Browse Topic: Computational fluid dynamics (CFD)
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.
This paper discusses the design of a 2000-lb manned eVTOL aircraft propelled by a novel cycloidal rotor propulsion system. To systematically evaluate the performance of the proposed configuration, a coupled trim model was developed to quantitatively evaluate the performance of the configuration across a range of forward flight speeds. The trim framework integrates an efficient physics-guided neural-network-based aerodynamic model for cycloidal rotor performance with a vehicle-level dynamic response model. This framework is used to conduct a systematic parametric study to identify key cycloidal rotor and airframe design parameters. The selected configuration is verified using high-fidelity CFD simulations, and a detailed structural design, powertrain design, and CAD model of the aircraft is developed. In addition to CFD validation, the proposed cycloidal rotor underwent structural optimization to confirm the validity of such a concept at this scale. The results demonstrate that the cycloidal rotors provide a viable propulsion alternative for eVTOL aircraft with strong potential to overcome limitations of existing configurations.
The present study explores the active vibration suppression of a lift-offset (L.O.) coaxial rotor system in high-speed forward flight by applying a multicyclic controller with individual blade control (IBC) actuation. A high-fidelity vibration analysis is conducted through a loose-coupling (LC) framework that combines a compressible 3D (three-dimensional) CFD (Computational Fluid Dynamics) solver with a comprehensive aeromechanics (CA) method. Since the upper and lower rotor experience different aerodynamic environments, an asynchronous or different IBC actuation for each rotor is applied to achieve greater vibration reduction performance than the usual synchronous or identical actuation. Open-loop control results indicate that the asynchronous actuation suppresses the vibratory loads more effectively, from which the best actuation inputs are identified and subsequently incorporated into the more involved closed-loop control. The validity of closed-loop controller is verified across all harmonic combinations indicating that the closed-loop controller yields additional vibration reduction as compared with the open-loop control case, overall resulting in up to 87.8% compared with the uncontrolled case. In addition, the asynchronous multicyclic IBC offers improved vibration reduction of about 3.1% as opposed to the counterpart synchronous actuation. The airloads distributions over the rotor disk before and after the optimum IBC actuation is also investigated to gain insights leading to the vibration reduction of the rotor system.
The induced and profile power of a hovering rotor was evaluated using experimental and computational methods. Momentum theory principles were coupled with experimental measurements over a range of thrust conditions to characterize the induced and profile power consumption at low Reynolds number conditions ∼ 105. An empirical induced power factor, κi, was extracted to quantify the non-ideal losses. Results show that these losses increase as the Reynolds number reduces, and nearly twice the power is required at Retip = 0.27×105 than the ideal momentum theory prediction. These results were compared with high-fidelity computational fluid dynamics simulations using the partial-pressure field (PPF) force/power decomposition to extract the induced and profile power contributions of the rotor. The PPF method decomposes the static pressure field of a numerical Reynolds-averaged Navier-Stokes solution into Euler and dissipative partial pressure fields. Simulations were performed across a range of thrust conditions, from which the induced power factor and profile drag coefficient,Cd0 , were computed for each simulation from the extracted power contributors.
A high-fidelity computational study is conducted to investigate the aerodynamic behavior and flight response of an electric Vertical Take-Off and Landing (eVTOL) multirotor configuration using unsteady computational fluid dynamics (CFD) framework. Four simulation cases are considered to examine the vehicle aerodynamics under both prescribed and fully coupled conditions. Prescribed hover and forward-flight cases isolate rotor aerodynamics and rotor-airframe interactions under constrained kinematics. Six-degree-of-freedom (6-DoF) free-flight maneuvering simulations capture the coupled evolution of aerodynamic loads, vehicle attitude, and translational motion. The results demonstrate that the high-fidelity unsteady CFD framework, coupled with rigid-body dynamics, effectively resolves the tightly coupled aerodynamic–dynamic interactions inherent to eVTOL configurations. This work provides a foundation for future investigations into trim strategies, control modeling, and expanded flight envelopes.
Helicopter blades are often modeled as one-dimensional (1D) beams and considered to undergo medium-to-large deformations. The degree of nonlinearity that a beam theory can handle greatly affects prediction accuracy. In this work, quantitative evaluations are made for static and dynamic behavior of beams and blades using the classic moderately-large deformation beam (MLB) model as reference to a geometrically exact beam (GEB) model. A rotorcraft aeromechanics analysis framework is constructed to incorporate both beam models. The framework contains various solution procedures such as trim, blade response, loads, and vibrations while allowing external interface to high-fidelity computational fluid dynamics (CFD) analysis. A validation study is performed to examine the extent of the accuracy in the large deformation behavior of benchmark beam problems in static and dynamic conditions. Next, the HART (Higher-harmonic Aeroacoustic Rotor Test) II rotor is applied to evaluate the relative prediction accuracy of the beam models. Numerical simulations indicate that the MLB model shows poor correlation with the test data or fails to reach stable solutions as the level of loads and geometric twist angles become increased substantially. In case of HART II rotor, however, the MLB-based predictions show reasonable correlations with the measured data while indicating good agreements with the GEB results. The centrifugal stiffening by the rotor rotation causes limitations in the blade deformation which essentially lead to comparable solutions between the two beam results.
The Rotor Blown Wing (RBW) is a tailsitter Vertical Takeoff and Landing (VTOL) Unmanned Aerial System (UAS) configuration that leverages cutting-edge autonomous flight controls through Sikorsky's MATRIX™ technology to create a highly capable, efficient, and scalable technology platform. By combining the benefits of fixed- and rotary-wing aircraft, the RBW configuration eliminates the need for traditional UAS launch and recovery infrastructure. This paper describes the RBW-5 prototype, a 100-pound, dual 5-foot diameter proprotor demonstrator, and discusses the comprehensive evaluation of its design and operability through a combination of flight tests, wind tunnel experiments, and computational fluid dynamics (CFD) simulations. The results demonstrate the maturity of the UAS and highlights key accomplishments of the RBW-5 program, including successful autonomous takeoff and landing and transitions between hover and forward flight, the extraction of critical "blown-physics" underlying RBW aerodynamics, and the validation of CFD models against unpowered and powered wind tunnel data.
