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Application of Udwadia-Kalaba method for rotorcraft analysis
No full textModeling a rotorcraft as a system of interconnected bodies requires a systematic procedure to assemble the equations of motion of the individual components and enforce the constraints between them. Such systematic treatment often involves the use of a redundant set of coordinates to describe the motion of the system so that the modeling process could easily be automated. The kinematic constraints relating these redundant coordinates are then enforced through Lagrange multipliers, resulting in a system of differential algebraic equations (DAEs). However, higher the index of the DAEs, more arduous methods need to be employed for their numerical integration. Udwadia-Kalaba (U-K) approach allows for the reduction of the equations of motion of a constrained multibody system to a set of ordinary differential equations (ODEs) even when redundant coordinates are employed. Thus, U-K approach improves the ease of implementation of the multibody simulation engine as a computer program and allows for the use of explicit time integration schemes. In this work, the U-K approach is first used to model the benchmark problem of a double four bar mechanism. The benchmarking results indicate that the U-K method provides a better accuracy than the augmented Lagrangian formulation. Subsequently, the U-K method is used to model a system of interest to the rotorcraft community, a rigid rotating flapping blade. In this case, the simulated results from the U-K method are shown to be in good agreement with those from the minimal coordinates approach. This proves that the U-K method could be used for automatic development of the equations of motion in a comprehensive rotorcraft analyses package
Aluminium alloy mmc 2009 - 15% SiCP high fatigue performances aluminium alloy for helicopters application
Get PDFIn aeronautical field, material selection is the result of regular trade-offs between different criterions that may be conflicting: weight, cost, strength performance, susceptibility to flaws, repairability,… The importance of these criterions will obviously change depending on the application (component) targeted, keeping in mind that light weighted structures remain a basic criterion for flying products. In order to minimize the weight while improving strength characteristics, AIRBUS Helicopters has developed for many years with its suppliers a metallic matrix composite, a 2xxx Aluminum alloy reinforced with 15% Silicium carbide particles, and proposes today many certified components manufactured with this material, on the full range of helicopters. For decades now, AIRBUS Helicopters has gathered a huge manufacturing and in-service experience, showing the relevance of this very specific and particular material for the retained applications, and confirming the pertinence of the choice
Compound helicopter load alleviation in forward flight
Get PDFThe potential benefits of using redundant controls on a compound helicopter for manoeuvre load alleviation whilst optimizing handling qualities are investigated. The research is focused on forward flight lateral (roll) manoeuvres and hub loads. The main control effectors of interest are the ailerons and lateral cyclic pitch. A nonlinear simulation model of a compound version of the UH-60A Black Hawk helicopter is developed. Open-loop responses of the simulation model indicate that both the ailerons and lateral cyclic pitch can be used effectively for lateral control. However, these controls have a profoundly different effect on the hub loads. An analysis of the control strategy in trim, including other redundant controls such as compound thrust, reveals that the trim strategy has a major impact on steady state loads and the required power. Hence, there is an opportunity to optimize the control allocation strategy for load alleviation purposes. System identification techniques are used to obtain accurate linear models of the lateral dynamics for control law design. Roll attitude command, attitude hold (ACAH) control laws are developed for different gearing ratios between lateral cyclic pitch and aileron deflection. The control laws are optimized for handling qualities. The minimization of hub loads is the secondary objective. It is demonstrated that compared to conventional helicopter mode control or fixed-wing mode control, predicted handling qualities can be improved and hub loads can be reduced significantly (in the order of 30%-50%) for moderate to large amplitude manoeuvres if the gearing ratio between lateral cyclic pitch and aileron deflection is carefully selected
A new comprehensive analysis tool for the preliminary design and design evaluation of helicopters - the CORAL project
Get PDFDesigning modern VTOL vehicles is a complex task that demands the interaction of disciplines such as flight mechanics, structural analysis, aerodynamics, dynamics and control, aeroelasticity, power systems and avionics. In addition, aeroacoustics has also become crucial as the certification authorities enforce to the manufacturers more restrictive environmental noise impact concerns. The project CORAL (Comprehensive Rotorcraft Analyses Lab) is concerned with the integration of all these disciplines, except for power systems and avionics, into a single easy-to-adapt and user-friendly simulation tool that allows for comprehensive analyses of not only conventional helicopter configurations but also nonconventional VTOL vehicles
Numerical simulations of active flow control for rotor vibration reduction at moderate to high advance ratios
No full textThe vibration reduction capability of active flow control (AFC) jets installed on the blades of a helicopter rotor is examined using comprehensive aeroelastic simulations. The simulations represent a four-blade hingeless rotor operating at several different advance ratios in the range of 0.20??????0.35. The closed-loop control scheme for reducing vibrations is based on the higher-harmonic control (HHC) algorithm, subject to actuator saturation constraints. Dynamic stall strongly influences vibrations at ????=0.35. Furthermore, the control sen-sitivity matrix used in the HHC algorithm varies significantly depending on the advance ratio. This produces different weightings of the 2-, 3-, 4-, and 5/Rev control harmonics used for vibration reduction. The overall level of vibrations at the rotor hub is consistently reduced by 70% to 80% below the baseline for all advance ratios considered. The reduction of in-plane shear force vibrations improves as the advance ratio increases, indicat-ing that the adverse effects of dynamic stall are alleviated by AFC. The performance penalty associated with vibration reduction is also calculated. The additional rotor power required increases with advance ratio, due to increased drag penalty associated with fluidic actuation
