Neural Network Research on Validating Experimental Tilt-Rotor Performance

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

This study examines hexacopter performance in hover and forward flight with single rotor failure. A classical configuration (adjacent rotors spinning in opposite directions) and an alternate configuration are considered. The simulation model uses calculates rotor forces and moments by using blade element theory coupled with a finite-state dynamic inflow model. For the alternate hexacopter, if one of four rotors (out of six) fails, trim is possible in hover and forward flight, and the aircraft remains fully controllable. Recovery from the failure of one of the other two rotors is impossible. For the classical hexacopter, if any of the forward-facing rotors fail, the aircraft can be trimmed in forward flight and is fully controllable. If an aft rotor fails, trim is not possible but the aircraft could be turned around to orient the failed rotor forward. Although the classical hexacopter can be trimmed in hover by turning off the rotor diametrically opposite to the failed rotor, the aircraft is not independently controllable about all axes. In the event of a rotor failure, the classical hexacopter could cruise and land, but it cannot maintain a sustained hover. Power penalties of up to 23% are observed in the event of failure due to increased drag on the operational rotors.

This paper addresses the acoustic and aerodynamic characteristics of small rotor configurations, including the influence of the rotor–rotor interactions. For this purpose, a Rotor/Rotor/Pylon configuration is chosen for both the test and numerical simulations. The wind tunnel experiments on various rotor configuration were performed in DLR’s Acoustic Wind Tunnel Braunschweig (AWB). The experiments involve isolated rotors, and rotors in tandem and coaxial configuration in hover and forward flight. For numerical simulations, an unsteady free wake 3D panel method (UPM) is used to account for aerodynamic non-linear effects associated with the mutual interference among the Rotor/Rotor/Pylon configurations. The effect of the pylon is simulated using potential theory in form of a panelized body. Finally, the sound propagation into the far field is calculated with DLR’s FW–H code APSIM, using UPM blade surface pressure as input. The validation effort is supported by CFD TAU steady simulations on selected hover test cases. The experiments and numerical results indicate that the noise at the blade passing frequency (BPF) and its higher harmonics is the dominant source of the noise for the present rotor selection. The extra subharmonics between two BPFs appearing in the results are caused by the small geometric discrepancy between the blades as well as the motor noise. Broadband noise is also observed in the experiment, but its contribution to the overall sound pressure is very small and can be neglected. The simulation of the acoustic scattering from the rotor support system for the isolated rotor cases indicated an influence about 1–3 dB on the overall sound pressure of the polar microphones. In both the coaxial and the tandem configuration, the acoustic interferences are particularly well visible in the numerical simulations and cause a more complex noise directivity. There is almost no change in time-averaged inflow by applying phase angles. In the coaxial condition, in hover, the phase delay between rotors does not change the maximum noise level. In forward flight, the phase delay can influence the maximum level of the noise radiation. In both coaxial and tandem configuration, the position of the downstream rotor is key for the noise radiation, and therefore, avoiding the interaction with upstream wake can reduce the noise radiation.

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, 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

This paper describes design optimization of a rotor blade for variable pitch quadrotor unmanned air vehicle (UAV) to ensure optimal performance in hover and forward flight. In order to optimize the blade profile to maximize hover power loading, a modified Blade Element Theory based analysis is developed and validated using experimental measurements for sets of symmetric-untwisted rectangular blade and cambered-twisted variable chord blade. The blade twist and chord distribution is parametrized using fifth order polynomial functions and the BEMT analysis is coupled to Matlab optmization toolbox to maximize the power loading for an operational thrust of approximately 3N. It is observed that use of rotor blade with non-linear twist and non-linear chord variation results in significant improvement in hover performance for the variable pitch quadrotor UAV. The optimized blade profile and chord distribution with GOE-744 airfoil gives approximately 4% higher power loading than the COTS cambered and twisted blade and 17% higher power loading than the untwisted rectangular blade with symmetric airfoil. The forward flight performance of the optimized blades is compared with the baseline blades using a Blade Element Theory and Drees inflow model based trim analysis. It is observed that the optimized blade profile doesn't incur any significant penalty during forward flight due to high twist and its performance is similar to that of untwisted blade.

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.

