Browse Topic: Wind tunnel tests
Despite advances in CFD, wind tunnel testing remains indispensable for aerodynamic validation, correlation, and homologation. Increasing configuration complexity, shortened development cycles, and stringent result robustness and documentation requirements demand a shift from isolated facilities to integrated, data-driven ecosystems within the overall development and company-wide test processes. We present a software-centric approach integrating wind tunnel operations into a strategic element of the Digital Thread. By orchestrating test planning, execution, data acquisition, and documentation within a unified framework, experimental data becomes reusable across projects and traceable for compliance and homologation. The interaction between CFD and physical testing is important. Such approach systematically improves simulation models with wind tunnel tests. And CFD results guide efficient test matrix definition. Extended measurement methodologies include automated actuation of active aerodynamic components in test sequences, while BEVs introduce further aerodynamic and thermal aspects for range and efficiency. Thus, extended and automated test definition down to the step-level of test sequences is introduced. Within such integrated environment, AI can be a supporting engineering tool to enhance testing. AI-based methods can assist in identifying relevant test points within complex parameter spaces and in correlating experimental and simulated results, assisting but not replacing established engineering judgment. Also, for the operating department, analyzing process data for maintenance predictions and efficiency optimizations can be assisted by AI-based methods and supporting AI-agents. The approach boosts efficiency by reducing test effort and tedious manual tasks, leading to shorter development cycles, supporting improved time-to-market. Structured workflows and standardized data handling enhance data quality, improve comparability of results, and ensure robust documentation for reliable audit trails. By combining physical testing, simulation, and intelligent processing, the wind tunnel becomes a reproducible, innovation-enabling element in modern product development, positioning software as the backbone of efficient, future-proof aerodynamic testing.
For analysing flow and acoustic induced structural vibration, a fully run time coupled framework combining a hybrid CFD-CAA approach with a modal response simulation was validated and presented at the ISVNH 2022 (SAE Technical Paper 2022-01-0938). In this paper i We apply this CFD–CAA–modal coupling method to a series-representative bonnet geometry and demonstrate its capability to capture flow and aeroacoustically driven vibration with two-way coupling. ii We analyse the modal properties of the bonnet and show that confined air volumes beneath the bonnet can introduce significant fluid loading effects, which are already embedded in experimentally validated FE modal models and must therefore be treated carefully in two-way coupled simulations. iii We validate the fully coupled aeroelastic simulation against wind-tunnel measurements with undisturbed inflow, show close agreement with the measured vibration response and analyse that the dominant excitation is in this case from below the bonnet due to acoustic pressure fluctuations.
In recent years, especially in high-performance spark-ignition engines, the thermal stress of pistons has gradually increased due to the implementation of various technologies, aimed at meeting emission reduction and specific power increase requirements. If the heat is not properly dissipated, cracking and plastic deformation of the material as well as formation of hot spots triggering pre-ignition in the combustion chamber mixture can occur. This last aspect is even more true considering innovative fuels such as hydrogen. To overcome these problems, one or more jets of oil are directed towards the piston under-crown region, impacting at high speed. This technique ensures immediate cooling and allows the engine performance to be increased without compromising the useful life. In order to optimize the oil jet effectiveness, 3D-CFD can be proficiently adopted. In this regard, the aim of this work is to define a robust numerical methodology able to simulate oil jet impingement and piston thermal field. In particular, a 3D-CFD Volume-of-Fluid (VoF) simulation is used to numerically assess the oil jet impact and provide a map of heat transfer coefficients, which, in turn, is adopted in a 3D-CHT model to estimate the piston thermal field. The proposed methodology is validated against experimental data on a high-performance engine piston. In particular, a pair of oil jets is investigated and the resulting heat transfer coefficient map is exploited to obtain the thermal field of the piston, which is finally compared to the available experimental temperature measurements. The results show that the predicted temperatures agree with the experimental data within an error lower than 2.5%.
