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This paper discusses the development of handbook techniques for the preliminary assessment of V/STOL aircraft performance while hovering in the proximity of the ground. In particular, the suckdown forces due to entrainment are described for configurations using a single jet. Corrections to current methods are suggested from an analysis of recent large-scale test data.
Some fundamental aspects of rotor noise generation and radiation are presented. Data on which the study is based were obtained during a hover test at the NASA Ames Outdoor Aerodynamic Research Facility of a one-sixth-scale, four-bladed, helicopter rotor. The test site provided an open environment ideal for acquiring good acoustic data. Information is presented on the delineation between the acoustic near-field and far-field, and on the effect of a simple boundary-layer trip device. Data obtained at near-zero thrust conditions indicate that it is possible to isolate thickness noise with minimum contamination by loading noise effects. An abrupt change in the nature of the emitted noise at high-thrust conditions is investigated and is related to unsteady wake/support interaction.
New method for analyzing free wake of hovering-helicopter rotor produces more reliable results and requires less computer time. Copes with wake instabilities, both physical and numerical, that afflict usual time-stepping analysis methods. Tip-vortex position, in coordinates moving with rotor blade, updated in relaxation procedure. Wake-airflow solution reached when updates converge to steady wake shape.
A piloted simulator experiment was conducted to investigate directional axis handling qualities requirements for low speed and hover tasks performed by a Scout/Attack helicopter. Included were the directional characteristics of various candidate light helicopter family configurations. Also, the experiment focused on conventional single main/tail rotor configurations of the OH-58 series aircraft, where the first-order yaw-axis dynamic effects that contributed to the loss of tail rotor control were modeled. Five pilots flew 22 configurations under various wind conditions. Cooper-Harper handling quality ratings were used as the primary measure of merit of each configuration. The results of the experiment indicate that rotorcraft configurations with high directional gust sensitivity require greater minimum yaw damping to maintain satisfactory handling qualities during nap-of-the-Earth flying tasks. It was also determined that both yaw damping and control response are critical handling qualities parameters in performing the air-to-air target acquisition and tracking task. Finally, the lack of substantial yaw damping and larger values of gust sensitivity increased the possibility of loss of directional control at low airspeeds for the single main/tail rotor configurations.
Analysis and testing were conducted in the Langley Transonic Dynamics Tunnel to investigate the aeromechanical stability of a soft inplane hingeless rotor model. Rotor stability data were obtained in hover and in forward flight up to an advance ratio of 0.35. Model rotor parameters evaluated were blade sweep and droop, pre-cone of the blade feathering axis, and blade pitch-flap coupling. Data obtained during these tests are presented without analysis.
The hovering performance predictions of the TFAR1 and OPLIN codes and the experimental data are discussed. The TFAR1 program solves the full-potential equation in a rotor-fixed coordinate system by use of the line relaxation method. The OPLIN program calculates the positions of wake vortices and rotor performance using the influence-coefficient-and-lifing-line method. The two programs are combined by adding the induced velocities from the OPLIN code to the near flow-field of the rotor from the TFAR1 code. Results show that the TFAR1 program converges better with the downwash-coupling method than the twist correction method to include the wake downwash.
An experimental and theoretical investigation of rotor/wing aerodynamic interactions in hover is described. The experimental investigation consisted of both a large-scale and small-scale test. A 0.658-scale, V-22 rotor and wing was used in the large-scale test. Wind download, wing surface pressure, rotor performance, and rotor downwash data from the large-scale test are presented. A small-scale experiment was conducted to determine how changes in the rotor/wing geometry affected the aerodynamic interactions. These geometry variations included the distance between the rotor and wing, wing incidence angle, and configurations both with the rotor axis at the tip of the wing (tilt rotor configuration) and with the rotor axis at the center of the wing (compound helicopter oonfiguration). A wing with boundary-layer control was also tested to evaluate the effect of leading and trailing edge upper surface blowing on the wing download. A computationally efficient, semi-empirical theory was developed to predict the download on the wing. Finally, correlations between the theoretical predictions and test data are presented.