The advanced air mobility (AAM) sector is using novel aircraft configurations and distributed electric propulsion to revolutionize aviation. These concepts require rotors that are efficient in vertical and forward flight. A concept that shows potential for this application is the slotted, natural-laminar-flow (SNLF) airfoil due to its high lift and low drag characteristics. This work explores the impacts of using an SNLF airfoil on an AAM rotor. Comparisons are made with blade element momentum theory (BEMT) method and computational fluid dynamics (CFD) to study the impact on the performance of an isolated rotor in hover. It is found that the rotational speed of the SNLF rotor can be reduced by 8% while still maintaining the necessary thrust for trim. A rotor broadband noise prediction shows that the slower SNLF rotor is 1-2dB quieter in terms of overall sound pressure level. Comparison of both rotors in forward flight indicates that the SNLF rotor consistently has a 1-2% higher propeller efficiency.
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.
In this work, comparisons between simulations & measurements in flight are proposed for different low-speed flight conditions out of ground effect on an Airbus Helicopters H175 PT1 rotorcraft equipped with a 5-bladed Spheriflex® rotor. Numerical results have been obtained by full-helicopter unsteady simulations relying on a single-rotor loose coupling approach between the Computational Structure Dynamics& Computational Fluid Dynamics parts, assuming blade elasticity and six degrees-of-freedom trim. One flight condition is tackled with both rigid-blade and elastic-blade modelling so as to highlight the influence of the blade softness on the results. The paper showcases good agreement between the simulation results & flight-test measurements regarding variations of main-rotor collective pitch, airframe attitude angles, rotor power & rotor loads with true airspeed. Airframe download is also numerically analysed.
The NASA Revolutionary Vertical Lift Technology project aims to support and guide the development of vertical flight vehicles for the benefit of the U.S. rotorcraft community and to increase the quality of life of the public. As part of this effort, the Multirotor Test Bed (MTB) – designed and built by NASA – has been tested twice at the U.S. Army 7- by 10-Foot Wind Tunnel at NASA Ames Research Center in 2019 (MTB1) and 2022 (MTB2). This study utilizes MTB2 experimental data for sensitivity studies on rotor aerodynamic performance of a quadrotor configuration using two mid-fidelity tools, the Comprehensive Hierarchical Aeromechanics Rotorcraft Model (CHARM) as well as Blade Element Theory based disk modeling in the OVERFLOW CFD solver. Additionally, this study leverages analyzing computational rotor performance predictions with experimental data to help identify future test configurations for the upcoming MTB3 test in the National Full-Scale Aerodynamics Complex 40- by 80-Foot Wind Tunnel.
This paper presents the development and application of a computational fluid dynamics (CFD) modeling approach for the Dragonfly rotorcraft lander, a NASA New Frontiers mission to study prebiotic chemistry on Titan, Saturn's largest moon. The primary CFD approach uses Siemens Digital Industries Simcenter STAR-CCM+ to generate a large database of aerodynamic loads for various flight phases, including Preparation for Powered Flight (PPF), Transition to Powered Flight (TPF), and surface flights. The mid-fidelity CFD approach relies on a steady-state Reynolds Averaged Navier Stokes (RANS) and Virtual Disk Blade Element Momentum Theory (BEMT) model to produce the aerodynamic loads for more than 3000 flight conditions. The CFD was used with Gaussian Process Regression (GPR) to create a surrogate model for predicting aerodynamic loads, aerodynamic performance, handling qualities and control margins; the surrogate is queried over 10 billion times during flight dynamics analyses. Higher fidelity CFD runs, using CREATE-AV Helios, were conducted to build confidence in the midfidelity approach and understand its limitations. The results demonstrate the effectiveness of the CFD approach in supporting the design and maturation of the Dragonfly lander from Preliminary Design to Critical Design and provide valuable insights into the complex aerodynamics of the vehicle during various flight phases.
This paper presents findings from a joint computational-experimental venture that seeks to advance the physical understanding and validation-quality database for a model-scale generic tractor proprotor–wing system during the tiltrotor conversion maneuver. This study evaluates the interactions in a quasi-static manner for various proprotor tilt angles (θ) across the tiltrotor conversion maneuver. Independent experimental measurements of the wing and proprotor loads accompany synchronous wing surface pressure measurements along with stereoscopic particle image velocimetry flow field measurements at discrete spanwise locations. High-fidelity computational fluid dynamics simulations leverage the multi-disciplinary rotorcraft simulation tool CREATE™-AV Helios to assess the interactional aerodynamics of the proprotor–wing configuration across the tiltrotor conversion maneuver. Computational simulations use a newly implemented Helios module to trim to the experimental proprotor thrust. Validation of the computational model, along with a mesh refinement study, is shown via comparison of the integrated proprotor and wing loads as well as wing surface pressure distributions. The validated computational model is used to examine the proprotor-to-wing aerodynamic interactions further. For low proprotor tilt angles (0° ≤ θ ≤ 30°), interactions between the wing boundary layer and proprotor wake results in the breakdown of the wake along the wing chord. For moderate proprotor tilt angles (45° ≤ θ ≤ 60°), the interactions are dominated by the proprotor tip vortices curling around the wing leading edge. Approaching edgewise flight (75° ≤ θ ≤ 90°), the predominant proprotor–wing interactions stem from standing vortices on the wing upper surface, producing strong low-frequency responses on the wing. Overall, this work introduces new insights into the complex proprotor–wing aerodynamic interactions across the tiltrotor conversion.
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.
