Helicopter In Forward Flight Research Articles

Identification of Lateral/Directional Model for a UAV Helicopter in Forward Flight

A lateral/directional state-space model of the Raptor 50 helicopter is identified in forward flight condition. A frequency sweep test is properly designed and executed to collect a well-suited time-history database. The measured accelerometers are transferred from the sensor location to the gravity, then, a complete set of non-parametric input-to-output frequency responses that fully characterizes the coupled helicopter dynamics are extracted from the frequency sweep test data. With the help of small perturbation theory, a linearized dynamic model for the stability and control derivative is extracted from linearization of a nonlinear mathematical model. A nonlinear search based on secant method is conducted for the linear lateral/directional state-space model that matches the frequency-response data set. The identified model is verified in the time domain, which indicates that the model matches the flight data well.

적합직교분해를 이용한 로터 블레이드의 차수축소모델 구축 및 공력특성 분석

본 연구에서는 전진 비행하는 헬리콥터 로터 블레이드 표면의 압력장에 대한 공력 특성 분석 및 차수축소모델 구축을 위해 적합직교분해 (POD) 방법을 이용하였다. 에너지가 큰 특정 모드를 기반으로 전진 비행하는 비정상 로터 블레이드에 대한 공기역학적 특성을 분석하였으며, CFD 계산 결과의 검증을 위해 제자리비행에 대한 실험 결과와 비교하였다. 수렴속도를 향상시키기 위해 Multi-grid 기법을 사용하였으며, 회전하는 로터 블레이드 주위의 비정상 유동을 모사하기 위해 슬라이딩 격자를 이용하였다. 그 결과 240개의 Snapshot에 대해 에너지율 99% 이상을 포함하는 지배적인 POD 모드 7개가 선정되었으며, POD 모드와 전개 계수를 이용하여 차수축소모델을 성공적으로 구축하였다. The proper orthogonal decomposition (POD) method can identify principal modes that optimally capture the energy content from large multi-dimensional data set. In this study unsteady pressure fields on the rotor blade surface of a helicopter in forward flight are expressed by a reduced order model based on the POD method. Special modes containing high energy are analyzed to investigate the aerodynamic characteristics in more efficient way. The CFD simulation of flowfields around helicopter rotor blade in hovering motion is also conducted to validate its prediction with experimental result. In the process 7 modes containing energy ratio 99% from 240 snapshots information are identified and utilized to construct a reduced order model.

NUMERICAL SIMULATION OF UNSTEADY FLOW FIELD AROUND HELICOPTER IN FORWARD FLIGHT USING A PARALLEL DYNAMIC OVERSET UNSTRUCTURED GRIDS METHOD

A parallel Navier-Stokes solver based on dynamic overset unstructured grids method is presented to simulate the unsteady turbulent flow field around helicopter in forward flight. The grid method has the advantages of unstructured grid and Chimera grid and is suitable to deal with multiple bodies in relatively moving. Unsteady Navier-Stokes equations are solved on overset unstructured grids by an explicit dual time-stepping, finite volume method. Preconditioning method applied to inner iteration of the dual-time stepping is used to speed up the convergence of numerical simulation. The Spalart-Allmaras one-equation turbulence model is used to evaluate the turbulent viscosity. Parallel computation is based on the dynamic domain decomposition method in overset unstructured grids system at each physical time step. A generic helicopter Robin with a four-blade rotor in forward flight is considered to validate the method presented in this paper. Numerical simulation results show that the parallel dynamic overset unstructured grids method is very efficient for the simulation of helicopter flow field and the results are reliable.

Simulation and prediction of main rotor, tail rotor, and engine parameters from flight tests

In the framework of this research project, the main rotor torque, tail rotor torque, engine torque, and main rotor speed of a helicopter in forward flight are estimated by using a state-space model from flight tests data. The state-space model inputs are non-linear terms made of combinations of pilot controls and helicopter states. The model simulates the helicopter outputs while knowing the states and controls at all times. It was also implemented as a prediction tool, for possible use in an envelope protection flight control system in which the states, controls, and outputs are known at the present time, and predict the future helicopter states and controls following to pilot controls time history. The state-space model parameters are identified by using the subspace identification method, a relatively recent non-iterative algorithm, which constructs an observability matrix from input and output data and uses this matrix to obtain the state-space matrices. The obtained parameters are then optimized with the Levenberg—Marquardt output-error method. A comparison of the results with and without optimization is also conducted. The results show that the subspace method provides a good estimate of the outputs within the FAA tolerance bands and that these results can further be improved by use of the minimization algorithm. The generated model using the subspace method is found to be very good for prediction applications, which makes it a promising model for flight control simulator applications.

