Browse Topic: Fuselages
A challenge in establishing rotor performance map for sizing tool during design cycle is the rotor performance uncertainty for full vehicle. Sometimes, simplified tests at different setup/scale are conducted to guide performance map, but this introduces another uncertainty due to configuration difference from full vehicle. To aid insights, validated computational fluid dynamics simulations (using CREATE-AV™ Helios) were carried out to examine hovering rotor performance prediction variations at different design stages, or different modeling/testing setup with identical blade design. Quantitative rotor figure of merit differences has been demonstrated along with descriptions of underlying physical reasons. The examined model setup includes isolated rigid blades with and without flapping, elastic blades, model-scale blades, whirl-tower conditions, blades installed on fuselage, and full-vehicle including tail rotor. Both fully turbulent flow and laminar-turbulence transition flow assumptions were simulated. Rigid blades showed a negative impact on performance due to the lack of nose-down elastic twist. Model scale suffered from lower Reynolds number effects but took advantage of delayed laminar-turbulence transition. Whirl-tower blockage and altered vortex trajectories reduced peak figure of merit but delayed stall. Fuselage installation increased performance with partial in-ground effect. However, the tail rotor disrupted the main rotor vortex system and caused a substantial figure of merit drop. The figure of merit variation summary from current study can provide qualitative trends and rough estimate of the rotor performance change along different analysis or test condition during new design process. Well validated computational fluid dynamics simulations can be used as a risk-reduction approach by comparing with simplified-model results (used during fast design cycle) with the expected full-vehicle model results.
This study evaluates the capability of Simcenter™ Flightstream™, a viscous surface-vorticity compressible-flow panel method, for predicting aerodynamic performance of rotorcraft configurations. Simulations are performed on the ROBIN-mod7 fuselage, PSP rotor, and combined rotor-fuselage system under conditions consistent with available experimental data. Results are compared against experiments and high-fidelity CFD methods, including DES, URANS, and IBM-ASM. For the isolated fuselage, Simcenter Flightstream accurately captures surface pressure distributions, particularly in attached flow regions. For the isolated rotor, thrust, torque, and figure of merit trends show strong agreement with reference data. In the rotor-fuselage configuration, the solver successfully captures interaction effects and predicts performance within experimental uncertainty. Notably, Simcenter Flightstream achieves these results with one to two orders of magnitude lower computational cost compared to high-fidelity approaches. The findings demonstrate its effectiveness as a rapid, cost-efficient tool for early-stage rotorcraft aerodynamic design and analysis.
This study investigates the dynamics and associated vibratory loads of an underactuated swashplate-less rotor and its impact on the flight dynamics of a small-scale helicopter powered by this rotor via a combined experimental and computational approach. Unlike prescribing cyclic pitch using a swashplate, here the pitch is a response to the 1/rev cyclic rotor speed input. This is enabled on the current two-bladed rotor using a skewed lag hinge that utilizes the cyclic speed variation to produce lagging motion and subsequently pitching the blades in a cyclic fashion (ƍ4 coupling) for generating the pitch and roll control moments. One of the key dynamic characteristics that distinguishes this rotor from a conventional swashplate-controlled rotor is that the two blades have dissimilar pitch and flap responses leading to high fixed-frame vibratory loads. Results show that a large 1/rev vertical shear force is transferred to the fuselage resulting in half-peak-to-peak loads of +/−0.67g. The flap responses of the two blades being out of phase caused a net inertial force, which was the dominant contributor to the vibratory vertical shear force. Upon removal of the flap hinge, the vertical shear force was reduced by almost 85% while increasing the control moment authority by 50%. The inertial moment about the rotor axis generated by cyclic lagging nearly balances the moment due to the angular acceleration (Ω ) of the rotor, which significantly reduced the dynamic torque requirement from the motor.
