Browse Topic: Fuselages
A 1/5th scale powered coaxial rotor and propeller system has been developed and tested in the National Full Scale Aerodynamic Complex (NFAC) 40x80 ft Wind Tunnel. Test conditions include airspeeds in excess of 250 kts, the highest recorded for a rotor in edgewise flight at the NFAC. The system was studied in four configurations: a powered coaxial rotor, a powered coaxial rotor with a propeller wake rake, a powered coaxial rotor with a powered propeller, and a bare hub rotor with a propeller wake rake. The high-quality data from the test included propeller, fuselage and main-rotor performance; aerodynamic-interactions between the rotors, fuselage, empennage, and propeller; acoustics and handling-qualities attributes. These results have been used to validate physics-based rotorcraft modeling tools and enhance the quality of full-scale X2 Technology® aircraft designs. Innovative solutions to test measurement challenges included rotor shaft strain gages, balance thermal control systems
The paper presents a general framework for building an aeromechanic model in FLIGHTLAB, suitable for high fidelity, pilot-in-the-loop simulator. The focus is on aerodynamic modeling of AW609 tiltrotor in Airplane Mode flight regime. The framework can be extended to helicopter and conversion modes with additional considerations for rotors-airframe aerodynamic interference. It can also be adapted to different tiltrotor geometries, with some adjustments depending on their peculiarities. The model uses Blade Element Theory loads evaluation of lifting surfaces, corrected with tabulated distributed loads to tune FLIGHTLAB predictions against high-fidelity aerodynamic references. Bluff bodies are modeled using force and moment tabulated data. Verification was conducted against reference data in wind tunnel mode and against flight data in trim analysis. The proposed method allowed to match lift distribution on slender bodies, as well as lift and drag integral loads, with aerodynamic references
The influence of ground, wall, and corner boundaries on multirotor vehicle performance was investigated through a series of controlled flight tests. Changes in rotor inflow profiles were represented by near-field rotor pressure measurements captured by a custom Kiel probe wake rake. Ground effect was characterized by reduced thrust and power requirements, primarily driven by the vehicle fuselage, which induced regions of reduced pressure and increased flow unsteadiness around the airframe. Operating near a wall boundary was found to restrict airflow into the portion of the rotor disk closest to the wall, leading to increased power requirements to maintain hover and a consequent reduction in performance. While vehicle orientation had minimal impact on overall rotor performance, it did influence local rotor inflow behavior near the wall, depending on the relative position of the interaction region formed with adjacent rotors. As the vehicle descends from the isolated wall effect into
Mid-fidelity computational techniques have long been sought after in the engineering community to expedite the generation of high-quality engineering data. As digital engineering gains prominence, the demand for faster computational methods continues to grow. Within the rotorcraft community, actuator line and immersed boundary methods play a crucial role as mid-fidelity tools for modeling full helicopters. This study investigates the efficacy of mid-fidelity immersed boundary and actuator line methods using the HPCMP CREATETM-AV Helios ROAM model in predicting the fuselage download of the ROBIN wind tunnel model. Predictions from these methods are compared against both high-fidelity computations and available wind tunnel data. The study also examines the impact of combining mid-fidelity and high-fidelity elements on the results and the time required for solution. The findings indicate that employing mid-fidelity rotor and fuselage models yields sufficiently accurate trends in fuselage
NRC developed a higher-order mathematical model structure of coupled rotor-body flapping dynamics for inflight control applications. The hybrid (rigid body fuselage state and rotating hub rotor state) 8DOF model was developed utilizing explicit measurements from a novel rotor hub state measurement system enabling estimation rotor blade dynamics. The method identified second-order rotor flap dynamics, attitude-rate and rotor flap dynamics response correlation, and response lead of rotor flap dynamics over rigid body dynamics. Reducing implementation resource burdens of past approaches, this novel rotor state measurement and modelling methodology may prove useful in applied development cycles across a spectrum of needs for articulated (helicopter) and non-articulated rotor (tiltrotor, eVTOL) aeromechanics, modelling, monitoring, and operations.
