whirl flutter predictions are synthesized by coupling Project Chrono, lifting line aerodynamics, and FuselageSolver, as well as an RCAS model without the wing/pylon structure to FuselageSolver. Results indicate that the new synthesis approach is viable and accurate.
The Enhanced Tiltrotor blade, also known as the RGF3 blade, represents a major milestone in Leonardo Helicopters Division's pursuit of advanced rotorcraft technology. Developed at the Yeovil facility in the United Kingdom as part of a dedicated program and in collaboration with the European Clean Sky 2 initiative, it is a key enabler for the Next Generation Civil Tiltrotor Technology Demonstrator. Leveraging the AW609 airframe, the NGCTR integrates a new lateral rotor control system and a V-tail with ruddervators to expand maneuverability and control authority. The RGF3 blade combines aerodynamic efficiency with manufacturability, cost effectiveness, and certification readiness. Innovations include advanced airfoil families, highly swept anhedral tips, dual-redundant anti-ice systems, and full compatibility with legacy components. A comprehensive test campaign—covering structural loads, lightning and bird strikes, icing, and wind tunnel validation—confirmed its robustness and performance. The RGF3 blade embodies Leonardo's vision for high-speed, sustainable, and reliable next-generation rotorcraft.
Shrouded rotor configurations provide aerodynamic efficiency benefits that are valuable for aerial vehicles requiring high endurance in hovering flight. The effect of the shroud on the loads produced by a vehicle in response to cyclic pitch are not well known in literature. This paper aims to understand the magnitude and phasing of the steady hub loads produced by shrouded coaxial hingeless rotors in hover due to cyclic pitch. A shrouded coaxial rotor vehicle weighing 0.934 kg and with a rotor radius 0.138 m is used as the test platform. The rotors are operated at a tip Reynolds number of 70,000 and a tip Mach number of 0.18. First, the origins of steady hub loads due to cyclic pitch on a hingeless rotor are described using theory. To quantify the effect of the shroud, different configurations of rotors were tested by giving cyclic pitch to only one rotor. For the open hingeless rotor, cyclic variation in drag force at blade sections contributed to 55% of the steady lateral force. In response to cyclic pitch input, the ducted configurations produced 35% larger lateral forces and 60% (coaxial rotors) to 100% (single rotor) larger lateral hub moments than those produced by the unshrouded configurations. The presence of the shroud also changed the phase angle between the forces and moments. While the lateral force and moment acted in similar directions for the open configurations, the angle between them increased by more than 60° when enclosed in a shroud.
In this study, a novel rotorcraft comprehensive analysis (CA) framework is constructed for the aeromechanics analysis of various rotorcraft configurations such as single main/tail rotors, coaxial rotors, and tilt rotors. The framework incorporates numerous solution procedures that include trim, blade response, loads, and vibration with external interfaces to computational fluid dynamics (CFD) analysis. The structural module is characterized by a displacement-based geometrically exact beam (GEB) model and multibody formulations while the internal aerodynamics module is developed using a lifting-line representation of blades with simple inflow models. A series of validation on static benchmark examples showed excellent correlations with published records, particularly in the context of large deformation behavior of heavily loaded beams. The rotorcraft aeromechanics analyses of various configuration rotors such as single main rotor and lift-offset coaxial rotor types are performed next to evaluate airloads, blade motions, and structural loads. Numerical results indicate good to excellent agreements with their respective test data accurately capturing unsteady aerodynamics effects due to blade-vortex interaction or blade-to-blade cross-over events, which demonstrate the reliability and versatility of the present rotorcraft CA system.
This paper introduces an eigenvalue-based whirl flutter prediction method accounting for aerodynamic interactions between a wing and propeller. The linearized unsteady vortex lattice method was utilized to model fixed-wing aerodynamics while the linearized viscous vortex particle method was utilized to model rotary-wing aerodynamics. The complete aerodynamics model was then coupled with computational structural models to demonstrate the capabilities of the model to predict whirl flutter using an eigenvalue-based method. Two computational structural models were used: the first being an analytical propeller model affixed to a rigid wing via root springs and dampers, and the second being the University of Michigan's Nonlinear Aeroelastic Simulation Toolbox. These models demonstrate the capabilities of the linearized aerodynamics model in predicting instability with structural models of different fidelities, both considering and not considering aerodynamic interactions. The linearized aerodynamics model predicts a reasonable aeroelastic solution when coupled with the structural models, but requires more investigation as to whether aerodynamic interactions are being sufficiently captured.
