Evaluation of High-Fidelity Multidisciplinary Sensitivity-Analysis Framework for Multipoint Rotorcraft Optimization

This paper describes the aeroelastic stability test of the small-scaled 'Next-generation Blade(NRSB)' with NRSH (next-generation hub system) and HCTH hingeless hub system in hover and forward flight conditions. Excitation tests of rotor system installed in GSRTS (general small-scale rotor test system) at KARI (Korea Aerospace Research Institute) were carried out to get lead-lag damping ratio of blades with flexures as hub flexure. MBA(moving block analysis) technique was used for the estimation of lead-lag damping ratio. First, NRSB-1F blades with HCTH hub system, then NRSB- 1F with NRSH hub system were tested. Second, NRSB-2F blades with NRSH hub system were tested. Tests were done on the ground and in the wind tunnel according to the test conditions of hover and forward flight, respectively. Non-rotating natural frequencies, non-rotating damping ratios and rotating natural frequencies were showed similar level for each cases. Estimated damping ratios of NRSB-1F, NRSB-2F with HCTH and NRSH were above 0.5%, and damping ratio increased by collective pitch angle increasement. Furthermore damping ratios of NRSB-2F were higher than damping ratios of NRSB-1F in high Pitch angle. It was confirmed that the blade design for noise reduction would give observable improvement in aeroelastic stability compared to paddle blade and NRSB-1F design.

The study of the fixed-pitch small-scale propeller in ground/ceiling effect is extended to forward flight conditions. Force-based experiment and smoke flow visualization is conducted at the University of Dayton Low Speed Wind Tunnel. The propeller is mounted at different incidence angles to simulate the transition to forward flight from hover and forward flight conditions. The effect of propeller incidence angle seems to be negligible at J=0. The power required at constant thrust shows an increasing trend after a threshold advance ratio for each h/D and then decreases steeply upon further increase in advance ratio. Smoke flow visualization revealed the presence of three distinct zones separated by the onset of the fountain vortex and the onset of the ground vortex at each ground distance. At advance ratios below the onset of the fountain vortex lies the optimal ground effect zone. An increase in advance ratio beyond this threshold will result in power increment due to the fountain vortex. Further increase in advance ratio initiates the onset of the ground vortex where the ground effect is significantly reduced. The onset of the fountain vortex and the ground vortex occurs at lower advance ratios for a propeller at an incident angle to the ground.

The design of new helicopter rotor blades is a challenging task. The individual blade sections undergo very different flow conditions during the various flight regimes of the helicopter. In forward flight, the advancing side operates in a transonic regime where potentially shock waves can occur, while on the retreating side little flow velocities at high angle of attacks are seen up to reverse flow. In hover, the oncoming tip vortex of the previous blade drastically influences the inflow on the rotor. This paper joins the classical blade shape design with numerical optimization techniques. Opposing to the direct solution process of directly modifying the blade shape, the rotor blade is conventionally considered as a set of airfoils, a planform and a twist distribution. First, rotor airfoils are found through numerical optimization and then placed on the reference rotor. The planform and twist of this rotor are then also numerical optimized. The obtained trade-off blade for hover and forward flight drastically improved the performance over the reference HART-II blade (11% in forward flight, 5% in hover flight). In order to arrive at industrial relevant blades, further work including more disciplines becomes necessary.

In this paper, the flight dynamic model of Vertical Take-Off and Landing (VTOL) aircraft which equips a propeller-wing-flap system that produces slipstream is established. The complete aircraft is modeled by combining experimental results, theoretical calculation, and empirical data. Firstly, the propeller deflected slipstream effects and the parameters of ducted fan are obtained by bench test. The DATCOM are used for generating the aerodynamic coefficients. Then, the 6-DoF nonlinear flight dynamic model in both hover and forward flight conditions are deducted by employing the assumption of linear superposition. The gyro effect and gravity are considered. Finally, specific trim conditions in hovering, transitional and level flight states are determined, and dynamic characteristics at trim are studied by both full model simulation and linear system analysis obtained through small-perturbation theory. It is found that the aircraft is critically stable in hovering condition. The damping ratio of short-period mode is relatively small, and the phugoid mode divergence in the transition state. The aircraft is less well damped in the level flight condition. The built model accurately reflects the dynamic characteristics of the VTOL aircraft and provides help for the design of the flight control system afterward.KeywordsSlipstreamVertical Take-off and Landing (VTOL) aircraftBench testDynamic modeling

