Browse Topic: Helicopters
The Sikorsky S-92® helicopter fleet, representing more than 300 aircraft and 2.6 million flight hours, is relied upon to support a large range of important missions across the globe. In previous efforts, a high-fidelity CFD-CSD based full-aircraft simulation methodology, co-simulated with production FCS, was developed and applied to model both coaxial aircraft and single main/tail rotor configurations (Refs. 1-5). The CFD solver is based on the CREATE™-AV HELIOS toolset (Ref. 6) and the CSD solver is based on Rotorcraft Comprehensive Analysis System (RCAS) (Ref. 7). The current paper further correlated the CoSim methodology (Ref. 1) with the S-92® helicopter flight-test database at both hover, cruise and edge-of-envelope maneuver flight conditions. The consistent correlations for flight dynamics, static and fatigue component loads at conditions across the flight envelope demonstrate the reliable predictive capability of the high-fidelity CoSim methodology to be-used as a virtual digital flight test and to support advanced design at early stage.
In this study, a multifidelity aeroelastic framework is presented for predicting trim conditions in rotary-wing aircraft, with the main focus placed on the DUST implementation and its application to helicopters and quadrotors. The methodology combines aerodynamic and structural solvers of different fidelity, specifically DUST and the multibody dynamics solver MBDyn, through the preCICE coupling interface to enable direct comparison with rigid and coupled aeroelastic solutions. The trim problem is formulated from the six degree of freedom rigid body equilibrium equations in a helical turn reference frame, naturally covering both steady and maneuvering flight. Although the same formulation can be extended to fixed-wing configurations, the present paper is focused on rotorcraft applications. The framework is first applied to the SA330 Puma helicopter, chosen for the availability of validated flight test data. The methodology is then extended to a multirotor derived from a NASA quadrotor, demonstrating that the same trim strategy can be transferred to distributed-lift rotorcraft. Results highlight the potential of the proposed approach to provide physically consistent and computationally affordable predictions of helicopter and multirotor equilibrium states.
This paper presents results of flight tests conducted on a coaxial ultralight helicopter. An automated flight test evaluation method is presented and exemplified through its application to steady horizontal flight. The results shown include pilot controls, helicopter attitude angles, power, thrust and torque distribution between the rotors, rotor harmonic thrust components, and teeter angles, along with their rotor harmonic components across varying flight speeds. This study focuses on the dependencies of these parameters on center of gravity position and sideslip angle.
The current work presents a methodology to estimate the mission and performance capabilities of a generic rotorcraft configuration, to satisfy the need of evaluating the integration of a full electric powertrain in the aircraft design. To include all the design steps, two different approaches are proposed. For the preliminary phase, the "Analytic Method" is considered, which exploits a purely resistive model. Conversely, a method based on look-up tables called "Table Method" is intended to be used in more advanced phase, when the battery pack is defined. Both approaches are tested by evaluating a reference mission and a hover chart. Finally, a verification of the presented methodology is carried out by comparing the mission results with a commercial software, specialized in the evaluation of the cell discharge when a given power spectrum is provided.
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.
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 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.
Recent flight tests and simulations have suggested that the outwash from eVTOL air-taxis could be larger than conventional helicopters of equal weight and thus pose greater safety issues for their operation than previously anticipated. This has prompted interest in the analytical and experimental study of the aerodynamics related to multi-rotor aircraft outwash. This paper will describe work investigating some of the related issues, specifically (1) how wake models and wake model parameters impact outwash predictions in comprehensive rotorcraft analyses and (2) considerations when scaling results from model scale to full scale. This work will also compare outwash predictions for conventional and multi-rotor VTOL aircraft obtained with a Lagrangian free-vortex wake model and with an Eulerian velocity-vorticity grid based wake model.
With the flights of the Ingenuity Mars helicopter completed and the development work on the Titan Dragonfly rotorcraft/lander proceeding, it is now time to consider aerial flight on Venus. Challenges of developing aerial explorers for Venus are discussed along with past and present conceptual design vehicles. A summary of the scientific impact and necessary instrumentation to understand Venus’s climate and geographical makeup is provided. This paper presents possible aerial-vehicle-assisted approaches to exploring Venus, with an emphasis on rotary-wing vehicles/systems. Aerial conceptual design vehicles are presented in three categories that include flying: above the clouds (altitudes greater than 60 km), below the clouds (altitudes less than 50 km), and near the surface.
NASA's successful demonstration of powered flight on Mars through the Ingenuity Helicopter, as part of the Mars 2020 Perseverance rover mission, has led to the development of next generation Martian rotorcraft. The future of Martian rotorcraft has evolved to include high payload-carrying vehicles to possibly contribute to planetary science missions, which will require improved flight dynamics and rotor aerodynamic performance to fly at nominally high forward flight speeds and at higher flight altitudes. To ensure the feasibility and viability of successful mission performance, it is also critical to mature the structural design for advanced Martian rotorcraft to bridge the gap between the best practices of the spacecraft and aircraft communities. This paper focuses on the structural analysis of a Mars Science Helicopter (MSH) blade using finite element methods. Multiple loading conditions including launch and operational flight were applied to investigate the blade’s structural integrity. The blade’s modal natural frequencies were also analyzed to investigate the blade's dynamic behavior.
