Collective Pitch Angle Research Articles

Understanding Ducted Rotor Antitorque and Directional Control CharacteristicsPart II: Unsteady Simulations

Understanding the dynamic relationship between the antitorque thrust moment and the applied collective pitch angle is crucial, especially for directional control sensitivity analyses. Although there are many studies in the literature on the steady-state behavior of the FANTAIL TM , little is known about the transient response and thrust buildup, which is the primary focus of this paper. Computational fluid dynamics is used for the solutions here because it provides a more complete flowfield prediction, especially in low-power, near edgewise conditions. The flowfield is assumed to be inviscid, and the Euler equations are solved with a blade-element model for the FANTAIL. The main rotor is excluded in this study. Solutions are obtained by modifying the computer code PUMA2 (Parallel Unstructured Maritime Aerodynamics) and using an unstructured grid of 2.8 million cells. The code was run on Beowulf PC clusters. Dynamic fan thrust and moment response to applied collective pitch in hover and forward flight are presented and discussed. Nomenclature Cp = pressure coefficient L = total length of the helicopter N =y awing moment Tfan =f an thrust t = time V = freestream velocity v � =a verage induced velocity y = helicopter spanwise station θ.75 = collective pitch angle

비정렬격자계를 이용한 헬리콥터 덮개 꼬리 로터의 제자리 비행 유동 해석

본 연구에서는 비정렬격자계를 사용하여 덮개 꼬리 로터의 제자리 비행에 대한 압축성, 비정성 유동을 해석하였다. 유동계산을 위한 수치적 기법으로는 셀 중심에 기초한 유한체적법과 내재적인 시간적분법을 사용하였다. 계산은 로터의 한 블레이드에 대해 수행되었으며, 블레이드와 블레이드 사이에는 주기적 경계조건을 설정하였다. 덮개가 없는 로터 형상에 대한 성능은 실험 결과와 잘 일치함을 보였다. 덮개를 포함하는 로터 형상에 대한 계산은 비교된 실험 형상의 불확실성을 고려하여 추력이 일치하는 피치각을 가지는 경우에 대해 수행하였으며, 자세한 유동은 실험결과와 잘 일치함을 확인할 수 있었다. 그 결과로부터 본 방법이 블레이드 끝단간극을 포함하는 복잡한 3차원 덮개 꼬리 로터 형상 해석에 매우 유용하게 사용될 수 있음을 알 수 있었다. Detailed flow of a shrouded tail rotor in hover is studied by using a compressible inviscid flow solver on unstructured meshes. The numerical method is based on a cell-centered finite-volume discretization and an implicit Gauss-Seidel time integration. Numerical simulation is made for a single blade attached to the center body and guide by the duct by imposing a periodic boundary condition between adjacent rotor blades. The results show that the performance of an isolated rotor without shroud compares well with experiment. In case of a shrouded rotor, correction of the collective pitch angle is made such that the overall performance matches with experiment to account for the uncertainties of the experimental model configuration. Details of the flow field compare well with the experiment confirming the validity of the present method.

Predicting Aerodynamic Rotor-Fuselage Interactions by Using Unstructured Meshes

A numerical method has been developed for the analysis of interactional aerodynamics between helicopter rotor and fuselage in forward flight. A 3-D steady compressible Euler solver is used to compute time-averaged interactional effects between the rotor and the fuselage. The rotor is modeled as an actuator disk with zero thickness carrying pressure jump across it. Collective and cyclic pitch angles are calculated to satisfy the rotor trim condition. Unstructured meshes are used to model complex rotor-fuselage configurations. Calculations are performed for two generic helicopter fuselage geometries. The results are compared with available experimental data for validation.

Aeroelastic Stability Investigation of a Composite Hingeless Rotor in Hover

The aeroelastic stability of a composite hingeless rotor with elastic couplings is investigated in hover. A two-bladed, Froude-scaled, soft in-plane hingeless rotor model with various isotropic and graphite/ epoxy composite flexures was designed and then tested on a hover stand. The rotor with an isotropic flexure was tested first to examine the experimental procedures. The effects of elastic flapwise bending-torsion and chordwise bending-torsion structural coupling on the aeroelastic stability were then investigated using elastically coupled composite flexures. Lag-mode stability test results are compared with the theoretical predictions obtained using the comprehensive aeroelastic rotor analysis, UMARC (University of Maryland Advanced Rotorcraft Code). The introduction of negative chordwise bending-torsion coupling has a stabilizing effect on the lag damping for positive collective pitch angles. Predictions agree with measured data.

