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At least 109 records · Page 6

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NASA NTRS2000

Application of Three-Component PIV to a Hovering Rotor Wake

The key to accurate predictions of rotorcraft aerodynamics, acoustics, and dynamics lies in the accurate representation of the rotor wake. The vortical wake computed by rotorcraft CFD analyses typically suffer from numerical dissipation before the first blade passage. With some a priori knowledge of the wake trajectory, grid points can be concentrated along the trajectory to minimize the dissipation. Comprehensive rotorcraft analyses based on lifting-line theory rely on classical vortex models and/or semi-empirical information about the tip vortex structure. Until the location, size, and strength of the trailed tip vortex can be measured over a range of wake ages, the analyses will continue to be adjusted on a trial and error basis in order to correctly predict blade airloads, acoustics, dynamics, and performance. Using the laser light sheet technique, tip vortex location can be acquired in a straightforward manner. Measuring wake velocities and vortex core size, however, has been difficult and tedious using point-measurement techniques such as laser velocimetry. Recently, the Particle Image Velocimetry (PIV) technique has proven to be an efficient method for acquiring velocity measurements over relatively large areas and volumes of a rotor wake. The work reported to date, however, has been restricted to 2-component velocity measurements of the rotor wake. Three-component velocity measurements of a hovering rotor wake were acquired at NASA Ames Research Center in May 1999. This experiment represents a major step toward understanding the detailed structure of a rotor wake. This paper will focus primarily on the experimental technique used in acquiring this data. The accuracy and limitations of the current technique will also be discussed. Representative velocity field measurements will be included.

Yamauchi, Gloria K.↗

NASA NTRS1987

Analysis of induced velocities in the wake of a hovering helicopter rotor

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.

Light, Jeffrey S.↗

NASA NTRS2016

Coaxial Rotor Flow Phenomena in Forward Flight

Coaxial rotors are nding use in advanced rotorcraft concepts. Combined with lift offset rotor technology, they offer a solution to the problems of dynamic stall and re-verse ow that often limit single rotor edgewise forward ight speeds. Lower tip speed means reduced high-speed impulsive noise. The need for an anti-torque tail rotor is eliminated, a major boon during operation in conned areas. However, the operation of two counter-rotating ro-tors in close proximity generates many possibilities for aerodynamic interactions between rotor blades, blades and vortices, and between vortices. The parameter de-sign space is very large, and requires efcient computations as well as basic experiments to explore important physics to determine performance, loads, and acoustics. Computations are done on the classic HarringtonDingeldein rotor test case from the 1950s using the ROTUNS Navier Stokes code. Two regimes are explored: very low advance ratio as a perturbation from hover, and high advance ratio. Flow eld properties from RotUNS are used with 2-D OVERFLOW computations to capture blade crossing effects including those of higher subsonic Mach numbers. Bladeblade and bladevortex intersection events are captured using a MatLab-based predictor.

Flow Phenomena↗

NASA NTRS1989

Computational and experimental evaluation of helicopter rotor tips for high speed forward flight

A computational and experimental method for the evaluation of helicopter rotor tips in high-speed forward flight is presented which uses an unsteady full-potential solver on the advancing side of the rotor disk. Forces and moments measured during angle-of-attack sweeps reveal the soft-stall phenomenon of the double-swept planforms as well as the delayed stall of the single-swept hyperbolic tip. Double-swept planforms are shown to exhibit the most favorable performance, creating counter-rotating vortices that augment the lifting capabilities of the blade at high angles of attack and delay the onset of stall.

Sigl, Dave↗

NASA NTRS1995

Turbomachinery design and tonal acoustics computations

The objective of this research was two-fold. The first objective was to complete the three-dimensional unsteady calculations of the flow through a new transonic turbine and study the effects of secondary flows due to the hub and casing, tip clearance vortices, and the inherent three-dimensional mixing of the flow. It should be noted that this turbine was and is still in the design phase and the results of the calculations have formed an integral part of the design process. The second objective of this proposal was to evaluate the capability of rotor-stator interaction codes to calculate tonal acoustics.