This study introduces a structured methodology for identifying Control-Equivalent Turbulence Input (CETI) models using rotorcraft flight dynamics simulations. A new Moving Spatial Turbulence Field (MSTF) model was developed to generate input datasets, enabling CETI model identification for four distinct aircraft configurations: a generic utility helicopter resembling the H-60, and three small-scale multi-rotor UAS types—a quadcopter, hexacopter, and octocopter. The CETI models were validated in hover using frequency-domain analysis, with flight-derived CETI models serving as the benchmark. To further assess model performance in forward flight, CETI models for the H-60 were identified at airspeeds ranging from 0 to 140 knots in 40- knot increments. Results indicated that the MSTF-based CETI models for the H-60 effectively captured key spectral features of the flight-test data, though some deviations were observed, potentially due to variability in atmospheric conditions. In contrast, the CETI models for the multi-rotor UAS configurations showed closer alignment with flight data, likely reflecting more accurately captured atmospheric inputs during simulation. These findings highlight the promise of simulation-based CETI modeling for both rotorcraft and UAS, offering valuable insights into turbulence effects across diverse platforms and flight regimes.
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.
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.
During helicopter air-to-air refueling the rotor of the helicopter might enter the slipstream of the tanker aircraft's propeller. Based on blade element momentum theory, the impact of the accelerated air within the propeller slipstream on rotor blade aerodynamics (thrust, rolling and pitching moments) can be solved analytically. Also, DLR's comprehensive rotorcraft code has been used with the Pitt-Peters induced inflow plus rotor-rotor interference model. Additionally, DLR's free-wake code was used for both the propeller and the helicopter main rotor, including mutual wake-wake-interactions. The helicopter rotor's collective and cyclic controls needed for disturbance rejection are computed with all these models for a typical air-to-air refueling scenario without and with blade flapping motion. A propeller wake affecting the retreating side of the rotor requires much larger control inputs to retrim than an impingement on the advancing side. The results of all modelling approaches are compared to each other and generally a reasonably good agreement is found, despite their very different levels of complexity and computational effort.
In this paper the time accurate coupling between the high fidelity CFD code FLOWer and the multi-body dynamics code SIMPACK is presented. To facilitate this coupling a socket-based data exchange was developed and used to exchange aerodynamic forces and kinematic data. Two flight states were investigated: a hover and a forward flight. To obtain a reasonable initial flight state a previously obtained, trimmed solution was taken as the base. This study shows the feasibility of the strong coupling approach with the direct influence of the helicopter motion on the flow field and vice-versa. As expected, the factor limiting the overall performance is the runtime of the CFD simulation. The effort of running the flight mechanics simulation and the data exchange necessary for the strong coupling is negligible compared to this runtime.
The paper presents a general framework for building an aeromechanic model in FLIGHTLAB, suitable for high fidelity, pilot-in-the-loop simulator. The focus is on aerodynamic modeling of AW609 tiltrotor in Airplane Mode flight regime. The framework can be extended to helicopter and conversion modes with additional considerations for rotors-airframe aerodynamic interference. It can also be adapted to different tiltrotor geometries, with some adjustments depending on their peculiarities. The model uses Blade Element Theory loads evaluation of lifting surfaces, corrected with tabulated distributed loads to tune FLIGHTLAB predictions against high-fidelity aerodynamic references. Bluff bodies are modeled using force and moment tabulated data. Verification was conducted against reference data in wind tunnel mode and against flight data in trim analysis. The proposed method allowed to match lift distribution on slender bodies, as well as lift and drag integral loads, with aerodynamic references. Trim analysis has shown satisfying accordance with experimental data for all the comparison parameters. However, rotor aerodynamic modeling is still being investigated to improve correlations in torque prediction. Additionally, installation effects, such as those related to the interaction between the wing and fuselage with the rotors, are under examination to address biases in flap motion evaluation.
In April of 2024, Sikorsky flight tested an open loop Higher Harmonic Control system on an S-97® helicopter. The S-97® helicopter is a prototype aircraft, based on Sikorsky's X2 Technology™, that first flew in May 2015. It has contra-rotating, stiff in-plane main rotors with fly-by-wire controls, and a pusher propeller. This paper describes the HHC design, how it was implemented on the aircraft, how it was tested, and what the test results were.
Several efforts have been made to develop Flight Test Maneuvers for Handling Qualities evaluations, aimed at quantifying the effects of vehicle characteristics and assistance systems on a Helicopter Air-to-Air Refueling mission profile. However, these Flight Test Maneuvers have not achieved widespread adoption, likely due to the substantial logistical challenges associated with tanker deployment. Depending on a tanker aircraft not only incurs significant costs but also requires extensive organizational effort and prior testing, before Handling Qualities can be evaluated for the aerial refueling capabilities of a new rotorcraft design. Additionally, these available Flight Test Maneuver setups are not standardized or widely applied to the same degree as Mission Task Elements of the Aeronautical Design Standard, which limits repeatability and comparability. A new approach is proposed to address these limitations by introducing a repeatable, standardized method to reveal Handling Qualities deficiencies considering a worst-case situation of Helicopter Air-to-Air Refueling. This approach involves analyzing drogue motion to create a synthetic, deterministic target forcing function, based on the summation of several sine waves. Resulting laws of motion are applied to a target tracking task replicating a drogue chasing scenario by projecting all required references into the pilots' field of view. Piloted simulator studies conducted at the Air Vehicle Simulator (AVES) of the German Aerospace Center (DLR) demonstrate a high degree of similarity in pilot control behavior between the proposed Flight Test Maneuver and actual simulated Helicopter Air-to-Air Refueling.