This paper introduces a comprehensive model, specifically developed to inherently capture interactional effects. Due to the high computational cost associated with the large analysis matrix including variations in angle of attack, angle of sideslip, velocity, and weight, a surrogate model is used in creating aerodynamic databases. This database, which reflects interactional effects under a wide range of flight speed, angle of attack, angle of sideslip, and weight configuration, is integrated into a rotorcraft analysis tool. Simulations are performed, and results are compared against flight test data for the T625 Gökbey, covering low-speed, high-speed, rightward and climb conditions. The results highlight the impact of interactional aerodynamics on flight characteristics and load predictions. Overall, the study emphasizes the importance of including interactional effects to ensure accurate and reliable rotorcraft design in the early design stages without requiring flight test data.
Quenching is the most critical step in the sequence of heat-treating operations, aiming to preserve the solid solution formed at the solution heat-treating temperature by rapidly cooling the material to near room temperature. Currently, there is no reliable, performance-informed quenching process that can consistently reduce the high scrap rate of airframe aluminum forging parts, which often suffer from significant residual stress and distortion. This limitation stems from the complex interactions between temperature, phase transformations, and stress/strain behavior—each influenced by the evolving temperature distribution and microstructural state of the workpiece. Conventional modeling techniques for quenching processes typically lump these multiscale, multi-physics phenomena into a simplified heat transfer coefficient (HTC). However, determining the spatial and temporal variations of HTC through experiments is both prohibitively time-consuming and costly. To address this challenge and enable rapid process tailoring for reduced distortion, we have developed and validated a digital twin-based Quenching Laboratory Software (QLAB) tool. QLAB integrates a thermal multi-phase computational fluid dynamics (CFD) model, sequentially coupled with a Mechanical Threshold Stress (MTS) model and a precipitation model. The thermal CFD component captures turbulent flow, multi-phase transformations, and the complex heat transfer stages of quenching including vapor blanket formation, nucleate boiling, and convection—to accurately predict temperature evolution. The MTS-precipitation model quantifies the effects of microstructural precipitates on the material's mechanical response under thermal loading. QLAB has been thoroughly validated using representative aluminum airframe components, including aluminum bars with pockets and Lcorner parts. We demonstrate the tool's predictive accuracy by comparing its output against experimentally measured temperature and distortion fields. Finally, we apply the validated QLAB to conduct a virtual quenching test on a simplified aluminum airframe structure, showcasing its potential for performance-informed process optimization.
Preparation for Powered Flight (PPF) is a critical phase for Dragonfly, the National Aeronautics and Space Administration (NASA) mission to Saturn’s moon Titan. During PPF the descending Lander is lowered below the Backshell and uses its rotors to remove or “despin” any residual yaw motion of the vehicle. A 1/2-scale model of the Dragonfly PPF configuration was tested in the National Full-Scale Aerodynamics Complex (NFAC) 80 by 120-foot wind tunnel to measure aerodynamic loads and surface pressures on the Lander and Backshell. The results were used to improve understanding of the complex aerodynamic interactions and provide validation data for the Computational Fluid Dynamics (CFD) simulations used to develop the aerodynamic databases for full-scale, Titan conditions. Configurations tested in the wind tunnel included Lander-alone-no-rotors (L), Lander-alone-with-rotors (LR), and Lander-with-Rotors-and-Backshell (LRB). Both LR and LRB configurations were tested at multiple descent velocities and pitch and roll attitudes. The eight powered rotors were operated in nine combinations of active and inactive rotors, using steps of increasing RPM from idle to maximum. The results were used to identify conditions that generated the desired yaw moment for despin as well as those that generated reduced or opposite sign moment. In parallel, CFD simulations applied the STAR-CCM+ toolset to the experimental geometry. Time-averaged virtual disks represented the rotors by adding spatially varying axial and rotational momentum to the flow. There was generally good agreement between experiment and CFD, but conditions having strong interactions between the rotors, Lander, and Backshell were found to be improved by use of more sophisticated time accurate and discrete blade models. The lessons from the experiment are being applied to the CFD for the full-scale Lander under Titan conditions.
Blade–wake interaction (BWI) is a significant source of broadband noise and is often dominant in rotors with high blade counts. Accurately capturing the resulting unsteady blade loading is computationally expensive and, therefore, drives the cost of BWI noise calculation. To address this challenge, a low-fidelity BWI noise prediction tool was developed using aerodynamic data from the blade element momentum theory (BEMT) and the lattice Boltzmann method (LBM) for a series of rotor configurations with medium to high solidity. Starting from a six-bladed baseline rotor, 13 additional configurations were generated by varying blade twist, taper, root collective, solidity, and blade count. The relationship between vortex miss distance and blade loading unsteadiness was quantified to construct a semi-empirical BWI noise model. The model predicted BWI noise with a root mean square error of 3.9 dBA and a mean absolute percentage error of 1%. It was subsequently integrated into a BEMT framework to produce aerodynamic and acoustic data for training a tandem neural network (TNN) that was employed to optimize two rotor geometries. The optimized designs achieved up to a 7% reduction in BWI noise and a 7% improvement in performance. Additional geometric modifications—including blade tip anhedral, forward sweep, and a mixed configuration—were also assessed using LBM, each demonstrating notable noise reduction.
Accurate simulation of fluid-structure interactions (FSI) is critical for designing aircraft systems, particularly for applications involving fuel tank sloshing and large deformations. Traditional added mass methods often fail to capture the nonlinear and frequency-dependent behavior of these coupled systems. This study applies the Finite Pointset Method (FPM), a mesh-free computational fluid dynamics (CFD) technique, coupled with an explicit finite element solver, to predict complex FSI phenomena. Validation is performed using benchmark experiments, including a harmonic tank sloshing test and a guided plate ditching scenario, with results demonstrating strong agreement with measured pressures and structural responses. Additional validation on a composite fuel tank drop impact test confirms FPM's ability to model large deformations and rupture under dynamic loading. The findings highlight FPM's robustness and adaptability for aerospace FSI problems, offering a powerful alternative for virtual prototyping and certification workflows where conventional methods are insufficient.