In this paper, the potential flow equations are converted to ordinary differential equations through a Galerkin approach in which velocity and pressure potential functions are expanded in terms of closedform solutions to Laplace's equation. Because the method gives differential equations for the flow in terms of a relatively few generalized coordinates (that represent modes of the flow field) the resultant equations can be used effectively in preliminary design, real-time simulations, and dynamic eigenvalue analysis for aeroelasticity. This new theory is more general than the Peters-He dynamic wake model since it has a more rigorous derivation and includes inflow modes previously neglected in the Peters-He model. Results are presented in the frequency domain for simple harmonic motion. The axial velocity components are obtained by this new methodology on and off the disk, for axial and skewed flows, and for different pressure distributions and are compared with the Peters-He model and with an exact solution obtained by a convolution integral. INTRODUCTION Dynamic inflow models go back over 50 years in terms of their development and application. Sissingh first formulated a mathematical model that described how gradients in inflow could result from cyclic pitch changes or from fuselage pitch and roll rates. He further showed how these inflow gradients could create cyclic changes in blade angle of attack that significantly changed the actual rotor response. Curtiss and Shupe refined this idea into a lift deficiency function for rotor cyclic variations. Ormiston and Peters showed how the formulation could be generalized into an inflow model * AIAA Fellow Copyright © 2002 The American Institute of Aeronautics and Astronautics Inc. All rights reserved. for both uniform and side-to-side variations in inflow. Peters generalized these concepts to the unsteady case by including apparent mass terms to account for the time delay involved in the development of the rotor flow field. Comparisons with static data in hover and forward flight showed that the previous methodologies were inadequate in forward flight. Pitt and Peters used principles of potential flow theory to show that these concepts could be developed from first principles without ad hoc assumptions on time delay or on the effects of forward flight. They obtained an unsteady flow model from potential functions that gave spectacular correlation with wind tunnel response data throughout the frequency and advance ratio range. However, despite the unquestioned success of this inflow model, it still remained a theory with only three inflow degrees of freedom: 1) uniform, 2) fore-to-aft, and 3) side-to-side. Later, Peters and He showed how the Pitt-Peters ideas could be truly generalized to a theory with an arbitrary number of inflow harmonics and an arbitrary number of radial shape functions per harmonic. They showed how the general theory could reduce in special cases to the old Pitt-Peters model as well as to Loewy theory and Prandtl tip-loss theory. This new theory was used to correlate unsteady wind tunnel data from the NASA Langley wind tunnel for various planforms, thrust coefficients, and advance ratios. The match in both steady and unsteady distributions was excellent The Peters-He model represents a mature dynamic inflow model that is now used in many production codes, including 2GCHAS (Government code), FLIGHTLAB (Advanced Rotorcraft Technology), COPTER (Bell helicopter), ONERA-DFVLR (European Community) and many university applications. More recent extensions of the theorv include applications to the off-axis coupling problem, and to ground effect in flight simulation. 1 American Institute of Aeronautics and Astronautics (c)2002 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. The inflow models described in the above paragraphs are formulated such that the states of the model represent induced flow distributions on the rotor disk (expanded radially as polynomials and azimuthally as Fourier coefficients). The first-order differential equations for these states depend on the free-stream impingement angle, flight speed, and thrust level. However, they are all in closed form in terms of a mass matrix and a damping matrix, such that they are easily assembled either into a comprehensive code, an eigenvalue analysis, or a flight simulator. This makes the model extremely efficient. However, in many of the applications, such as the ground effect work, it is necessary to find not just the normal flow at the disk, but all three components of flow off of the disk as well. Peters and Morillo presented a consistent methodology for computing all three components of the flow in axial flow, both on and off of the rotor disc within the context of a finite-state model. The model is formulated in a manner fairly similar to previous work (in that the potential functions in ellipsoidal coordinates are used). However, in contrast with the previous work, the states represent velocity potentials rather than individual flow components. In addition, all potential functions are considered (not just the ones that have a pressure discontinuity across the disk); and the derivation of the equation coefficients is done in a more consistent and rigorous manner than in the earlier derivations. The result is a simpler derivation and a more complete inflow theory for the velocity both off and on the disk. Previous dynamic inflow models in axial flow are shown to be special cases of the new model when offdisk coupling is neglected. This new methodology based on a Galerkin approach provides the exact solution on as well as off the disk for the axial velocity component in axial flow. In this paper, the concepts introduced by Peters and Morillo are extended to analyze skewed flow cases. FLUID DYNAMICS EQUATIONS The three-dimensional potential flow equations (momentum and continuity equations) for the pressure and velocity fields P and v , with a free-stream velocity V^are f;-Jp-^ (1) V • v = 0 (2) These equations have been non-dimensionalized by defining P as pressure divided by pVj, v as induced velocity divided by V^ and time as a reduced time T, (i.e., time multiplied by VJR.) The variable £ is the non-dimensional coordinate along the free-stream line, positive upstream. All lengths are divided by the rotor radius R. Figure 1 shows the coordinate system. Fig.l Coordinate system. From continuity, Eq. (2), it is observed that v can be expressed by a velocity potential that satisfies Laplace's equation. It can also be shown that P satisfies Laplace's equation. Therefore, P can be expressed as a summation of pressure potentials, 0; and v can be expressed as a summation of the gradient of velocity potentials, V. NEW FORMULATION Pressure Potentials* and Velocity Potentials To transform Eqs. (1) and (2) by a Galerkin method, it is required to expand the pressure potential, 4>, and the velocity potentials, *P, in terms of complete set of functions that satisfy Laplace's equation. In addition, they have to fulfill the boundary conditions for pressure in the case of <&, and for velocity in the case of V. The boundary conditions for pressure are given by a discontinuity across the rotor disk. The use of an ellipsoidal coordinate system, has the advantage that any odd function in v, will allow a representation of a discontinuity across the rotor disk. An additional advantage of using an ellipsoidal coordinate system is that an analytical solution of Laplace's equation is known and can be expressed as,