Time–accurate Modular CFD‐CSD Coupling for Aeroelastic Rotor Simulations

AbstractThis paper addresses the timewise accuracy of different coupling approaches applied to instationary aeroelastic simulations of rotors in forward flight. Two different approaches which are widely discussed in literature are examined: the tight or strong coupling, and the fully integrated or monolithic coupling.Strong coupling means an exchange of fluid loads and structural deformations at each time step which is effectuated in a fully modular manner. We will address aspects of conservativity and time‐accuracy, and will present results for a helicopter forward flight scenario. However, objections concerning the correct solution of the global non‐linear three field problem – structure, grid deformation, aerodynamics – remain.These objections are normally rejected by the monolithic approach. Here, a common set of partial differential equations is derived and solved in a single code. However, a truly monolithic system of equations is only needed for stability analysis, and it can be decomposed in a three field problem respecting appropriate boundary conditions for each domain. Thus, modularity can be maintained, conceiving a quasi‐monolithic procedure, when both domains are simultaneously solved in a common non‐linear iteration loop on a per time‐step basis. First results will be shown for a 2D flutter testcase.

Helicopter Vibration Reduction through Cyclic Variations in Rotor Blade Root Stiffness

This study analytically examines the influence of cyclic variations in flap-, lag-, and torsion-stiffness of the blade root region (at harmonics of the rotational speed), for reduction of vibratory hub loads of a helicopter in forward flight. The results indicate that considerable reduction in hub vibrations is possible using small-to-moderate amplitude cyclic variations in stiffness (no greater than 15% of the baseline stiffness value). Torsion stiffness variations produced moderate reductions in vertical hub force, lag stiffness variations produced substantial reductions in all hub forces and the hub yaw moment, and flap stiffness variations produced very significant reductions in all hub forces and the hub roll and pitch moments. The amplitude of the cyclic stiffness variations required generally increase with increasing forward speed, for comparable reductions in vibration. At any given forward speed, if the amplitudes of cyclic stiffness variation are too large, the hub vibrations can actually increase. The stiffness variations that reduce the vibratory hub loads could produce increases in certain vibratory blade root load harmonics. Vibration reductions are achieved due to a decrease in the inertial contribution to the hub loads, or a change in relative phase of various contributions.

Periodic control of helicopter rotors for attenuation of vibrations in forward flight

The paper is devoted to the design of a control system for the attenuation of vibrations in the main rotor of a helicopter in forward flight. The problem can be formulated as the one of rejecting a periodic disturbance of known frequency acting at the output of a linear time-periodic system. A solution inspired to optimal control design methods is proposed. A major preliminary issue is the estimation of the characteristics of the disturbance, for which two methods are worked out. The performance of the obtained control system is assessed by means of several simulation trials.

Flow Phenomena on Rotary Wing Systems and their Modeling

Flow fields around rotary wings, especially around helicopter rotors, are extremely complex due to strong three-dimensional and viscous effects. Moreover, in yawed flow conditions of propellers and wind turbines as well as in helicopter forward flight, they are highly unsteady. Fluid-structure interactions require the treatment of the rotor blades as aeroelastic systems. Furthermore, helicopter and propeller flows may become transonic in some regions. Flow interaction effects and problems regarding noise generation must also be taken into account. Accordingly, the solution of the basic equations governing these problems is very difficult and usually only feasible with the help of numerical methods, since systems of nonlinear partial differential equations must be analyzed, A numerical investigation requires appropriate turbulence modeling for most technical applications. The paper presents several theoretical and numerical models for the simulation of the aforementioned flow phenomena.