An internal layout design framework for a medium-class rotorcraft fuselage is attempted to build based on the idea of the energy-based load-transfer index. Load-transfer index will quantify the way in which the flight loads are distributed among the fuselage internal structural members. The static load-transfer analysis will identify an inefficient transfer region in the baseline fuselage configuration, and the resulting layout refinement will lead to a more unified load-transfer pattern and allow an additional weight reduction in the subsequent thickness-optimization stage. For a UH-60A aircraft, the existing literature provides well-established information for an airframe layout, finite-element modeling guideline, and ground vibration test correlation.
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 effects of hover operations near a partial boundary structure were assessed for a free-flying quadrotor platform under both wind-off and wind-on conditions. The partial boundary structure was selected to replicate a building facade or urban vertiport environment, providing a realistic operational context for these free-flight tests. Test points were chosen to investigate operations near the partial boundary wall and edge, and across a range of partial ground effect conditions to capture the progressive onset of ground effect characteristics. Regions of degraded vehicle performance, quantified primarily by rotor thrust coefficient (CT ) and power requirements, emerged near the partial boundary edge. These performance trends were attributed to localized changes in rotor inflow profile, characterized by near-field rotor pressure measurements. Partial ground effect was found to not resemble full ground effect until much of the vehicle had traversed over the partial boundary, with the vehicle airframe and fuselage serving as the primary factor driving the onset and development of ground effect characteristics. Under wind-on conditions, the interaction between the wind freestream and the partial boundary structure significantly shaped vehicle performance and handling qualities. The most adverse handling qualities coincided with regions of peak flow vorticity, highly unsteady flow, and large velocity gradients at the shear layer above the partial boundary.
T-tail architectures show potential for enhancing vertical tail-efficiency and lowering fuselage download and hub load cycles during low-speed transition. However, a horizontal stabilizer is principally susceptible to rotor wake impingement during cruise flight, which, in unfavorable conditions, could induce dynamic loads along with associated vibrations and structural fatigue. Predicting this phenomenon is challenging due to the complex aerodynamics and sensitive structural dynamics involved. This paper demonstrates the capabilities of a mid-fidelity simulation methodology for predicting empennage structural loads and vibrations. The approach utilizes mid-fidelity interactional aerodynamics modeling, building upon previously published Vortex-Lattice Model (VLM) results and extending them to include a Viscous Vortex Particle Wake (VVPM) analysis, coupled with a modal structural dynamics model of the fuselage. The study extends the simulation model's validation against experimental data across various flight states and sensors, incorporating a sensitivity analysis of the aerodynamic modeling. Additionally, the work presents flight state sensitivities, as well as design sensitivity studies examining the influence of main rotor blade number and T-tail planform. The results indicate that the mid-fidelity tool chain is a valuable industrial asset supporting the aeroelastic and aeromechanical design of airframe tailplanes affected by interactional effects. It allows efficient analyses with adequate numerical accuracy over a large range of operating conditions on the one hand and covering a variety of architectural choices on the other hand in view of vibratory loads and tailplane vibrations. The sensitivity study demonstrates the advantages of a high number of main rotor blades and a swept T-tail planform design for reducing vibratory loads, considering both, aerodynamic excitation and structural response.
A 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.
Between the 1920s and 1930s, aluminum started replacing wood as the primary material in aircraft construction and soon became the backbone of modern aviation. Its popularity stemmed from a combination of properties, high strength-to-weight ratio, corrosion resistance, and ease of forming that made it ideal for demanding aerospace applications. Throughout much of the 20th century, high-strength aluminum alloys dominated aircraft design, accounting for 70-80 percent of commercial airframes and more than half of many military aircraft. Even after the introduction of fiber-polymer composites in the early 2000s, aluminum has remained a critical material because it continues to offer the strength, lightness, and versatility needed for modern aviation. Industry forecasts predict that commercial air travel will double in the next 25 years, which means more pollution will be released into the atmosphere. One way to help reduce these emissions is by building airplane fuselages and wings with lighter and stronger materials.