The Sikorsky BLACK HAWK® is the primary medium lift helicopter for the U.S. Army performing a wide range of missions that encompass Air Assault, MEDEVAC, CSAR, Command and Control, and VIP transport. The Multimission UH-60M is one of the latest in the BLACK HAWK helicopter product family, more capable, more survivable, more maintainable, more powerful, and more effective than its predecessors. In previous efforts, a high-fidelity CFDCSD based full-aircraft trim and maneuvering simulation methodology was developed and applied to model both coaxial aircraft and single main/tail rotor configurations (Refs. 1-4). The CFD solver is based on the CREATE™-AV HELIOS toolset (Ref. 5) and the CSD solver is based on Rotorcraft Comprehensive Analysis System (RCAS) (Ref. 6). The current paper further enhances the previously developed 6-DOF CFD-CSD full-aircraft trim methodology to robustly handle the trim solution for the single main/tail rotor configurations. The enhanced methodology was applied to rotor control loads and aircraft performance calculations at both steady level flight and aggressive Angle-of-Bank (AOB) turn maneuver conditions, including comparisons with available flight test data. The simulation results demonstrate a consistent correlation with flight test, including at stall conditions. The high-fidelity full-aircraft CFDCSD simulation approach provides an efficient, accurate and robust methodology for aircraft performance, control loads and stall boundary predictions to support advanced design at early stage.
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.
The biography of Henrich Focke is well known and documented. During a small period from October 1954 to February 1956 he held lectures at the Technical University of Stuttgart during the winter semester. In the summer period he returned to Brazil for continuation of his contract work on the "Convertiplane" (a quad-tiltrotor aircraft) and the "Bei-jaflor" (a small single rotor helicopter). The topic of Focke's lecture in the winter semester 1954-55 was "Design of Fixed-Wing Aircraft", but the lecture manuscript of it is unavailable. In the following period 1955-56 Focke lectured about "Helicopter Design" and the manuscript was recently found in the central archive of DLR. It covers 123 pages of text with sketches and graphs and provides deep insights into the helicopter design philosophy of Henrich Focke.
Neonatal patients in need of specialized care may require transport by rotary-wing air ambulances. These patients are subjected to environmental stressors during transport, including elevated levels of mechanical vibration. Aircraft vibration is transmitted through the transport system and incubator to the patient. The unique vibration profile is dependent on vehicle model and phase of flight. To improve safety for these patients, we aim to evaluate the vibration exposure across this complex system. The purpose of this paper is to present and evaluate the methods used for aircraft data collection and replication of aircraft vibration profiles in a laboratory setting. Our current focus is on neonatal transportation in Ontario, Canada, where Leonardo AW139 helicopters are used for patient transport. AW139 field data were collected and processed to generate excitation profiles for discrete phases of flight. The vehicle data were used to drive a series of laboratory shaker-table experiments, in three axes, to evaluate the response of different configurations of the transport system. We present the methods used to simulate transport conditions, from vehicle data collection to laboratory shaker experimentation, and evaluate the behavior of the test apparatus. The simulated motion has been verified against the aircraft data to identify sources of error in the experimental setup. Some limitations in the shaker and control system present inherent differences in the input and response; however, it was found that the greatest spectral error occurred outside the frequency range of interest (>80 Hz), and that the shaker controller successfully replicates the energy levels recorded in the aircraft. The shaker experiment results, such as the response of the transport system and incubator, will be analysed in future work to identify equipment configurations and/or modifications which can reduce neonatal patient vibration exposure during rotary-wing transportation.
A cooperative flight test campaign between the US Army and NASA was performed. This test sought to characterize the acoustic emissions of a fully instrumented MD530F helicopter using a snapshot array and a phased array of microphones. The snapshot array of microphones aimed to provide even coverage across the surface of a hemisphere, providing an acoustic emission hemisphere in a single 'snapshot' of time. The phased array of microphones was designed to provide enough resolution to determine noise sources from each individual blade as well as perform source separation from main rotor and tail rotor emissions. Test conditions for the characterization effort were chosen using a traditional one-factor-at-a-time approach as well as three design of experiment approaches. Characterization conditions included constant speed level flight, descent, and ascent conditions. Transient maneuver conditions were also captured over the snapshot array. The vehicle instrumentation included measurements of pilot controls, optical sensors to measure blade azimuth locations, pitch link loads, along with strain gauges to measure structural loads, blades and fuselage. This report will provide an overview of the test, document the data acquired, and provide some initial results.
This study numerically investigates the relationship between airspeed, drop height, and ground water coverage during helicopter-based aerial firefighting. With the effect of global warming and human activities the threat of forest fires has increased and finding optimal water dumping strategies for effective suppression is a crucial part of the firefighting operations. How varying airspeed and water drop height influence water dispersion and ground coverage has been analyzed utilizing numerical simulations with the VOF model in STAR-CCM+. Findings show that to maximize firefighting efficiency, balancing two contradicting phenomena is essential. These are, minimizing ineffective mist formation due to high drop height/high airspeed and fueling of the fire from rotor downwash due to low height/low airspeed passing by over the fire zone.
The next generation of Mars rotorcraft may involve an increase in scale and number of rotors. A key focus area that has been identified is to increase the fidelity of rotor wake modeling, including its impact on flight dynamics. To that end, this paper pursues the use of a Viscous Vortex Particle Method (VVPM) for mid-fidelity rotor wake predictions in Mars atmospheric conditions. Simulated aerodynamic hover performance, as well as control efforts in trimmed forward flight, of the Ingenuity Mars Helicopter with a VVPM wake is shown to correlate well with available experimental data. Qualitative and quantitative coaxial wake effects for Ingenuity-type rotors in hover and forward flight as predicted with VVPM are studied. Utilizing VVPM to evaluate rotor-rotor interference effects in a large-scale Mars hexacopter across a wide range of flight conditions showcases the capability to comprehensively model the induced wake of complex multi-rotor configurations within feasible computational cost. An approach to emulate interference effects derived from the VVPM data that is compatible with finite-state dynamic inflow models is demonstrated, exemplifying how VVPM can be adopted to inform flight dynamics analysis and control system design of next-generation Mars rotorcraft.