Nonlinear autopilot control design for a 2-DOF helicopter model

The paper illustrates a robust control scheme for application to helicopters in vertical flight mode (both take-off and landing) to guarantee altitude stabilisation. A nonlinear helicopter model is used to derive the proposed control in which Lyapunov's direct method is used to establish the overall system stability. A recursive design technique is applied to design a nonlinear robust controller using the highly coupled system structure. It will be shown that the proposed robust controller provides semiglobal stabilisation of uniform ultimate boundedness for achieving the desired altitude. That is, the control design is valid for all values of the helicopter's collective pitch angle away from zero. The vertical flight application demonstrates a unique robust control mechanisation for helicopter and V/STOL autopilot augmentation systems.

Results from laser sheet visualization of a periodic rotor wake

A laser sheet has been applied to the visualization of a periodic rotor wake produced by a model rotorcraft in forward flight. A stiff, two-bladed teetering rotor is mounted independently of an idealized airframe consisting of an instrumented circular cylinder with a hemispherical nose. The model was installed in a wind tunnel, and tests were carried out at three different advance ratios (forward flight speeds) at one value of rotor collective pitch angle. The laser sheet technique enables detailed flow visualization to be carried out. Quantitative data regarding the location of the rotor tip vortex can be obtained. Results from the flow visualization confirm that the rotor wake is a complex flowfield not amenable to simple analysis. However, the details of the tip vortex geometry are both periodic and repeatable. The motion of the rotor wake in close proximity to the airframe is examined. The behavior of the tip vortex is strongly affected by the presence of the airframe. In particular, the interaction of the rotor tip vortex with the airframe results in the destruction of the vortex core; apparently the vorticity is diffused, and the vortex loses its definition. No vortex filaments appear below the airframe as identifiable flow features. Examples of airframe surface pressure data are presented and are seen to correlate well with the results from the flow visualization.

Euler calculations for flowfield of a helicopter rotor in hover

Aerodynamic loads on a multibladed helicopter rotor in hovering flight are calculated by solving the threedimensional Euler equations in a rotating coordinate system on body-conforming curvilinear grids around the blades. Euler equations are recast in the absolute flow variables so that the absolute flow in the far field is uniform but the relative flow is nonuniform. Equations are solved for the absolute flow variables employing Jameson's finite-volume explicit Runge-Kutta time-stepping scheme. Rotor-wake effects are modeled in the form of a correction applied to the geometric angle of attack along the blades. This correction is obtained by computing the local induced downwash with a free-wake analysis program. The calculations are performed on a CRAY X/MP-48 for a model helicopter rotor in hover at various collective pitch angles. The results compared with experimental data.

Application of a Mixed Variational Approach to Aeroelastic Stability Analysis of a Nonuniform Blade*

ABSTRACT A mixed variational approach is used to investigate the coupled flap-lag-torsional aeroelastic stability of a nonuniform helicopter blade. This approach is particularly suited to treating rotating blades with sharp variations or discontinuous changes in their properties or cross-sectional dimensions. Numerical results by this approach are presented in hovering conditions, for both uniform and nonuniform blades with low torsional stiffness, and compared with the results by the usual Rayleigh-Ritz method. It is shown that the present approach has good convergence properties for both static equilibrium position and critical collective pitch angles.

Performance degradation of a model helicopter rotor with a generic ice shape

An experimental program using a commercially available remotely controlled model helicopter in the Texas A&M University (TAMU) subsonic wind tunnel has been conducted to investigate the performance degradation resulting from the simulated formation of ice on the leading edge of the main rotor blades in both hover and forward flight. The rotor blades utilized a NACA 0012 airfoil with a 2.5-in. constant chord. A generic ice shape derived from a predetermined natural ice condition was applied to the 53.375-in.-diameter main rotor, and thrust and torque coefficients were measured for the main rotor as functions of velocity, main rotor rpm, fuselage angle of incidence, collective pitch angle, and spanwise extent of icing. The model helicopter test exhibited significant performance degradation of the main rotor when generic ice was added. An increase of approximately 150 percent in torque coefficient to maintain a constant thrust coefficient was noted when generic ice had been applied to the 85 percent rotor radial location. Also, considerable additional degradation occurred when generic ice was applied to the 100 percent rotor radial location, as compared with the 85 percent simulated ice performance values, indicating the sensitivity of the rotor tip region.