A comprehensive numerical study was conducted to reduce helicopter rotor hub vibratory loads and fuselage vibrations using the Higher Harmonic Control (HHC) technique. A CAMRAD II model of a medium utility helicopter was developed for aeromechanical simulation, and a linear system model representing both hub vibratory load and fuselage vibration characteristics was identified offline. Optimal control inputs were then computed to minimize vibration responses under different weightings on hub vibratory load and fuselage vibration in the objective function. The predicted performance was verified through CAMRAD II simulations. Additionally, a closed-loop HHC system incorporating actuator amplitude limitations was investigated. A control algorithm regulated actuator amplitudes while maintaining phase consistency, dynamically adjusting control inputs after each iteration. The results demonstrate that the amplitude-limited closed-loop control limits excessive pitch link loads while maintaining effective vibration suppression, thereby maintaining structural safety and ensuring practical feasibility for real-world helicopter applications.
The current effort presents novel investigations of rotor-wake–surface interactions for the Dragonfly lander, NASA's rotorcraft lander to explore Titan. The numerical framework couples unsteady RANS with blade-element and virtual disk rotor models and a coupled Lagrangian particle tracking method to examine rotor–ground interactions and brownout. Simulations span a range of complexity, from isolated rotor benchmarks and rotor pairs to full eight-rotor configurations without a fuselage and the eight-rotor configuration with a simplified Dragonfly fuselage. To quantify model fidelity and near-ground shear, blade-resolved simulations of the isolated rotor are performed using Spalart–Allmaras and Reynolds Stress turbulence models with vorticity confinement, demonstrating that virtual blade models under-predict tip-vortex strength and local inflow distortion but reproduce wall shear reasonably well, whereas blade-resolved RSM solutions yield higher peak shear levels relevant to brownout prediction. These findings improve understanding of planetary rotorcraft aeromechanics and sediment transport in ground-effect while supporting ongoing efforts to assess environmental risks for Dragonfly operations and inform multi-rotor VTOL design for terrestrial applications.
Rotor-rotor and rotor-boundary aerodynamic interactions of a quadrotor system without a fuselage in ground effect and ceiling effect for varying rotor-boundary distances and hub spacings were investigated. A GPU-accelerated Lattice-Boltzmann Method (LBM) was coupled to new unsteady actuator disk method (ADM) and actuator slice method (ASM) based rotor models for this purpose. Validation was conducted against experiments for both performance and particle image velocimetry flow field data. The trends in thrust and power were accurately predicted by both actuator methods, with high computational efficiency. Interactional flow physics were resolved, causing the consistent performance benefits very close to the ground, the performance penalties caused by the fountain flow effect between rotors occurring over a limited range of ground distances, and the persistent performance augmentation in ceiling effect. The ASM rotor model, with its individual blade representation, was found to predict rotor-rotor interactions more accurately than the ADM. However, it generally overpredicted aerodynamic loads, which was attributed to the Gaussian regularization applied along the blades. The results indicate that the GPU-accelerated LBM coupled with the introduced actuator methods is capable of computationally efficient prediction of multirotor aerodynamic performance and flow fields in ground and ceiling effects.
The paper presents a general framework for building an aeromechanic model in FLIGHTLAB, suitable for high fidelity, pilot-in-the-loop simulator. The focus is on aerodynamic modeling of AW609 tiltrotor in Airplane Mode flight regime. The framework can be extended to helicopter and conversion modes with additional considerations for rotors-airframe aerodynamic interference. It can also be adapted to different tiltrotor geometries, with some adjustments depending on their peculiarities. The model uses Blade Element Theory loads evaluation of lifting surfaces, corrected with tabulated distributed loads to tune FLIGHTLAB predictions against high-fidelity aerodynamic references. Bluff bodies are modeled using force and moment tabulated data. Verification was conducted against reference data in wind tunnel mode and against flight data in trim analysis. The proposed method allowed to match lift distribution on slender bodies, as well as lift and drag integral loads, with aerodynamic references. Trim analysis has shown satisfying accordance with experimental data for all the comparison parameters. However, rotor aerodynamic modeling is still being investigated to improve correlations in torque prediction. Additionally, installation effects, such as those related to the interaction between the wing and fuselage with the rotors, are under examination to address biases in flap motion evaluation.