This study presents subscale wind tunnel experiments investigating the transient aerodynamic interactions of a rotor during a continuous relative-wind approach toward the landing deck of the NATO Generic Destroyer. Time-resolved rotor loads and stereoscopic particle image velocimetry measurements were used to characterize the interacting ship-rotor flow field under headwind and quartering wind-over-deck conditions. The measurements captured the evolving influence of ship airwake, ground effect, and superstructure-induced recirculation as the rotor moved from downstream to the final hover position over the deck. The results show that rotor thrust, rolling moment, and pitching moment underwent distinct changes throughout the approach, with the loading trends varying significantly with wind-over-deck angle. Time-frequency analysis further reveals that the unsteady rotor response was concentrated in a limited band of frequencies associated with various coherent flow structures shed from the ship superstructure. Spectral proper orthogonal decomposition was used to identify the dominant airwake features responsible for these fluctuations at specific frequencies, including large-scale structures originating from the radar and hangar region. These findings demonstrate that the dynamic approach resolved both slow- and fast-changing transient aerodynamic effects along the approach path that cannot be captured with static hovering measurements.
This study presents the development and evaluation of two multi-fidelity surrogate models for predicting the first blade-passage frequency acoustic directivity of a propeller-wing configuration across a parametric sweep of wing leading-edge positions. The configuration follows the experimental setup at NASA Langley Research Center, comprising a three-bladed Mejzlik propeller operating upstream of a NACA 632-215 MOD B wing at 40 discrete leading-edge positions spanning the horizontal and vertical parameter space. Medium-fidelity predictions from the Rotorcraft Comprehensive Analysis System (RCAS), using the Viscous Vortex Particle Method, serve as the low-fidelity input, while high-fidelity predictions from NASA's OVERFLOW solver coupled with PSU-WOPWOP, both serve as the training datasets. QR decomposition is employed in both frameworks to identify the most informative subset of wing positions for high-fidelity simulation. The first surrogate model, a linear regression model, requires 15 high-fidelity snapshots to outperform the RCAS predictions, achieving l∞ norm, l2 norm, and RMSEs below the RCAS baselines of 32.98%, 8.56%, and 5.67 dB, respectively, over the entire parameter space. The second surrogate model, a Co-Kriging surrogate model implemented via the Surrogate Modeling Toolbox 2.0, requires only seven high-fidelity snapshots and achieves an RMSE of 4.45 dB and an R² value of 0.67. Both surrogate models accurately reproduce the high-fidelity directivity at wing positions where the CFD and RCAS trends are in reasonable agreement, but struggle at positions where the two solvers predict substantially different trends, highlighting the sensitivity of multi-fidelity methods to solver discrepancies and the potential need for denser parameter space sampling in regions of high acoustic variability.
The paper presents the successful drag reduction of the Racer demonstrator's rotor head through its innovative full fairing, based on a robust de-risking methodology leveraging 2D Robust Design Optimization (RDO) for airfoils, 3D CFD analysis with multiple fidelity levels, and experiments. We provide a unique end-to-end comparison across the full development cycle, correlating simulation predictions with both experimental and flight-test data. The fully faired architecture achieves a significant 42% reduction in rotor-hub form drag. At the full-vehicle level, flight tests confirm a 10% net drag reduction, including complex interactions with the airframe. This real-world measurement correlates highly with dynamic URANS predictions (11-12%), while effectively contextualizing the more optimistic 16% gains observed during static wind-tunnel and steady RANS evaluations. These findings provide a comprehensive validation of the low-drag fairing concept, offering valuable insights for the aerodynamic design of future high-speed rotorcraft.