The present work develops a computational framework for simulating the two-way coupled ship-helicopter dynamic interface using large-eddy simulation. The Simple Frigate Shape 2 geometry is modeled using the immersed boundary method, and baseline simulations under both uniform inflow and neutral atmospheric boundary layer (ABL) conditions are validated against wind tunnel measurements for two wind-over-deck angles. Rotor modeling techniques, including the actuator line model (ALM) and actuator disk model (ADM), are verified and validated across several configurations: the Knight and Hefner rotor, ONERA HAD-1 propeller, and NASA Dragonfly Phase B* coaxial rotor. The lower-fidelity ADM captures wake characteristics consistent with the ALM with up to a 9× speedup. The ADM maintains strong agreement with experimental and numerical results for integrated performance metrics and is suitable for two-way coupled simulations. The developed framework is applied to a rotor-obstacle configuration based on a GARTEUR 22 experimental survey and subsequently to a fully-coupled ship-rotor-ABL system.
The validity of using comprehensive analysis (CA) tools coupled with computational fluid dynamics (CFD) to predict the aeromechanics of classical rotor blades was proven in the literature. This paper aims to enhance this validation for the complex double-swept planform ERATO blade under high-thrust level-flight condition. In order to do so, HOST comprehensive analysis tool and elsA/HOST high-fidelity loose coupling are compared to the results of the experimental campaign of the ERATO rotor carried out by ONERA in 1998 at the S1MA transonic wind tunnel. Trim commands and airloads are reviewed and enhanced with respect to a previous publication and structural loads (flap bending moment, chord bending moment and torsion moment) are used to validate the numerical simulations. The results highlight the need for high-fidelity methods in order to improve the accuracy of both the aerodynamic and structural responses.
This paper presents the development of a ballistically launchable, 360 g coaxial Micro Air Vehicle (MAV) utilizing a two-axis gimbal thrust-vectoring system for pitch and roll control. A combined experimental and analytical framework is employed to characterize aircraft performance across the full ballistic profile, from the initial projectile phase to the transition to hover. Wind tunnel experiments are first used to quantify passive stability during the projectile phase. Subsequently, a nonlinear six-degree-of-freedom (6-DOF) flight dynamics model is developed by synthesizing the mechanical and aerodynamic models of the individual aircraft subsystems. These subsystem aerodynamic models are obtained by combining experimental lookup tables with analytical models. The flight dynamics model is then leveraged to analyze the transition phase, characterize the effectiveness of the thrust-vectoring mechanism, and establish performance guarantees across the ballistic flight profile. A key finding of this work is that the effectiveness of the two-axis gimbal thrust-vectoring mechanism is fundamentally limited by the axial velocity experienced by the propellers at the point of transition. This operational limit is quantified for the coaxial MAV, and the vehicle's ability to successfully transition to hover during a ballistic launch is validated through a single-degree-of-freedom (1-DOF) experimental testing.
A methodology was developed and validated to predict total noise from a 1/5th scale eVTOL rotor in hover by coupling 2D-RANS airfoil simulations with the comprehensive code, CHARM, for rotor loading, the acoustic code, PSU-WOPWOP, for tonal noise, and the broadband noise code, UCD-QuietFly, for broadband noise prediction. Improved sectional 2D aerodynamic inputs obtained from 2D-RANS simulations were used throughout the framework, replacing XFOIL-derived inputs to enhance the prediction accuracy of low Reynolds-number effects such as transition and laminar separation bubbles. Predictions were validated against measurements at Virginia Tech. Results indicate that turbulence model selection influences local boundary-layer development and sectional aerodynamic loading, producing modest differences in tonal noise at 4000 RPM (first BPF) and spectral differences of approximately 4 dB in broadband noise, while integrated OASPL shows slight sensitivity to turbulence model choice. At 2000 RPM, broadband noise dominates the total SPL, with transitional models predicting thinner boundary layers and reduced high-frequency trailing-edge noise. In addition, high-fidelity 3D-DDES simulations performed in OpenFOAM capture the mid-frequency noise content that is underpredicted by CHARM. Overall, the combined mid-fidelity broadband noise results and high-fidelity mid-frequency range noise results provide improved agreement with experimental spectra and demonstrate an effective approach for eVTOL rotor noise prediction.