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, and a dual arm rotating rake to measure inflow at the propeller plane. Main rotor hub and blade system performance trends are presented as functions of speed, thrust, and body attitude. Time-averaged and unsteady pressure measurements on the fuselage, empennage, and propulsor plane characterized the interactions between the fuselage, rotor hubs and blades, empennage, and propeller over a wide range of operating conditions.
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 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 corner effect, created by the intersection of the wall and ground, an exchange between the dominating ground and wall effects is observed, with corresponding ground heights and wall distances identified.
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.
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 download across various advance ratios and thrust levels. Additionally, combining a mid-fidelity rotor with a high-fidelity fuselage offers a favorable balance between download accuracy and simulation efficiency.
This study explores the best vibration reduction using a multicyclic controller through an individual blade control (IBC) actuation scheme for a lift-offset coaxial helicopter in high-speed flight. The rotorcraft dynamics model consists of coaxial, three-bladed counter-rotating rotors and a finite element fuselage stick model constructed based on the measured data of the XH-59A helicopter. The two-way coupled rotor-body vibration analysis results exhibit excellent correlations with the test data for rotor hub loads and airframe vibrations. The best actuation scenarios are sought for the minimum vibration of the vehicle using either open- or closed-loop control scheme. It is shown that the IBC actuation effectively reduces the vibrations at both locations of the rotorcraft. The co-reduction of 3P (per rotor revolution) and 6P vibration of the rotorcraft is achieved using the multicyclic control with offline system identification. A multicyclic harmonic IBC actuation enables to suppress the rotorcraft vibration by 81.4% and 3P pilot seat vibration by 92%, respectively, as compared to the uncontrolled case, leading to a significantly reduced vibration level (below 0.05g) of the rotorcraft. The closed-loop multicyclic control using the identified system and the aircraft model shows a good correlation ensuring the suitability of the present optimal control simulations.
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.
AAM concepts use multiple distributed electric motors driving propellers and rotors to augment or directly generate lift and propulsive forces. Several current concepts incorporate separate drive systems for providing vertical lift, for takeoff and landing, and propulsive thrust for wing-borne cruising flight. Measurement of loads and performance on these rotating systems is very important in both the design and development stage, as well as for certification use and ultimately supporting HUMS monitoring. However, providing instrumentation in the rotating frame and extracting their associated measurements is often problematical, as it requires some means for both power and signals to bridge the rotating interface between the blade of the rotor/propeller and the fixed frame (fuselage) system. This paper describes work conducted to leverage prior CDI development of a novel optical telemetry/instrumentation system to create a prototype unit that can support ground and flight tests, allowing for multiple installations on the many rotors that constitute current AAM configurations. The resulting hardware was designed to expand the capabilities developed previously in types and rates of data collected, on-board processing, and user configuration options, supporting NASA and commercial organizations in their testing activities.
Hub drag reduction and understanding the resulting interactional aerodynamic impact on performance is a critical need for high speed rotorcraft designs. Wind tunnel test results are presented for rotorcraft hub models and aerodynamic interaction with a generic fuselage model. Testing included a defeatured baseline hub model, which has been the subject of previous experimental and numerical studies, and a new, low-profile, low drag hub design. Separate load measurements with a rotor test stand balance and fuselage frame load cells allow for determining component contributions of the hub and fuselage models to total drag of each tested configuration. Wake measurements using a hot-wire survey allows for comparisons of velocity deficit, turbulence, and harmonic content from different hub configurations with the fuselage. Surface pressure measurements and flow visualization on the aft section of the fuselage model are used to document the significant changes of the fuselage flow field due to each hub configuration.