A challenge in establishing rotor performance map for sizing tool during design cycle is the rotor performance uncertainty for full vehicle. Sometimes, simplified tests at different setup/scale are conducted to guide performance map, but this introduces another uncertainty due to configuration difference from full vehicle. To aid insights, validated computational fluid dynamics simulations (using CREATE-AV™ Helios) were carried out to examine hovering rotor performance prediction variations at different design stages, or different modeling/testing setup with identical blade design. Quantitative rotor figure of merit differences has been demonstrated along with descriptions of underlying physical reasons. The examined model setup includes isolated rigid blades with and without flapping, elastic blades, model-scale blades, whirl-tower conditions, blades installed on fuselage, and full-vehicle including tail rotor. Both fully turbulent flow and laminar-turbulence transition flow assumptions were simulated. Rigid blades showed a negative impact on performance due to the lack of nose-down elastic twist. Model scale suffered from lower Reynolds number effects but took advantage of delayed laminar-turbulence transition. Whirl-tower blockage and altered vortex trajectories reduced peak figure of merit but delayed stall. Fuselage installation increased performance with partial in-ground effect. However, the tail rotor disrupted the main rotor vortex system and caused a substantial figure of merit drop. The figure of merit variation summary from current study can provide qualitative trends and rough estimate of the rotor performance change along different analysis or test condition during new design process. Well validated computational fluid dynamics simulations can be used as a risk-reduction approach by comparing with simplified-model results (used during fast design cycle) with the expected full-vehicle model results.
This study develops an efficient framework coupling the Lattice Boltzmann Method with the Actuator Line Method to evaluate the unsteady downwash/outwash of eVTOL aircraft. By incorporating a modified Prandtl loss function with geometric smearing correction, the framework accurately predicts velocity profiles and preserves unsteady vortices at lower computational costs than conventional RANS-based simulations. Analyzing three distinct eVTOL configurations sized for identical payload missions reveals that higher disk loading and multi-rotor interactions generate highly asymmetric, localized jet-like outwash structures, contrasting with the symmetric ground-level footprint of single rotor designs. Utilizing a 95th percentile velocity metric, transient peak hazards breach the regulated vertiport Safety Area boundary, extending up to 1.65 times the prescribed baseline limit. These findings demonstrate that time-averaged metrics underestimate physical hazards, highlighting the necessity for future guidelines to mandate configuration specific evaluations utilizing high-resolution unsteady flow data and robust statistical processing.
The National Research Council of Canada and International Test Pilot School collaborated in the experimentation with Supervised Deep Learning methods for prediction of high-order rotor dynamics and blade loads. Using an intentionally limited flight dataset (431 samples, 6.7 seconds) from the NRC Bell 412 Advanced Systems Research Aircraft (ASRA), an evaluation framework was developed to expose limitations of conventional validation practices. Latin Hypercube Sampling is used to improve training data coverage, restoring temporal generalization and yielding positive (R2) across all eleven rotor quantities, including hub dynamics and blade bending at two spanwise stations. Peak results include beam bending at R2 = 0.94, (CNN, (r=0.97)) and inboard chord bending at R2 = 0.81, (r = 0.90). Ensemble averaging further improves temporal-split performance from R2 = 0.33 to R2 = 0.64 at no additional cost. Results provide guidance for rotor load estimation under data-constrained conditions.
A generalized turbulence model for rotorcraft, analogous to the Dryden or von Kármán models commonly used for fixed-wing aircraft, does not yet exist. The closest available formulations are Control-Equivalent Turbulence Input (CETI) models, which reproduce the portion of the vehicle response attributable to atmospheric disturbances through equivalent control inputs applied at the inceptor or effector level. While the underlying concept is broadly applicable, these models are highly configuration dependent, making their broader generalizability uncertain. This paper adopts a recently developed methodology to extract CETI models directly from simulation and extends it to the identification of State-Equivalent Turbulence Input (STETI) models, which are Dryden-like in form and inject turbulence-equivalent excitations directly into the state dynamics. The approach is applied to six conventional main-tail rotor helicopters spanning trainer, light, intermediate, medium, heavy, and ultra-heavy classes, from the Robinson R-22 to the Mil Mi-26, and to four tiltrotors spanning the current range of tiltrotor weight classes, from the Bell XV-15 to the Bell Boeing V-22. The resulting CETI and STETI models are used to examine cross-vehicle trends, assess the extent to which they generalize across configurations, and evaluate whether they can be parameterized in terms of fundamental rotorcraft properties, including gross weight and rotor radius, as well as derived parameters such as disk loading.