This paper presents the implementation and validation of a state-space free-vortex wake model with a vortex lattice near-wake formulation, developed for rotorcraft applications. The model is expressed in state-variable form as a nonlinear time-periodic (NLTP) system in first-order structure, enabling linearization and simplification through time-invariant systems theory and model-order reduction techniques. It is applied to a UH-60-like rotor and evaluated in hover, forward flight, and vortex ring state (VRS) conditions. Two configurations are considered: one with tip vortex dynamics only, and another incorporating both tip vortex dynamics and a near-wake vortex lattice model. A parametric study is conducted to determine optimal parameters for solution accuracy. Validation in hover and forward flight is performed against high-fidelity computational fluid dynamics (CFD) results and available experimental data. Validation in VRS includes comparisons with experimental measurements and momentum theory. Overall, the state-space free-vortex wake model accurately predicts blade loads and induced velocities across all three validation cases, with the exception of a slight underprediction observed in fully developed VRS conditions. The inclusion of the near-wake model is essential for accurate blade load prediction in hover and contributes to improved predictions in forward flight.

Systematic design methodology for development and flight testing of a variable pitch quadrotor biplane VTOL UAV for payload delivery

The Aerodynamic Shape Optimizer for Rotor Blades (ADSORB) framework was used to perform four case studies optimizing the UH-60A blade airfoils along the blade span. The framework is built within Galaxy Simulation Builder. It uses CFD to generate data and construct a surrogate model for airfoil coefficient predictions. Dakota and a multiobjective genetic algorithm are then used to optimize the blade airfoils for four distinct segments of the blade. RCAS was then used to predict the rotor power in hover and forward flight (advance ratio 0.3), as well as the rotor L/DE in a high load forward flight condition (CT/σ ≈ 0.13). Each case study compared different combinations of objective functions to determine the most desirable traits. Ultimately, a design was found that increased the high load L/DE by 6.7%, while decreasing the hover and forward flight rotor power by 2.0% and 2.7%, respectively.

In the recent decades, the accuracy and robust of finite state inflow model to provide the velocity on the disk have been greatly improved. It is adequate to compute the thrust, pitch and roll moment real time for normal flight conditions. For lifting rotor flying close to a non-penetrable surface, the wake will interact with the inflow of the lifting rotor. Additionally, for co-axial rotors the wake of the upper rotor influences the lower rotor. Therefore, it is necessary to develop a model to compute the flow field in the wake since the Peters-He model cannot give the induced velocity below the rotor disk. In this paper, the improved finite state model will be introduced to compute the axial induced velocity everywhere in the flow field including wake and downstreamregions. The adjoint theorem will be introduced, and the computational efficiency of the convolution method and numerical algorithm for computing the adjoint varibales with time delay will be compared. With the parameters of a Harrington rotor, the results in the time domain for such a rotor in hover and forward flight conditions will also be illustrated.

A high-fidelity multidisciplinary analysis and gradient-based optimization tool for rotorcraft aero-acoustics is presented. Tightly coupled discipline models include physics-based computational fluid dynamics, rotorcraft comprehensive analysis, and noise prediction and propagation. A discretely consistent adjoint methodology accounts for sensitivities of unsteady flows and unstructured, dynamically deforming, overset grids. The sensitivities of structural responses to blade aerodynamic loads are computed using a complex-variable approach. Sensitivities of acoustic metrics are computed by chain-rule differentiation. Interfaces are developed for interactions between the discipline models for rotorcraft aeroacoustic analysis and the integrated sensitivity analysis. The multidisciplinary sensitivity analysis is verified through a complex-variable approach. To verify functionality of the multidisciplinary analysis and optimization tool, an optimization problem for a 40% Mach-scaled HART-II rotor-and-fuselage configuration is crafted with the objective of reducing thickness noise subject to aerodynamic and geometric constraints. The optimized configuration achieves a noticeable noise reduction, satisfies all required constraints, and produces thinner blades as expected. Computational cost of the optimization cycle is assessed in a high-performance computing environment and found to be acceptable for design of rotorcraft in general level-flight conditions.

