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At least 1,351 records · Page 75
All sourcesNASA NTRSDOE OSTI
NASA NTRS1981

A shock wave approach to the noise of supersonic propellers

To model propeller noise expected for a turboprop aircraft, the pressure ratio across the shock at the propeller tip was calculated and compared with noise data from three propellers. At helical tip Mach numbers over 1.0, using only the tip shock wave, the model gave a fairly good prediction of the noise for a bladed propeller and for a propeller swept for aerodynamic purposes. However for another propeller, which was highly swept and designed to have noise cancellations from the inboard propeller sections, the shock strength from the tip over predicted the noise. In general the good agreement indicates that shock theory is a viable method for predicting the noise from these supersonic propellers but that the shock strengths from all of the blade sections need to be properly included.

Dittmar, J. H.↗
NASA NTRS1981

An experimental investigation of the flow past a finite circular cylinder at a low subcritical Reynolds number

Results of hot wire measurements made in the near wake at a Reynolds number of 9955 are reported. The measurements include the mean velocity profiles, root mean square values of the velocity fluctuations, frequency spectra, and velocity cross correlations. The mean velocity profiles were used to determine the wake width, whose variation in the downstream and spanwise directions was examined. It is observed that close to the cylinder, the wake is narrower toward the free end than it is away from it, while further downstream the wake is wider toward the tip than it is away from it. It is found that the flow over the span can be characterized by four regions: a tip region where vortex shedding occurs at a lower frequency than that prevalent for away from the tip; an intermediate region adjacent to the first one where a frequency component of a nonshedding character is present; a third region characterized by a gradually increasing shedding frequency with increasing distance from the tip; and a two dimensional region where the shedding frequency is constant.

Budair, M.↗
NASA NTRS1981

The plasma wave and quasi-static electric field instrument /PWI/ for dynamics Explorer-A

It is explained that the Plasma Wave Instrument (PWI) on Dynamics Explorer-A measures both plasma wave phenomena and quasi-static electric fields. The quasi-static electric fields are measured parallel to the spin axis of the spacecraft in a range of 2 mV/m to 2 V/m and perpendicular to the spin axis 0.5 mV/m to 2 V/m at 16 samples/s. The ac electric field sensors include a 200-m tip-to-tip long wire antenna and a 0.6-m short electric antenna, both of which are perpendicular to the spin axis, and a 9-m tip-to-tip tubular antenna parallel to the spin axis. AC electric wave fields are measured over a frequency range of 1 Hz to 2 MHz and over an amplitude range of 0.03 microvolt/m to 100 mV/m.

Shawhan, S. D.↗
NASA NTRS1982

Model helicopter rotor low frequency broadband noise

The results of an experimental investigation of low frequency broadband noise (LFBN) radiated from model helicopter rotors are presented. The results up to tip Mach number of 0.50 suggest that the peak sound pressure level (SPL) of LFBN appears to follow tip Mach number to a fourth power law and rms velocity of turbulence to a second power law. The experimental results on the effect of tip speed and advance ratio on the peak SPL of LFBN can be explained on the basis of a simple scaling law. However, the experimental results on the effect of blade loading on the peak SPL of LFBN is still not clearly understood. A simple peak SPL scaling law for noise from a helicopter in forward flight encountering a sinusoidal gust is also developed. The trends predicted by the scaling law with the experimental results are found satisfactory for the cases of variation of the peak SPL of LFBN with tip speed and advance ratio.

Humbad, N. G.↗
NASA NTRS1983

Noise of the 10-bladed 60 deg swept SR-5 propeller in a wind tunnel

Noise generated by supersonic helical tip speed propellers is a possible cabin environment problem for future airplanes powered by these propellers. Noise characteristics of one of these propellers, designated SR-5, are presented. A matrix of tests was conducted to provide as much acoustic information as possible. During aerodynamic testing it was discovered that the propeller had an aeroelastic instability which prevented testing the propeller at its design advance ratio of 4.08 at axial Mach numbers over 0.7. Plots of the variation of the maximum blade passage tone with helical tip Mach number indicate that, at higher helical tip Mach numbers, the propeller operated on sharply increasing portion of the noise curve; therefore, extrapolations to the design condition would not be accurate. A possible extrapolation indicated that SR-5 at its design point should be quieter than SR-3 at its design point. Directivity plots at the higher helical tip Mach numbers indicate a lobed directivity pattern as was observed previously on the SR-3 propeller.