Surface Pressure Measurements at the Tip of a Model Helicopter Rotor in Hover

Surface pressures were measured near the tip of a hovering single-bladed model helicopter rotor with two tip shapes. The rotor had a constant-chord, untwisted blade with a square, flat tip which could be modified to a body-of-revolution tip. Pressure measurements were made on the blade surface along the chordwise direction at six radial stations outboard of the 94 percent blade radius. Data for each blade tip configuration were taken at blade collective pitch angles of 0, 6.18 and 11.4 degrees at a Reynolds number of 736,000 and a Mach number of 0.25 both based on tip speed. Chordwise pressure distributions and constant surface pressure contours are presented and discussed.

Experimental aerodynamcis of a rotor entry vehicle.

An unpowered entry vehicle model has been tested at Mach numbers 0.10 to 3.54 at angles of attack from 15° to 90°. The tests included variations of blade collective and cyclic pitch angles. A summary of the autorotative equilibrium speed, forces, and moments is presented. The results indicate that adding the to the capsule improved vehicle performance throughout most of the Mach number range. Maximum lift drag ratio was substantially improved at low Mach numbers. Large pitching moments were produced by the rotor, requiring the use of flaps for longitudinal trim. Control phasing was unusual compared to the predicted response of an ideal rotor but did not change grossly with forward flight speed.

Investigation and prediction of helicopter rotor noise part I. Wessex whirl tower results

This report includes noise measurements taken from a three-bladed Wessex rotor of 56 ft diameter, mounted on a rotor whirl tower. The blades used for the measurements were of constant chord and made to a N.A.C.A. 0012 profile section. The range of noise measurements taken was concentrated in the speed range from 180 to 250 rev/min and at collective pitch angles (measured at the cuff) from 7° to 17°. These measurements have been compared with theoretical predictions of broad-band noise. The magnitude of the higher harmonics of the ordered noise found from the tests appears to be in good agreement with the narrow pressure pulse width acting on the blade surface. This is discussed in the report and illustrated in Figure 10. The broad-band or “vortex” noise has been found to obey a reasonable relationship with the sixth power of the blade tip velocity and also to be dependent upon C L 1·66, where C L is a tip-referred lift coefficient. The limitations on the test results imposed by the use of the whirl tower are discussed in section 6 and suggestions for further work on other rotor configurations are indicated.

High Offset Flapping Pin Rotor Analysis

The theory of rotor dynamics given in Ref. 1 is extended to include the effects of coupling between feathering and flapping (δ3 angle) and flapping hinge offset. Both introduce considerable modification to the classic equations, and instead of simple explicit equations for flapping amplitudes, coning angle, collective pitch and inflow angles, five simultaneous equations have now to be solved. Data sheets have been constructed which enable this to be done quickly and accurately for any design of linearly tapered and twisted blade. It is suggested that the intelligent use of such data sheets is of great assistance in a design office, not only because of the very considerable time savings achieved, but also because they eliminate the most fruitful sources of error in numerical calculation. It is shown that a high offset rotor enables much higher speeds to be achieved with a conventional helicopter—an effect which has already been fairly well publicized. A penalty is paid for this in the form of hub pitching moments which have to be balanced out externally; either by the use of two rotors, offset C.G., aerodynamic surfaces, or inclination of the mechanical axis. These effects will be considered in detail in a further article. Finally, equations are developed for a convenient method of calculating blade elemental angle of attack which is claimed to be superior to classic methods for design office purposes.

A Type of Lifting Rotor with Inherent Stability

The more important of the forward flight aerodynamic characteristics of a rotor type are presented where the collective blade pitch angle is kinematically coupled with the collective flapping or coning angle, while no coupling is provided between cyclic flapping and cyclic pitch angles. The rotor with pitch-cone change may be designed to have positive static stability as compared to the negative static stability of the conventional fixedpitch rotor; partial blade stall during pull-ups or upward gusts may be avoided in spite of a high and economic thrust coefficient in steady flight, and no adjustment of the collective pitch control is required for the transition from powered flight to autorotation. The rotor with pitch-cone change, because of its inherent stability, promises, without recourse to external stabilizing devices, appreciably improved flying qualities of the helicopter.