This paper tests and validates an electric rotor-propeller phase-locking system to emulate a mechanical gearbox for lift- and thrust-compound helicopter configurations. A comprehensive control scheme is designed and integrated into the UMD compound rig to rotate the main rotor and pusher propeller at desired gear ratios with known azimuthal positions. Wind tunnel testing is completed with the system to validate the mechanically decoupled, phase-locked rotor and propeller using time-accurate camera imaging. Flowfield measurements are collected to examine longitudinal velocity variation at multiple rotor phases and gear ratios. The objective of this study is to demonstrate the feasibility of an electric gearbox-less coupled rotor-propeller system for high-speed compound helicopter wind tunnel testing.
A 5.5-ft diameter dynamically-scaled hingeless rotor was tested at high advance ratios (μ) up to 1.4, representing the first aeroelastic stability characterization of a hingeless rotor at high-μ. This paper describes the wind tunnel test setup, hover and forward flight stability data, and comprehensive analysis predictions. A novel rotating frame piezoelectric actuator-based perturbation system located in-line with the pitch links is developed to excite the blade pitch. Damping is identified via the matrix pencil method, which is shown to outperform the moving block method for the highly damped flap mode. Hover data shows constant flap damping until stall onset, where a drop in damping is observed and captured by the University of Maryland Advanced Rotorcraft Code (UMARC). UMARC has been modified to solve for elastic blade stability using linearized perturbation equations in conjunction with Floquet transition matrix theory. Elastic blade modeling is shown to be required for accurate high-μ stability predictions. In forward flight, three rotor speeds are tested corresponding to three different flap frequencies and comprehensive analysis predictions are compared against wind tunnel data. Scatter in extracted damping values is addressed using Kernel Density Estimation (KDE) to robustly identify the mode of the extracted damping distribution across repeated trials. This work validates the capability of a novel perturbation and damping extraction system to experimentally characterize rotor aeroelastic stability at high advance ratios.
The Vertical Lift Proprotor Noise Test (VLPNT) was a wind tunnel testing campaign in the NASA Langley Research Center's 14- by 22-Foot Subsonic Tunnel (14x22) aimed at measuring the aerodynamic performance and acoustic behavior of proprotors operating at conditions representative of vectored thrust UAM vehicles with articulating propulsors. This was a continuation of a proprotor performance test conducted previously in the same facility. A secondary objective of the VLPNT was to perform a scaling investigation on a subset of the tested proprotor geometries in an effort to identify low-Reynolds number impacts on performance and acoustics. It is hoped that the results of the VLPNT effort will provide modelers and vehicle developers with critical knowledge of the aerodynamic and acoustic behavior of proprotors during the transitional operating modes between vertical and axial forward flight.
The aerodynamics of propeller--wing interactions during a dynamic tiltrotor conversion maneuver were experimentally studied. This investigation builds upon previous work studying the conversion maneuver as a series of discrete tilt angles. This study varied the freestream velocity, rotational frequency, number of proprotors, proprotor spacing, and conversion time period. Wing loads, surface pressures, and particle image velocimetry were used to investigate tiltrotor aerodynamics. For the multi-proprotor configuration, as the conversion period decreased, wing performance increasingly deviated from quasi-static measurements. Dynamic effects decreased as the freestream velocity increased. Minimal dynamic effects were observed when only one proprotor was used. The greatest dynamic wing performance effects resulted from proprotor-proprotor interactions in proximity to the wing. Several nondimensional parameters including the Transition Number and reduced frequency were evaluated to assess how the dynamic effects observed in the wind tunnel may scale to full-size aircraft.
