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The purpose of this paper is to provide an approach to predict rotor thrust and hub moments under in ground effect (IGE) in transient flight. Target of the research is developing a new integrated methodology that can be applied in the simulation of rotor flow field IGE.

Free-wake model and panel method are two methods used to predict ground influence on rotor flow field. However, these methods can result in unphysical phenomena, such as wake vortex moving below the ground during simulation and fluctuation taking place on vortices near ground, which is named noise problem. Thus, a new tactic called “constant volume rectification” is developed to rectify the unreal vortex location, and a third-order time-stepping algorithm called CB3D (Center difference and Backward difference 3rd-order scheme with numerical Dissipation) with strengthened stability is proposed to replace the existing second-order time-stepping algorithm CB2D (Center difference and Backward difference 2nd-order scheme with numerical Dissipation) to inhibit the development of discrete error.

The new free-wake model is effective and stable in predicting the characteristics of the rotor flow field in steady and transient flights under IGE. The newly developed CB3D scheme is more stable and more suitable for wake prediction of rotor under IGE than the CB2D scheme. At different advance ratios, the predicted flow regimes of recirculation and ground vortex agree well with the test images. In the accelerating condition, the predicted variations of rotor thrust and hub moments with advance ratios are consistent with the corresponding experimental results. It is found that the slow movement of wake geometry with advance ratio in the accelerating condition is the cause of the delay in the variation tendency of rotor forces compared to that in steady condition.

The proposed model can be used in rotor designing and helicopter flight dynamics simulation because of its favorable stability and relatively low computational cost.

This paper proposes several new methods to make the time-stepping wake model highly appropriate for rotor aerodynamics prediction under IGE. These methods provide new perspectives in solving the unstable and unphysical problems that often arise in vortex–ground interaction in rotor free-wake prediction.

The purpose of this paper is to present and validate an efficient time‐marching free‐vortex method for rotor wake analysis and study the rotor wake dynamics in transient and maneuvering flight conditions.

The rotor wake is represented by vortex filament elements. The equations governing the convection, strain and viscous diffusion of the vortex elements are derived from incompressible Navier‐Stokes equations based on the viscous splitting algorithm. The initial core size of the blade tip vortices is directly computed by a vortex sheet roll‐up model. Then, a second‐order time‐marching algorithm is developed for solving the governing equations. The algorithm is formulated in explicit form to improve computing efficiency. To avoid the numerical instability, a high order variable artificial dissipative term is directly introduced into the algorithm. Finally, the developed method is applied to examine rotor wake geometries in steady‐state and maneuvering flight conditions. Comparisons between predictions and experimental results are made for rotor wake geometries, induced inflow distributions and rotor transient responses, to help validate the new method.

The algorithm is found to be numerically stable and efficient. The predicted rotor responses have good agreement with experimental data. The transient behavior of the wake dominates the rotor responses following rapid control inputs in hover. The wake curvature effect induced by rotor pitching or rolling rate significantly changes the rotor off‐axis response.

This method should be further validated using experimental measurements of full‐scale helicopter rotors.

The paper presents a new time‐marching free‐vortex wake method, which is suitable for application in helicopter flight simulation.

A new method for predicting rotor wake in low speed and hovering flight is described to investigate the motion of the helical tip vortex. Beginning with the generalized wake model, a semi‐empirical correction for the vortex core effect on rotor wake is made and free wake calculation is carried out. As an example of its engineering application, the calculated downwash velocity field along the rocket launch line is presented and simply analysed. In terms of theory, the method developed here may provide of a referable basis for further study the formation mode of the tip vortex and vortex core interior structure.

The purpose of this paper is to evaluate the prediction capability of comprehensive structural dynamics (CSD) analysis codes for the higher harmonic control aeroacoustic rotor test (HART) II data.

A nonlinear flexible multibody dynamics analysis code DYMORE, as well as the comprehensive analytical model of rotorcraft aerodynamics and dynamics (CAMRAD) II, are used to perform the task. The predicted results on rotating free vibration analysis, airloads, blade elastic motions, and structural moments are correlated with the measured data for the baseline, minimum noise, and minimum vibration cases.

The DYMORE analysis results with a free wake model show a good performance in capturing blade vortex interaction peaks in the prediction of section normal forces but apparently with a phase shift problem. The high‐frequency behavior in the airloads signal does not affect much on the aeroelastic response and structural moments of the rotor.

The present approach uses two separate CSD codes to systematically validate the HART II data. The accuracy of each code on structural dynamic aspects of HART II rotor is assessed using a consistent set of inputs. The effects of blade tip deflections on the interaction of blades and their trailed vortices leading to a reduced noise emission are also investigated.

With the increasing sophistication of modern helicopter designs the problems arising from the interactional aerodynamic flow field around the helicopter has become more acute. Interactional aerodynamics are, by origin, of utmost complexity, because many of the interactions involve viscous processes, the flow usually is unsteady and the interactions are strongly interdependent.