This research paper addresses the challenge of helicopter vibrations, which have significant implications for passenger comfort, mission effectiveness, and structural integrity. The paper introduces a new tool to enhance the existing tool chain used at Airbus Helicopters. The tool integrates advanced computational structural dynamics (CSD) with computational fluid dynamics (CFD) to improve the prediction of the dynamic response of the airframe to unsteady loads. The proposed method couples well-established computational tools in a loose manner. Specifically, the coupling involves the comprehensive CAMRAD II code and the CFD solver FLOWer. The authors propose a novel method that distributes the airframe loads through an artificial node, allowing unsteady loads based on CFD predictions to be incorporated into the CAMRAD domain. Additionally, the elastic deformations predicted by CAMRAD can be integrated into the CFD domain, thereby enhancing the physical representation of the system within CFD simulations. Furthermore, any motion calculated by CAMRAD, including rigid body motions of the helicopter, can now be included in the CFD, enabling a more comprehensive analysis.
Aeroelastic stability prediction is critical to the successful design, development and flight testing of rotorcraft. As configurations reach higher speeds, new challenges in high Mach number unsteady aerodynamic modeling need to be addressed, especially for higher frequency aeroelastic modes with significant coupling. In this paper, Linear Unsteady aerodynamics and Leishman-Beddoes attached flow models are applied and compared to 2D CFD (airfoil) and 3D CFD/CSD (rotor) analysis for operating conditions of interest. The Leishman-Beddoes model demonstrates improved agreement with CFD data. In the 2D assessment, RCAS is used to model a representative airfoil undergoing prescribed pitch and heave oscillations. CFD results are presented to compare each model (Linear Unsteady and Leishman-Beddoes). In the 3D assessment, a full rotor CFD/CSD test case is evaluated for aeroelastic stability and compared to RCAS standalone analysis. The RCAS rotor structural model is coupled with the HELIOS CFD code and a swashplate cyclic pitch input is used to excite a lightly-damped rotor mode. The transient response based on RCASHELIOS is compared to the standalone RCAS internal aerodynamic result for both Linear Unsteady and Leishman- Beddoes unsteady aerodynamics. This study demonstrates that the Leishman-Beddoes model can produce similar stability results to the computationally expensive coupled CFD/CSD approach of RCAS-HELIOS, but at a lower computational cost, even for critical high-speed conditions.
Dragonfly is a rotary-wing lander, and its mission is to explore Titan. It will make multiple flights over several years to explore different sites on Titan. There is limited information on the chemical processes that led to life on earth. Among the other places in the solar system, Titan is the most like the early earth and therefore exploring its organic surface chemistry will help to better understand our own prebiotic history. During Titan flight the rotor induced unsteady aerodynamic loads, as well as the interactional aerodynamic loads due to the rotor to rotor and rotor to lander interferences drive the structural vibrations. Therefore, robust and accurate predictions of Dragonfly structural loads and vibrations are essential for designing a vehicle that can successfully perform its mission. This paper presents the structural loads and vibration predictions of the Dragonfly lander using Rotorcraft Comprehensive Analysis System (RCAS) coupled with the Viscous Vortex Particle Method (VVPM) inflow model as well as the high fidelity coupled Computational Fluid Dynamics/Computational Structural Dynamics (CFD/CSD) simulations. Elastic fuselage modal model, generated by the NASTRAN FEA, is also used to represent the dynamics of the flight lander in the simulations. Hub loads and the lander vibrations simulations are provided at multiple flight conditions. The paper also demonstrates how the comprehensive analysis can be used to provide insight on the design changes that can result in significant reduction of the structural loads and vibrations.
Blade Vortex Interaction (BVI) noise primarily occurs in rotorcraft when tip vortices generated by the blades interact with other blades. When BVI noise occurs, it dominates at mid and high blade passing frequency harmonics. To mitigate BVI noise, we employ leading-edge serrations on the OLS rotor between 75% blade span and the tip. High-fidelity computational fluid dynamics simulations, using delayed detached eddy simulation, combined with an acoustic analogy, are conducted to analyze various leading-edge serration geometries with different serration height and wavelength parameters. The results show that rotor BVI noise is reduced by up to 5 dB at the rear of the vehicle when serrations are applied, with higher serration height-to-wavelength ratios proving more effective. The findings demonstrate that when vortices directly impinge on the rotor blades, the serrations disrupt the vortices and generate a fluctuating pressure field on the blade surface, leading to destructive phase interference at the far-field microphone locations.
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.
This paper presents the implementation and validation of a state-space free-vortex wake model with a vortex lattice near-wake formulation, developed for rotorcraft applications. The model is expressed in state-variable form as a nonlinear time-periodic (NLTP) system in first-order structure, enabling linearization and simplification through time-invariant systems theory and model-order reduction techniques. It is applied to a UH-60-like rotor and evaluated in hover, forward flight, and vortex ring state (VRS) conditions. Two configurations are considered: one with tip vortex dynamics only, and another incorporating both tip vortex dynamics and a near-wake vortex lattice model. A parametric study is conducted to determine optimal parameters for solution accuracy. Validation in hover and forward flight is performed against high-fidelity computational fluid dynamics (CFD) results and available experimental data. Validation in VRS includes comparisons with experimental measurements and momentum theory. Overall, the state-space free-vortex wake model accurately predicts blade loads and induced velocities across all three validation cases, with the exception of a slight underprediction observed in fully developed VRS conditions. The inclusion of the near-wake model is essential for accurate blade load prediction in hover and contributes to improved predictions in forward flight.
This study presents an integrated optimization framework for rotor blade design that combines aerodynamic shape optimization and internal structural design within a unified multidisciplinary process. A variable fidelity modeling (VFM) approach is employed to efficiently optimize the blade outer geometry for improved figure of merit (FM) in hover and lift-to-drag ratio (L/Dq) in forward flight. Based on the optimized aerodynamic shapes, internal structural optimization is subsequently performed using a surrogate model for predicting cross-sectional properties, ensuring dynamic feasibility while minimizing blade vibration and weight. Final aeroelastic performance is evaluated through high-fidelity CFD/CSD loose coupling simulations. Optimization results show that individual designs achieve up to 6.5% improvement in FM or up to 6.6% improvement in L/Dq compared to the baseline HART II rotor. Furthermore, cross-validation comparing blades independently optimized by Seoul National University (SNU) and DLR reveals similar aerodynamic trends and performance, demonstrating the robustness and general applicability of the proposed framework across different simulation setups.