Rangwalla, Akil A.↗

NASA NTRS1974

Laser velocimeter measurements of the helicopter rotor-induced flow field

The use of a two-color laser velocimeter to measure the flow velocities in the wake of a helicopter rotor is discussed, including methods for obtaining two components of both instantaneous and time-averaged velocities. Results are presented from an experiment using a 2.13 m diameter model helicopter rotor operating at a tip speed ratio of 0.18 in a wind tunnel. The location of the tip vortex from the preceding blade was determined on the advancing side, and the diameter of the vortex core was found to be 15 percent of the blade chord (1.5 percent of the radius). The effects of the airfoil's bound vorticity were observed in the velocity distributions vry near the blade. These effects suggest that the laser velocimeter may be used to determine the aerodynamic loading (circulation) at a spanwise station on the blade. Also, the structure and boundary of the time-averaged wake were investigated.

Biggers, J. C.↗

NASA NTRS2002

Revolutionary Concepts for Helicopter Noise Reduction: SILENT Program

As part of a NASA initiative to reduce helicopter main rotor noise, a Phase 1 study has been performed of candidate noise reduction concepts. Both conventional and novel design technologies have been analyzed that reduce the community impact of helicopter operations. In this study the noise reduction potential and design implications are assessed for conventional means of noise reduction, e.g., tip speed reduction, tip shapes and airfoil tailoring, and for two innovative design concepts: modulated blade spacing and x-force control. Main rotor designs that incorporate modulated blade spacing are shown to have reduced peak noise levels in most flight operations. X-force control alters the helicopter's force balance whereby the miss distance between main rotor blades and shed vortices can be controlled. This control provides a high potential to mitigate BVI noise radiation. Each concept is evaluated using best practice design and analysis methods, achieving the study's aim to significantly reduce noise with minimal performance degradation and no vibration increase. It is concluded that a SILENT main rotor design, incorporating the modulated blade spacing concept, offers significantly reduced noise levels and the potential of a breakthrough in how a helicopter's sound is perceived and judged. The SILENT rotor represents a definite advancement in the state-of-the-art and is selected as the design concept for demonstration in Phase 2. A Phase 2 Implementation Plan is developed for whirl cage and wind tunnel evaluations of a scaled model SILENT rotor.

Edwards, Bryan↗

NASA NTRS1989

The prediction of blade wake interaction noise based on a turbulent vortex model

Blade wake interaction is defined as the broadband noise generated by the interaction of helicopter rotor blades with their own wake. Experimental observations have shown that this is a strong function of advance ratio and tip path plane angle. This paper describes how this noise source can be associated with the blade vortex interactions in the forward sector of the rotor. Measured levels of turbulence in the vortex core are used to predict the broadband noise levels with some success. However, more detailed information on the turbulence spectrum and the trajectory of the shed vortices is required before more accurate noise predictions can be made.

Glegg, Stewart A. L.↗

NASA NTRS1983

Rotor-vortex interaction noise

A theoretical and experimental study was conducted to develop a validated first principles analysis for predicting noise generated by helicopter main-rotor shed vortices interacting with the tail rotor. The generalized prediction procedure requires a knowledge of the incident vortex velocity field, rotor geometry, and rotor operating conditions. The analysis includes compressibility effects, chordwise and spanwise noncompactness, and treats oblique intersections with the blade planform. Assessment of the theory involved conducting a model rotor experiment which isolated the blade-vortex interaction noise from other rotor noise mechanisms. An isolated tip vortex, generated by an upstream semispan airfoil, was convected into the model tail rotor. Acoustic spectra, pressure signatures, and directivity were measured. Since assessment of the acoustic prediction required a knowledge of the vortex properties, blade-vortes intersection angle, intersection station, vortex stength, and vortex core radius were documented. Ingestion of the vortex by the rotor was experimentally observed to generate harmonic noise and impulsive waveforms.