ABSTRACT Wind-tunnel tests of a heavy-class helicopter model were carried out to evaluate the effectiveness of passive flow control system in alleviating the fuselage parasite drag. An array of counter rotating vortex generators was selected to reduce/remove the flow separation occurring on the rear loading ramp responsible of the high pressure drag. Different technical solution for the VGs design and location were selected with respect to previous work. The basic fuselage geometrically scaled 1:7 of a heavy class helicopter was investigated with and without passive flow control system. The comprehensive experimental campaign involved the use of different measurement techniques. Indeed, pressure measurements and stereo particle image velocimetry surveys were performed to gain a physical insight about the results of load measurements. This paper addresses the promising results obtained during the wind-tunnel campaign, since significant drag reduction was achieved for a wide range of
ABSTRACT In this work, iced rotors are studied to develop insight in the potential of acoustics-based ice detection. Based on the HMB CFD solver, approximate iced shapes are used and results are analyzed using the FW-H method. Several candidate monitoring positions are assessed for acoustic sensors to be placed on the helicopter fuselage. The influence of ice on the aero-acoustic characteristics of a rotor is calculated, and parameters such as the ice amount and the icing position on the blade are quantified.
ABSTRACT Accurate prediction of aeroelastic coupling between rotor wake and structure remains a key challenge to the development of advanced rotorcraft. Limitations of existing analysis tools to predict such aeroelastic interactions, notably empennage buffeting effects, have resulted in costly late-cycle design changes in multiple rotorcraft development programs, including the UH-60A and AH-64A. Aeromechanical phenomena involving interactions of the fuselage and rotor wake are complex, interdisciplinary, and three-dimensional in nature. For this reason, full vehicle CFD/CSD coupled analysis is essential to accurately capture the mutually dependent interactions between the aerodynamic loads and the aeroelastic response associated with these phenomena. The current state-of-the-art in rotorcraft analysis involves CFD/CSD coupled analysis of aeroelastic rotors and wings, but rigid representations of the fuselage and empennage structures (Ref. 1). To address this limitation, an elastic
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
The AW609 tiltrotor features a unique high-mounted wing with rotatable nacelles positioned at the wing tips, it is capable of operating both in airplane and vertical flight mode. To achieve suited protection of the occupants during emergency landing, the wing - which is particularly stiff in order to sustain the heavy weights at the tips, where rotors, engines and transmissions are positioned - implements a controlled failure mechanism at root, so that during emergency landings it breaks and unloads the fuselage of the weight of wingbox and nacelles, thus avoiding catastrophic collapse. As the effectiveness of such mechanism was never demonstrated under impact conditions, certification agencies requested an empirical validation through experimental testing. The test was carried out July 2022 at Polytechnic of Milan, Italy; the present work details the Test activity, from its preliminary phases to the Test Day, to the analyses of its outcomes.
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A new measurement capability was created by combining photogrammetry and metrology techniques to accurately measure one half of the XV-15 Tilt Rotor Research Aircraft at the Smithsonian’s Udvar-Hazy museum. The challenges imposed by the fuselage and surrounding environment at Udvar-Hazy were overcome by careful application of photogrammetry and metrology techniques. Data analyses and processing included the use of multiple reverse engineering programs to accurately generate a complete 3-dimensional water-tight geometry of the aircraft and rotor blade. This paper describes the photogrammetry and metrology measurement systems, technology and hardware set-up, data analysis and processing methods, future work, and lessons learned. In addition, selected measurement results of the fuselage and rotor blade are presented.
High fidelity code-to-code comparisons have been made between the University of Glasgow HMB3 code and the HPCMP CREATE™-AV Helios code under The Technical Cooperation Program collaboration project, Next Generation Rotor Blade Design. The comparisons are made for two model-scale rotors - Langley baseline (LBL) rotor and Pressure Sensitive Paint (PSP) rotor. Hover and forward flight performance results are compared against test data. For the LBL rotor, hover performance is in a good agreement between the test data and HMB3 results over a full range of CT. However, the comparison between the HMB3 and Helios results at a CT of 0.0084 shows the difference in Figure of Merit (FM) by approximately 2 counts (2.2-3.2%). In forward flight, the HMB3 and Helios performance results overpredict the test data at the low advance ratios but improve the predictions at the high advance ratios. At an advance ratio of 0.31, the code-to-code comparison indicated that the Helios torque was lower by 2.8-3.1
A new hardware-in-the-loop (HIL) dynamic wind tunnel setup is used to study the behavior of a slung load at high speeds and methods of stabilizing problematic loads. The main element of the setup is a movable cargo hook. In addition the cable angles, model spatial attitude, and hook force are measured continuously. All the measurements are fed into a computer that calculates the cargo hook resultant motion in real-time by summing the rotorcraft angular motion effects (not used in the current study) and the hook motion relative to the rotorcraft fuselage. The computer output includes motion commands to the hook. The slung loads are two configurations of an M119 howitzer: folded and ready for firing. Initial wind tunnel studies showed that these loads exhibit significant LCO (Limit Cycle Oscillations) and severe instabilities at high speeds. Frequency sweep tests are used to derive dynamic models of the slung loads. These models are used to develop two controllers based on an Active
An aeroelastic coupling framework is applied to the UH-60A platform to examine aerodynamic-induced vibrations at four advance ratios spanning the flight envelope. Both one-way and two-way aeroelastic coupling results are examined at each condition. The two-way coupled results are observed to generally predict closer values to measured flight test data on the lifting surfaces of the empennage, and a less pronounced effect is seen in stiffer, nonlifting structure. The effect of aeroelastic coupling subiterations is examined, and they are found to further refine the two-way coupled results, generally improving prediction quality.
