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At least 577 records · Page 32
All sourcesNASA NTRSDOE OSTI
NASA NTRS1978

Performance of a short annular dump diffuser using suction-stabilized vortices at inlet Mach numbers to 0.41

A short, annular dump diffuser was designed to use suction to establish stabilized vortices on both walls for improved flow expansion in the region of an abrupt area change. The diffuser was tested at near ambient inlet pressure and temperature. The overall diffuser area ratio was 4.0. The inlet height was 2.54 cm and the exit pitot-static rakes were located at a distance from the vortex fence equal to two or six times the inlet height. Performance data were taken at near ambient temperature and pressure for nominal inlet Mach numbers of 0.18 to 0.41 with suction rates of 0 to 18 percent of the total inlet airflow. The exit velocity profile could be shifted toward either wall by adjusting the inner- or outer-wall suction rate. Symmetrical exit velocity profiles were unstable, with a tendency to shift back to hub- or tip-weighted profile. Diffuser effectiveness was increased from about 47 percent without suction to over 85 percent at a total suction rate of about 14 percent. The diffuser total pressure losses at inlet Mach numbers of 0.18 and 0.41 decreased from 1.1 and 5.6 percent without suction to 0.48 and 5.2 percent at total suction rates of 14.4 and 5.6 percent, respectively.

Smith, J. M.↗
NASA NTRS1977

Performance of high-area-ratio annular dump diffuser using suction-stabilized-vortex flow control

A short annular dump diffuser having a geometry conductive to formation of suction stabilized toroidal vortices in the region of abrupt area change was tested. The overall diffuser area ratio was 4.0 and the length to inlet height ratio was 2.0. Performance data were obtained at near ambient temperature and pressure for inlet Mach numbers of 0.18 and 0.30 with suction rates ranging from 0 to 18 percent of total inlet mass flowrate. Results show that the exit velocity profile could be readily biased toward either wall by adjustment of inner and outer wall suction rates. Symmetric exit velocity profiles were inherently unstable with a tendency to revert to a hub or tip bias. Diffuser effectiveness was increased from about 38 percent without suction to over 85 percent at a total suction rate of 10 to 12 percent. At the same time diffuser total pressure loss was reduced from 3.1 percent to 1.1 percent at an inlet Mach number of 0.3.

Juhasz, A. J.↗
NASA NTRS1978

Performance characteristics of two annular dump diffusers using suction-stabilized vortex flow control

Test results are described for two abrupt area change annular diffusers with provisions for maintaining suction stabilized toroidal vortices at the area discontinuity. Both diffusers had an overall area ratio of 4.0 with the prediffuser area ratio being 1.18 for diffuser A and 1.4 for diffuser B. Performance was evaluated at near atmospheric pressure and temperature for a range of inlet Mach numbers from 0.18 to 0.41 and suction rates from 0 to 18%. Static pressure recovery improved significantly as the suction rate was increased to approximately 11%. Results obtained with diffuser A were superior to that obtained with diffuser B. Flat radial profiles of exit velocity were not obtained since the flow showed preferential hub or tip attachment at moderate suction rates. At high suction rates the diffuser exit flow became circumferentially nonuniform and unstable.

Juhasz, A. J.↗
NASA NTRS2023

Simulation of a Single Rotor and a Quadrotor Hovering in Full and Partial Ground Effect

With the development of urban air mobility, the safety of multirotor vehicle operations in proximity to vertiports must be assessed. This paper examines high-fidelity blade-resolved OVERFLOW simulations and medium-fidelity vortex particle–mesh simulations that capture the aerodynamics of a rotorcraft in full and partial ground effect. Results obtained with each CFD method are compared to verify the consistency of the models introduced in the latter, which include an immersed lifting line for the blade aerodynamics and various boundary conditions to account for solid surfaces such as ground and walls. Satisfactory agreement is obtained for the blade loads and rotorwash velocity profiles in the case of a single rotor, except for small discrepancies at the blade tip. Finally, a quadrotor air taxi is simulated to quantify the load imbalance when the vehicle hovers above a rooftop edge. The combined use of high- and medium-fidelity methods is envisioned to accelerate the workflow when a large number of configurations need to be tested.