The Dragonfly relocatable lander was selected as NASA's New Frontiers mission in 2019 to explore the organic-rich surface of Titan, Saturn's largest moon. The coaxial quadrotor vehicle will fly to multiple geologic sites covering a distance of over 50 miles near the Titan equator. At each site, Dragonfly will sample materials, determine the surface composition, and investigate how far prebiotic chemistry has progressed on Titan. Upon arrival, the lander will enter the Titan atmosphere protected inside an aeroshell, which will descend and decelerate with parachutes. At an altitude of approximately 1 km above the ground, the lander will separate from the backshell and perform a controlled transition to powered flight. Prior to separation from the backshell and after the heatshield has been ejected, the Preparation for Powered Flight (PPF) sequence will be initiated, which ensures the lander is in a safe and stable state for autonomous descent. A critical element of PPF is the de-spin maneuver, where diagonally opposing rotors rotate at maximum speed to reduce any residual angular momentum by creating a yaw moment in lander body axes. This paper presents high-fidelity computational fluid dynamics simulations of the Dragonfly rotorcraft lander during the PPF sequence. Aerodynamic performance predictions are compared with test data from the National Full-Scale Aerodynamics Complex to validate the simulations and build confidence in the PPF simulation results. Blade-resolved simulations capture the unsteady and complex flow behavior in Titan's dense, low-temperature atmospheric conditions during PPF. The results are analyzed, providing insight into aerodynamic performance and the aerodynamic moments critical for mission success.
Current paper summarizes a correlation study of two flow solvers (CREATETE-AV Helios and Simcenter STAR-CCM+), routinely used at Sikorsky, with multiple model-scale wind-tunnel tests. The Helios modeling approach was aiming for a high-fidelity accurate simulation, whereas the STAR-CCM+ modeling approach was aiming for a fast turn-around time with reasonable solution accuracy with a relatively coarse mesh and simplifications. The two solvers generally agreed well with the test data within reasonable accuracy and captured the airloads and flowfield trends. The calculations presented herein show the impact of the turbulence model on component loads, the aerodynamic interactions among components, and the effect of transition modeling on rotor performance. The Reynolds-Averaged Navier-Stokes CFD model generally delayed separation and resulted in lower drag. By modeling the airframe supporting structure in CFD simulations, an improvement on correlation for inflow on the propeller plane was shown. Additionally, improvements in the rotor system L/De correlation were realized by including a turbulence-transition model, which reduces the rotor drag.
This study investigates the aerodynamic behavior of lift rotors in a representative lift+cruise electric vertical takeoff and landing (eVTOL) configuration using high-fidelity Computational Fluid Dynamics (CFD) simulations. As lift+cruise concepts gain prominence for Urban Air Mobility (UAM) applications due to their operational simplicity, flight performance, and reduced cruise noise, a detailed understanding of rotor aerodynamics during transition and cruise is critical. CFD analysis was conducted for both slowed rotors at high advance ratios and fully stopped rotors, where traditional predictive tools become inaccurate. Results show that lift rotors operating at advance ratios approaching three exhibit quasi-steady behavior similar to stopped rotors. The influence of rotor lock orientation on aerodynamic loads was characterized, with a freestream-aligned lock angle minimizing drag and asymmetry. A rotor hub fairing was found to reduce blade root separation and drag, though at the cost of slightly increased hub moments. The sensitivity of axial loads to vehicle pitch angle was quantified, highlighting the need to account for effective angle of attack in cruise load predictions. These findings inform future modeling strategies, control system design, and airframe integration for advanced eVTOL vehicles concepts.
This study explores the aerodynamics of aerial screws, drawing inspiration from Leonardo da Vinci's visionary 16th-century designs. Using high-fidelity Computational Fluid Dynamics (CFD) analysis, the research identifies a vortex-based thrust generation mechanism, centered on the formation of a "da Vinci vortex"—a coherent helical structure critical to performance. Systematic investigations into the effects of pitch and taper variations reveal nuanced strategies for optimizing thrust efficiency. Bilinear modifications to these parameters achieve up to a 13% improvement in figure of merit over linear designs. Detailed wake analyses further elucidate the importance of early vortex anchoring, swirl minimization, and coherent wake contraction for maximizing aerodynamic efficiency. This work not only deepens the understanding of aerial screw physics but also lays the groundwork for their application in next-generation Vertical Take-Off and Landing (VTOL) vehicles.
This paper describes the development process of a comprehensive pilot-in-the-loop simulation framework suitable for preliminary feasibility, and on-deck handling qualities assessment of the Leonardo AW609 civil tiltrotor, when operating with the Italian Navy aircraft carrier Cavour. A pilot-in-the-loop engineering simulator was used for simulations in which steady, quasi-unsteady, and fully unsteady ship airwakes were created using Computational Fluid Dynamics (CFD) and experimental data. A dedicated analysis of the simulation environment provided a strong agreement with various pilot inputs and aircraft response parameters when compared with flight data. Snapshot CFD simulations taken from a simulated lateral entry on ship deck allowed a comparison of airframe loads predicted by the aeromechanical model. While there are some good agreements and matched trends, development is ongoing to improve these aspects. Back-to-back piloted simulator approaches found a relatively good representativeness in terms of pilot feel and workload of both the quasi-unsteady and unsteady airwakes compared to flight.