Computations were performed to assess the effect of fluidically-oscillating jets on a ROBIN-mod7 helicopter fuselage. The simulations utilize previously experimentally validated methodologies that rely on a new boundary condition formulation at the actuator throats, based on phase-averaged flow variables, which obviates the need to resolve the internal cavities simultaneously with the outer flow. Predictions of the base flow past the helicopter fuselage were validated against experimental and computational data available in the literature. The fluidic oscillator characteristics were then evaluated at different scales and pressure ratios, and invariant quantities were identified. In the flow control evaluation, flow separation was significantly reduced and, in some cases, suppressed. However, drag reduction was not obtained, indicating the sensitivity of the actuation location and operating conditions to the vehicle design and flight orientation.
ABSTRACT T-tail configurations are a promising approach to increase vertical tail efficiency, reduce fuselage download and hub load cycle amplitudes in low speed transition. However, the horizontal tail can be subject to rotor wake impingement in cruise flight which might lead to high dynamic loads and structural fatigue. The involved aerodynamics are in addition highly complex and hence difficult to be predicted by simulation. In this work a simulation approach for empennage structural loads and vibration prediction is established based on free-wake analysis and modal fuselage approximation, focusing on the expectedly most dominant aerodynamic interaction effects at the T-tail. The results are compared to flight test data to evaluate the approach, and sensitivities of the framework are assessed. The results indicate that the motion of the horizontal tail is characterized only by a few modeshapes, predominantly driven by rotor wake influence, rather than rotor loads via the structural
Forward-flight predictions of a rotor-fuselage system are made using a computational fluid dynamics approach. Solutions are generated using the NASA OVERFLOW solver with laminar-turbulent transition modeling enabled. Simulations are based on a hover tip Mach number of 0.57 at an advance ratio of 0.30 with prescribed collective and cyclic pitch angles. The configuration is based on a three-bladed Pressure Sensitive Paint rotor with a Rotor Body Interaction mod7 fuselage that was tested in the NASA Langley 14x22 subsonic tunnel. Grid generation and computational methods are described. The rotor and fuselage flowfield is analyzed and the effect of the transition model is assessed. Predicted forward flight performance is compared with measurements obtained at NASA Langley.