The aeromechanics of a full-wing lift-compounded slowed-rotor rotorcraft were investigated experimentally at the Glenn L. Martin Wind Tunnel, characterizing the effects of rotor shaft tilt, wing configuration, and advance ratio on performance, blade structural loads, and hub vibratory loads. Measurements were obtained across advance ratios up to μ=0.7, three shaft tilt angles (-4°, 0°, and 4°), and three wing configurations, including an asymmetric wing arrangement. The results were used to validate the University of Maryland Advanced Rotorcraft Code (UMARC) coupled rotor-wing analysis. Rearward shaft tilt and increased wing lift sharing improved lift-to-drag ratio, reduced blade structural loads, and decreased hub vibratory loads due to the rotor being placed in a descent state and being partially unloaded. Rearward shaft tilt alone yielded a 5% improvement in lift-to-drag ratio and a 32% reduction in steady rotor flap bending moment relative to the forward tilt configuration at an advance ratio of 0.5 and 4° symmetric wing incidence. A peak combined rotor and wing lift-to-drag ratio of 9.6 was achieved at an advance ratio of 0.7 at rearward shaft tilt and with an asymmetric wing incidence configuration, demonstrating the potential of lift compounding for efficient high-speed edgewise rotorcraft flight.
A wind tunnel investigation to assess the impact of rotor-fuselage spacing on the development of the Vortex Ring State and flow topology is presented. Particle Image Velocimetry was utilised to investigate flow mechanisms across a range of rotor-fuselage spacings and descent ratios, which were compared to that of an isolated rotor configuration. Mean flow data was used to identify coherent flow structures, whilst flow unsteadiness was investigated through statistical analysis of the velocity fluctuations. It was found at cases of Vortex Ring State onset, the presence of the fuselage delays the development of the Vortex Ring State for all rotor-fuselage separation distances tested. Furthermore, certain cases of rotor-fuselage spacings display a rotor-fuselage aerodynamic interaction that results in an increased effective descent ratio.
This paper presents the design, development, and successful demonstration of the first-ever ballistically tube-launched tailsitter unmanned aerial system. The vehicle expands upon the capabilities of existing tube-launched systems by simultaneously integrating the hovering capability of a rotary-wing aircraft with the efficiency and speed of a fixed-wing aircraft. To achieve this, the platform's design incorporates a novel coaxial thrust-vectoring propeller system for control in vertical flight and a unique foldable wing design for ultra-compact storage in the launch tube. The aeromechanics of the foldable wings during deployment are studied through a combination of wind tunnel experiments and flight dynamics model simulations, and the results are used to formulate a methodology for executing the ballistic launch. Simulations are also performed to characterize the robustness of the system against asynchronous deployment of the left and right wings. Experimental data, collected from flying a prototype in vertical, horizontal, and transitioning flight, demonstrate the aircraft’s flying performance. The study ultimately culminates with a demonstration of the prototype being rapidly launched from a tube at 25 m/s (56 mph) and autonomously unfolding, stabilizing, and transitioning into self-powered cruising flight.
Helicopter tail shake constitutes a significant limitation to both passenger comfort and aircraft stability. Under powered descent conditions, elevated Angle of Attack (AoA) cause flow separation around the rotor hub and engine cowling, leading to the development of an unsteady wake dominated by large-scale turbulent structures. To support the helicopter tail shake phenomenon investigation, a dedicated Particle Image Velocimetry (PIV) experimental setup was designed in this work, together with four aerodynamic devices aimed at mitigating tail shake. These components were then tested through a wind tunnel campaign with the PIV setup. The proposed aerodynamic components were conceived to either deflect the hub wake away from the tail empennages or to decrease the Turbulent Kinetic Energy (TKE) within the wake. To achieve these objectives, a dorsal fin, a horse-collar, and two spoiler configurations inspired by automotive applications were designed and experimentally evaluated. The devices were tested both as standalone solutions and in combined arrangements on a scaled helicopter wind tunnel model featuring a rotating hub and blade shanks. The vertical velocity component, was used as an indicator of wake deflection, and the Turbulent Kinetic Energy was used as an indicator of wake turbulence. The Horse Collar and the Large Spoiler showed a reduction in both indicators suggesting possible tail shake mitigating capabilities, and additional improvements were achieved when the two devices were deployed in combination.