HEC↗
NASA NTRS1989

Numerical simulation of rotorcraft

The objective of the research is to develop and validate accurate, user-oriented viscous CFD codes (with inviscid options) for three-dimensional, unsteady aerodynamic flows about arbitrary rotorcraft configurations. Unsteady, three-dimensional Euler and Navier-Stokes codes are developed, adapted, and extended to rotor-body combinations. Flow solvers are coupled with zonal grid topologies, including rotating and nonrotating blocks. Special grid clustering and wave-fitting techniques were developed to capture low-level radiating acoustic waves. Significant progress was made in computing the propagation of acoustic waves due to the interaction of a concentrated vortex and a helicopter airfoil. The need for higher-order schemes was firmly established in relatively inexpensive two-dimensional calculations. In three dimensions, the number of grid points required to capture the low-level acoustic waves becomes very large, so that large supercomputer memory becomes essential. Good agreement was obtained between the numerical results obtained with a thin-layer Navier-Stokes code and experimental data from a model rotor. In addition, several nonrotating configurations that are sometimes proposed to simulate rotor blade tips in conventional wind tunnels were examined, and the complex flow around the radical tip shape of the world's fastest helicopter is under investigation. These studies demonstrate the flexibility and power of CFD to gain physical insight, study novel ideas, and examine various possibilities that might be difficult or impossible to set up in physical experiments. As a prelude to studies of rotor-body aerodynamic interactions, a preliminary grid topology and moving-interface strategy were developed. A new Euler/Navier-Stokes code using these techniques computes the vortical wake directly, rather than modeling it, as in most previous rotorcraft studies. Several hover cases were run for conventional and advanced-geometry blades. Numerical schemes using multi-zones and/or adaptive grids appear to be necessary to simulate the complex vortical flows in rotor wakes.

Mccroskey, William J.↗
NASA NTRS1972

Application of the leading edge suction analogy to prediction of longitudinal load distribution and pitching moments for sharp edged delta wings

The leading-edge-suction analogy of Polhamus has been used to develop the longitudinal load distribution of the vortex lift for delta wings. This distribution is shown to be similar in shape to that of the potential-flow longitudinal loading for delta wings having aspect ratios of 2 or less. The totals of the two theoretical distributions for delta wings with an aspect ratio near 1 are in good agreement with the experimentally determined loadings over the angle-of-attack range from 0 to 30 deg. The corresponding predicted pitching moments show slightly more stability than those measured, because of loss of lift near the wing tips.

Snyder, M. H., Jr.↗
NASA NTRS1989

Induced drag - Historical perspective

Induced drag is associated with the shedding of vorticity along the span of a finite wing, especially its tip region; for most subsonic aircraft configurations, induced drag constitutes about 50 percent of total aircraft drag throughout the flight envelope. NASA and the U.S. aircraft industry have aggressively studied induced-drag reduction methods. The state-of-the-art CTOL commercial aircraft wing is as a result of these efforts virtually optimal, with a total induced drag lying within a percent of the theoretical minimum. Many of the devices currently under study for induced drag reduction are added to wingtips, yielding benefits through their effects on the wake vortex as well as through forces generated in the flowfield.

Henderson, William P.↗
NASA NTRS1985

Vortex generators as a means for increasing rotor performance

Field tests on a horizontal axis wind turbine (HAWT) have shown that Vortex Generators (VGs) can increase the efficiency of large propeller type (horizontal axis) wind turbines. VGs are devices which are attached to the surfaces of an aerodynamic body to influence the boundary layer behavior. It is pointed out that VGs were originally developed for delaying stall on aircraft wings. An investigation was conducted regarding the possibility to employ VGs also for the improvement of the performance of an intermediate size HAWT with a diameter in the range from 24 to 46 meters. This investigation included wind tunnel tests involving a rotor blade tip section, and field tests. The wind tunnel tests showed that VGs can improve the peak lift capabilities of the section while only slightly increasing the drag. The field tests showed that VGs can increase the rotor power in winds above 6 m/s.