This paper utilizes a combined experimental and modeling approach to investigate techniques for improving the forward-flight roll-control authority of a Quadrotor Biplane Tailsitter (QBiT). QBiT is a mechanically simple, efficient hover/cruise aircraft whose roll authority in forward flight is traditionally limited by differential propeller-torque-based control. The two roll-control enhancement techniques investigated are propeller canting and the use of ailerons. A 2-kg instrumented QBiT platform was developed and flight tested to collect high-fidelity flight data across multiple flight regimes including hover, transition, cruise, and coordinated turns. A flight dynamics model was developed and validated using wind tunnel measurements and flight-test data. Flight tests showed that the cant-only configuration exhibited limited roll authority during coordinated turns due to motor control saturation, whereas the cant-plus-aileron configuration provided improved roll performance. Using test data from forward-flight roll excitation maneuvers, roll-control authority was evaluated both in the time domain and frequency-domain. The results showed that adding ailerons increased forward-flight roll-control authority by about 2.3 times from analyzing the flight data and by up to 2.6 times based on the flight dynamics simulations.
This study evaluates the capability of Simcenter™ Flightstream™, a viscous surface-vorticity compressible-flow panel method, for predicting aerodynamic performance of rotorcraft configurations. Simulations are performed on the ROBIN-mod7 fuselage, PSP rotor, and combined rotor-fuselage system under conditions consistent with available experimental data. Results are compared against experiments and high-fidelity CFD methods, including DES, URANS, and IBM-ASM. For the isolated fuselage, Simcenter Flightstream accurately captures surface pressure distributions, particularly in attached flow regions. For the isolated rotor, thrust, torque, and figure of merit trends show strong agreement with reference data. In the rotor-fuselage configuration, the solver successfully captures interaction effects and predicts performance within experimental uncertainty. Notably, Simcenter Flightstream achieves these results with one to two orders of magnitude lower computational cost compared to high-fidelity approaches. The findings demonstrate its effectiveness as a rapid, cost-efficient tool for early-stage rotorcraft aerodynamic design and analysis.
Historical rotor designs for Earth and Mars have typically landed at thrust-weighted solidities of ∼0.1-0.15 as a best compromise of performance and weight. Comprehensive analysis predicts that high solidity rotor designs of more than twice this range have the potential to significantly increase the lift capability of future Mars explorers severely limited by packaging and weight. However, there is limited existing experimental data of high solidity rotor designs at representative densities to quantify the efficiency impact and verify models of the aerodynamic environment. Therefore, the Mars Exploration Program (MEP) funded a joint test campaign between NASA's Jet Propulsion Laboratory, NASA Ames Research Center, and AeroVironment, Inc.to validate performance predictions for low- and high- solidity rotor variants at Mars pressures. Experimental setup, test matrix, data processing, data quality, and performance results for the High Solidity Test (HST) campaign are presented and discussed.
Wake measurements were performed on a 2-m diameter rotor in forward flight at an advance ratio of 0.2 undergoing sinusoidal collective and cyclic inputs using 2D-3C phase-resolved PIV. Input frequencies of 0.05/rev, 0.1/rev, 0.2/rev, and 0.4/rev were tested. The goal of this study was to characterize the time-varying rotor wake and extract Pitt-Peters dynamic inflow model parameters for the lateral cyclic inflow state. In both input cases, the effects of the pitch inputs manifested as modulation of the local upwash and downwash of the trailing tip vortices near the tip path plane. It was found that additional azimuthal measurements are necessary to improve the extracted value of the steady Pitt-Peters term. However, the extracted mass term was within 12% of the Pitt-Peters value, demonstrating the ability of the presented analysis to resolve the dynamics of the wake with a limited number of azimuthal measurements.