This study presents subscale wind tunnel experiments investigating the transient aerodynamic interactions of a rotor during a continuous relative-wind approach toward the landing deck of the NATO Generic Destroyer. Time-resolved rotor loads and stereoscopic particle image velocimetry measurements were used to characterize the interacting ship-rotor flow field under headwind and quartering wind-over-deck conditions. The measurements captured the evolving influence of ship airwake, ground effect, and superstructure-induced recirculation as the rotor moved from downstream to the final hover position over the deck. The results show that rotor thrust, rolling moment, and pitching moment underwent distinct changes throughout the approach, with the loading trends varying significantly with wind-over-deck angle. Time-frequency analysis further reveals that the unsteady rotor response was concentrated in a limited band of frequencies associated with various coherent flow structures shed from the ship superstructure. Spectral proper orthogonal decomposition was used to identify the dominant airwake features responsible for these fluctuations at specific frequencies, including large-scale structures originating from the radar and hangar region. These findings demonstrate that the dynamic approach resolved both slow- and fast-changing transient aerodynamic effects along the approach path that cannot be captured with static hovering measurements.
The Vortex Ring State (VRS) is an intriguing phenomenon where rotary wings are trapped in their own wake. It is inherently difficult to model with the classic momentum theory due to the breakdown of slipstream assumptions. In practice, it is still a critical safety concern for helicopters and emerging multi-rotor platforms. Despite extensive wind tunnel tests, flight tests, and modelling over the past decades, our quantitative understanding of the underpinning flow details is still limited, because of limitations in measurements and modelling resolution. First-principles-based, high-resolution simulations could uncover the flow details, but the modelling is still rare and challenging due to complexities in the flow and flight physics, and particularly the associated high computational costs. Nevertheless, in this work, a series of high-resolution simulations of the VRS phenomenon are presented. Fully blade-resolved and unsteady simulations of an isolated helicopter rotor within the VRS were carried out for over 20 revolutions. The simulations were performed using the Helicopter Multi-Block 3 (HMB3) CFD framework developed at the University of Glasgow, solving the fully compressible and unsteady Reynolds-averaged Navier-Stokes (RANS) equations with a Scale-Adaptive Simulation (SAS) closure. The modelling revealed the formation of the large vortex ring stemming from the discrete tip vortices, and its evolution over the 20 revolutions. The blade loading distribution and evolution were analysed and compared with test data. Moreover, we extracted the inflow features from the CFD results via direct extraction and inverse Blade Element Theory (BET). The results highlighted the strong induction of the vortex ring, and the influence of secondary flow features besides the inflow. It was also noted, with the correct inflow information, the BET was able to reconstruct the VRS loading with reasonable accuracy. These high-fidelity results provide unique insight into the flow and flight physics underpinning the VRS, and contribute to the ongoing GARTEUR AG28 collaboration on multi-rotor VRS investigation.
This paper presents a wind tunnel investigation on the interactional aerodynamics of a slowed-rotor lift- and thrust-compounded helicopter model in high-speed forward flight. A systematic configuration study was conducted to isolate the aerodynamic contributions of the main rotor, wings, fuselage, and pusher propeller to the aft flowfield, measured using phase-resolved 2D-3C particle image velocimetry. Measurements were acquired at an advance ratio of 0.5 across multiple rotor thrust levels, lift offset trim states, and propeller rotational speeds. The fuselage induces a streamwise velocity deficit of nearly 50% of the freestream near the tail boom due to oncoming flow blockage. This deficit is modulated by the main rotor and wing configurations. The rotor slipstream partially alleviates the deficit by convecting high-speed freestream flow downwards. Lift offset in the asymmetric half-wing configuration suppresses the rotor wake influence, deepening the velocity deficit relative to a conventional rotor trim state. The pusher propeller partially recovers the streamwise velocity deficit, with propeller performance improving in proportion to the magnitude of the deficit due to reduced climb inflow and increased blade sectional angle of attack. These findings demonstrate that the velocity gradient aft of a compound rotorcraft impacts pusher propeller performance, highlighting the importance of flowfield-informed propeller design for improved performance in high-speed forward flight.