A method to define a scalable turbulence response model for the ship-helicopter dynamic interface
Get PDFMaritime helicopters are often required to operate in the turbulent flow or air wake generated in the lee of a ship’s superstructure. This turbulent air wake disturbs the aircraft motion and therefore compensatory control inputs need to either be provided by a pilot or a control system for accurate station keeping. This paper presents the method used to understand how aircraft size, expressed as either maximum take-off weight or rotor disc loading, affects the turbulent response of an aircraft operating in a ship’s air wake. Details of the simulation implementation and investigation are given. The turbulence responses of the aircraft are analyzed using frequency domain system identification techniques and low order equivalent systems are identified. The results of this analysis were then captured using a simple scaling law to provide an approximation of an aircraft’s turbulence response using the ambient wind-speed and aircraft disc loading. The derived scalable turbulence model consists of a transfer function whose standard deviation and break frequency relate to the ambient wind velocity and the rotor disc loading. ‘Conservative’, ‘Standard’ and ‘Optimistic’ versions of the model were created. For the standard deviation model parameter, the ‘Standard’ version of the model appears to be the best fit, except for the pitch and roll axes of stiffer hingeless or teetering rotor head vehicles, where the ‘Conservative’ fit appears to be the better model. The ‘Standard’ model appears to be the best fit for all rotorcraft types for the break frequency parameter of the model
Experimental flight test evaluation of the effects of rotor state measurements and feedback control on variable stability helicopters
No full textThe National Research Council of Canada Flight Research Laboratory (NRC-FRL) has been experimenting with high-order rotorcraft state measurements based on main rotor hub mounted sensor systems installed on the NRC Bell 412 Advanced Systems Research Aircraft (ASRA). Utilizing the NRC Bell 205A Airborne Simulator (AS), a model following flight test investigation of the effects of rotor state measurements and feedback on flying, handling and ride qualities was performed. Using an eight degree-of-freedom mathematical model of the Bell 412 ASRA that included rotor flapping dynamics, several control system designs were developed. Desktop simulation was used to investigate controllers designed using Root Locus Method (RLM), Classical Multivariable Control (CMC), and Eigenstructure Assignment Control (EAC) algorithms featuring rigid-body and rotor state dynamics and feedback. The research concluded that rotor state feedback of longitudinal and lateral disc tilt dynamics significantly improves inter-axis decoupling, disturbance rejection characteristics, rotor response dynamics, command tracking accuracy, and rigid-body bandwidth performance
Optimal trajectory and tracking control system for a helicopter safe landing in autorotation
No full textIn case of engine failure autorotation represent a safe maneuver for helicopter emergency landing. In autorotation, the rotor is no longer engine-powered but keeps rotating thanks to the impinging flow. During the descent, rotor angular velocity is constant for a given setting of collective pitch and the power dissipated by its rotation is compensated by a loss of potential energy during the descent. A steady descent is then possible down to a fixed altitude where, exploiting rotor residual energy, it is possible to perform a flare bringing the helicopter to minimum touch down velocity. For small scale helicopters an autonomous control system for autorotation can reduce damages and economic losses, saving the rotorcraft and its payload in case of an engine failure. In this paper, a preliminary investigation on steady descent conditions in autorotation, and the design of a possible flare maneuver has been made through a model-based design approach. Also, a closed loop control system has been developed, to perform the two main phases of autorotation maneuver. Simulations results show the suitability of the proposed approach for a wide range of initial conditions (altitude and advancing velocity)
Conceptual design and performance analysis software tool for rotorcraft systems
Get PDFThe availability of a free open source tool for the design and performance evaluation of rotary wing aircrafts is scarce, this work intends to provide a solution for engineering students, aircraft engineers or anyone interested in aircraft design with such a tool. The tool applies the Momentum theory and the Blade Element theory to execute the analysis of the rotor performance and presents the results in a clear and simple way (power against airspeed plots, rotor disk distribution plots of the calculated variables, as well as numerical results explicitly identified). A small data base with airfoils (and their aerodynamic performances) is included in the tool. The tool’s results validation is done and presented to assure the user of the reliability of the analyses done
High-fidelity CFD simulation and optimisation of ducted propellers
No full textThis paper presents numerical analyses and optimisation of aerodynamics of ducted/un-ducted propellers, based on high-fidelity CFD methods with adjoint formulations. A ducted propeller test case by NASA is chosen as the validation case and baseline design. High-fidelity CFD simulations are performed using the in-house solver HMB3. An adjoint-based aerodynamic optimisation framework is proposed and applied to studies of the separate and coupled duct shape and blade twist designs for the ducted propeller, with blades resolved. A new parametrisation method for the duct shape is proposed, allowing variations in curve shapes and leading/trailing edge point positions. Sensitivities of the performance with respect to design variables of the duct shape and blade twist are solved using the HMB3 adjoint solver and are analysed in detail. Performance comparisons of the optimised designs are presented and detailed