Schlinker, R. H.↗

NASA NTRS1983

Rotor-vortex interaction noise

A theoretical and experimental study is carried out to develop a first-principles analysis for predicting noise generated by helicopter main-rotor shed vortices interacting with the tail rotor. The generalized prediction procedure calls for a knowledge of the incident vortex velocity field, the rotor geometry, and the rotor operating conditions. The analysis encompasses compressibility effects and chordwise and spanwise noncompactness and treats oblique intersections with the blade planform. In assessing the theory, a model rotor experiment which isolates the blade-vortex interaction noise from other rotor noise mechanisms is conducted. An isolated tip vortex, generated by an upstream semispan airfoil, is convected into the model rotor, and acoustic spectra, pressure signatures, and directivity are measured. Ingestion of the vortex by the rotor is experimentally observed to generate harmonic noise and impulsive waveforms.

Schlinker, R. H.↗

NASA NTRS1977

Fan inlet disturbances and their effect on static acoustic data

There is evidence that measurements of fan-rotor inlet noise taken during static test situations are at variance with aircraft engine flight data. In particular, static tests generally yield a significantly higher tone at blade passing frequency than that measured during flight. An experimental program was carried out to investigate this discrepancy. Inlet ground vortices and large-scale inlet turbulence were generated intentionally in an anechoic test chamber. Far-field acoustic measurements and inlet flow field hot film mappings of a fan rotor were then carried out to study the influence of such inlet disturbances at a static test facility. Experimental results indicate that the acoustic effect of such disturbances appears to be less severe for supersonic than for subsonic tip speeds. Further, a reverse flow that occurs on the exterior cowl in static test facilities appears to be an additional prime candidate for creating inlet disturbances and causing variance between flight and static acoustic data.

Bekofske, K. L.↗

NASA NTRS1992

Forward rotor vortex effects on counter rotating propeller noise

Three configurations of a model counter rotating propeller manipulate the blade tip flow by: placing the CRP at angle of attack, installing shrouds, and turning the upstream blades to provide forward sweep. Flow visualization and flow measurements with thermal anemometry show no evidence of a tip vortex; however, a leading edge vortex was detected on aft swept blades. The modifications served to alter the strength and/or path of the leading edge vortex. The vortical flow is eliminated by forward sweep on the upstream propeller blades. Far field acoustic data from each test indicate only small influences on the level and directivity of the BPFs. The interaction tone at the sum of the two BPF's was significantly altered in a consistent manner. As the vortex system varied, the interaction tone was affected: far field noise levels in the forward quandrant increased and the characteristic noise minimum near the plane of rotation became less pronounced and in some cases were eliminated. If the forward propeller leading edge vortex system does not impact the rear propeller in the standard manner, a net increase in the primary interaction tone occurs for the model tested. If the leading edge vortex is removed, the interaction tone increases.

Laur, Michele↗

NASA NTRS1991

Three-dimensional analysis of the Pratt and Whitney alternate design SSME fuel turbine

The three dimensional viscous time-mean flow in the Pratt and Whitney alternate design space shuttle main engine fuel turbine is simulated using the average passage Navier-Stokes equations. The migration of secondary flows generated by upstream blade rows and their effect on the performance of downstream blade rows is studied. The present simulation confirms that the flow in this two stage turbine is highly three dimensional and dominated by the tip leakage flow. The tip leakage vortex generated by the first blade persists through the second blade and adversely affects its performance. The greatest mixing of the inlet total temperature distortion occurs in the second vane and is due to the large leakage vortex generated by the upstream rotor. It is assumed that the predominant spanwise mixing mechanism in this low aspect ratio turbine is the radial transport due to the deterministically unsteady vortical flow generated by upstream blade rows. A by-product of the analysis is accurate pressure and heat loads for all blade rows under the influence of neighboring blade rows. These aero loads are useful for advanced structural analysis of the vanes and blades.