Corrigan, R. D.↗
NASA NTRS2011

The Effects of Ambient Conditions on Helicopter Rotor Source Noise Modeling

A new physics-based method called Fundamental Rotorcraft Acoustic Modeling from Experiments (FRAME) is used to demonstrate the change in rotor harmonic noise of a helicopter operating at different ambient conditions. FRAME is based upon a non-dimensional representation of the governing acoustic and performance equations of a single rotor helicopter. Measured external noise is used together with parameter identification techniques to develop a model of helicopter external noise that is a hybrid between theory and experiment. The FRAME method is used to evaluate the main rotor harmonic noise of a Bell 206B3 helicopter operating at different altitudes. The variation with altitude of Blade-Vortex Interaction (BVI) noise, known to be a strong function of the helicopter s advance ratio, is dependent upon which definition of airspeed is flown by the pilot. If normal flight procedures are followed and indicated airspeed (IAS) is held constant, the true airspeed (TAS) of the helicopter increases with altitude. This causes an increase in advance ratio and a decrease in the speed of sound which results in large changes to BVI noise levels. Results also show that thickness noise on this helicopter becomes more intense at high altitudes where advancing tip Mach number increases because the speed of sound is decreasing and advance ratio increasing for the same indicated airspeed. These results suggest that existing measurement-based empirically derived helicopter rotor noise source models may give incorrect noise estimates when they are used at conditions where data were not measured and may need to be corrected for mission land-use planning purposes.

Schmitz, Frederic H.↗
NASA NTRS1991

Investigation of rotor wake interactions with a body in low speed forward flight

Wind tunnel tests are reported in which the unsteady aerodynamic interactions between rotor wake and a body are examined for the case of low-speed flight. A wide-field shadowgraph is employed to visualize the rotor-tip vortices during the interactions, and the pressure signatures are measured and correlated with wake trajectories. Consideration is given to unsteady potential-flow theory to estimate the pressure loads from measured wake trajectories. The pressure signatures are very different upstream of, downstream of, and at the wake-impingement point, at which point large transient loads are transferred to the body. Rotor wake is characterized, and the wake filaments impinge on the advancing side of the body first. The pressure signatures present a complex and varying range of patterns, but it is determined that vortex surface interactions can be modeled using potential-flow analysis.

Bi, Nai-Pei↗
NASA NTRS1990

Wake Geometry Effects on Rotor Blade-Vortex Interaction Noise Directivity

Acoustic measurements from a model rotor wind tunnel test are presented which show that the directionality of rotor blade vortex interaction (BVI) noise is strongly dependent on the rotor advance ratio and disk attitude. A rotor free wake analysis is used to show that the general locus of interactions on the rotor disk is also strongly dependent on advance ratio and disk attitude. A comparison of the changing directionality of the BVI noise with changes in the interaction locations shows that the strongest noise radiation occurs in the direction of motion normal to the blade span at the time of interaction, for both advancing and retreating side BVI. For advancing side interactions, the BVI radiation angle down from the tip-path plane appears relatively insensitive to rotor operating condition and is typically between 40 and 55 deg below the disk. However, the azimuthal radiation direction shows a clear trend with descent speed, moving towards the right of the flight path with increasing descent speed. The movement of the strongest radiation direction is attributed to the movement of the interaction locations on the rotor disk with increasing descent speed.