A solution for transonic flow past a twisted rotor blade in hover is obtained using a modified version of the full-potential code ROT22 and a circular wake. The flow is also evaluated for a fixed-wing-type straight wake. The solutions for the straight wake and circular wake, and the circular wake and a two-dimensional wake are compared. The data reveal that the circular wake and the general two-dimensional wake solutions have similar characteristics.
A finite element formulation is used to investigate ground and air resistence in hover for a bearingless rotor. Aerodynamic forces are studied using quasi-steady strip theory, and unsteady aerodynamic effects are introduced through an inflow dynamics model. Reasonable correlation was found between predicted ground and air resonance results and data obtained from measurements using a 1/8th Froude-scaled dynamic model. Systematic parametric studies of the effects of various design parameters were performed, and lag frequency was found to significantly influence ground resonance stability, whereas pitch-lag coupling, blade sweep and pitch link stiffness had powerful effects on air resonance stability.
A simple analytical method has been developed to predict the induced velocities at the tip vortex in the wake of a hovering helicopter rotor. The wake model in the analysis includes the blade bound vortex as well as a tip vortex and vortex sheet trailed from each blade. Standard, prescribed wake models were used to describe the geometry of the tip vortex and the vortex sheet. Curved vortex elements were used to model the tip vortices and the vortex sheets. The analysis was used to study the contribution from each wake component to the axial-induced velocity of the tip vortex. Results show a highly variable induced velocity at the tip vortex both before and after the first blade passage. However, the prescribed tip vortex models require a constant axial velocity of the tip vortex. This inconsistency between the prescribed axial velocity of the tip vortex and the velocity that the wake geometry induces on the tip vortex points to a limitation of the current prescribed tip vortex geometry models.
The effect of Control/Display interaction in the hover task is analyzed using an optimal control approach to modeling pilot control behavior. The control/display configurations considered are those previously evaluated in a flight research program. The experimental data-base is reviewed and the procedure for modeling the task and the displayed information is presented in detail. All model-based results, time-domain as well as frequency-domain, are found to correlate extremely well with the subjective pilot ratings and comments. Time-domain measures consist of root mean-square errors and control inputs, attention allocation to displayed quantities, and magnitudes of task objective function. Frequency-domain measures include bandwidth, stability margins, and pilot phase compensation. Results are also shown to agree with previous findings on task interference in multi-axis tasks.
A small scale, 1.92 m diam, torsionally soft, hingeless helicopter rotor was investigated in hover to determine isolated rotor stability characteristics. The two-bladed, untwisted rotor was tested on a rigid test stand at tip speeds up to 101 m/sec. The rotor mode of interest is the lightly damped lead-lag mode. The dimensionless lead-lag frequency of the mode is approximately 1.5 at the highest tip speed. The hub was designed to allow variation in precone, blade droop, pitch control stiffness, and blade pitch angle. Measurements of modal frequency and damping were obtained for several combinations of these hub parameters at several values of rotor speed. Steady blade bending moments were also measured. The lead-lag damping measurements were found to agree well with theoretical predictions for low values of blade pitch angle. The test data confirmed the predicted effects of precone, droop, and pitch control stiffness parameters on lead-lag damping. The correlation between theory and experiment was found to be poor for the mid-to-high range of pitch angles where the theory substantially overpredicted the experimental lead-lag damping. The poor correlation in the mid-to-high blade pitch angle range is attributed to low Reynolds number nonlinear aerodynamics effects not included in the theory. The experimental results also revealed an asymmetry in lead-lag damping between positive and negative thrust conditions.