Kirtley, K. R.↗

NASA NTRSDate not supplied

LES-Based Analysis of Film Cooling in a High-Pressure Turbine Rotor Including Effects of Purge and Tip Flow

Accurate modeling of coolant airflows, which form protective films over turbine blades, is essential for designing fuel-efficient and environmentally sustainable gas turbine engines. Excessive coolant reduces thermal efficiency, while insufficient coolant leads to blade overheating and causes damage. Therefore, precise prediction of flow field interactions with cooling air is critical for optimizing turbine performance. This study numerically investigates the cooling effectiveness of purge and film cooling flows within a high-pressure turbine (HPT) rotor using Large Eddy Simulation (LES). The study utilizes NASA Glenn Research Center’s Glenn-HT solver. The simulation models ethe conditions of the Penn State University START rotating rig. A high-fidelity structured mesh comprising up to 800 million cells is employed to resolve high-Reynolds number flow (Re ≈ 350,000) and to capture intricate secondary flow structures, including tip leakage and purge-induced vortices. Film cooling effectiveness computations are highly sensitive to boundary conditions at the cooling holes and to grid resolution. Even with well-resolved grids and included plena, strong mixing challenges traditional eddy viscosity models. To address this, a simplified configuration is simulated: a truncated row of shaped holes on the suction side near the leading edge and a row on the pressure side, both fed from internal plena while the purge slot and tip clearance are also modeled. Two isothermal LES cases are conducted at two distinct wall temperatures, which yield the adiabatic wall temperature and the heat transfer coefficient. The definition and means of computation of the effectiveness is discussed in this paper. The simulations reveal detailed three-dimensional unsteady flow features, including coherent vortical structures and secondary flows originating from the purge cavity. Film cooling effectiveness and Nusselt number distributions are presented for both the blade surface and tip, highlighting regions of elevated heat transfer and complex thermal behavior. These findings underscore the importance of high-resolution LES and realistic boundary conditions in capturing the dynamics of purge and film cooling, offering valuable insights for improving turbine blade design and thermal management strategies.

Secondary Flows↗

NASA NTRSDate not supplied

A GPU-based Approach for Turbomachinery Application

Gas Turbine↗

NASA NTRS1987

Prediction of tip vortex self-induced motion parameters in terms of rotor blade loading

The self-induced motion of curved vortex filaments at the tip of a helicopter rotor blade is investigated analytically. The derivation of a method for inviscid roll-up (IRU) is presented in detail, with attention to the cutoff-distance reformulation of the problem, a control-volume analysis of the self-induction properties of the vortex core, the adaptation of the roll-up method of Betz (1932) to IRU in three dimensions, the treatment of linear and elliptic loading, vortices with turbulent central cores, and the effect of cutoff distance on helicopter free-wake computations. Numerical results showing the significant effect of core properties on the geometry of a rotor wake in hover are presented in graphs and briefly characterized.

Bliss, Donald B.↗

NASA NTRS1977

Effect of inlet disturbances on fan inlet noise during a static test

Measurements of fan rotor inlet noise taken during static test situations are at variance with aircraft engine flight data. In particular, static tests generally yield a significantly higher tone at blade passage frequency than that measured during flight. To explain this discrepancy, the extent of the influence of inlet ground vortices and large-scale inlet turbulence on the forward-radiated fan noise measured at a static test facility was investigated. While such inlet disturbances were generated intentionally in an anechoic test chamber, far-field acoustic measurements and inlet flow-field hot-film mappings of a fan rotor were obtained. Experimental results indicate that the acoustic effect of such disturbances appears to be less severe for supersonic than for subsonic tip speeds. Further, a reverse flow that occurs on the exterior cowl in static test facilities appears to be an additional prime candidate for creating inlet disturbances and causing variance between flight and static acoustic data.

Bekofske, K. L.↗

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.↗