This study investigates the active vibration reduction of a coaxial rotor system in high-speed forward flight using an individual blade control (IBC) scheme with multicyclic control strategy. A high-fidelity simulation framework is developed based on a loosely coupled analysis between a compressible three-dimensional computational fluid dynamics (CFD) solver and a comprehensive aeromechanics (CA) analysis program. The rotorcraft analysis model adopts an isolated rotor model and is validated against flight test data, resulting in reasonable agreements in predicting the rotor hub vibratory loads. Through the open-loop analysis, dominant control frequencies are identified which are used to design a closed-loop multicyclic controller. The closed-loop controller implemented using the identified system is shown to significantly reduce the rotor hub vibration. A maximum reduction up to 84.7% in vibratory hub loads is achieved in reference to the uncontrolled case. The results reveal that the multicyclic IBC actuation, integrated through using CFD/CA-based analysis, provides an accurate and robust solution for the suppression of the rotor hub vibration for a lift-offset (L.O.) coaxial rotor system operating in high-speed flights.
The emergence of three-dimensional Computational Structural Dynamics for helicopter rotors warrants the development of a higher fidelity fluid-structure interface that can replace the one-dimensional sectional airload interface commonly used for coupled analysis with Computational Fluid Dynamics. Three methods of progressively higher fidelity are examined for imposing airloads onto the structure. These are defined as level-III, II, and I, based on fluid stresses, patch forces, and sectional airloads (baseline), respectively. A model problem investigating a 3-D cylindrical shell with large deformations near the boundaries is used to verify the methods. The patch force interface (level-II) approaches the stress interface (level-III) when the mesh is highly refined. Level-I (baseline) produces no solution at all (or zero solution). Level-II is then applied to a UH-60A-like rotor and compared with level-I. Only a forced response was carried out, not a full-fledged trim solution. For this practical problem, severe mismatches of the fluid and structural meshes required a special algorithm to impose the patch forces onto the structure. The key conclusion was that a higher fidelity interface is indeed feasible for rotors, and level-II appears to be the most convenient. Even though axial (bending) stresses showed approximately 5% difference from the baseline airloads interface, the errors in shear stresses were dramatic, rising to 100% and higher inside the main load bearing parts of the rotor across its mid-span.
Advanced Air Mobility (AAM) faces operational challenges because a significant portion of AAM flight operations are likely to occur within the atmospheric boundary layer (ABL). In particular, terminal flight paths within the ABL roughness sublayer will involve flying through building wakes that will likely result in a considerable increase in significant dynamic and vibratory loads on the vehicle, affecting flight safety and ride quality. A new representative environmental method (REM) has been developed that provides real-time estimates of the unsteady wind environments, including the roughness sublayer. The approach has numerous advantages over computational fluid dynamics solutions of any fidelity, as no meshing is required and it can easily be modified to evaluate the sensitivity of different environmental factors on operations or design. This approach is explained, verified, and validated using computational and experimental data.
Rotors and propellers in edgewise flight typically encounter reverse-flow on the retreating blade, especially when operating at low rotational speeds and high speed flight. This phenomenon is well known and has been observed in rotorcraft and vertical take-off and landing (VTOL) applications, with impacts on vehicle performance and aerodynamic loads. Reverse flow is characterized by flow incident to the trailing edge of an airfoil with an angle of attack (AoA) of around 180°. Aerodynamic coefficients for reverse flow conditions are difficult to find in literature, and wind tunnel measurements often focus on the normal operating range of airfoils. This study investigates the fundamental aerodynamic characteristics of airfoils in reverse flow using high fidelity computational fluid dynamics, and analyzes the impact of using accurate aerodynamic coefficients on comprehensive rotorcraft analysis. Although the effect on flight performance is well understood, for applications on lift rotors of eVTOL configurations, reverse flow phenomena are expected to have a significant impact on rotor loads, especially during flight transition from VTOL to fixed-wing modes.
A fixed-pitch speed-controlled coaxial rotor system (Dragonfly Phase B*) was tested in the NASA Langley Transonic Dynamics Tunnel (TDT). The rotors have a diameter (D) of 1.35 meters and an inter-rotor spacing of 0.3375 meters, or D/4. The primary objective of the TDT test was to experimentally measure rotor performance of a candidate full-scale flight rotor for the Dragonfly program, NASA's 4th New Frontiers Mission, in an atmosphere as close as possible to that on Saturn's largest moon Titan. The TDT heavy gas (HG) test setup provided Mach scaled data at one-third chord-based Reynolds number when compared to Titan condition. These data serve as a validation anchor for computational fluid dynamics (CFD) performance tables used by the Dragonfly team to predict rotor performance on Titan. The present work provides a thorough CFD validation study of coaxial rotor performance estimation with accuracy of order 5-10% over the primary flight envelope using an efficient hybrid BEMT-URANS flow solver, RotCFD. Airfoil lookup tables (i.e. C81 tables) were generated in OVERFLOW, run fully-turbulent using the Spalart-Allmaras turbulence model. The full CFD performance matrix consisted of more than 1,500 cases total, including hover, climb, edgewise flight, descent, vortex ring state (VRS), turbulent wake state (TWS), and windmill brake state (WBS). Significant effort was devoted to quantitative comparisons between experimental data and CFD results, with emphasis on uncertainty quantification and confidence levels of performance predictions. This work has been instrumental in establishing the hybrid BEMT-URANS methodology to provide mean coaxial rotor performance data for Dragonfly Mobility Closed-Loop Simulations.
This paper presents a meshless large eddy simulation approach for rotorcraft wake prediction, using a vortex particle method accelerated on GPUs. The solver couples a rotor model with a vortex particle wake model, employing the Fast Multipole Method for computational efficiency and implementing viscous diffusion through Particle Strength Exchange and Core Spreading Methods. GPU acceleration achieves speed-ups of up to 10x compared to CPU execution. The solver’s predictions are validated against experimental data, showing excellent agreement. Effects of time step size, numerical integration schemes, viscous models, and particle overlap factors on simulation accuracy and computational cost are systematically analyzed. This GPU-based vortex particle framework provides a fast, accurate, and scalable tool for rotorcraft wake simulations.