Dynamic rollovers represent a major hazard for helicopters during near-ground operations, often resulting in significant aircraft damage and passenger injuries. To improve safety in operations, recent studies have focused on developing a Helicopter Flight Data Monitoring framework to provide data-driven insights on operational safety. This work contributes to that effort by proposing an approach to identify precursors to dynamic rollovers. According to NTSB reports, approximately 60% of such incidents occur during in-flight phases like hover, hover-taxi, or landing. To capture the complex non-linear dynamics of helicopters, physics-based simulations were conducted to estimate a first hitting time metric, defined as the time until blade-ground contact, across a wide range of initial conditions for an inflight initial state of the helicopter. Eight parameters were identified as driving the first hitting time, and a probabilistic model was created to predict the distribution of that metric for different values of those parameters. Based on the predicted distributions, a risk-based metric was derived to robustly assess the risk of dynamic rollover and identify safer operational boundaries.
A typical helicopter drive system consists of a multi-stage gearbox with highly loaded dynamic components such as gears, shafts, and bearings, crucial for safe flight and landing. Planetary reduction stages are commonly used in the final reduction stage of rotorcraft main gearboxes due to their ability to handle high torques at high gear ratios within a compact envelope. The planet gear, a critical component in this arrangement, is subjected to significant loads on both flanks of its teeth and must meet stringent weight and assembly requirements, leading to a thin rim design with integrated bearing races. This design makes the planet gear susceptible to relevant reduction of its fatigue life. This paper explores analysis methods to evaluate the damage resistance of the planetary stage assembly, focusing on the planet gear. The study aims to assess the "growth" or "no growth" condition of the planet gear against defined flaw defects. An iterative calculation loop determines the critical length and position of a crack that may lead to full crack propagation and, in worst cases, to system jamming. Initial crack propagation simulations use NASGRO software, with stress fields derived from a non-linear FEM of the planet gear availing of detailed Transmission3D model for teeth meshing forces evaluation. Further additional analysis can involve a dedicated FE model of the crack, iteratively updating its geometry. The impact of crack propagation on the remaining components of the planetary stage assembly is also addressed, considering the unbalanced load conditions caused by stiffness loss in the planet gear. The dissertation object of this paper aims to outline a comprehensive, effective and efficient procedure to determine the maximum allowable defect size for "no growth" condition and the operational hours until failure, providing a robust approach to support the strength substantiation of the involved components.
Leveraging lessons learned from NASA's Ingenuity Mars helicopter and concepts such as the Mars Sample Recovery Helicopter, and Mars Science Helicopter has enabled partners at NASA's Jet Propulsion Laboratory (JPL), NASA Ames, and AeroVironment, Inc. to mature a hexacopter vehicle concept (Chopper) with the ability to support a wide range of mission scenarios. This work focuses on the critical aeronautics-related challenges encountered transitioning from an Ingenuity-size vehicle to a much larger vehicle (˜15 times the mass) and discusses engineering efforts to address these challenges. Critical upgrades include optimized airfoils, higher solidity blades, and higher fidelity computational models. Because multiple rotors are required to lift the heavier vehicle, increased understanding of the impact of rotor-to-rotor interactions is also necessary. Rotors have been designed that are tailored to more demanding missions and will be validated in a joint test campaign between the partners. While the Chopper concept will be utilized to illustrate these maturation efforts, the lessons learned are applicable to other heavier next generation Mars rotorcraft platforms also.
Acoustic flight testing of rotorcraft often involves generating noise source hemispheres to gain an understanding about the aircraft's acoustic emissions. However, aerodynamically complex Urban Air Mobility and Future Vertical Lift vehicles may not maintain a steady aerodynamic state during flight, making source hemispheres measured using traditional linear arrays unreliable or difficult to interpret. To address this challenge, all emission angles need to be measured simultaneously. This has lead to the concept of the two dimensional 'snapshot' array layout. A mathematically defined microphone distribution was utilized to achieve uniform coverage on the source hemisphere. Within the chosen distribution, two lower microphone count distributions are embedded, allowing for a comparison of the effects of number of microphones. The array was deployed as part of a joint Army/NASA acoustic research flight test in July of 2024. Data were collected using an MD530F helicopter as the test vehicle, executing both steady and unsteady flight. Analysis of the array resolution is used to determine adequate channel count and spatially varying sensitivity to array density.
This paper expands on a previous exploratory investigation into the safety implications of helicopter operations at hospital landing sites. The paper analyses the interaction between rotor downwash, the turbulent wake shed from nearby buildings and the effect of varying windspeed and aircraft position. A RANS CFD method has been used to compute the mean airflow in the vicinity of a hospital helipad with a helicopter, representative of a Bell 412, hovering at three different positions around the site. The main rotor of the aircraft was modelled using a Virtual Blade Model, enabling a coupled solution between the airflow around nearby structures and the helicopter. The study examines the resulting airflow patterns and velocity magnitudes around the site for two incoming windspeeds and three varying aircraft positions. Results presented are focussed on areas where the rotor downwash is present and likely to impact pedestrians. The findings show that windspeed can affect how the downwash from the rotor is distributed through the local environment and highlights that, in certain areas, calculated air velocities are found to be at levels considered hazardous to pedestrians.