The Enhanced Tiltrotor blade, also known as the RGF3 blade, represents a major milestone in Leonardo Helicopters Division's pursuit of advanced rotorcraft technology. Developed at the Yeovil facility in the United Kingdom as part of a dedicated program and in collaboration with the European Clean Sky 2 initiative, it is a key enabler for the Next Generation Civil Tiltrotor Technology Demonstrator. Leveraging the AW609 airframe, the NGCTR integrates a new lateral rotor control system and a V-tail with ruddervators to expand maneuverability and control authority. The RGF3 blade combines aerodynamic efficiency with manufacturability, cost effectiveness, and certification readiness. Innovations include advanced airfoil families, highly swept anhedral tips, dual-redundant anti-ice systems, and full compatibility with legacy components. A comprehensive test campaign—covering structural loads, lightning and bird strikes, icing, and wind tunnel validation—confirmed its robustness and performance. The RGF3 blade embodies Leonardo's vision for high-speed, sustainable, and reliable next-generation rotorcraft.
This paper utilizes a combined experimental and modeling approach to investigate techniques for improving the forward-flight roll-control authority of a Quadrotor Biplane Tailsitter (QBiT). QBiT is a mechanically simple, efficient hover/cruise aircraft whose roll authority in forward flight is traditionally limited by differential propeller-torque-based control. The two roll-control enhancement techniques investigated are propeller canting and the use of ailerons. A 2-kg instrumented QBiT platform was developed and flight tested to collect high-fidelity flight data across multiple flight regimes including hover, transition, cruise, and coordinated turns. A flight dynamics model was developed and validated using wind tunnel measurements and flight-test data. Flight tests showed that the cant-only configuration exhibited limited roll authority during coordinated turns due to motor control saturation, whereas the cant-plus-aileron configuration provided improved roll performance. Using test data from forward-flight roll excitation maneuvers, roll-control authority was evaluated both in the time domain and frequency-domain. The results showed that adding ailerons increased forward-flight roll-control authority by about 2.3 times from analyzing the flight data and by up to 2.6 times based on the flight dynamics simulations.
This study examines the aerodynamic interactions between rotors in quadrotor vehicles and their impact on forward-flight stability and performance. Through wind tunnel testing of plus and cross configurations, individual rotor forces and moments were measured across varying hub spacings and advance ratios. Results indicate that rotor-rotor interference significantly alters thrust distribution, inducing unintended rolling and pitching moments. Furthermore, Particle Image Velocimetry (PIV) identified asymmetrical inflow distributions as the primary physical driver of these interactions. While increased hub spacing was found to mitigate aerodynamic coupling. These findings highlight the importance of accounting for aerodynamic interactions in multirotor vehicle design and control, particularly for trimming and optimizing forward-flight performance.
The design, testing, and analysis of a Guided Autorotative Delivery System (GADS) for suppression of incipient wildfires is described. The GADS consists of an unpowered 1 m diameter rotor, a control unit, and a payload of 2.2 kg of fire suppressant powder. On release from a fixed-wing UAV, the rotor passively deploys and enters autorotation, decelerating the payload and allowing precise delivery of the suppressant using cyclic pitch control. A numerical model of the system was developed to calculate the trajectory of the GADS during rotor deployment and descent, in the presence of ambient wind and cyclic pitch inputs. A reduced-scale model of the rotor was tested in a wind tunnel, and an uncontrolled full-scale, 1.5 kg prototype of the GADS was fabricated and tested by dropping from a hovering quadcopter as well as a fixed-wing UAV. The full-scale drop experiments validated the deployment and autorotation stability of the system, and demonstrated that the GADS maintains descent velocities suitable for incipient fire suppression (≈ 5 m/s). Numerical predictions indicate that the GADS descent trajectory can be controlled with cyclic pitch in an ambient crosswind of at least 5 m/s (10 kts). Measurements captured during the drop tests using onboard instrumentation show good qualitative agreement with numerical predictions. Future work will include drop tests with remotely controlled cyclic pitch, followed by fully autonomous controlled descent. The study establishes design guidelines for guided autorotative systems and illustrates their potential for scalable UAV-based wildfire suppression or emergency response.