This study investigates the dynamics and associated vibratory loads of an underactuated swashplate-less rotor and its impact on the flight dynamics of a small-scale helicopter powered by this rotor via a combined experimental and computational approach. Unlike prescribing cyclic pitch using a swashplate, here the pitch is a response to the 1/rev cyclic rotor speed input. This is enabled on the current two-bladed rotor using a skewed lag hinge that utilizes the cyclic speed variation to produce lagging motion and subsequently pitching the blades in a cyclic fashion (ƍ4 coupling) for generating the pitch and roll control moments. One of the key dynamic characteristics that distinguishes this rotor from a conventional swashplate-controlled rotor is that the two blades have dissimilar pitch and flap responses leading to high fixed-frame vibratory loads. Results show that a large 1/rev vertical shear force is transferred to the fuselage resulting in half-peak-to-peak loads of +/−0.67g. The flap responses of the two blades being out of phase caused a net inertial force, which was the dominant contributor to the vibratory vertical shear force. Upon removal of the flap hinge, the vertical shear force was reduced by almost 85% while increasing the control moment authority by 50%. The inertial moment about the rotor axis generated by cyclic lagging nearly balances the moment due to the angular acceleration (Ω ) of the rotor, which significantly reduced the dynamic torque requirement from the motor.
Flight simulations are critical for aerial firefighting training, but realistic modelling of aircraft-atmosphere interactions within fire scenarios is particularly challenging. To this end, a two-way-coupled flight simulation system, the Daedalus I framework, has been developed at the University of Glasgow for helicopter firefighting research applications. This paper presents the initial results from flight experiments conducted with different coupling schemes between the rotorcraft model and the GPU-accelerated Lattice Boltzmann atmosphere model within the system. The two-way coupling scheme was first validated using an isolated, transient rotor case. To quantify differences in pilot control and strategy between the two-way, fully-coupled rotor-atmosphere method and two (2) one-way, superposition-based coupling methods, a series of flight experiments were conducted using the bimodal modification of the McRuer pilot model representing human pilot controls, in conjunction with objective performance metrics. The results highlighted noticeable changes in helicopter behaviours and pilot responses across the three tested coupling methods, with the two-way coupling method showing the most resistance to the generated fire disturbance.
This paper investigates the impact of aerodynamic interactions on the dynamic aeroelastic stability of a wing-propeller configuration, with emphasis on whirl flutter. The wing structural dynamics are modeled using linear Euler-Bernoulli beam finite elements, while the propeller is represented using Reed's two-degree-of-freedom model. Baseline stability analyses neglecting aerodynamic interactions employ strip theory for the wing and the Houbolt-Reed formulation for the propeller. Analyses that account for aerodynamic interactions are then performed by coupling the wing and propeller structural models with the unsteady vortex-lattice method. Whirl flutter points are identified from transient simulations under both thrusting and windmilling conditions. Results show that three-dimensional aerodynamic effects increase the whirl flutter speed, whereas wing-propeller aerodynamic interactions play a slightly destabilizing role. Thrusting conditions produce a lower critical speed than the wind-milling case. The results demonstrate the viability of the unsteady vortex-lattice method as a unified aerodynamic framework for aeroelastic stability analysis of wing-propeller systems with mutual aerodynamic interactions. In addition, they reinforce findings from previous work that highlighted the destabilizing role of wing-propeller aerodynamic interactions.
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.
The Vortex Ring State (VRS) is an intriguing phenomenon where rotary wings are trapped in their own wake. It is inherently difficult to model with the classic momentum theory due to the breakdown of slipstream assumptions. In practice, it is still a critical safety concern for helicopters and emerging multi-rotor platforms. Despite extensive wind tunnel tests, flight tests, and modelling over the past decades, our quantitative understanding of the underpinning flow details is still limited, because of limitations in measurements and modelling resolution. First-principles-based, high-resolution simulations could uncover the flow details, but the modelling is still rare and challenging due to complexities in the flow and flight physics, and particularly the associated high computational costs. Nevertheless, in this work, a series of high-resolution simulations of the VRS phenomenon are presented. Fully blade-resolved and unsteady simulations of an isolated helicopter rotor within the VRS were carried out for over 20 revolutions. The simulations were performed using the Helicopter Multi-Block 3 (HMB3) CFD framework developed at the University of Glasgow, solving the fully compressible and unsteady Reynolds-averaged Navier-Stokes (RANS) equations with a Scale-Adaptive Simulation (SAS) closure. The modelling revealed the formation of the large vortex ring stemming from the discrete tip vortices, and its evolution over the 20 revolutions. The blade loading distribution and evolution were analysed and compared with test data. Moreover, we extracted the inflow features from the CFD results via direct extraction and inverse Blade Element Theory (BET). The results highlighted the strong induction of the vortex ring, and the influence of secondary flow features besides the inflow. It was also noted, with the correct inflow information, the BET was able to reconstruct the VRS loading with reasonable accuracy. These high-fidelity results provide unique insight into the flow and flight physics underpinning the VRS, and contribute to the ongoing GARTEUR AG28 collaboration on multi-rotor VRS investigation.
