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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1961

Flight Investigation of Effects of Selected Operating Conditions on the Bending and Torsional Moments Encountered by a Helicopter Rotor Blade

Flight tests have been conducted with a single-rotor helicopter to determine the effects of partial-power descents with forward speed, high-speed level turns, pull-outs from autorotation, and high-forward-speed high-rotor-speed autorotation on the flapwise bending and torsional moments of the rotor blade. One blade of the helicopter was equipped at 14 percent and 40 percent of the blade radius with strain gages calibrated to measure moments rather than stresses. The results indicate that the maximum moments encountered in partial-power descents with forward speed tend to be generally reduced from the maximum moments encountered during partid-power descents at zero forward speed. High-speed level turns and pull-outs from auto-rotation caused retreating-blade stall which produced torsional moments (values up to 2,400 inch-pounds). at the 14-percent-radius station that were as large as those encountered during the previous investigations of retreating-blade stall (values up t o 2,500 inch-pounds). High-forward- speed high-rotor-speed autorotation produced flapwise bending moments (values up to 7,200 inch-pounds) at the 40-percent-radius station which were as large as the flapwise bending moments (values up to 7,800 inch-pounds) a t the 14-percent-radius station encountered during partial - power vertical descents. The results of the present investigation (tip-speed ratios up to 0.325 and an unaccelerated level-flight mean lift coefficient of about 0.6), in combination with the related results of at zero forward speed produce the largest rotor-blade vibratory moments. However, inasmuch as these large moments occur only during 1 percent of the cycles and 88 percent of the cycles are at moment values less than 70 percent of these maximum values in partial-power descents, other conditions, such as high-speed flight where the large moments are combined with large percentages of time spent,must not be neglected in any rotor-blade service-life assessment.

Ludi, LeRoy H.↗
NASA NTRS1959

Flight Investigation of Effects of Transition, Landing Approaches, Partial-Power Vertical Descents, and Droop-Stop Pounding on the Bending and Torsional Moments Encountered by a Helicopter Rotor Blade

Flight tests have been conducted with a single-rotor helicopter, one blade of which was equipped at 14 percent and 40 percent of the blade radius with strain gages calibrated to measure moments rather than stresses, to determine the effects of transition, landing approaches, and partial-power vertical descents on the rotor-blade bending and torsional moments. In addition, ground tests were conducted to determine the effects of static droop-stop pounding on the rotor-blade moments. The results indicate that partial-power vertical descents and landing approaches produce rotor-blade moments that are higher than the moments encountered in any other flight condition investigated to date with this equipment. Decelerating through the transition region in level flight was found to result in higher vibratory moments than accelerating through this region. Deliberately induced static droop-stop pounding produced flapwise bending moments at the 14-percent-radius station which were as high as the moments experienced in landing approaches and partial-power vertical descents.

Ludi, LeRoy H.↗
NASA NTRS1986

Helicopter rotor and engine sizing for preliminary performance estimation

Methods are presented for estimating some of the more fundamental design variables of single-rotor helicopters (tip speed, blade area, disk loading, and installed power) based on design requirements (speed, weight, fuselage drag, and design hover ceiling). The well-known constraints of advancing-blade compressibility and retreating-blade stall are incorporated into the estimation process, based on an empirical interpretation of rotor performance data from large-scale wind-tunnel tests. Engine performance data are presented and correlated with a simple model usable for preliminary design. When approximate results are required quickly, these methods may be more convenient to use and provide more insight than large digital computer programs.

Talbot, P. D.↗
NASA NTRS1982

High-speed noise of helicopter rotors

Various parameters of helicopter rotor noise were considered. Impulsive noise, flow regions of a helicopter rotor, noise prediction, aspect ratio, blade tip shape and speed, and blade-vortex interaction noise were among the topics addressed. Recommendations were also given.

Shenoy, K. R.↗
NASA NTRS1975

Wind tunnel investigation of helicopter-rotor wake effects on three helicopter fuselage models

The effects of rotor wake on helicopter fuselage aerodynamic characteristics were investigated in the Langley V/STOL tunnel. Force, moment, and pressure data were obtained on three fuselage models at various combinations of windspeed, sideslip angle, and pitch angle. The data show that the influence of rotor wake on the helicopter fuselage yawing moment imposes a significant additional thrust requirement on the tail rotor of a single-rotor helicopter at high sideslip angles.

Wilson, J. C.↗
NASA NTRS1985

An evaluation of three helicopter rotor sections

Three helicopter rotor sections were tested in the NASA Ames Research Center 2- by 2-Foot Transonic Wind Tunnel over a Mach range from 0.2 to 0.88. The sections tested had maximum thickness/chord ratios of 0.078, 0.09, and 0.10. The thickest section was of early technology and had been tested previously in other wind tunnels. This section was included in the investigation to establish a basis for comparing the two thinner sections, which were of recent design. The results of the investigation showed that the pitching-moment characteristics for the three airfoil sections were acceptable. The drag divergence Mach numbers for the three sections were 0.80, 0.825, and 0.845 in order of decreasing thickness.