In this paper, an optimal controller is investigate d based on the Linear Quadratic Control method. The controller simulation is execut ed based on a small-scale autonomous model helicopter, Yamaha R-50. Genetic algorithms are employed to generate weighting matrices by optimizing the performance index of the control system. As a result, this method produces optimal weighting matrices which further improve the Linear Quadratic Controller. A full state feedback approach is used in the control design process. A Kalman observer is integrated into the controller to predi ct full state variables, since only a limited number of state variables are measured. Finally, th e performance of the controller is evaluated in the time domain with and without disturbances when the model helicopter is in hovering flight and forward flight. I. Introduction utonomous helicopters offer special advantages for major operations in both civilian and military sect ors, particularly in cases that are considered either in accessible or dangerous for human beings, due to th eir vertical take-off and landing (VTOL) capabilities. The contr ol system of autonomous helicopters plays a key rol e in its ability to carry out the assigned task effectively. This has lead to a growing interest in the design and analysis of the control system of autonomous helicopters. In 1990s, classic control theory was widely used for autonom ous helicopter control systems, which is mainly a succe ssive loop closure approach. The control gains are selected separately using tools such as Root locus, Bode and Nyquist plots. The controller parameters are usual ly tuned based on the controller performance measured either in ti me domain or in frequency domain. This design procedure becomes increasingly difficult as more and more loo ps are added, specifically, when the dynamics invol ve multiple inputs, multiple outputs, or multiple feedback loop s with complex performance criteria. To control an autonomous helicopter modeled as a complex Multiple Input Multiple Output (MIMO) system, the performance criteria can not be easily met by the approach based on classis cont rol theory. Several researchers have developed a number of diff erent ways for the control system design of autonom ous helicopters in the literature. Shim et al 1 presented three different control methods using li near robust, fuzzy logic, and nonlinear tracking controls, for an autonomous helicopter in hover and low speed flight. Johnson a nd Kannan 2

Aerodynamic characteristics of rigid coaxial rotor by wind tunnel test and numerical calculation

In this work, optimization methodologies of the kinematics of flapping wings in forward and hover flights are considered. Particularly, local and global optimization algorithms are combined with the unsteady vortex lattice method (UVLM) to determine the most efficient kinematics. In the first problem, the kinematic optimization of a three-dimensional flapping wing in forward flight with active shape morphing is aimed at maximizing the propulsive efficiency under lift and thrust constraints. Results show that due to the quasi-concavity of the objective function associated with the flapping kinematics employed in forward flight, local gradient-based optimizers perform very well in terms of accuracy and computational cost. In the second problem, the objective is to identify the optimized kinematics of a twodimensional hovering wing that minimize the aerodynamic power under lift constraint. The results show that using a hybrid optimization algorithm (combination of local and global schemes) accelerates the convergence to optimal points and avoids being trapped at local minimum points. While the efficiency of using a local optimizer, a global optimizer, or a hybrid of the two depends on the nature of the problem and associated design space, it is determined that hybrid optimization is best suited for determining local minima.

This work conducts a validation study for the XH-59A helicopter using a rigid coaxial rotor system in order to establish the techniques of the conceptual design and performance analysis for the lift-offset compound rotorcraft. As a tool for conceptual design and performance analysis, NDARC (NASA Design and Analysis of Rotorcraft) is used for the present study. An assumed mission profile is considered for the conceptual design of the XH-59A. As a validation result of the design, the dimensions and weight of the XH-59A are appropriately designed when compared to the target values since the relative error is less than 0.5%. Then, performance analyses are conducted for the designed XH-59A model with and without auxiliary propulsion in hover and forward flight conditions. The present analyses show good validity since the prediction results compare well with both the flight test and previous analyses. Therefore, the techniques for the conceptual design and performance analysis of the lift-offset compound helicopter are overall considered to be appropriately established. In addition, this study investigates the influence of the lift-offset on the rotor effective lift-to-drag ratio of the XH-59A helicopter with auxiliary propulsion. As a result, the improvement of the rotor effective lift-to-drag ratio can be obtained by appropriately increasing the lift-offset in high-speed flight.

To analyze the dynamic stability of helicopters in real flight environments and understand the influence of key flight state parameters on dynamic stability, the six-degree-of-freedom equations of helicopter flight dynamics and the small disturbance principle are utilized. The flight test is used to obtain the variations in motion parameters after the helicopter is disturbed. A parameter identification method is then employed to solve the motion equations, yielding the oscillatory mode parameters that characterize the helicopter’s dynamic stability. Sensitivity analysis of the dynamic stability parameters under different flight states is conducted based on this method. The results indicate that a helicopter’s dynamic stability is closely related to its flight weight, center of gravity, speed, and flight mode. Under hover and forward flight conditions, the influence on longitudinal and lateral dynamic stability varies.

Rotor wake causes unsteady aerodynamics of rotor blade. So, accurate prediction of wake is very important and vortex method is good solution for this problem. Aerodynamic force of the rotor blade is calculated by potential panel method and the rotor wake is simulated by vortex particle method. The vortex particle method is easier to treat wake-body interaction and has better performance to expect the effect of ground and fuselage interaction. Rotor in hovering and forward flight condition is simulated through these methods. Thrust and surface pressure of rotor are compared with experiment data.