Dittmar, J. H.↗
NASA NTRS1982

KC-135 wing and winglet flight pressure distributions, loads, and wing deflection results with some wind tunnel comparisons

A full-scale winglet flight test on a KC-135 airplane with an upper winglet was conducted. Data were taken at Mach numbers from 0.70 to 0.82 at altitudes from 34,000 feet to 39,000 feet at stabilized flight conditions for wing/winglet configurations of basic wing tip, 15/-4 deg, 15/-2 deg, and 0/-4 deg winglet cant/incidence. An analysis of selected pressure distribution and data showed that with the basic wing tip, the flight and wind tunnel wing pressure distribution data showed good agreement. With winglets installed, the effects on the wing pressure distribution were mainly near the tip. Also, the flight and wind tunnel winglet pressure distributions had some significant differences primarily due to the oilcanning in flight. However, in general, the agreement was good. For the winglet cant and incidence configuration presented, the incidence had the largest effect on the winglet pressure distributions. The incremental flight wing deflection data showed that the semispan wind tunnel model did a reasonable job of simulating the aeroelastic effects at the wing tip. The flight loads data showed good agreement with predictions at the design point and also substantiated the predicted structural penalty (load increase) of the 15 deg cant/-2 deg incidence winglet configuration.

Montoya, L. C.↗
NASA NTRS1984

Apparatus for disintegrating kidney stones

The useful life of the wire probe in an ultrasonic kidney stone disintegration instrument is enhanced and prolonged by attaching the wire of the wire probe to the tip of an ultrasonic transducer by means of a clamping arrangement. Additionally, damping material is applied to the wire probe in the form of a damper tube through which the wire probe passes in the region adjacent the transducer tip. The damper tube extends outwardly from the transducer tip a predetermined distance, terminating in a resilient soft rubber joint. Also, the damper tube is supported intermediate its length by a support member. The damper system thus acts to inhibit lateral vibrations of the wire in the region of the transducer tip while providing little or no damping to the linear vibrations imparted to the wire by the transducer.

Angulo, E. D.↗
NASA NTRS1984

An experimental study of the compressor rotor blade boundary layer

The three-dimensional turbulent boundary layer developing on a rotor blade of an axial flow compressor was measured using a miniature 'x' configuration hot-wire probe. The measurements were carried out at nine radial locations on both surfaces of the blade at various chordwise locations. The data derived includes streamwise and radial mean velocities and turbulence intensities. The validity of conventional velocity profiles such as the 'power law profile' for the streamwise profile, and Mager and Eichelbrenner's for the radial profile, is examined. A modification to Mager's crossflow profile is proposed. Away from the blade tip, the streamwise component of the blade boundary layer seems to be mainly influenced by the streamwise pressure gradient. Near the tip of the blade, the behavior of the blade boundary layer is affected by the tip leakage flow and the annulus wall boundary layer. The 'tangential blockage' due to the blade boundary layer is derived from the data. The profile losses are found to be less than that of an equivalent cascade, except in the tip region of the blade.

Pouagare, M.↗
NASA NTRS1984

Helicopter noise as predicted by three-dimensional monopole and quasi-steady full-potential dipole sources

A three-dimensional quasi-steady full-potential rotor-flow-analysis program, called ROT22 developed at NASA Ames Research Center is run in conjunction with Farassat's (1981) helicopter-noise-prediction code to assess the thickness and loading noises made by a rotor blade. As a model example, the case of a 1/7th UH-1H NACA-0012-profile straight nonlifting rotor blade in hover is studied. Results for tip Mach numbers ranging from .4 to .962 are presented. Also presented is the effect of blade-tip loading on the overall noise and of profile curvature on the thickness noise. The study confirms that the blade volume displacement is a dominant source of helicopter noise and further concludes that the blade tip makes only a small noise contribution that steadily decreases with increasing tip Mach numbers, and that the inclusion of blade profile curvature tends to improve the negative peak amplitudes overpredicted by Schmitz and Yu (1983).