Knowing a rotorcraft blade’s elastic deflections along the entire span, including twist, during flight would provide great insight into rotor dynamics, provide validation of structural codes, monitor blade structural integrity and provide blade clearance feedback relative to other blades or the fuselage. A full span SEEDIT sensor based on elongated strain gages is proposed, developed and validated for simultaneous flap, lag and twist elastic motion. Equations to convert the sensor output to deflections are derived from first principles. A custom software code designs the sensor for an arbitrary blade planform and structure and then produces the CAD files needed for a flexible circuit manufacturer. A full span sensor with a length of 54 inches (1.372m) and width of 3 inches (0.0762m) was manufactured for a Mach scale rotor blade. The sensor consists of a layer of Constantan resistive sensing traces and a layer of copper leads with Kapton insulation layers for a total thickness of 10.4 mil (0.264mm). The SEEDIT sensor was embedded in a Mach scale composite blade with a cambered airfoil, twist, sweep and varying spanwise chord. The sensor has 45 leads and is delivered to the blade manufacturer with the connector already secured to the leads for simple integration into the blade layup process. Fatigue testing demonstrated sensor life on par with the resistive material. The sensor was calibrated using motion capture cameras on the benchtop. For dynamic arbitrary motion up to 8Hz, the sensor accuracy at the blade tip was 4 mils (0.102mm) for lag, 70 mils (1.78mm) for flap and 0.23° for twist for blade elastic motions of ±500 mil (12.7mm) for lag, ±2000 mil (50.8mm) for flap and ±4.2° for twist. This is a first for flexible circuits with layers of different materials connected and for embedding a large full-span sensor within a rotorcraft blade. The sensor adds negligible complexity to the blade manufacturing process.
Magnesium alloy, known for its high strength and lightweight properties, finds widespread utilization in various technical applications. Aerospace applications, such as fuselages and steering columns, are well-suited for their utilization. These materials are frequently employed in automotive components, such as steering wheels and fuel tank lids, due to their notable corrosion resistance. The performance of magnesium alloy components remains unimproved by normal manufacturing methods due to the inherent characteristics of the material. This work introduces a contemporary approach to fabricating complex geometries through the utilization of Wire-Electro Discharge Machining (WEDM). The material utilized in this study was magnesium alloy. The investigation also considered the input parameters associated with the Wire Electrical Discharge Machining (WEDM) process, specifically the pulse duration and peak current. The findings of the study encompassed the material removal rate and surface roughness. Taguchi successfully employed a single aspect optimization technique to ascertain the ideal combination that would enhance the efficacy of the WEDM process. The findings of the investigation revealed that the proposed technique significantly improved the efficiency of the WEDM approach.
In-flight icing significantly influences the design of large passenger aircraft. Relevant aspects include sizing of the main aerodynamic surfaces, provision of anti-icing systems, and setting of operational restrictions. Empennages of large passenger aircraft are particularly affected due to the small leading edge radius, and the requirement to generate considerable lift for round out and flare, following an extended period of descent often in icing conditions. This paper describes a CFD-based investigation of the effects of sweep on the aerodynamic performance of a novel forward-swept horizontal stabilizer concept in icing conditions. The concept features an unconventional forward sweep, combined with a high lift leading edge extension (LEX) located within a fuselage induced droplet shadow zone, providing passive protection from icing. In-flight ice accretion was calculated, using Ansys FENSAP-ICE, on 10°, 15° and 20° (low, intermediate, and high) sweep horizontal stabilizers, with the tail as attached to the full-scale aircraft, for a 45-minute holding pattern. The aerodynamic analyses of the iced tails were carried out using Ansys Fluent Aero with angle of attacks from 0° to -15°. All three sweeps show reduction in lifting performance under icing conditions, due to early separation. However, comparisons of lift coefficient versus angle of attack (CL-α) curves, spanwise lift distributions, and three-dimensional flow fields show that increasing sweep reduces the early separation tendency along the span, but also reduces LEX lifting performance. This leads overall to reduced icing associated lift degradation at higher sweep. This effort has been undertaken as part of the Cleansky 2 IMPACT project (GA no. 885052).