Periodic LQG/H∞ Control of Helicopter Rotors for Attenuation of Vibrations in Forward Flight

This paper is devoted to the design of a control system for the attenuation of vibrations in the main rotor of a helicopter in forward flight. The problem can be formulated as the one of rejecting a periodic disturbance of known frequency acting at the output of a linear time-periodic system. A solution inspired to H∞ control design methods is proposed and compared with a previously developed LQG solution. The performance of the obtained control system is assessed by means of several simulation trials

Application of the Wide‐Field Shadowgraph Technique to Helicopters in Forward Flight

A test was conducted to examine the feasibility of using the wide-field shadowgraph method for visualizing the wake of a small-scale helicopter rotor in forward flight. A wide range of test conditions were examined, including thrust and forward speed variations. Shadowgraphs from the side and top of the rotor were obtained. The visibility of the tip vortices using the shadowgraph method was documented for advance ratios up to 0.175. It was demonstrated that shadowgraphs of the rotor wake in forward flight can be used to obtain qualitative and quantitative information about wake geometry, wake/body interactions, and blade/vortex interactions. Video cameras were used for the first time to record the shadowgraphs. These provided several advantages compared to still cameras and greatly enhanced the capabilities of the shadowgraph method.

Comment on 'Induced drag based on leading-edge suction for a helicopter in forward flight'

Comment on 'Induced drag based on leading-edge suction for a helicopter in forward flight'

Helicopter vibration reduction using structural optimization with aeroelastic/multidisciplinary constraints - A survey

This paper presents a survey of the state-of-the-art in the field of structural optimization when applied to vibration reduction of helicopters in forward flight with aeroelastic and multidisciplinary constraints. It emphasizes the application of the modern approach where the optimization is formulated as a mathematical programming problem, the objective function consists of the vibration levels at the hub, and behavior constraints are imposed on the blade frequencies and aeroelastic stability margins, as well as on a number of additional ingredients that can have a significant effect on the overall performance and flight mechanics of the helicopter. It is shown that the integrated multidisciplinary optimization of rotorcraft offers the potential for substantial improvements, which can be achieved by careful preliminary design and analysis without requiring additional hardware such as rotor vibration absorbers of isolation systems.

Induced drag based on leading edge suction for a helicopter in forward flight

Estimating induced drag for a helicopter in forward flight is a three-dimensional, unsteady aerodynamic problem complicated by fluid compressibility and wake geometry. Based on an acceleration potential approach, the chordwise velocity and the derivative d$/dt of the velocity potential at the leading edge of a thin rotor blade in subsonic flow were re-examined in this paper to assess unsteady and compressibility effects on the induced drag using a leading-edge suction model. The chordwise velocity was shown to have a singular and a continuous component. The d$/dt was shown to be continuous and hence does not contribute to induced drag. The induced drag calculated from the leading-edge suction model and the more traditional model to be referred to as the induced angle model were compared to quantify the differences in the two approaches. The results show that variations can be significant. While these variations cannot substantiate the validity of either approach, it is clear that the leading edge suction model is simpler to apply with fewer assumptions. b A A d

Comment on "Calculation of Rotor Impedance for Articulated-Rotor Helicopters in Forward Flight"

References Kuhlman, J.M. and Warcup, R.W., Investigation of Jet-Induced Loads on a Flat Plate in Hover Out-of-Ground Effect, NASA CR-159004, Feb. 1979. Margason, R.J., Review of Propulsion-Induced on Aerodynamics of Jet/STOL NASATN D-5617, Feb. 1970. Gentry, G.L. and Margason, R.J., Induced Lift Losses on VTOL Configurations Hovering In and Out of Ground Effect, NASA TN D-3166, Feb. 1966. Kuhlman, J.M. and Warcup R.W., Effects of Jet Decay Rate on Jet-Induced Loads on a Flat Plate, Journal of Aircraft, Vol. 15, May 1978. pp. 293-297. Kuhlman, J.M., Ousterhout, D.S., and Warcup, R.W., Investigation of Effect of Jet Decay Rate on Jet-Induced Pressures on a Flat Plate, NASA CR-2979, April 1978. Kuhlman, J.M., Ousterhout, D.S., and Warcup, R.W., Investigation of of Jet Decay Rate on Jet-Induced Pressures on a Flat Plate: Tabulated Data, NASA CR-158990, Nov. 1978. Keffer, J.F. and Baines, W.D., The Round Turbulent Jet in a Cross-Wind, Journal of Fluid Mechanics, Vol. 15, Pt. 4, April 1963, pp. 481-497. Kamotani, Y. and Greber, I., Experiments on a Turbulent Jet in a Cross Flow, NASA CR-72893, June 1971. Fearn, R.L. and Weston, R.P., The Induced Pressure Distribution of a Jet in a Crossflow, NASA TN D-7916, July 1975. Fearn, R.L., and Weston, R.P., Vorticity Associated with a Jet in a Cross Flow, AIAA Journal, Vol. 12, Dec. 1974, pp. 1666-1671. 11 Ousterhout, D.S., An Experimental Investigation of a Cold Jet Emitting from a Body of Revolution into a Subsonic Free Stream, NASA CR-2089, Aug. 1972. Ziegler, H. and Wooler, P.T., Analysis of Stratified and Closely Spaced Jets Exhausting into a Crossflow, NASA CR132297, Nov. 1973. Moller, P.S. and Elliott, R.L., Base Drag of a Thick Annular Jet, Journal of Aircraft, Vol. 9, July 1972, pp. 451-455.