This paper presents research addressing the technological gap in the predictive capabilities of modern computational fluid-structures interaction (FSI) in the context of the US Navy's requirement for accurate loads prediction for both critical rotorhead and fuselage components in the context of a real aircraft’s fatigue life tracking program, with the ultimate objective of loads accuracy and performance robust enough to support near real-time lifing assessments across the full flight regime. This research, funded via US Navy STTR N17A-T009, touches on a broad range of innovative research areas, starting with application of the Load Confluence Algorithm (LCA) to a coupled main rotor-fuselage/tail rotor model for the UH-60A. This paper will document research to date, extending beyond traditional approaches such as the main rotor analyzed in isolation or a fuselage model affected only by the inflow generated via an actuator disk model of the main rotor. This includes the development of
Rotorcraft with a teetering rotor design are susceptible to a phenomenon known as "mast bumping" or “excessive flapping” which can lead to severe shaft structural damage followed by total separation of the rotor from the vehicle and a potential incursion of the rotor blade into the fuselage. Mast bumping accidents are nearly always fatal and are generally unavoidable once specific flight conditions are met. Certain teetering rotor vehicles are prohibited from specific maneuvers that may lead to mast bumping events. However, specific incidents indicate that certain causes of mast bumping may have not yet been determined, and the extreme danger of the phenomenon makes studies using flight testing impossible. This research uses the Rotorcraft Comprehensive Analysis System (RCAS) to create a physics-based, parameterized model of a nominal teetering rotor helicopter to simulate and assess the mast bumping risk of various level flight conditions and specific maneuvers. This data is used to
Experimental measurements of the unsteady flow fields generated by a scale model rotor, hub, and fuselage, plus the unsteady loads generated on a horizontal stabilizer, have been used as the basis for comparison to two computational fluid dynamics (CFD) simulations. The STAR-CCM+ commercial solver and CREATETM-AV HELIOS using the KCFD and SAMCART solver were applied to a series of seven test cases. The configurations were fuselage and hub with blades-on and blades-off for velocity fields, as well as the stabilizer in two locations for unsteady normal forces. The quantities examined included time averaged rotor, hub, fuselage, and tail forces and moments, time averaged, unsteady, and periodic velocities, and stabilizer forces. Overall for the forces and velocities, both codes did well for the time averages, and captured the trends and qualitative features of the unsteady quantities. Cases driven by a strong tip vortex – stabilizer interaction were modelled well, the key issue being
2016, to form their 34th Student Design Competition around designing an aircraft capable of hovering for a duration of 24 hours with a 176 lb (80 kg) payload. This paper presents the preliminary design of the Swarm, a configuration of the Electric Powered Reconfigurable Rotor (EPR2 ) concept, drafted by the joint Universite de Sherbrooke and Georgia Tech team. The Swarm, comprised of a central fuselage tethered to three electrically propelled unmanned fixed-wing aircraft flying in a circular trajectory, achieves its goal through rotor reconfiguration. Throughout the flight, the fixed-wing aircraft modify their trajectory and airspeed to optimize performance and minimize fuel burn rate. A hybrid-electric powerplant in the fuselage supplies power via conductive tethers to the fixed-wing aircraft. The design and reconfiguration were optimized using a 2-step time-marching optimization. The designed system greatly exceeds the competition requirements by accomplishing 31 hours of hover
The effect of uncertainty in Reynolds-Averaged Navier–Stokes (RANS) simulation is determined through application of Uncertainty Quantification (UQ). In the present study, the sensitivity of aerodynamic coefficients to uncertainty in freestream turbulence intensity (FSTI) and surface roughness is computed for both a rotorcraft fuselage (ROBIN- Mod7) and an airfoil (SC1095). Laminar-turbulent transition model has been extended to account for roughness- induced transition by incorporating an existing roughness-induced transition model. Current UQ analysis is based on the Monte Carlo method with Gaussian distributions of uncertain input parameters. The use of a surrogate model allows for incorporating the results from intensive RANS simulations into a Monte Carlo method. The surrogate model is generated using either a cubic interpolation for a single uncertain parameter or a radial basis function (RBF) for multiple uncertain parameters. The stochastic standard deviation is measured as an
The HPCMP CREATE™-AV Helios is a rotorcraft aeromechanics simulation framework that combines higher fidelity Computational Fluid Dynamics (CFD) based aerodynamic modeling with rotorcraft comprehensive analysis. Vehicles with multiple rotors and differing rotor frequencies require several enhancements to the traditional CFD/CSD coupling approach used for rotorcraft aeromechanics predictions. Additionally, CFD/CSD coupling modifications are also required for non-rotor aerodynamic entities such as fuselages and wings whose aerodynamic loads are influenced by the interference and impingement of multiple rotor wakes. Methodologies for consistently coupling multiple rotors with non-harmonic arbitrary frequencies are developed. The developed technique also allows for variations in rotor frequencies at each CFD/CSD coupling iteration. The methodology is successfully demonstrated to achieve full vehicle trim for the UH-60A with aerodynamic loading on all main wetted surfaces, i.e main-rotor