Considerable amounts of water accumulate in aircraft fuel tanks due to condensation of vapor during flight or directly during fueling with contaminated kerosene. This can result in a misreading of the fuel meters. In certain aircraft types, ice blocks resulting from the low temperatures at high altitude flights or in winter time can even interfere with the nozzles of the fuel supply pipes from the tanks to the engines. Therefore, as part of the maintenance operations, water has to be drained in certain intervals ensuring that no remaining ice is present. In the absence of an established method for determining residual ice blocks inside, the aircraft operator has to wait long enough, in some cases too long, to start the draining procedure, leading potentially to an unnecessary long ground time. A promising technology to determine melting ice uses acoustic signals generated and emitted during ice melting. With acoustic emissions, mainly situated in the ultrasonic frequency range, a very high number of events can be recorded to characterize stress relaxation processes that occur during conversions from ice to water. In the present paper, in addition to the case of the fuel tank, the icing of a fuselage panel is also considered. The results obtained provide evidence that it is possible to determine the moment when all ice has melted. However, it is not possible to give exact figures on the amount of ice remaining or melted, which is not a limitation in practice.
This SAE Aerospace Information Report (AIR) provides various graphical displays of atmospheric variables related to aircraft icing conditions in natural clouds. It is intended as a review of recent developments on the subject, and for stimulating thought on novel ways to arrange and use the available data. Included in this Report is FAR 25 (JAR 25) Appendix C, the established Aircraft Icing Atmospheric Characterization used for engineering design, development, testing and certification of civilian aircraft to fly in aircraft icing conditions.
This document describes a practical system for a user to determine observer-to-aircraft distances. These observer-to-aircraft distances can be either closest point of approach (CPA) distances during field measurements or overhead distances during acoustic certification tests. The system uses a digital camera to record an image of the subject aircraft. A method of using commercial software to obtain the distance from such an image is presented. Potential issues which may affect accuracy are discussed.
Computational fluid dynamics simulations of the flow around the ROBIN-mod7 fuselage with PSP rotor are conducted using an immersed boundary method and an actuator surface model in OpenFOAM. The ROBIN-mod7 fuselage is represented by the immersed boundary method, while the unsteady rotor is modeled using the actuator surface model. The integration of the immersed boundary method and actuator surface model is straightforward; there is no fundamental reason to be conflicted with each other in both theory and practice. A comprehensive analysis of the generic helicopter configuration is carried out for the hovering flight condition; the isolated fuselage is simulated to provide its baseline aerodynamics, and the isolated rotor and rotor-fuselage cases are studied to measure the rotor performance in hover and the fuselage effect on the performance. The validation of each test case is conducted against both experimental measurements and computational data from the literature. The surface pressure data from the isolated fuselage case shows good agreement with the experimental measurements. Also, the rotor performance predicted on the isolated and installed rotors (rotor-fuselage) has excellent agreement with the reference data; in particular, the performance data on the installed rotor agrees with the experimental data better than the previous numerical study does. The fuselage effect has been analyzed by comparing the isolated rotor and rotor-fuselage data sets. The computational effort for different grid levels of each test case is provided. Overall, the results have demonstrated an equivalent level of accuracy compared to the previous high-fidelity simulation results at their fraction of setup and computational expenses.
The paper discusses the synthesis of linear and nonlinear observers to estimate rotor states from fuselage state measurements alone. First, the paper reviews two forms of the classical Luenberger linear observer applied to the rotor state estimation problem and identifies some limitations thereof. Thereafter, the paper proposes a new robust nonlinear discontinuous observer based on the sliding mode theory to simultaneously estimate rotor flapping and lead-lagging states from fuselage state measurements. For this new nonlinear observer, the paper presents stability analyses to determine conditions that guarantee rotor state estimation accuracy despite unknown but bounded turbulence input. The nonlinear observer also lends itself to the online estimation of the unknown turbulence input. Simulation results in calm and turbulent air conditions highlight the efficacy and performance of the nonlinear discontinuous observer. Such rotor state observers could provide an independent source of online rotor states estimates to complement or supplement in situ rotor state measurement apparatus for various flight control and health monitoring applications.