Full-scale static test (FSST) is a key test program for the certification of new helicopter airframe. The strength and deformation requirements in airframe certification are substantiated by full-scale tests of the airframe structures. It provides experimental evidence that the structure is able to support limit loads without detrimental permanent deformation and carry ultimate loads for at least three seconds. In design stage, the total number of flight and ground limit load conditions is around 500. In FSST, the number of test load cases should be remarkably reduced. However, the selected load scenarios should cover all of the critical design load scenarios. In this paper, test load generation procedures in FSST of a light utility helicopter is explained. The comparison of design load envelope and static test load envelopes are provided.
The paper describes a method for optimal design of a helicopter tail shaft that considers rotordynamic effects from long shaft assembly. The tail shaft transmits power from the main gearbox (MGB) to the tail rotor of the helicopter and operates at high speeds that may exceed 6000 rpm. While higher speeds allow for weight reduction, they also pose risks associated with supercritical operation, necessitating careful design optimization. The objective of the optimization is to maximize the first three transverse natural frequencies with the constraint of the safety parameter (avoidance of the resonance/critical zone) while minimizing the weight of the system. A Non-Dominated Sorting Genetic Algorithm (NSGA-II) is used to obtain the solution to this multiobjective optimization problem, which involves shaft design variables such as length, outer diameter, and wall thickness. In addition, the optimization framework also incorporates system related design variables, including the stiffness of tail shaft bearing supports, the location of bearings, and coupling characteristics, to comprehensively evaluate and enhance the system performance. A rotordynamics model utilizing Timoshenko beam finite elements is deployed to predict natural frequencies.
A follow-on study to the 2024 paper by Kottapalli, Silva, and Boyd is presented with improved acoustics tools to examine whether the Vertical Aviation International (VAI) Fly Neighborly operational recommendations that are designed for single main rotor/tail rotor configurations will hold for non-conventional UAM rotorcraft with multiple rotors. The 6-occupant quadrotor concept vehicle designed under the NASA Revolutionary Vertical Lift Technology (RVLT) Project is studied. The tip speed is 550 ft/sec, with three blades per rotor. Predictions are made for three steady maneuvers: level turns, descending turns, and climbing turns. The RVLT Toolchain is exercised using CAMRAD II, pyaaron/AARON/ANOPP2 and AMAT (ANOPP2 Mission Analysis Tool). Quadrotor noise trends are analyzed using Sound Exposure Level (SEL) ground maps because it is anticipated that the upcoming updated Fly Neighborly recommendations will involve SEL maps. Importantly, unlike conventional helicopters with a single main rotor, quadrotor noise can increase in a direction opposite to the turn direction; this finding may need to be taken into consideration in formulating the upcoming updated Fly Neighborly recommendations. The recommendation Level turns are quieter than descending turns is followed by this quadrotor under the flight conditions examined. Turning away from the advancing blade … is quieter than turning into the advancing blade is not relevant for symmetric aircraft such as the quadrotor. Straight flight is quieter than turning flight is not predicted to be best for this quadrotor under the flight conditions examined – turning flight is quieter than straight flight for the quadrotor in this study.
Helicopters' Vertical Take-Off and Landing (VTOL) capabilities are essential for maritime operations, especially for small-deck naval vessels. Unmanned Aerial Vehicles (UAVs) offer a cheaper, expendable, and efficient alternative for certain tasks, such as reducing pilot risk and lowering fuel consumption. While the procedures to approach and land on (moving) ships are standardized and bound to established operational limits in the case of crewed helicopters, UAVs lack such guidelines. This study investigates optimal rotary-wing UAV approach trajectories to a moving ship, for varying wind conditions and relative initial positions, and for different objectives. The goal is to provide preliminary guidelines for maritime UAV recovery operations, and a preliminary estimation of performance-based operational limits. The optimal trajectories are obtained using a global path-performance optimization framework based on Optimal Control Theory. The trajectories are compared to each other and to reference cases using the Longest Common SubSequence (LCSS) similarity measure, revealing how the unmanned helicopter adjusts its path to exploit the wind direction and profile for more efficient ground speeds. The violation of performance and/or geometric constraints is used to preliminarily indicate the presence of operational boundaries. The control effort and energy consumption are used to identify optimal starting positions for the helicopter approach phase for a given wind profile and intensity.
Traditional safe-life methodologies for rotorcraft structural components often result in overly conservative life estimates, increasing maintenance costs and reducing aircraft availability. This study explores the integration of digital twin concepts with probabilistic modeling and machine learning to enhance structural life assessment, demonstrated through a practical case involving the Royal Canadian Air Force CH-146 Griffon helicopter. A probabilistic fatigue model determines a fatigue life distribution by incorporating material variability and uncertain operational loads inferred directly from flight data. Unlike conventional approaches, this method dynamically estimates load spectra, including uncertainty instead of relying on conservative assumptions. Monte Carlo simulations are used to quantify structural risk and assess the impact of load and material uncertainties. Sensitivity analyses highlight these uncertainties’ contributions to failure probability. The proposed approach provides probabilistic life predictions, supporting risk-based maintenance strategies to potentially optimize operational efficiency. The long-term goal is to develop an adaptive digital twin model that continuously updates with new operational flight data, enhancing predictive accuracy for helicopter fleet management.