ABSTRACT The Bell 525 main rotor Pylon Bipod Mount is a flight critical structural component that must meet the fatigue tolerance requirements of §29.571. The complex flight load environment of the Pylon Bipod Mount involves coupling the rotor loads to the fuselage through interactions that factor roof beam displacement and the unique stress state induced by the elastomeric inner member of the Bipod. Simulating the full flight load structural response of the Bipod in a physical component test is not only extremely time consuming but also runs the risk of being inaccurate. As an alternative, an experimentally correlated finite element model is used to analyze the Bipod under the flight load environment. The approach incorporates the installation stress on the Bipod induced by its elastomeric component. The associated finite element models will be validated using state of the art experimental techniques of strain survey/mapping. Results from the correlated analytical models will be
ABSTRACT This work investigates rotorcraft-pilot coupling phenomena in tiltrotors. A detailed tiltrotor model, representative of the Bell-Boeing XV-15, has been built. Biomechanical models of the pilot, acting on the power lever and on the centre stick, are included in feedback loop to define the Pilot-Vehicle System. Pilot-Assisted Oscillation phenomena are investigated on the overall conversion corridor using Nyquist's criterion. Pilot-in-the-loop analyses demonstrate that a critical parameter is detected in the vertical fins geometry. Due to an asymmetric flaperons deflection the wing's wake impacts on the vertical fins, producing a side force. The pulsating tail-side-force makes the fuselage to yaw and excites the asymmetric wing chord mode coupled with the lateral pilot's biomechanics, leading to a reduction, or even a loss, of stability. No unstable event is detected about the longitudinal direction. Conversely, a resonance between the pilot's biomechanics and the aircraft poorly
ABSTRACT This study is focused on the improvement of UH-60A blade structural loads correlation with flight and wind-tunnel test data. The blade airloads prediction has been proven to be reasonably good from past studies. However, the blade structural loads, especially the edgewise bending moment, were poorly predicted, and these are the subject of the current study. Several variations of modeling effort have been examined. This includes the drive-train model, refined grid, lag-damper model variation, hub impedance with test stand, tunnel wall, and dis-similar blades. The drive-train and refined grid showed noticeable improvement in edgewise bending moment, but not enough to close the gap between measured data and prediction. A lag damper study indicated that the current nonlinear damper seems to be adequate, and variations to the lag damper had a limited impact to inboard region only. The influence of dis-similar blades provides further insight into the sensitivity of the edgewise
ABSTRACT Rotor cant is simulated on an SUI Endurance quadcopter. Two types of rotor cant, flapwise and torsional cant, are defined, and multirotor coordinates are used to define four aircraft-level modes of cant for each type. Collective flapwise cant causes an increase in collective control and power required, and a positive correlation exists between collective flapwise cant and pitch control. It also causes the longitudinal and lateral poles to retreat from the origin. Postive longitudinal flapwise and negative lateral torsional cant cause a reduction in nose-down attitude in forward flight, reducing drag and negative lift on the fuselage by 13% and 31% at 15 m/s, which reduces power required by 6% while increasing hover power by only 0.5%. Lateral flapwise cant and longitudinal torsional cant affect the roll attitude, though no power savings is available. Differential flapwise cant causes forward speed to impose a net rolling moment, which is compensated by roll control
ABSTRACT A feedback controller is designed and implemented for a regular hexacopter based on the AeroQuad Cyclone ARF kit. This controller is designed with an inner loop control law as a set of parallel PID controllers for aircraft altitude, pitch, roll, and yaw attitudes, as well as an outer loop for control over aircraft body velocities. Rotor failure is modeled in the dynamic simulation by setting the rotor force and moment output to be zero regardless of the commanded control input to that rotor, the feedback controller utilizes no knowledge of this fault during simulation. Various trajectories are commanded to examine the performance of the baseline feedback controller in the event of forward rotor failure, including hover, forward flight, and more complex maneuvers. The controller is demonstrated to recover the aircraft states after the transient effects of the rotor failure, as well as complete the defined state trajectory, demonstrating tolerance to single rotor failure.
ABSTRACT The US Army's Aviation Development Directorate (ADD) has successfully collaborated with its industry partners to reduce system parasitic weight for aviation platforms through multifunctional structures technology development. In short, this can be generalized as achieving weight savings by replacing the combination of aircraft structure and an independent, add-on mission enabler with a singular system that performs the functions of both structure and mission enabler. This extensive multifunctional technology development for aviation structural applications has yielded significant weight savings over parasitic designs. Technologies demonstrating this structural multifunctionality for weight reduction include integrally armored helicopter floor, lightweight integrally armored helicopter floor, lightning-protected structure, structural antenna aperture, helicopter empennage antenna structure, combat tempered aft fuselage, blast attenuating aircraft structure, and highly durable
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