Based on the FW-H equations, the CFD method and the time domain equivalent source method (TDSEM), the rotor aeroacousitc scattering characteristics considering the fuselage aerodynamic configuration parameters are calculated and analyzed. First, a set of CFD/FW-H/TDESM hybrid analysis method for rotor/fuselage aeroacoustic characteristics is developed. The aeroacoustic characteristics of the UH-1H rotor in hover and a point acoustic source nearby a rigid sphere are calculated, and the employed numerical analysis method is validated through comparisons with reference data. Then, the aeroacoustic characteristics of the Bo-105 in hover (main rotor/fuselage) are analyzed, and the scattered noise is discussed in detail. Finally, the aeroacoustic characteristics of the AH-64 helicopter in forward flight is calculated. In addition, parameters, such as the distance between fuselage and rotor, are quantified, and some conclusions are obtained. The fuselage will influence the directivity of rotor noise, and an appropriate fuselage configuration will reduce the rotor noise at the azimuth of 250-degree(1dB in the present case).
This paper describes a new mid-fidelity, steady RANS based, method to predict the aerodynamic interactions between rotating and fixed parts of the compound rotorcraft RACER of Airbus Helicopters. The method models main rotor and side propellers with source terms, hence avoiding near-body grids and unsteady overset grid techniques. The method is then applied to the analysis of the flight mechanics of the RACER in hover flight under cross-wind conditions, as part of the de-risking of the coming flight testing. The method is validated by comparison with higher-fidelity, URANS based, results taken from the literature. Trends on aircraft controls and pitch and roll attitudes are found to be consistent. The new mid-fidelity methodology provides physical insight on how the aerodynamic interactions between main rotor, fuselage, wings, rear parts, side propellers and cross wind impact aircraft controls and attitude to maintain a stabilized hover under cross-wind.
This paper investigates the application of K-Means Clustering algorithms to traditional aircraft conceptual-level weight estimation techniques. As a proof of concept demonstration, application was narrowed to fuselage basic weight estimation with expansion to additional component weights as a planned follow on activity. A variety of weight sources were parsed and curated to produce a large, diverse dataset consisting of 82 separate aircraft with a corresponding new universal baseline regression to compare against. A K-Means Clustering algorithm was then employed that sorted aircraft into groupings based on configuration as well as topology and created an associated regression for each grouping. Configuration-based groupings utilized information such as a high-level abstraction of the structural layout as well as whether the aircraft is a fixed-wing or rotary-wing vehicle. Topology-cased groupings utilized information such as landing gear location and possession of a cargo ramp or wing. The configuration-based groupings showed modest improvement compared to the baseline regression while the topology-based groupings consistently outperformed both the baseline regression as well as the configuration-based regressions. Under all conditions, a subset of the data associated with fixed-wing aircraft was shown to be an outlier in regards to error as a result of a large range of weight and speed scales, as well as possible secondary pressurization impacts. Special treatment of the winged dataset led to further reduction in error based on unique design features, presenting an overall fuselage weight estimation methodology that leverages machine learning algorithms that can improve and inform existing best practices.
A higher harmonic control simulation along with actuators designed by the physics-based approach is attempted in this paper. The object rotorcraft used in the simulation is UH-60A Black Hawk and a multibody dynamics analysis program DYMORE is used for the simulation. The three actuators are located upon the non-rotating swashplate, and represented by the prismatic joints. Pitch angles of the rotor blades are adjusted by the combination of linear motion of the actuators. The rotor system is verified by the comparison against the references via the modal and trim analysis. The response of the fuselage is reflected regarding its entire hardware by the order reduction according to Herting's method. The fuselage is finally modeled as the beam element. A higher harmonic control with the transformation from the harmonic coefficients to the displacements of the servo actuators is to be simulated. By LQG based algorithm that is proposed by the authors, the vibration reduction capability of the present control algorithm will be verified.
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