This paper explores the dynamics of rotating Tuned Vibration Absorbers (TVAs), focusing on the phenomena arising from gyroscopic effects. Some products from Leonardo Helicopters (LH) can have a TVA fitted in the rotor mast, counteracting the in-plane vibratory loads of the rotor directly at their source, implying that the absorber rotates with the rotor itself. Although gyroscopic effects are negligible for most of the LH TVAs, specific design choices may have notable impacts on tuning and performance. An analytical model is implemented, demonstrating that the gyroscopic terms influence the dynamics causing a frequency displacement of the anti-resonance evaluated without considering this effect. Additionally, a regression analysis investigates the interplay between this phenomenon and the physics of the system, revealing how to optimize the design to mitigate gyroscopic effects. Finally, the performance of the TVA is analyzed as a coupled problem, showing that the anti-resonance effectiveness is reduced by the orthogonal contribution introduced by the gyroscopic matrix.
Heavy wind and high sea states pose challenges to operating unmanned rotorcraft on-board a naval ship, in particular the recovery phase. A novel autonomous landing strategy for unmanned rotorcraft is proposed and investigated. The new landing strategy makes use of a prediction of the future deck motion based on a sensor on the ship deck. The study is based on a nonlinear simulation environment which includes the dynamics of a 100 kg unmanned helicopter and the dynamics of an ocean-going patrol vessel of the Royal Netherlands Navy. The performance of the autonomous landing strategy is evaluated for a wide variety of environmental conditions (sea state) and operational conditions (ship speed and heading). The results clearly indicate that the environmental conditions have a strong influence on the landing performance in terms of touchdown velocity and landing accuracy. Furthermore, the autonomous landing strategy is effective in reducing the mean and peak value of the touchdown velocity compared to a standard automatic landing strategy. A reduction of as much as 60% in landing impact is observed for the worst case environmental and operational condition considered. The results confirm and underline the potential of the novel landing strategy.
The oil cooling fan of a Main Gearbox (MGB) is a mechanically-driven component whose purpose is to force an air flow through an air cooled oil cooler; its performance is crucial in ensuring that the MGB oil temperature does not exceed a predefined threshold, set to alert the crew in case of an abnormal situation. The design and the certification of a cooling fan is a process involving several steps and multiple disciplines; mechanical design, aerodynamic analysis, dedicated tests carried out both on rigs and at aircraft level need to be exploited as complementary tools to assess the correct aero-mechanical behavior of the system. The aerodynamic assessment is associated to performance, measured in terms of MGB oil temperature: considering a comparison between two cooling fans, one outperforms the other if the resultant MGB oil temperature is lower, keeping the same boundary conditions (engine torque, wind speed, ambient temperature, etc.). The correct mechanical behavior is instead associated to multiple requirements that need to be satisfied: target fatigue life, high and low temperature limits, blade loss containment, maximum speed, etc. The process connecting the very first blank page to the certification of the part is usually linear and well defined, and the experience gathered by Leonardo Helicopters over the years has allowed to follow this path easily in most cases. This paper discusses an intriguing situation: the goal of the activity herein discussed was to replace an obsolete oil cooling fan with a new one. The latter was expected to outperform the first both from the aerodynamic point of view (i.e. lower MGB oil temperature) and from the mechanical point of view (i.e. longer life). The preliminary CFD analyses and the rig tests proved that the new fan was able to outperform the legacy one in regards of the aerodynamic requirement; nevertheless, during the last steps of verification carried out at aircraft level, the installation of the new fan resulted in a higher MGB oil temperature. This triggered dedicated troubleshooting: a sequence of investigative tests performed by means of dedicated rigs, targeted at properly identifying the aerodynamic operative point of the cooling fan and at understanding the reason behind the unexpected outcomes obtained. The relevance of this contribution does not lie in the description of the certification process per se, rather in the nonlinear dynamics characterizing the test sequence once the flight activities pointed out an unpredicted behavior.
This paper provides an overview on the contributing phenomena to unanticipated yaw described in the FAA Helicopter Flying Handbook. Trimmed aerodynamic - flight-mechanic - coupled simulations with a validated model of the BK117 C-2 capture the relevant interactions for weathervaning, main rotor-to-tail rotor interactions and vortex ring state effects at the tail rotor. An investigation of the impact of the main rotor downwash on the vortex ring state at the tail rotor in sideward flight and yaw turn is provided, concluding that the presence of the main rotor effectively inhibits the occurrence of a fully developed deep vortex ring state at the tail rotor. The consequent limited impact of the incipient tail rotor vortex ring state on the helicopter trim is estimated. Further, maneuver simulations of the BK117 C-2 are provided, describing the typical entry in unanticipated yaw turn and the exit to stop the yaw motion by means of pedal inputs of different magnitude and input speeds.
eVTOL aircraft are a stable part of nowadays rotorcraft industry, gathering attention and investments throughout all geographies. The challenge of designing such a vehicle is the necessity to combine transformative flight and distributed lifting systems. This paper presents a methodology developed within Leonardo Helicopters Division (LHD) to perform the preliminary design of eVTOLs, following an approach that starts from hovering flight, investigating the design permutations able to satisfy certain criteria of performance, maneuverability, and safety.