The paper describes a methodology to extract linearized state-space dynamic inflow model from a generic and high fidelity aerodynamic analysis for single and coaxial helicopter rotor systems. Both hover and forward flight are studied including maneuvering and the results are demonstrated by a free-wake analysis. The proposed methodology which is founded on a Single Frequency Analysis is computationally efficient and provide the entire frequency-domain behavior of the wake inflow states. Besides being able to incorporate any advanced computational tool, the present methodology takes advantage of exact wake states and thus adequately encompasses high frequency responses. The paper also extends the methodology to the case where states of higher harmonics are required and to the case of heaving, rolling and pitching rotor disc. For a single rotor system the results resemble the well known classical dynamic inflow model. For coaxial rotor system in hover, the high mutual couplings between the rotors are clearly demonstrated and affect the resulting system matrices. An important feature of the present methodology lays in the fact that the entire formulation is non-dimensional and the in-flow analysis is completely separated from the loads analysis, which enables a consistent generalization of the results.

In the present study, the parametric study was performed to analyze the effect of the tab on the aerodynamic performance and characteristics of rotor blades. Also, the tab shape was designed to improve the aerodynamic performance of rotor blades. A computational fluid dynamics solver based on three-dimensional Reynolds averaged Navier–Stokes equation using an unstructured mesh was used for the parametric study and the tab design. For airfoils, the effect of length and angle of a tab was studied on the aerodynamic characteristics of airfoils. In addition, including those parameters, the effect of a span of a tab was studied for rotor blades in hovering flight. The results of the parametric study were analyzed in terms of change of the aerodynamic performance and characteristics to understand the effect of a tab. Considering the analysis, the design of tab shape was conducted to improve the aerodynamic performance of rotor blades. The simply attached tab to trailing edge of the rotor blades increases the thrust of the rotor blades without significant changing of aerodynamic characteristics of the rotor blades in hovering and forward flight.

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.

Many helicopters use elastomeric lag dampers to prevent ground resonance and aeromechanical instability in hover and forward flight. Recent experimental results have shown that when the damper motion occurs at two superimposed frequencies, which is characteristic of operation in forward flight, the damper properties are not well-predicted by a superposition of the damper properties at each of the motion frequencies. This paper presents experimental data obtained with an elastomeric damper while it was undergoing single- and dual-frequency motion. The effect of the nonlinear dual-frequency damper characteristics on predicted rotor aeromechanical stability in forward-flight operation in a wind tunnel was evaluated using the comprehensive rotorcraft analysis program called CAMRAD.

Compound rotorcraft can be a viable option to achieve moderately high speeds of 240 knots. However, there is no systematic wind tunnel test data in the public domain to understand the complex interactional aeromechanics involved. This work is part of a test campaign to conduct multiple high-speed wind tunnel tests of various compound configurations. The compound configuration consists of a single main rotor, a single wing on the retreating side for lift augmentation, and a single rear propeller for propulsive augmentation. In this paper, hover test results are presented. Hover tests were conducted for various configurations and at different tip Mach numbers with the maximum value of 0.5. Rotor and wing performances are measured separately using individual balances. Rotor blades are equipped with strain gauges for measuring flap, lag, and torsional moments at radial stations. The wing location is below the main rotor, and the measured wing download is only 2.8% of the rotor thrust. The wing provides a partial ground effect, making the rotor more efficient and almost compensates for the download with less than 1% reduction in net figure of merit. Blade structural loads correlate satisfactorily between the analysis and measurement except for lag bending moment. This discrepancy is a large 2/rev lag bending moment measured in the test, which needs further investigation. Propeller hover test was conducted up to 2500 revolutions per minute (tip Mach number of 0.25) at a single collective. Prediction correlates well and captures the low-Reynolds-number effect for lower revolutions per minute.

Design and experimental study of a new flapping wing rotor micro aerial vehicle

Unmanned aerial vehicles, particularly drones, have been increasingly deployed for different tasks in the community. They have become an important part of the economic and social benefits that society is exploiting from modern technology development. However, efforts are still required to further develop technologies which can mitigate the negative impacts. Among them, drone noise is considered a major health concern for the community. The present study undertakes an experimental investigation of the effectiveness of blade modifications on drone noise in an aeroacoustic wind tunnel facility. A quadcopter drone is programmed to operate in both hover and forward flights. Three modified blade configurations, including trailing-edge serrations combined serration–finlets, and an unmodified (baseline) blade, are manufactured. The far-field noise signals are recorded by two polar microphone arrays to quantify both the magnitude and directivity. The results show that all modified blades are able to reduce the drone noise at mid-to-high frequencies in both hover and forward flights, and this leads to a noticeable reduction in the overall sound pressure level. More importantly, the combined serration–finlet configuration outperforms all the other blades. Psychoacoustic analysis is also performed using the far-field acoustic time series. Interestingly, only the serration–finlet combination demonstrates a consistent reduction in the psychoacoustic annoyance levels, suggesting that it is important to use metrics from both acoustic and psychoacoustic analysis when developing noise mitigation strategies in the socio-economic context.