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The average velocity of helicopter blades relative to the air is greater than that of airplane wings. The helicopter may turn out to be more economical than the airplane wing for extreme velocities of horizontal flight, the airplane then requiring a very great speed range.
Experimental data on broadband noise from airfoils are compared, together with analytical methods, in order to identify the mechanisms of noise emission. Rotor noise is categorized into discrete frequency, impulsive, and broadband components, the last having a continuous spectrum originating from a random source. The results of computer simulations of different rotor blade types which produce broadband noise were compared with experimental data and among themselves in terms of predictions of the spectra obtained. Consideration was given to the overall sound pressure level, unsteady turbulence forces, rotational forces, inflow turbulence, self-generated turbulence, and turbulence in the flow. Data are presented for a helicopter rotor and light aircraft propeller. The most significant source was found to be inflow turbulence induced lift fluctuations in helicopter rotors and boundary layer trailing edge noise on large wind energy conversion systems
In order to properly validate experimental data, a theoretical model must be generated and converge with the results. Providing possible results for future testing can help experimenters better understand the results they are getting. Certain effects cannot be shown during experimentation, and must therefore be quantified through another means. The Mars Helicopter is currently being tested in the N242 wind tunnel at NASA Ames Research Center, which can reach a density and pressure similar to Mars. A test propeller was analyzed in a computational fluid dynamics solver. The propeller blades were previously modeled, and were simulated in the full experimental setup, which includes the blast shields, rotor stand and tunnel walls, to thoroughly generate the effects of the true testing conditions. Simulations were run using both an isolated hover condition and with the full setup, and results for thrust and total power were quantified. The data generated was used to quantify the impact of the facility on the propeller. These calculations will ultimately be used to help separate these effects from the actual thrust the propeller will generate, this making the data more accurate.
Report presents a theory which gives a complete picture and an exact quantitative analysis of the whole phenomenon of the working of blade screws, but also unites in a continuous whole the entire scale of states of work conceivable for a blade screw. Chapter 1 is devoted to the establishment of the system of fundamental equations relating to the blade screw. Chapter 2 contains the general discussion of the 16 states of work which may establish themselves for a blade screw. The existence of the vortex ring state and the whirling phenomenon are established. All the fundamental functions which enter the blade-screw theory are submitted to a general analytical discussion. The general outline of the curve of the specific function is examined. Two limited cases of the work of the screw, the screw with a zero constructive pitch and the screw with an infinite constructive pitch, are pointed out. Chapter 3 is devoted to the study of the propulsive screw or propeller. (author)
Two screw propellers, a normal airplane propeller with a pitch/diameter ratio H/D = 0.5 and a helicopter propeller H/D = 0.2 were tested in the large wind tunnel of the Gottigen Aerodynamic Institute. With both propellers the angle between the propeller axis and the direction of the wind was varied from 0 to 90 degrees. In addition to the three force and the three moment components in a wind-fast coordinate system, the corresponding components in a propeller-fast coordinate system, a total of six force and six moment components, together with the propulsive efficiency, were measured or calculated from measurements. The results are given in the form of diagrams and briefly discussed. The propeller models and method of experimentation are described and the symbols are defined.
Methodology developed to predict the performance degradation of rotating systems in natural icing conditions is described and discussed. Theoretical studies of the increments performance degradation due to icing involving the propeller, helicopter in hover and forward flight, and XV-15 propulsion modes are summarized. Related experimental studies on the NACA 0012 airfoil and model helicopter with/without generic ice shapes are reviewed. The results of these experimental and theoretical studies are used to suggest refinements to current methodology.
A numerical method and the results of a computer program are presented for solving an exact, three dimensional, full potential equation that models rotating and nonrotating inviscid, absolutely irrotational, homentropic flows. Besides calculating the flows through an arbitrarily shaped rotor or stator blade row mounted on an axisymmetric hub and confined in an axisymmetric duct, the computer program is also capable of analyzing flow fields about arbitrarily shaped wing body combinations, propellers, helicopter rotors in hover, and wind turbine rotors. The governing equation is solved numerically in a fully conservative form by using an artificial time concept, a finite volume technique, rotated type dependent differencing, successive line overrelaxation, and sequential boundary conforming grid refinement. An artificial viscosity is added in fully conservative form, and an initial guess for the potential field is applied, as determined by a two dimensional cascade analysis.
This paper presents a numerical method and the results of a computer program for solving an exact, three-dimensional, full-potential equation that models rotating and nonrotating inviscid, absolutely irrotational, homentropic flows. Besides calculating the flows through an arbitrarily shaped rotor or stator blade row mounted on an axisymmetric hub and confined in an axisymmetric duct, the computer program is also capable of analysing flow fields about arbitrarily shaped wing-body combinations, propellers, helicopter rotors in hover, and wind turbine rotors. The governing equation is solved numerically in a fully conservative form by using an artificial time concept, a finite volume technique, rotated type-dependent differencing, successive line overrelaxation, and sequential boundary-conforming grid refinement. An artificial viscosity is added in fully conservative form; and an initial guess for the potential field is applied, as determined by a two-dimensional cascade analysis.
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The governing equation and computing technique for the prediction of helicopter rotor and propeller noise are described. The method which gives both the acoustic pressure time history and spectrum of the noise includes the thickness and the loading noise. It was adapted to computers resulting in a new capability in noise prediction by removing many of the restrictions and limitations of previous theories. The capability results from the fact that the theory is developed entirely in the time domain. The formulation and the technique used are not limited to compact sources, steady level flight or to the far-field. In addition, the inputs to the computer program are normally available or are amenable to experimental measurements. This program can be used to study rotor and propeller noise with the aim of minimizing the radiated noise to reduce annoyance to the public. Several examples demonstrating the features and capability of the computer program are presented.
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The autorotative performance of an assumed helicopter was studied to determine the effect of inoperative jet units located at the rotor-blade tip on the helicopter rate of descent. For a representative ramjet design, the effect of the jet drag is to increase the minimum rate of descent of the helicopter from about 1,OO feet per minute to 3,700 feet per minute when the rotor is operating at a tip speed of approximately 600 feet per second. The effect is less if the rotor operates at lower tip speeds, but the rotor kinetic energy and the stall margin available for the landing maneuver are then reduced. Power-off rates of descent of pulse-jet helicopters would be expected to be less than those of ramjet. helicopters because pulse jets of current design appear to have greater ratios of net power-on thrust to power-off, drag than currently designed rain jets. Iii order to obtain greater accuracy in studies of autorotative performance, calculations in'volving high power-off rates of descent should include the weight-supporting effect of the fuselage parasite-drag force and the fact that the rotor thrust does not equal the weight of the helicopter.