This study investigates the evolution of axial and radial velocities in the downwash-outwash region of a counter-rotating coaxial rotor hovering in-ground effect (IGE). The presence of the ground deflects the axial flow of the rotor wake radially outward, with mean radial velocities reaching approximately 2Vh along the ground. Based on the observed velocity profiles, the wake was classified into three distinct regions: the downwash region characterized by maximum wake contraction, the transition region where flow turns from axial to radial, and the outwash region exhibiting wall jet behavior. Results show that increasing inter-rotor spacing d/R and rotor height above ground (z/R)l extends the downwash and transition regions, delaying the onset of radial outwash. Aerodynamic loads on personnel were estimated, showing maximum mean forces and moments of 120N and 120Nm, remaining within safety thresholds for untrained personnel. However, the loads exceeded these limits for heavy-category helicopters with higher disk loading. Instantaneous velocities revealed unsteady behavior driven by tip vortex convection, with peak velocities reaching 2.8Vh and peak loads approximately 50% higher than mean values, as identified using the 98th percentile approach. Frequency analysis confirmed that fluctuations at the rotor frequency dominated the unsteadiness in the rotor wake. These findings provide critical insights for safe rotorcraft operation near ground surfaces and contribute to informed vertiport design and personnel safety guidelines.
Flight test students must explore a wide range of helicopter dynamic responses to learn how to assess conditions ranging from good conditions operation to those approaching, or even experiencing, loss of control. To introduce this evaluation process, the Flight Test and Research Institute (IPEV) implemented a helicopter flight dynamics model. This model is stitched in the x-body velocity (u) and y-body velocity (v) to achieve more accurate simulation, combined with a Variable Stability Augmentation System to assess different conditions prior to experiencing them in real flight. The use of robust control, where a fixed controller is applied to flight control systems under various operating conditions, presents an alternative to the traditional gain scheduling technique commonly used in aeronautical systems. This paper explores the potential to reduce controller design complexity while evaluating the impact on the helicopter’s full flight envelope through quantitative analysis and handling qualities evaluation by piloted simulation according to ADS-33E-PRF standard.
The Primary Author has been involved in Army Aviation Development and Acquisition since the Utility Tactical Transport Aircraft System (UTTAS), Advanced Attack Helicopter (AAH), Army Helicopter Improvement Program (AHIP), and Light Helicopter Experimental (LHX) Programs in the mid-1970s to the mid-1980s. The first three of these programs successfully made it to production aircraft, while the LHX became the RAH-66 Comanche and was canceled primarily due to technical problems and cost overruns. The initiation of the next phase by the Army Aviation Development (ADD) Directorate for Future Vertical Lift (FVL) did not occur until the beginning of the 2015-2000 timeframe. This was 35 years since the last Army Aviation Development in 1980. To help sustain this FVL development, the Primary Author led, oversaw, and helped conduct a program through the National Rotorcraft Technology Center (NRTC) in the 2015-2016 timeframe. It was called the Development Assurance Value-Based Acquisition (DAVBA) Program1. It included the following team members: Georgia Tech, University of Alabama Huntsville (UAH), Dassault Systèmes, and Clausewitz Technology. The Army ADD plan funded it for FY2015- 2016 through the NRTC. The objectives were to provide the Future Vertical Lift (FVL) Program with a Development Assurance for Airworthiness Qualification and a Value-Based Acquisition Overall Evaluation Criterion (OEC) for FARA and FLRAA concepts.. However, Army Aviation only funded the first phase in 2015, as FVL funds were then transferred to the new Army Futures Command. This paper will illustrate how DAVBA could have saved the Future Attack and Reconnaissance Aircraft Program (FARA) Program as well as providing a more cost effective Future Long Range Assault Aircraft (FLRAA) Program.
Survivability in the future operating environment is becoming more challenging as threat systems evolve and become more sophisticated. The ability to tailor and manage signatures will be one of the key methods to improve survivability, allowing operators to minimise detection and maximise the effectiveness of countermeasures. This paper presents the findings of an investigation into the application of classical Signal Detection Theory (SDT) to the aural detectability of helicopter noise signatures, considering human auditory capabilities. The paper has thus developed a novel methodology, applied it to both the experimental and numerical helicopter acoustics signatures of an LH platform, and used these results to infer the detectability characteristics of the aircraft, as well as how they are affected by the presence of background noise in different environments.
Developed in the frame of the European Clean Sky 2 program, the RACER High Speed Helicopter Demonstrator of Airbus performed its maiden flight on April 25th, 2024. In the continuity of the previous high-speed demonstrator X3 (1st flight in 2010) the RACER is a 7/8t (15000 / 18000 lb) class compound helicopter powered by two SHE Aneto-1X engines, including a wing and two propellers. The tail rotor is removed as the two propellers control the yaw axis by differential thrust. At flight 07, with its initial default settings, it reached a true airspeed of 227 kts in level flight, exceeding its objective of 220 kts.
The empennage of a helicopter is largely responsible for its stability in forward flight. Its performance is mainly determined by its aerodynamics. In this paper, the empennage of a CoAX 2D ultralight research helicopter is analyzed in detail. For this purpose, the helicopter was equipped with flow measurement devices and flight tests were performed, covering different flight conditions. Measurements from a nose boom as well as the pilot’s control inputs and helicopter's position are available for evaluation. For the empennage in particular, seven-hole flow probes were mounted on it and various cameras were used to record the movement of the surface tufts.