Rotor hover performance data were obtained for two full-scale rotor systems designed for the XV-15 Tilt Rotor Research Aircraft. One rotor employed the rectangular planform metal blades (rotor solidity = 0.089) which were used on the initial flight configuration of the XV-15. The second rotor configuration examined the nonlinear taper, composite-construction, Advanced Technology Blade (ATB), (rotor solidity = 0.10) designed to replace the metal blades on the XV-15. Variations of the baseline ATB tip and cuff shapes were also tested. A new six-component rotor force and moment balance designed to obtain highly accurate data over a broad range of thrust and torque conditions is described. The test data are presented in nondimensional coefficient form for the performance results, and in dimensional form for the steady and alternating loads. Some wake and acoustic data are also shown.
Three cases were selected for correlation from an experiment that examined the aeromechanical stability of a small-scale model of a hingeless rotor and fuselage in hover. The first case examined the stability of a configuration with 0 degree blade pitch so that coupling between dynamic modes was minimized. The second case was identical to the first except the blade pitch was set to 9 degrees which provides flap-lag coupling of the rotor modes. The third case had 9 degrees of blade pitch and also included negative pitch-lag coupling, and therefore was the most highly coupled configuration. Analytical calculations were made by Bell Helicopter Textron, Boeing Vertol, Hughes Helicopters, Sikorsky Aircraft, the U.S. Army Aeromechanics Laboratory, and NASA Ames Research Center and compared to some or all of the experimental cases. Overall, the correlation ranged from very poor-to-poor to good.
Three cases were selected for correlation from an experiment that examined the aeroelastic stability of a small-scale bearingless motor rotor in hover. The 1.8 m diameter model rotor included flap, lead-lag, and torsional degrees of freedom, but no body degrees of freedom. The first case looked at a configuration with a single pitch link on the leading edge, the second case examined a configuration with a single pitch link on the trailing edge, and the third case examined a configuration with pitch links on the leading and trailing edges to simulate a pitch link with shear restraint. Analyses from Bell Helicopter Textron, Boeing Vertol, Hughes Helicopters, Sikorsky Aircraft, and the U.S. Army Aeromechanics Laboratory were compared with the data, and the correlation ranged from poor to fair.
The equilibrium and stability of the coupled elastic lead/lag, flap, and torsion motion of a cantilever rotor blade in hover are addressed, and the influence of several higher-order terms in the equations of motion of the blade is determined for a range of values of collective pitch. The blade is assumed to be untwisted and to have uniform properties along its span. In addition, chordwise offsets between its elastic, tension, mass, and aerodynamic centers are assumed to be negligible for simplicity. The aerodynamic forces acting on the blade are modeled using a quasi-steady, strip-theory approximation.
A numerical solution technique is developed for computing the flow field around an isolated helicopter rotor in hover. The flow is governed by the compressible Euler equations which are integrated using a finite volume approach. The Euler equations are coupled to a free wake model of the rotary wing vortical wake. This wake model is incorporated into the finite volume solver using a prescribed flow, or perturbation, technique which eliminates the numerical diffusion of vorticity due to the artificial viscosity of the scheme. The work is divided into three major parts: (1) comparisons of Euler solutions to experimental data for the flow around isolated wings show good agreement with the surface pressures, but poor agreement with the vortical wake structure; (2) the perturbation method is developed and used to compute the interaction of a streamwise vortex with a semispan wing. The rapid diffusion of the vortex when only the basic Euler solver is used is illustrated, and excellent agreement with experimental section lift coefficients is demonstrated when using the perturbation approach; and (3) the free wake solution technique is described and the coupling of the wake to the Euler solver for an isolated rotor is presented. Comparisons with experimental blade load data for several cases show good agreement, with discrepancies largely attributable to the neglect of viscous effects. The computed wake geometries agree less well with experiment, the primary difference being that too rapid a wake contraction is predicted for all the cases.
A fully compressible method for determining helicopter rotor wake effects is described which computes the wake without requiring external specification of the wake, or separate computations for the wake and blade region. The method is a modification of a compressible finite volume Potential Flow technique, and it has been implemented in a program, HELIX I, for computing compressible rotor flow fields in hover with free wakes. Wake positions in substantial agreement with experiment have been calculated for cases including subsonic and transonic flows, high and low aspect ratios, and two- and four-bladed rotors.