This study presents computational analyses of coaxial rotor hub flows and validation against experimental data obtained from the fifth Rotor Hub Flow Prediction Workshop. Experiments were conducted in a 12-inch diameter water tunnel at Pennsylvania State Applied Research Laboratory, employing tomographic particle-image velocimetry (Tomo-PIV) and precise hub drag measurements. Three CFD codes (UMD Mercury, CREATETM-AV Helios, and OVERFLOW) utilizing hybrid Reynolds-Averaged Navier-Stokes (RANS) / Large Eddy Simulation (LES) modeling based on Spalart–Allmaras turbulence model, were applied to replicate and analyze hub flows. Counter-rotating coaxial rotor hubs under free-air condition was simulated as the simplest case and the hub drags are compared between the three CFD codes. The full water tunnel configuration, consisting of two hubs, a fairing, and shafts, was also simulated and compared to experimental results, with a focus on hub drag, wake velocity fields, and turbulence quantities. Results demonstrated that the computational frameworks effectively captured key flow physics, although some discrepancies in drag harmonics, wake velocity and turbulence intensity magnitudes were observed. Additionally, the study highlighted the impact of rotor hub geometry and installation of sail-fairing on drag and wake structures. These findings contribute to improve computational predictions, essential for designing high-speed rotor hub configurations.
Dynamic stall is an undesirable flow phenomenon that could occur on rotor blades of helicopters in forward flight due to azimuthal changes in local angle of attack resulting from blade motion, blade deformation and blade-vortex interactions. It is characterized by leading-edge vortex (LEV), or dynamic-stall vortex (DSV) shedding and significantly affects rotor performance and longevity. Therefore, the capability to predict dynamic stall, especially using rapid low-order approaches, is beneficial for vehicle design and flight-dynamics simulation. Recent work has resulted in the development of a theoretical parameter called leading-edge section parameter (LESP), which provides a measure of the suction force acting on the leading edge. It has been shown that the occurrence of dynamic stall on airfoils and finite wings corresponds to the time in an unsteady motion when the instantaneous LESP crosses a predetermined critical value. The current work shows that the critical LESP value, determined from relatively inexpensive 2D computational fluid dynamics (CFD) on an airfoil undergoing pitch and surge motions, can be used to predict the onset of dynamic stall on the section of a rotor blade in forward flight.
The CH-53K® King Stallion™ is the most advanced heavy lift helicopter developed by Sikorsky, a Lockheed Martin Company, to address the requirements of the United States Marine Corps. The aircraft was designed to support missions with a maximum design gross weight of 88,000 lbs and can carry external loads up to 36,000 lb. Performance flight tests for the CH-53K® have been completed as part of its System Design and Development (SDD) phase. Tethered hover and level forward flight performance measurements have been acquired that are used as a basis for Naval Air Training and Operating Procedures Standardization (NATOPS) flight manual performance charts. They were also used in the Key Performance Parameter (KPP) verification analysis, demonstrating that the CH-53K® exceeds its KPP for mission effectiveness. In addition to overview descriptions of the performance flight test program, the test results are herein compared with predictions from aircraft performance modeling tools that were largely based on earlier comprehensive model scale hover stand and wind tunnel tests. In some cases, the flight test results dictated that adjustments be made to the aircraft performance prediction tools so they consistently represent the overall vehicle flight performance characteristics. Recent computational fluid dynamics (CFD) simulations were employed to assist with the overall understanding of the complex aerodynamic flow field in both hover and forward flight. This includes modeling simulations using CREATE-AV™ Helios full-aircraft tool for rotor performance and interactional effects and the Simcenter STAR-CCM+ tool for isolated fuselage drag and fuselage variations. The insights gained from the CFD analysis, such as parasitic drag contributions, main rotor and tail rotor interactions, and other aerodynamics interactions, were used to make appropriate input adjustments to the aircraft performance modeling tools for enhanced correlation to the flight-test results.
Transition from hover to forward flight and vice versa represents the most critical flight phase of tiltwing aircraft. Despite its importance to ensure a safe operation, the aerodynamics of this maneuver are not sufficiently understood. This paper focuses on the study of transition flight for NASA's six-passenger tiltwing air taxi by means of high-fidelity computational fluid dynamics simulations. On the basis of a static trim solution, four points within the transition corridor are analyzed: transition mode at wing tilt angles of 60! and 44!, and airplane mode at airspeeds of 110 kt and 155 kt. We investigate the balance of forces and moments for rotor-borne and wing-borne regimes, and how rotor-on-rotor and rotor-on-wing interactions affect performance. The simulations indicate that during the early stages of transition, the vortices remain in close proximity to the proprotors, inducing large fluctuations on the order of the mean blade loading. Additionally, the blowing and swirling effects of the proprotor wakes delay flow separation over a portion of the wing. The mid-transition conditions appear to be critical, with extensive regions of separated flow over the wing. In airplane mode, proprotor-wing slipstream effects can be exploited to enhance lift generation. The results of this paper contribute to a deeper understanding of the complex aerodynamic interactions during transition flight to enable a safer and more efficient operation of tiltwing aircraft.
Vortex Ring State (VRS), a hazardous condition during rotorcraft descent, appears more problematic for high disc loadings aircraft like tiltrotors and eVTOL multi-rotors. Traditionally, semi-empirical models were calibrated using flight data to predict VRS. Recent advancements in modeling techniques and computational resources made it possible to predict the VRS by solving Navier-Stokes equations via Eulerian (VTM) or Lagrangian (VPM) approaches and conventional Computational Fluid Dynamics (CFD). The study aims to validate the capability of the open-source VPM code DUST to predict VRS boundaries and rotor loads and support the development of appropriate simulation models for the assessment of aircraft handling qualities. Finally, the effect of the rotor dynamics will be introduced to better simulate and design an escaping VRS strategy to enhance the correct training to increase future rotorcraft safety.