Martin, R. M.↗
NASA NTRS2015

LagRST Model Predictions of a Wingtip Vortex Floweld

The performance of the new LagRST models is compared with curvature corrected eddyviscosity RANS models. The Reynolds-stress models provide very good predictions of thisow without any additional curvature corrections. Along with comparison of predictionsof the vortex itself, some of the other measurements included in the oweld are comparedwith the predictions in the wake and over the wing. Dierences in predictions between thevarious RANS models in the wake region are small - not substantial, and all predictionsare in line with experimental measurements from 0.005c to 0.67c downstream of the trailingedge, at odds with what self-similar wake analysis would predict.

wing tip vortices↗
NASA NTRS1998

Effect of Apex Flap Deflection on Vertical Tail Buffeting

A computational study of the effect of vortex breakdown location on vertical tail buffeting is conducted. The position of the breakdown is modified by employing an apex flap deflected by an experimentally determined optimal angle. The delayed breakdown flow and buffeting response is then compared to the nominal undeflected case. This multidisciplinary problem is solved sequentially for the fluid flow, the elastic tail deformations and the grid displacements. The fluid flow is simulated by time accurately solving the unsteady, compressible, Reynolds-averaged Navier-Stokes equations using an implicit, upwind, flux-difference splitting finite volume scheme. The elastic vibrations of the tails are modeled by uncoupled bending and torsion beam equations. These equations are solved accurately in time using the Galerkin method and a five-stage Runge-Kutta-Verner scheme. The grid for the fluid dynamics calculations is continuously deformed using interpolation functions to disperse the displacements smoothly throughout the computational domain. An angle-of-attack of 35 deg.is chosen such that the wing primary-vortex cores experience vortex breakdown and the resulting turbulent wake flow impinges on tile vertical tails. The dimensions and material properties of the vertical tails are chosen such that the deflections are large enough to insure interaction with the flow, and the natural frequencies are high enough to facilitate a practical computational solution. Results are presented for a baseline uncontrolled buffeting case and a delayed breakdown case in which the apex flap has been deflected 15 deg. The flap was found to be very effective in delaying the breakdown, increasing the location from 50%c to 94%c, which resulted in a 6% increase in lift coefficient and pitching moment. However, the integrated buffet loads and tip responses were roughly equivalent for the two cases.

Massey, Steven J.↗
NASA NTRS1984

Helicopter model rotor-blade vortex interaction impulsive noise: Scalability and parametric variations

Acoustic data taken in the anechoic Deutsch-Niederlaendischer Windkanal (DNW) have documented the blade vortex interaction (BVI) impulsive noise radiated from a 1/7-scale model main rotor of the AH-1 series helicopter. Averaged model scale data were compared with averaged full scale, inflight acoustic data under similar nondimensional test conditions. At low advance ratios (mu = 0.164 to 0.194), the data scale remarkable well in level and waveform shape, and also duplicate the directivity pattern of BVI impulsive noise. At moderate advance ratios (mu = 0.224 to 0.270), the scaling deteriorates, suggesting that the model scale rotor is not adequately simulating the full scale BVI noise; presently, no proved explanation of this discrepancy exists. Carefully performed parametric variations over a complete matrix of testing conditions have shown that all of the four governing nondimensional parameters - tip Mach number at hover, advance ratio, local inflow ratio, and thrust coefficient - are highly sensitive to BVI noise radiation.

Splettstoesser, W. R.↗
NASA NTRS1985

Helicopter model rotor-blade vortex interaction impulsive noise - Scalability and parametric variations

Acoustic data taken in the anechoic Deutsch-Niederlaendischer Windkanal (DNW) have documented the blade vortex interaction (BVI) impulsive noise radiated from a 1/7-scale model main rotor of the AH-1 series helicopter. Averaged model scale data were compared with averaged full scale, inflight acoustic data under similar nondimensional test conditions. At low advance ratios (mu = 0.164 to 0.194), the data scale remarkable well in level and waveform shape, and also duplicate the directivity pattern of BVI impulsive noise. At moderate advance ratios (mu = 0.224 to 0.270), the scalig deteriorates, suggesting that the model scale rotor is not adequately simulating the full scale BVI noise; presently, no proved explanation of this discrepancy exists. Carefully performed parametric variations over a complete matrix of testing conditions have shown that all of the four governing nondimensional parameters - tip Mach number at hover, advance ratio, local inflow ratio, and thrust coefficient - are highly sensitive to BVI noise radiation.