The TiltRotor Aeroelastic Stability Testbed (TRAST) was developed to experimentally investigate whirl-flutter stability of tiltrotor aircraft. Previous wind-tunnel testing focused on configurations representative of current generation tiltrotors utilizing gimballed rotor hubs. The TRAST platform was also designed to support a hingeless rotor system to investigate whirl-flutter mechanisms representative of stiff proprotor configurations. This paper presents analytical whirl-flutter predictions for a hingeless rotor configuration of the TRAST model. Structural mode shapes derived from a NASTRAN finite-element model are combined with comprehensive aeroelastic analyses in CAMRAD II and RCAS. The results show that the dominant whirl-flutter mechanism differs from the gimballed configuration, with instability occurring through the wing in-plane mode rather than the wing vertical bending mode. Parametric studies examining rotor speed, pitch-spring stiffness, rotor flexibility, and diaphragm spring stiffness are conducted to evaluate the sensitivity of the predicted stability boundary. Results indicate that the hingeless configuration is significantly more stable than the equivalent gimballed configuration and exhibits different trends with rotor speed and structural stiffness. These predictions help identify configurations of interest for future wind-tunnel testing and provide insight into whirl-flutter mechanisms for hingeless tiltrotor systems.
A generalized turbulence model for rotorcraft, analogous to the Dryden or von Kármán models commonly used for fixed-wing aircraft, does not yet exist. The closest available formulations are Control-Equivalent Turbulence Input (CETI) models, which reproduce the portion of the vehicle response attributable to atmospheric disturbances through equivalent control inputs applied at the inceptor or effector level. While the underlying concept is broadly applicable, these models are highly configuration dependent, making their broader generalizability uncertain. This paper adopts a recently developed methodology to extract CETI models directly from simulation and extends it to the identification of State-Equivalent Turbulence Input (STETI) models, which are Dryden-like in form and inject turbulence-equivalent excitations directly into the state dynamics. The approach is applied to six conventional main-tail rotor helicopters spanning trainer, light, intermediate, medium, heavy, and ultra-heavy classes, from the Robinson R-22 to the Mil Mi-26, and to four tiltrotors spanning the current range of tiltrotor weight classes, from the Bell XV-15 to the Bell Boeing V-22. The resulting CETI and STETI models are used to examine cross-vehicle trends, assess the extent to which they generalize across configurations, and evaluate whether they can be parameterized in terms of fundamental rotorcraft properties, including gross weight and rotor radius, as well as derived parameters such as disk loading.
A new Adverse Environment Rotor Test Stand (AERTS II) facility was designed and constructed to investigate rotor blade icing phenomena and evaluate ice protection technologies under controlled rotating-frame conditions. The facility consists of a 350 hp (261 kW) direct-drive rotor and spray system in a walk-in freezer capable of producing controlled FAR Appendix C and Appendix O icing clouds for rotor diameters up to 22 ft (6.7 m). This work presents the facility configuration, calibration approach, and initial icing results. A liquid water content (LWC) calibration methodology based on rime ice accretion thickness was implemented to determine experimental LWC and establish repeatable operating procedures. Ice shapes from a NACA 0012 paddle blade were compared against LEWICE-based predictions, repeatability cases, and published reference wind tunnel ice shapes. Repeatability testing showed good agreement between runs, particularly in colder icing regimes, with maximum variations of 15.75% in cross-sectional ice area, 1.33% in impingement limits, and 15.12% in stagnation thickness. These results demonstrate AERTS II's ability to reproduce representative rotor icing conditions in a controlled indoor environment. The facility also demonstrated support for wirelessly controlled rotating-frame electro-thermal heater experiments, establishing AERTS II as a novel platform for rotor icing physics investigations and rotorcraft ice protection system evaluation.