The capabilities of two different laminar-turbulent transition models are evaluated for the prediction of the PSP rotor performance in hover. The first transition model originates on non-local semi-empirical transition criteria that are calculated on the basis of the history of boundary layer quantities along the wall streamlines. The second one is the Langtry-Menter model that consists in two additional transport equations based on a local transition criterion. The same numerical methods and same post-processing are used with the elsA CFD solver in order to have a fair comparison between the models. Both transition modeling technics provide a good agreement with the experimental measurements concerning the transition position on the upper side of the blade. On the lower side, the predictions are less satisfactory. Transition criteria approach gives good trends while Langtry-Menter results seem to be polluted by the tip vortex flow. A grid sensitivity study shows that Langtry-Menter model requires very fine grid in order to predict the expected behavior while transition-criteria approach is less affected by the grid resolution. In terms of rotor performance, both approaches predict the experimental measurements, Langter-Menter tending to slightly overestimate the Figure of Merit where transition criteria approach slightly underestimates its value.
ABSTRACT At the end of 2014, the Group for Aeronautical Research and Technology in EURope (GARTEUR) launched an action group (named AG22) in order to address both experimentally and numerically the issue of rotor wake interacting with obstacles. Within this group, several different experiments were set up and the results were provided to all the partners in order to compare and improve their numerical methods aimed at capturing interaction effects. In the present paper, we numerically investigate the experimental database provided by Politecnico di Milano (Polimi). A low fidelity method based on free wake approach and also CFD computations with different level of modeling are compared to experimental data. It shows that free wake approach is perfectly suitable to predict interaction effects on the rotor loads as long as there is no wake re-ingestion by the rotor. In other cases, the use of CFD is mandatory. However, computational cost can greatly be reduced using some approximation (no fuselage, immersed boundary method approach for the building, wall slip boundary condition on the ground) without significant loss in the rotor loads accuracy. Moreover, such approximation still gives acceptable results with regards to the effects of the rotor wake on the building itself and across the flowfield.
ABSTRACT Interactional aerodynamic interactions between various rotorcraft components can make a large contribution to steady and unsteady loads, performance, and vibration. Wind tunnel results from a powered model test have been analyzed to identify trends in the unsteady aerodynamic pressures on the horizontal stabilizer. Flow velocity measurements were also made behind the fuselage, rotor hub, and blades. Velocity components in all three directions were separated into time-averaged, periodic, and broadband components to identify factors contributing to unsteady tail loads and provide validation data for analysis. Computational Fluid Dynamics (CFD) has been applied to four configurations of the wind tunnel model. The calculated steady rotor and fuselage forces and the unsteady tail pressures have been compared to experiment. CFD has also been applied to a flight test configuration and the results compared to measured stabilizer accelerations. When all relevant components are included, the CFD analysis captures many key features, but there remains room for improvement in resolving the quantitative details.
A two-phase wind tunnel test was conducted to evaluate aerodynamic performance on a 1/5th scale model of the Sikorsky/Boeing X2™ technology representative aircraft for Future Vertical Lift (FVL). The test program provided valuable aerodynamic data for two important elements of the design: the faired coaxial hub system and the main inlet flow leading to the engine interface. Studies from previous X2™ technology aircraft have shown that hubs, pylons and sail fairings have strong interactions, and if well integrated can lead to low drag aircraft designs. Rotorcraft main inlets generally have aggressive turns; therefore, this inlet design was investigated for distortion and total pressure loss. Accuracy of modeling these aerodynamic interactions using Computational Fluid Dynamics (CFD) and other forms of computational aerodynamic assessment requires supporting empirical testing for validation. The two wind tunnel facilities used in Phase 1 and 2 offered different and unique advantages compared to each other, which allowed the Sikorsky/Boeing team to generate a large, complementary aerodynamic and propulsion database for comprehension as well as simulation validation.
Advancements in coupling techniques between computational fluid dynamics (CFD) and computational structural dynamics (CSD) codes have permitted highly-accurate computations of rotor aeromechanics. The computational fluid dynamics portion of these methods can be extraordinarily expensive, however, so the application of dual-solver hybrid codes is of interest. This work investigates best practices for the application of the dual-solver hybrid code OVERCHARM within the Helios framework in a CFD/CSD coupled simulation, and compares the accuracy and computational cost to conventional Helios CFD/CSD simulations in two different flight conditions. It was found that large time steps and a contiguous OVERFLOW domain are most appropriate for this application, and that Helios OVERCHARM simulations can provide aerodynamic and structural loads with the same level of accuracy as conventional Helios simulations at a 2-10X speedup.
As part of a US-France Project Agreement, the US Army and ONERA are investigating mid-fidelity computational approaches for rotorcraft aerodynamics. The approaches from both groups use immersed boundaries in the place of boundary-layer-resolved meshes and actuator lines in the place of rotor blades. Results are compared between the US Army and ONERA to assess strengths and limitations of the mid-fidelity algorithms. An isolated rotor case is first used to validate and compare actuator line wake structure against a high-fidelity result. Second, a static coaxial hub is used to compare immersed boundary algorithms. In the final application, immersed boundary methods and actuator lines are used together for the Dauphin 365N configuration in forward flight.
Installation effects of the Volocopter 2-X beam structures are studied by performing high-fidelity CFD simulations of a single and three-rotor configurations in hover. The studied cases are compared with simulations without airframe to investigate the installation effects. In addition, the noise emission of the configurations is simulated by using a Ffowcs Williams-Hawkings based CAA code. Scattering effects are also included by using a BEM code. The rotors are simulated at an identical RPM and are placed in their mounting position. Furthermore, an additional setup with individual rotor RPMs is simulated for the three-rotor configuration. The installation mainly affects the rotor wake, thrust and pressure fluctuations on the rotor, while the integral aerodynamic quantities remain almost unchanged. This resulted in additional oscillations in the acoustic pressure signal. Overall, the installation increases the OSPL by about 1.5 dB, but has a greater effect on the 3-20 harmonics. The simulation data were compared with the Volocopter measurements and showed good agreement. For the three-rotor configurations, the rotor-rotor interactions were found to be dominant for the aerodynamic and acoustic performance, while the installation effects only locally influenced the noise emission by 3-4dB.
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