Hicks, R. M.↗
NASA NTRS1993

Optical Shaft-Angle Encoder For Helicopter Rotor

Angular position of helicopter rotor blade determined precisely. Accomplished by use of optical shaft-angle encoder called "256 Ring" on rotor swashplate. Each 360 degree rotation of helicopter main rotor broken down into 256 reflective segments. As rotor rotates, beam of light reflected in turn from each segment into optoelectronic system. One of 256 segments reflects larger pulse than others do. Position of rotor determined by counting number of pulses after this reference pulse. While swashplate mounting requirements unique to each type of helicopter, concept applicable to all types of rotorcraft.

Golub, Robert A.↗
NASA NTRS1974

Wind tunnel investigation of helicopter rotor wake effects on three helicopter fuselage models

The effects of rotor downwash on helicopter fuselage aerodynamic characteristics were investigated. A rotor model for generating the downwash was mounted close to each of three fuselage models. The main report presents the force and moment data in both graphical and tabular form and the pressure data in graphical form. This supplement presents the pressure data in tabular form. Each run or parameter sweep is identified by a unique run number. The data points in each run are identified by a point number. The pressure data can be matched to the force data by matching the run and point number.

Wilson, J. C.↗
NASA NTRS1992

Optical Detection Of Ice On A Helicopter Rotor

Ice forming on helicopter rotor blades under adverse weather conditions detected via effect on polarization of reflected light, according to proposal. Ice-detection system based on this principal alerts helicopter pilot to need for corrective action, and/or turns on deicing equipment automatically.

Cornwell, Donald M., Jr.↗
NASA NTRS1979

Unsteady flow and dynamic response analyses for helicopter rotor blades

Research is presented on helicopter rotor blade vibration and on two and three dimensional analyses of unsteady incompressible viscous flow past oscillating helicopter rotor blades. A summary is presented of the two international research collaborations which resulted from the NASA project: the collaboration under the auspices of NATO between the University of Wisconsin-Milwaukee, University of Brussels, Belgium and the Aerodynamics Research Establishment in Goettingen, West Germany, and the collaboration under the auspices of the National Science Foundation between UWM and the University of Hamburg and the Ship Research Establishment in Hamburg, West Germany. A summary is given of the benefits from the NASA project to UWM, the College of Engineering and Applied Science, and the participants on the project.

Bratanow, T.↗
NASA NTRS1976

An investigation of flap-lag stability of wind turbine rotors in the presence of velocity gradients and helicopter rotors in forward flight

The flap-lag equations of motion of a torsionally rigid and centrally hinged spring-restrained rigid blade are developed using quasi-steady blade-element aerodynamic theory. These equations have periodic coefficients and are applicable to both wind-turbine rotors with velocity gradients and helicopter rotors in forward flight. By solving these equations both by the Floquet-Liapunov method and by an approximate method, the range of applicability of the latter method is established. Flap-lag and pure flapping stability boundaries illustrating the effects of flow and rotor parameters are presented. Finally, some discussion on the techniques for generating the Floquet transition matrix and on the relative merits of the coordinate axes system is presented.

Kaza, K. R. V.↗
NASA NTRS1986

Pitch Control for Helicopter Rotors

Pitch controller for helicopter rotors uses hub-mounted actuators located symmetrically between rotor blades. New controller designed for X-wing rotors requiring collective pitch control; pitch of all blades must be changed by same amount. Collective control allows rotor trimmed during variety of flight regimes, particularly during hovering and fixed-wing flight.

Jeffery, P.↗
NASA NTRS1997

Supersonic Quadrupole Noise Theory for High-Speed Helicopter Rotors

High-speed helicopter rotor impulsive noise prediction is an important problem of aeroacoustics. The deterministic quadrupoles have been shown to contribute significantly to high-speed impulsive (HSI) noise of rotors, particularly when the phenomenon of delocalization occurs. At high rotor-tip speeds, some of the quadrupole sources lie outside the sonic circle and move at supersonic speed. Brentner has given a formulation suitable for efficient prediction of quadrupole noise inside the sonic circle. In this paper, we give a simple formulation based on the acoustic analogy that is valid for both subsonic and supersonic quadrupole noise prediction. Like the formulation of Brentner, the model is exact for an observer in the far field and in the rotor plane and is approximate elsewhere. We give the full analytic derivation of this formulation in the paper. We present the method of implementation on a computer for supersonic quadrupoles using marching cubes for constructing the influence surface (Sigma surface) of an observer space- time variable (x; t). We then present several examples of noise prediction for both subsonic and supersonic quadrupoles. It is shown that in the case of transonic flow over rotor blades, the inclusion of the supersonic quadrupoles improves the prediction of the acoustic pressure signature. We show the equivalence of the new formulation to that of Brentner for subsonic quadrupoles. It is shown that the regions of high quadrupole source strength are primarily produced by the shock surface and the flow over the leading edge of the rotor. The primary role of the supersonic quadrupoles is to increase the width of a strong acoustic signal.

Farassat, F.↗
NASA NTRS1993

Sources of helicopter rotor hub inplane shears

Sources of helicopter rotor hub inplane shears are identified using simplified equations and the full aeroelastic analysis code, CAMRAD/JA (Johnson, 1988). Analytical results are obtained for an articulated rotor operating at moderate thrust and high airspeed. It is found that the blade chordwise inplane shear, which includes the aerodynamic component, the Coriolis contribution, and the inertial component, and the hub inplane shears are strongly dependent on the out-of-plane response. The sources of helicopter rotor hub inplane shears lie not only in the inplane response but depend on the flap and elastic flatwise responses/modes.

Kottapalli, Sesi↗