The front wing of a Formula 1 car is one of the most important aerodynamic components in design development. Particularly, as it is the first to interact with the upcoming airflow, the aerodynamic flow structures generated will have a strong interaction with the remainder of the car’s components. In 2026, the Fédération Internationale de l’Automobile will introduce new regulations that incorporate new aerodynamic philosophies for the front wing, including active aerodynamics. This paper presents a design methodology study for the development of a Formula 1 2026 front wing, compliant with Issue 9 of the technical regulations. A computational-based, structured optimisation series was conducted to enhance the aerodynamic performance of a front wing concept with a focus on improving downforce, maximising efficiency, and enhancing trailing flow for the remainder of the car. The final front wing concept at 40%, running at 30 m/s, generated 189 N of downforce and 19 N of drag. Active aerodynamics, aiming to reduce drag on high-speed straights, reduced downforce and drag by 64% and 62%, respectively. For paper novelty, a 40% scale low-cost wind tunnel model of the 2026 Formula 1 front wing was designed and assembled to test, which confirmed the computational results. The experimental campaign included an X-Mode sweep to gather load data for various flap deployment angles, a wake mapping analysis, tufting to reveal flow behaviours, and flow visualisation paint to map out regions of flow separation. Wind tunnel testing was used to validate the computational results, achieving a 5.1% error in downforce, a 13.3% error in drag, and an in-depth correlation in wake characteristics, vortex shedding, and flow-visualisation methods. Therefore, this paper intends to contribute to the technical literature through an aerodynamic investigation of the 2026 Formula 1 front wings’ aerodynamic philosophies before the implementation season, also sharing wind tunnel results to support further computational developments.
MSIL (Maruti Suzuki India Limited), India’s leading automotive manufacturer, offers a diverse range of SUVs (Sports Utility Vehicles) in its portfolio. Traditionally, SUVs are associated with an assertive stance and a commanding road presence; however, this bold design language often compromises aerodynamic drag performance. Over the past decade, demand for this segment has surged, while CAFE (Corporate Average Fuel Economy) regulations have become increasingly stringent. To address this growing market need, MSIL conceptualized a new SUV - Victoris - targeted to deliver best-in-class aerodynamic efficiency in MSIL SUV portfolio. This paper details the aerodynamic development process using CFD (Computational Fluid Dynamics) and full-scale WTT (Wind Tunnel Testing). Initially, the aggressive styling of Victoris negatively impacted drag performance. Strategic exterior surface refinements and integration of aero components enabled recovery of aerodynamic efficiency. Key interventions included redesigned front and rear bumpers, roof-end treatment, quarter spoiler profiling, bumper corner vents, and aero-cut alloy wheels. Victoris successfully retained its bold styling while meeting ambitious aerodynamic targets. The optimized aerodynamic silhouette comprising hood-to-windshield transition, roof contouring, backdoor spoiler, and side body profiling - along with underbody elements such as covers, strakes, and air curtains improved drag performance by 14% compared to the initial design. Achieving this improvement required significant shape modifications and aero part additions without compromising SUV identity. This paper explains how these measures were implemented and the resulting airflow characteristics.
Modern aeroacoustic wind tunnels are required to have flat axial static pressure distribution, very low background noise levels, and minimal low-frequency pressure fluctuations. These characteristics enable accurate measurement of aerodynamic forces acting on a vehicle as well as identification of noise sources. The collector of an open-jet or ¾ open-jet wind tunnel plays a critical role in achieving these goals. Collector self-generated noise contributes to the overall background noise level in the test section, and this contribution has become more significant as other noise sources, such as the main fan, have been addressed through improvements to acoustic treatment. Ever-increasing attention to detail is required to manage noise signatures as the overall facility noise floor is lowered. Furthermore, aspects of collector design that may be beneficial to aerodynamics or pressure fluctuation tend to be some of the worst offenders for noise generation. A new collector configuration was designed during construction of the Honda Automotive Laboratories of Ohio (HALO) Wind Tunnel. The collector design balances functional requirements for aerodynamics and acoustics, with development work making use of modern computational fluid dynamics techniques and sub-scale laboratory testing. The resulting collector design enabled a flat axial static pressure distribution, low background noise levels and helped minimize low-frequency pressure fluctuations. A previous paper describes the HALO wind tunnel’s overall features and commissioning results. This paper focuses specifically on the challenges, engineering approach, and trade-offs that went into the collector design.
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