Aggarwal, H. R.↗
NASA NTRS1985

Fundamental study of flow field generated by rotorcraft blades using wide-field shadowgraph

The vortex trajectory and vortex wake generated by helicopter rotors are visualized using a wide-field shadowgraph technique. Use of a retro-reflective Scotchlite screen makes it possible to investigate the flow field generated by full-scale rotors. Tip vortex trajectories are visible in shadowgraphs for a range of tip Mach number of 0.38 to 0.60. The effect of the angle of attack is substantial. At an angle of attack greater than 8 degrees, the visibility of the vortex core is significant even at relatively low tip Mach numbers. The theoretical analysis of the sensitivity is carried out for a rotating blade. This analysis demonstrates that the sensitivity decreases with increasing dimensionless core radius and increases with increasing tip Mach number. The threshold value of the sensitivity is found to be 0.0015, below which the vortex core is not visible and above which it is visible. The effect of the optical path length is also discussed. Based on this investigation, it is concluded that the application of this wide-field shadowgraph technique to a large wind tunnel test should be feasible. In addition, two simultaneous shadowgraph views would allow three-dimensional reconstruction of vortex trajectories.

Parthasarathy, S. P.↗
NASA NTRS1986

Full-scale tilt-rotor hover performance

The hover performance of three full-scale rotors was measured at the Ames Outdoor Aerodynamic Research Facility. The rotors, all designed for tilt-rotor aircraft, were the original metal blades for the XV-15 Tilt Rotor Research Aircraft, a set of composite, advanced technology blades for the XV-15, and a 0.658-scale model of the proposed V-22A Osprey (JVX) rotor. The composite advanced technology blades for the XV-15 were tested with several alternate blade root and blade tip configurations. This paper presents the performance of these three rotors, shows the effects of tip Mach number and root and tip configuration changes on rotor performance, and presents data on rotor wake velocity distributions and tip vortex geometry. Measured rotor performance is compared with theoretical predictions, and the discrepancies are discussed.

Felker, F. F.↗
NASA NTRS1986

Computation of transonic flow about helicopter rotor blades

An inviscid, nonconservative, three-dimensional full-potential flow code, ROT22, has been developed for computing the quasi-steady flow about a lifting rotor blade. The code is valid throughout the subsonic and transonic regime. Calculations from the code are compared with detailed laser velocimeter measurements made in the tip region of a nonlifting rotor at a tip Mach number of 0.95 and zero advance ratio. In addition, comparisons are made with chordwise surface pressure measurements obtained in a wind tunnel for a nonlifting rotor blade at transonic tip speeds at advance ratios from 0.40 to 0.50. The overall agreement between theoretical calculations and experiment is very good. A typical run on a CRAY X-MP computer requires about 30 CPU seconds for one rotor position at transonic tip speed.

Arieli, R.↗
NASA NTRS1987

Effect of planform taper on hover performance of an advanced AH-64 model rotor

The hover performance of a 27 percent scale model baseline rotor and advanced rotor with a 3:1 tapered tip (TR3) for the AH-64 attack helicopter was investigated and compared. Hover results from a previously tested advanced rotor with a 5:1 tapered tip (TR5) were also compared. Rotor thrust was varied over a range for two tip Mach numbers. The results indicated that the TR3 blades had improved performance compared with the TR5 blades, and both the TR3 and TR5 blades were superior to the baseline rotor. The additional margin in performance for the TR3 blades was likely due to an increase in blade area and Reynolds number in the tip region of the blades.

Kelley, Henry L.↗
NASA NTRS1987

Cruise noise of the 2/9th scale model of the Large-scale Advanced Propfan (LAP) propeller, SR-7A

Noise data on the Large-scale Advanced Propfan (LAP) propeller model SR-7A were taken in the NASA Lewis Research Center 8 x 6 foot Wind Tunnel. The maximum blade passing tone noise first rises with increasing helical tip Mach number to a peak level, then remains the same or decreases from its peak level when going to higher helical tip Mach numbers. This trend was observed for operation at both constant advance ratio and approximately equal thrust. This noise reduction or, leveling out at high helical tip Mach numbers, points to the use of higher propeller tip speeds as a possible method to limit airplane cabin noise while maintaining high flight speed and efficiency. Projections of the tunnel model data are made to the full scale LAP propeller mounted on the test bed aircraft and compared with predictions. The prediction method is found to be somewhat conservative in that it slightly overpredicts the projected model data at the peak.