This paper presents a wind tunnel investigation on the interactional aerodynamics of a slowed-rotor lift- and thrust-compounded helicopter model in high-speed forward flight. A systematic configuration study was conducted to isolate the aerodynamic contributions of the main rotor, wings, fuselage, and pusher propeller to the aft flowfield, measured using phase-resolved 2D-3C particle image velocimetry. Measurements were acquired at an advance ratio of 0.5 across multiple rotor thrust levels, lift offset trim states, and propeller rotational speeds. The fuselage induces a streamwise velocity deficit of nearly 50% of the freestream near the tail boom due to oncoming flow blockage. This deficit is modulated by the main rotor and wing configurations. The rotor slipstream partially alleviates the deficit by convecting high-speed freestream flow downwards. Lift offset in the asymmetric half-wing configuration suppresses the rotor wake influence, deepening the velocity deficit relative to a conventional rotor trim state. The pusher propeller partially recovers the streamwise velocity deficit, with propeller performance improving in proportion to the magnitude of the deficit due to reduced climb inflow and increased blade sectional angle of attack. These findings demonstrate that the velocity gradient aft of a compound rotorcraft impacts pusher propeller performance, highlighting the importance of flowfield-informed propeller design for improved performance in high-speed forward flight.
Electric Vertical Take-Off and Landing (eVTOL) vehicles are emerging as solutions for urban air mobility, but their operation can encounter hazardous aerodynamic conditions such as the Vortex Ring State (VRS), which causes thrust loss and intense vibrations. This study investigates VRS for the Archer Maker tilter propeller by combining numerical simulations using the mid-fidelity solver DUST and the high-fidelity solver OVERFLOW with prior experimental observations. Propeller performance is evaluated through thrust and torque evolution under various descent conditions, while flow fields in the propeller wake at different descent ratios around VRS conditions are evaluated and compared via 2D visualizations. A comparison of results reveals that both numerical approaches are capable of evaluating the performance degradation in correlation with vortex ring formation within specific descent regimes, showing slight discrepancies particularly regarding the descent ratio regime where VRS occurs. A flow field comparison with experimental data validates the multi-fidelity numerical approach, showing the capabilities of both numerical approaches to capture the flow physics mechanisms that lead to the generation of the vortex ring around the propeller disk.
This work evaluates the long-term fatigue life and structural compatibility of integrated optical fiber sensors (OFS) within an H145 (or BK117 D-3) helicopter flexbeam. Utilizing fiber Bragg grating (FBG) arrays, the study compares different deployment techniques under a 100,000-cycle fatigue test: embedded, surface-integrated, and surface-applied. A validated three-dimensional (3D) finite element model (FEM) was developed to reconstruct cross-sectional loads and correlate experimental strain data. Validation against conventional electrical strain gauges (SG) confirms that embedded FBGs significantly outperform SGs in durability, maintaining functionality beyond the operational limit of traditional sensors. Furthermore, the methodology successfully tracks global stiffness evolution and degradation throughout the fatigue life. Micro-computed tomography (µCT) scans verify that the integrated fibers do not compromise structural integrity. These findings demonstrate the potential of embedded OFS for continuous, in-service load monitoring and condition-based maintenance (CBM) of flight-critical rotorcraft components.
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.