Boxwell, D. A.↗
NASA NTRS2006

Thermal Challenge for Sun-Viewing Instruments on Solar Probe 3 R(sub s) from Sun

The Solar Probe science team at NASA GSFC is considering enhance science observations by adding the nadir viewing Plasma Spectrometer instrument and using heat shields to protect the Plasma Wave antennas to permit full extension at 3 R. The electrostatic mirror of the Plasma Spectrometer is thermally protected from the Sun by a miniature sunshield that consists of two carbon-carbon composite layers, and a much smaller carbon-carbon heat shield. The sun-viewing surface of the miniature sunshield has an Al2O3 thin film coating. The slope of the sunshield is minimized to the smallest possible extent so that it can still accommodate the elliptical electrostatic mirror. Two small MLI flaps minimize the heat radiation from the Plasma Spectrometer to the spacecraft bus. This design meets the thermal requirements. Each of the four Plasma Wave wire antennas is 5 m long, and thermally protected from the Sun by a sunshield. The sunshield consists of two carbon-carbon composite sheets. One sheet views the Sun, and one does not. The slope of most of the sun-viewing sheet is 23deg with the nadir axis. For mechanical reason, it gradually increases to 37.5deg at the tip. The exterior of the sun viewing surfaces has a thin film of Al2O3. The exterior of the shaded sheets has a thin film of vapor deposited rhodium. It minimizes the heat radiation to the spacecraft bus. A high temperature MLI blanket between the carbon-carbon sheets radiatively isolates them. The thermal design meets the temperature requirements, but exceeds the 20 W limit of heat radiation to the spacecraft bus by 20 W. Thermal options to reduce the heat radiation include reducing the alpha to epsilon ratio of the Al2O3 thin film and reducing the number of antennas to three. Testing in 2006 includes thermal model validation and materials testing. High temperature ovens and a high light intensity Vortex Lamp will be used.

Choi, Michael K.↗
NASA NTRS1987

Helicopter Model Rotor-Blade Vortex Interaction Impulsive Noise: Scalability and Parametric Variations

Acoustic data taken in the anechoic Deutsch-Niederlaendischer Windkanal (DNW) have documented the blade-vortex interaction (BVI) impulsive noise radiated from a 1/7-scale model main rotor of the AH-1 series helicopter. Averaged model-scale data were compared with averaged full-scale, in-flight acoustic data under similar non-dimensional test conditions using an improved data analysis technique. At low advance ratios (mu = 0.164 - 0.194), the BVI impulsive noise data scale remarkably well in level, waveform, and directivity patterns. At moderate advance ratios (mu = 0.224 - 0.270), the scaling deteriorates, suggesting that the model-scale rotor is not adequately simulating the full-scale BVI noise. Presently, no proved explanation of this discrepancy exists. Measured BVI noise radiation is highly sensitive to all of the four governing nondimensional parameters--hover tip Mach number, advance ratio, local inflow ratio, and thrust coefficient.

Boxwell, D. A.↗
NASA NTRS2010

A Computational Study of BVI Noise Reduction Using Active Twist Control

The results of a computational study examining the effects of active-twist control on blade-vortex interaction (BVI) noise using the Apache Active Twist Rotor are presented. The primary goal of this activity is to reduce BVI noise during a low-speed descent flight condition using active-twist control. Rotor aeroelastic behavior was modeled using the Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics code and the rotor noise was predicted using the acoustics code PSU-WOPWOP. The accuracy of the analysis was validated through comparisons with experimental acoustic data for the first generation Active Twist Rotor at an advance ratio of mu=0.14. The application of active-twist to the main rotor blade system consisted of harmonic actuation frequencies ranging from 2P to 5P, control phase angles from 0' to 360 , and tip-twist amplitudes ranging from 0.5 to 4.0 . The acoustic analysis was conducted for a single low-speed flight condition of advance ratio =0.14 and shaft angle-of-attack, c^=+6 , with BVI noise levels predicted on a flat plane of observers located 1.1 rotor diameters beneath the rotor. The results indicated reductions of up to 11dB in BVI noise using 1.25 tip-twist amplitude with negligible effects on 4P vertical hub shear.

Fogarty, David E.↗