Calculation of Rotor Impedance for Articulated-Rotor Helicopters in Forward Flight

A procedure is presented to calculate the loads transferred from an articulated flexible rotor to the fuselage when the hub is forced to oscillate sinusoidally. Blade motions are determined from a set of linear algebraic equations derived from equations of motion with periodic coefficients. The aerodynamic loads are based on two-dimensional quasisteady strip theory and the effect of preceding and returning wakes as well as the reversed flow are neglected. Sample calculations indicate that: 1) the major components of impedances with hub-forcing frequency predominate over those with interharmonic coupling frequencies; 2) the former impedances do not depend on the blade azimuth angle relative to the hub excitation phase; and 3) the former impedances are similar to those obtained in hovering flight.

Sensitivity Reduction in Aircraft Control Systems

A method for reducing trajectory sensitivity to parameter perturbations for linear feedback systems is described. Application of the method to the design of aircraft control systems, with special reference to a helicopter forward flight control system, is illustrated. The response of the system based on this method is found to be better than that obtained with fixed-feedback gain controllers, although the response may not be as good as that of an adaptive control scheme. The main advantage of the method is the simplicity of its implementation relative to that of the adaptive control scheme which requires an on-board computer.

In‐Flight Far‐Field Measurement of Helicopter Impulsive Noise

A new and highly successful method of collecting far‐field acoustic data radiated by helicopters in forward flight has been developed, utilizing a quiet aircraft flying in formation ahead of the subject helicopter. The lead aircraft, flown as an acoustic probe, was equipped with tape‐recording equipment and an external microphone. Spatial orientation of the helicopter with respect to the monitoring aircraft was achieved through visual flight reference. Far‐field acoustic data defining the impulsive noise radiation characteristics of the UH‐1H helicopter during high‐speed flight and partial ‐power descents have been gathered with this technique. Three distinct types of impulsive waveforms have been identified and correlated with helicopter steady operating conditions.

Numerical Calculation of Unsteady Transonic Potential Flow over Helicopter Rotor Blades

The small disturbance potential equation appropriate to a helicopter in forward flight is derived. This equation then is solved for the flow over a nonlifting transonic rotor blade, using a completely implicit scheme that is an extension of the Murman-Cole mixed difference technique. The flow in the tip region is most unsteady in the decelerating flow region, after the blade passes the \J/ - 90°azimuthal station. The unsteadiness appears to be caused by expansion and compression waves that move slowly upstream of the blade as the relative incident flow decelerates. The influence of aspect ratio, advance ratio, and Mach number on this process is discussed.

Some Approximations to the Flapping Stability of Helicopter Rotors

The flapping equation for a helicopter in forward flight are reported which have coefficients that are periodic in time, and this effect complicates the calculation of stability. A constant coefficient approximation which will allow the use of all the well known methods for analyzing constant coefficient equations are presented. The flapping equation is first transformed into the nonrotating coordinate frame, where some of the periodic coefficients are transformed into constant terms. The constant coefficient approximation is then made by using time averaged coefficients in the nonrotating frame. Stability calculations based on the approximation are compared to results from a theory which correctly includes all of the periodicity. The comparison indicates that the approximation is reasonably accurate at advance ratios up to 0.5.

Impulsive Sound from Helicopters Due to Compressibility Effects

The tip regions of the lifting rotor of a helicopter in forward flight may reach nearly sonic speeds with respect to the surrounding air. When this occurs, compressibility effects will cause an impulsive sound wave to be generated. Both lift and thickness distributions are responsible for this component of the sound. Calculations of the sound radiated from blade tips of various shapes and lift distributions have been made. Some results of the calculations and correlations with experimental data are presented.