Dittmar, James H.↗
NASA NTRS1984

Analytic investigation of helicopter rotor blade appended aeroelastic devices

Analytic evaluations of four different passive aeroelastic devices appended to helicopter rotor blades are presented. The devices consist of a passive tuned tab, a control coupled tab, an all-flying tip and a harmonic dilational airfoil tip. Each device was conceived for improving either aerodynamic performance or reducing vibratory control loads or hub shears. The evaluation was performed using a comprehensive rotor aeroelastic analysis (the G400PA code with appropriate modifications), together with data for a realistic helicopter rotor blade (the UH-60A Blackhawk), in high speed flight (90 m/s, 175 kts). The results of this study show that significant performance (L/(D sub e)) gains can be achieved with the all-flying free tip. Results from the harmonic dilational airfoil tip show the potential for moderate improvements in L/(D sub e). Finally, the results for the passive tuned tab and the control coupled tab, as configured for this study, show these devices to be impractical. Sections are included which describe the operation of each device, the required G400PA modifications, and the detailed results obtained for each device.

Bielawa, Richard L.↗
NASA NTRS1987

Cruise noise of the 2/9 scale model of the Large-scale Advanced Propfan (LAP) propeller, SR-7A

Noise data on the Large-scale Advanced Propfan (LAP) propeller model SR-7A were taken in the NASA Lewis Research Center 8 x 6 foot Wind Tunnel. The maximum blade passing tone noise first rises with increasing helical tip Mach number to a peak level, then remains the same or decreases from its peak level when going to higher helical tip Mach numbers. This trend was observed for operation at both constant advance ratio and approximately equal thrust. This noise reduction or, leveling out at high helical tip Mach numbers, points to the use of higher propeller tip speeds as a possible method to limit airplane cabin noise while maintaining high flight speed and efficiency. Projections of the tunnel model data are made to the full scale LAP propeller mounted on the test bed aircraft and compared with predictions. The prediction method is found to be somewhat conservative in that it slightly overpredicts the projected model data at the peak.

Dittmar, James H.↗
NASA NTRS1987

Rotor blade-vortex interaction impulsive noise source identification and correlation with rotor wake predictions

An acoustic source localization scheme applicable to noncompact moving sources is developed and applied to the blade-vortex interaction (BVI) noise data of a 40-percent scale BO-105 model rotor. A generalized rotor wake code is employed to predict possible VBI locations on the rotor disk and is found quite useful in interpreting the acoustic localization results. The highly varying directivity patterns of different BVI impulses generated at the same test condition are explained by both the localization results and predicted tip vortex trajectories. The effects of rotor tip-path-plane angle and advance ratio on the BVI source positions is studied. Decreasing tip-path-plane angle (at constant advance ratio) moves the general interaction region upwind on the rotor disk, significantly changing the interaction geometry. Increasing advance ratio (at constant tip-path-plane angle) shifts the general source region downwind on the rotor disk with the increased convection of the vortices until about 60 deg azimuth, where the BVI sources appear to become acoustically less effective. The region of strongest BVI sources lies between 60 and 70 deg azimuth and 80 and 90 percent radius for the moderate range of advance ratios studied.

Splettstoesser, W. R.↗
NASA NTRS1989

Arcjet cathode phenomena

Cathode tips made from a number of different materials were tested in a modular arcjet thruster in order to examine cathode phenomena. Periodic disassembly and examination, along with the data collected during testing, indicated that all of the tungsten-based materials behaved similarly despite the fact that in one of these samples the percentage of thorium oxide was doubled and another was 25 percent rhenium. The mass loss rate from a 2 percent thoriated rhenium cathode was found to be an order of magnitude greater than that observed using 2 percent thoriated tungsten. Detailed analysis of one of these cathode tips showed that the molten crater contained pure tungsten to a depth of about 150 microns. Problems with thermal stress cracking were encountered in the testing of a hafnium carbide tip. Post test analysis showed that the active area of the tip had chemically reacted with the propellant. A 100 hour continuous test was run at about 1 kW. Post test analysis revealed no dendrite formation, such as observed in a 30 kW arcjet lifetest, near the cathode crater. The cathodes from both this test and a previously run 1000 hour cycled test displayed nearly identical arc craters. Data and calculations indicate that the mass losses observed in testing can be explained by evaporation.

Curran, Francis M.↗