Boundary element method (BEM) techniques for the prediction of cavitating or ventilated flows around hydrofoils and propeller are summarized. Classical, supercavitating, and ventilated blade section geometries are considered. Recent extensions which allow for the modeling of cavities on either or both sides of the blade surface are presented. Numerical validation studies and comparisons with experimental measurements are shown.

Computational fluid dynamics (CFD)/computational structural dynamics (CSD) coupling analysis is an important method in the research of helicopter aeroelasticity due to its high precision. However, this method still suffers from some problems, such as wake dissipation and large computational cost. In this study, a new coupling method and a new air load correction method that combine the free wake model with the CFD/CSD method are proposed to maintain computational efficiency whilst solving the wake dissipation problem of the prior coupling methods.

A new coupling method and a new air load correction method that combine the free wake model with the CFD/CSD method are proposed. With the introduction of the free wake model, the CFD solver can adopt two-order accuracy schemes and fewer aerodynamic grids, thus maintaining computational efficiency whilst solving the wake dissipation problem of the prior coupling methods.

Compared with the predictions of the prior methods and flight test data, those of the proposed method are more accurate and closer to the test data. The difference between the two methods in high-speed forward flight is minimal.

Because of the chosen research approach, the research results may lack generalisability. Therefore, researchers are encouraged to test the proposed method further.

In this paper, a CFD/CSD/free wake coupling method is proposed to improve the computational accuracy of the traditional CFD/CSD coupled method and ensure the computational efficiency.

High autogyro accident rates prompted experimental investigation of this type of aircraft's low‐speed pitch characteristics. Pitch control is typically derived from main rotor tip‐path‐plane adjustment. Thus, autogyro designers often omit horizontal tails and pitch control surfaces. The purpose of this paper is to enable autogyro low‐speed pitch control by intentionally placing elevators in the propeller wake.

Wind tunnel tests were conducted on a 1:10 scale teetering rotor autogyro model. The model included a horizontal tail with elevators placed in the propeller wake. Straight‐and‐level flight conditions were estimated via a scaling scheme based on the main rotor diameter. At minimum flight speed, the pitching moment induced by 30° elevator deflection was measured. This process was repeated for a range of elevator positions behind the centre of the pitching rotation.

When placed in an autogyro propeller wake, deflected elevators induce significant pitching moments. If the elevator is shadowed from free stream flow by the autogyro cowling, the pitching moment remains unchanged regardless of the distance between elevators and centre of pitch rotation. However, if the elevator is immersed in the freestream, the pitching moment increases via deflection of both propeller wake and freestream flow.

Kinematic similarity ensures ratios between propeller wake, wind speed, and main rotor flows are representative of full scale. Without flow visualization, main‐rotor‐diameter‐based scaling does not ensure kinematic similarity. Results are therefore qualitative.

Elevators mounted in autogyro propeller wake are worthy of inclusion on all autogyros for pitch control at low speed.

Improved low‐speed pitch control arising from elevators mounted in autogyro propeller wake could potentially reduce accidents.

The purpose of this paper is to examine the cross‐coupled responses of a coupled rotor‐fuselage flight dynamic simulation model, including a finite‐state inflow aerodynamics and a coupled flap‐lag and torsion flexible blade structure.

The methodology is laid out based on model development for an articulated main rotor, using the theories of aeroelastisity, finite element and finite‐state inflow formulation. The finite‐state inflow formulation is based on a 3D unsteady Euler‐based concepts presented in the time domain. The most advantages of the model are the capability of modeling dynamic wake effects, tip losses and skewed wake aerodynamics. This is, in fact, a special type of the inflow model relating inflow states, to circulatory blade loadings through a set of first‐order differential equations. A non‐iterative solution of the differential equations has practically altered the model into a simple and direct formulation appending properly to the rest of the helicopter mathematical model. A non‐linear distribution of the induced velocity over the rotor disc is finally obtained by the use of both Legendre polynomials and higher‐harmonic functions. Ultimately, validations of the theoretical results show that the on‐axis response, direct reaction to the pilot input, has a good accuracy both quantitatively and qualitatively against flight test data, and the off‐axis response, cross‐coupled or indirect reaction to the pilot input are improved by this approach of modeling.

Improvements in dynamic prediction of both trim control settings and dynamic cross‐coupled responses of helicopter to pilot inputs are observed.

Further work is required for investigation of the augmented finite state inflow model, including the wake rotation correction factors to describe helicopter maneuvering flight characteristics.

The results of this work support the future researches on design and development of advanced flight control system, incorporating a high bandwidth with low‐phase delay to control inputs and also high levels of dynamic stability within minimal controls cross coupling.

This paper provides detailed characteristics on the mathematical integration problems associated with the advanced helicopter flight dynamics research.

An analytical method to predict the effects of rotor downwash on engine jet was proposed to investigate the flow pattern of helicopter engine jet. First, free wake analysis was carried out to calculate the rotor wake geometry and the downwash around the engine jet space. Then the engine jet path, cross section shape and physical variants were calculated and analyzed. Finally, an example was given to show the typical results for an insight into the influence of engine jet upon the temperature distribution around a helicopter. The method developed here can be used to predict rotor wake, engine jet and temperature distribution of a helicopter.