The effects of hover operations near a partial boundary structure were assessed for a free-flying quadrotor platform under both wind-off and wind-on conditions. The partial boundary structure was selected to replicate a building facade or urban vertiport environment, providing a realistic operational context for these free-flight tests. Test points were chosen to investigate operations near the partial boundary wall and edge, and across a range of partial ground effect conditions to capture the progressive onset of ground effect characteristics. Regions of degraded vehicle performance, quantified primarily by rotor thrust coefficient (CT ) and power requirements, emerged near the partial boundary edge. These performance trends were attributed to localized changes in rotor inflow profile, characterized by near-field rotor pressure measurements. Partial ground effect was found to not resemble full ground effect until much of the vehicle had traversed over the partial boundary, with the vehicle airframe and fuselage serving as the primary factor driving the onset and development of ground effect characteristics. Under wind-on conditions, the interaction between the wind freestream and the partial boundary structure significantly shaped vehicle performance and handling qualities. The most adverse handling qualities coincided with regions of peak flow vorticity, highly unsteady flow, and large velocity gradients at the shear layer above the partial boundary.
Cold spray deposition is a kinetic-based deposition method that uses an inert gas flow to accelerate particles, where kinetic energy causes plastic deformation upon impact with a substrate, as discussed in Reference 1. Cold spray has been investigated as a method to deposit metal coatings on polymer-based composites, such as aerospace carbon-fiber-reinforced plastics (CFRP's), as discussed in Reference 2. These methods also exhibit low deposition efficiency (15-45%) as shown in Reference 3. In this work, to achieve high deposition efficiency and create an erosion-resistant coating, we use metal-polymer composite powders for cold spray, to make polymer-on-polymer bonding the dominant and effective bonding mechanism; this method lowers impact velocities relative to pure metal deposition to avoid substrate damage. The polymer can also lower the effect of material mismatch, while the nickel can help enhance the erosion performance of the final coating above that of pure polymer. This paper discusses the development of the cold spray deposition process, sample post-processing, erosion testing, and characterization of the samples. The results validate that this method successfully creates a method of depositing a coating with erosion resistance approaching that of bulk titanium without substrate damage on PEEK substrates. Further work is needed to address the issues with depositing on CF-PEEK substrates.
Rotor Track and Balance (RTB) and the related shaft balance function are common helicopter maintenance practices. A smooth helicopter reduces high-cycle fatigue in components, improving operational readiness. For the flight crew, a smoother ride enhances safety and reduces fatigue. RTB algorithms attempt to provide blade adjustments to hub weights, pitch control rod links, or blade tabs to reduce vibration at aircraft sensors. Prior studies have described the mathematical formulations underlying RTB optimization with little attention given to implementation. This paper presents a Fourier-domain approach to RTB using the linear least-squares method for both vibration minimization and blade-track split reduction. The methodology is demonstrated across a range of operationally relevant scenarios, including rotors with varying blade counts, non-orthogonal blade spacing such as in the Bell 429 tail rotor, and unequal adjustment on the Bell 407 tail rotor system. Considerations such as adjustment directionality, input shaft balance coupling, and maintainer-centric implementation constraints are addressed. Rule-based expert systems leveraging engineer-curated knowledge are introduced for this narrow-domain application, with a brief discussion of their relationship to large language models (LLMs).
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.
A methodology was developed and validated to predict total noise from a 1/5th scale eVTOL rotor in hover by coupling 2D-RANS airfoil simulations with the comprehensive code, CHARM, for rotor loading, the acoustic code, PSU-WOPWOP, for tonal noise, and the broadband noise code, UCD-QuietFly, for broadband noise prediction. Improved sectional 2D aerodynamic inputs obtained from 2D-RANS simulations were used throughout the framework, replacing XFOIL-derived inputs to enhance the prediction accuracy of low Reynolds-number effects such as transition and laminar separation bubbles. Predictions were validated against measurements at Virginia Tech. Results indicate that turbulence model selection influences local boundary-layer development and sectional aerodynamic loading, producing modest differences in tonal noise at 4000 RPM (first BPF) and spectral differences of approximately 4 dB in broadband noise, while integrated OASPL shows slight sensitivity to turbulence model choice. At 2000 RPM, broadband noise dominates the total SPL, with transitional models predicting thinner boundary layers and reduced high-frequency trailing-edge noise. In addition, high-fidelity 3D-DDES simulations performed in OpenFOAM capture the mid-frequency noise content that is underpredicted by CHARM. Overall, the combined mid-fidelity broadband noise results and high-fidelity mid-frequency range noise results provide improved agreement with experimental spectra and demonstrate an effective approach for eVTOL rotor noise prediction.
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