The Main Gearbox of a helicopter is a crucial component that delivers the desired performance and ensures the highest possible level of safety of the aircraft; it includes several gears and bearings, which require to be continuously lubricated by a pressurized oil flow. Undesired circumstances may cause the oil to leak from the main circuit, hence reducing its pressure and consequently the oil flow rate targeted towards the rotating components; this modifies their friction coefficient, and subsequently leads to an overheating of the parts with the risk of degenerating in a catastrophic failure. During the design of a helicopter drive system, engineers need to take proper precautions and make sure that the MGB is fully equipped with the proper features to cope with a loss of lubrication event; specifically, the drive system is supposed to be able to run at least 30 minutes after the oil pressure drops to zero. A lot of effort has been put over the years at Leonardo Helicopters to find robust solutions to attain the longest performance of the drive system in no-oil conditions: the most important result is the certification of the AW189 for a 50-minutes “run dry” capability. Nevertheless, the dynamic environment typical of the rotorcraft industry pushes towards continuous innovation, and in the last few years the Transmissions Systems Design department of LH has been asked to investigate suitable ways to further augment the no-oil capabilities of the MGB: the main steps followed and entailed results are presented in this paper. The first part of the manuscript discusses the “state of the art” auxiliary lubrication system, currently flying on the AW189 drive system. The second part tackles the approach adopted to meet the novel requirements, unveiling both the methodology and the final design choice: the latter includes a metering element, able to tune the oil flow rate headed towards the component deemed the most critical in order to satisfy the requirement of longer no-oil performance. Numerical and experimental tools are exploited as complementary tools to properly crystallize the obtained results and corroborate the solution.
The vertical flight industry is on its way to a transformative era, with autonomous technologies set to alter aerial vehicle operations. While it seems certain that fully autonomous helicopters will eventually be deployed for a variety of missions, some high-stakes situations—like medical evacuations (MEDEVAC)—will for the foreseeable future demand human participation in the form of Emergency Medical Care-giving Crew. This study describes the testbed built to run and investigate hypothetical future situations in which a helicopter is autonomously piloted while a human medic with no aviation training, subjected to aviation and medical emergencies, manages patient care onboard. A total of 22 participants, with emergency medical technician certification, nursing or a medical board certification, were invited to run and evaluate the use of AI pilot (AP) in different scenarios of medical evacuation under the following emergencies: medical, empty fuel tank, pressure sensor miscalibration, and engine failure. A comprehensive evaluation of both objective and subjective performance metrics revealed that novice medical professionals could effectively execute medical evacuation operations in conjunction with an AI pilot, even during unforeseen circumstances. The analysis of response times unveiled distinct perspectives on how medics perceive and manage various emergency situations when an AP functions as a collaborative and effective team member.
By its seventh flight after the first take-off, the RACER (Rapid And Cost-Effective Rotorcraft) demonstrator smoothly reached the targeted 220kts speed in stabilized forward flight, validating the high-speed compound architecture developed by Airbus Helicopters in the frame of Clean Sky 2 programme. During the flight envelope exploration, the dynamic behavior of the main rotor was carefully assessed, by monitoring the vibratory loads and validating its aeroelastic stability. Particular care was taken to validate the predicted stability domain of the Dual Rotor phenomenon, a particular case of flap-lag coupling associated with high-speed flight conditions. This paper presents the most significant results shaping the success of RACER flight test campaign. After having introduced the theoretical background and the associated analytical equations, the simulation framework based on the comprehensive analysis tool STORM is presented to discuss the numerical resolution of the stability problem. Then, the rotor dynamics loads and airframe vibratory behavior of RACER are closely examined to demonstrate the absence of any sign of instability, in the various flight conditions offered by its rotor and wing configuration. At last the flight test results are compared to the computed stability domain to assess the margins and estimate the high-speed potential of the rotorcraft.
In January 1984, the Georgia Tech School of Aerospace Engineering (AE) hired Dr. Daniel P. Schrage as the Rotorcraft Design Professor and the Associate Director of the Georgia Tech Center of Excellence in Rotary Wing Aircraft Technology (CERWAT), one of the Army-sponsored Rotorcraft Centers of Excellence (RCOE). Dr. Schrage left St. Louis, MO at the end of 1983 as the Director for Advanced Systems (DAS) and the Associate Director of Army Aviation Science and Technology in the Aviation Research and Development Command (AVRADCOM). His departure was a tough decision for Dr. Schrage as he was designated to become the Technical Director of the new Aviation Systems Command (AVSCOM). However, the new AVSCOM motto was "Readiness Immediate and Development Eventual," which, as it turned out, meant that new Army Aviation Systems would not be developed for the next 35 years. Dr. Schrage was also recruited by Bob Lynn, Bell Helicopters VP Engineering, as the Bell Helicopters Director of Technology, with a promotion to VP Engineering upon Lynn's retirement in the next few years. This was one of the most challenging decisions of his life for Dr. Schrage and his family, as they had to move from their home in O'Fallon, IL - a great town and home for his growing family.. In addition, his wife Nancy's parents and his mother and sisters, all lived within a half hour from his home. His father, Albin, had passed away while we were moving to the St. Louis Area following his MSAE graduation from Georgia Tech in May 1974. When he arrived at Georgia Tech in January 1974, he started teaching the Rotorcraft Design Courses, which were similar in content to what he had developed in AVRADCOM and to what Dr. Dick Carlson taught at Stanford U. He also developed and taught design courses in his two weeks each year in the Department of Mechanics at USMA West Point during his reserve duty from 1978-1983. He continued in the U.S. Army Reserves and taught two weeks at USMA until 1996. He then retired as Colonel in the USAR in 2004. He used a computer-integrated environment and the top-down design decision support process taught at West Point. This provided a Framework for using Systems Engineering Methods for Product Design Driven and Quality Engineering Methods for Process Design Driven, as illustrated in Figure 1. Value-based Acquisition decision-making is a lost art in Digital Engineering and other attempts to produce effective and affordable systems. The Army FARA Contract termination in 2022 was primarily due to lack of Development Assurance Value Based Acquisition (DAVBA) (Ref. 1).
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