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At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS2010

Near Stall Flow Analysis in the Transonic Fan of the RTA Propulsion System

Turbine-based propulsion systems for access to space have been investigated at NASA Glenn Research center. A ground demonstrator engine for validation testing has been developed as a part of the program. The demonstrator, the Revolutionary Turbine Accelerator (RTA-1), is a variable cycle turbofan ramjet designed to transition from an augmented turbofan to a ramjet that produces the thrust required to accelerate the vehicle to Mach 4. The RTA-1 is designed to accommodate a large variation in bypass ratio from sea level static to Mach 4 flight condition. A key component of this engine is a new fan stage that accommodates these large variations in bypass ratio and flow ranges. In the present study, unsteady flow behavior in the fan of the RTA-1 is studied in detail with large eddy simulation (LES) and the numerical results are compared with measured data. During the experimental study of the fan stage, humming sound was detected at 100 % speed near stall operation. The main purpose of the study is to investigate details of the unsteady flow behavior at near stall operation and to identify a possible cause of the hum. The large eddy simulation of the current flow field reproduces main features of the measured flow very well. The LES simulation indicates that non-synchronous flow instability develops as the fan operates toward the stall limit. The FFT analysis of the calculated wall pressure shows that the rotating flow instability has the characteristic frequency that is about 50% of the blade passing frequency.

Hah, Chunill↗
DOE OSTI2022

An algorithmic approach to predicting mechanical draft cooling tower fan speeds from infrasound signals

Mechanical draft cooling towers (MDCTs) serve a critical heat management role in a variety of industries. For nuclear reactors in particular, the consistent, predictable operation of MDCTs is required to avoid damage to infrastructure and reduce the potential for catastrophic failure. Accurate, reliable measurement of MDCT fan speed is therefore an important maintenance and safety requirement. To that end, we have developed an algorithm for automatically predicting the rotational speeds of multiple, simultaneously operating fan rotors using contactless, infrasound measurements. The algorithm is based on identifying the blade passing frequencies (BPFs), their harmonics, as well as the motor frequencies (MFs) for each fan in operation. Using the algorithm, these frequencies can be automatically identified in the acoustic waveform’s short-time Fourier transform spectrogram. Attribution is aided by a set of filters that rely on the unique spectral and temporal characteristics of fan operation, as well as the intrinsic frequency ratios of the BPF harmonics and the BPF/MF signals. The algorithm was tested against infrasound data acquired from infrasound sensors deployed at two research reactors: the Advanced Test Reactor (ATR) located at Idaho National Laboratory (INL) and the High Flux Isotope Reactor (HFIR) located at Oak Ridge National Laboratory (ORNL). After manually identifying the MDCT gearbox ratio, the algorithm was able to quickly yield fan speeds at both reactors in good agreement with ground truth. Ultimately, this work demonstrates the ease by which MDCT fans may be monitored in order to optimize operational conditions and avoid infrastructure damage.

42 ENGINEERING↗
NASA NTRS2003

Noise Generation by Fans with Supersonic Tip Speeds

Fan noise continues to be a significant issue for commercial aircraft engines and there still exists a requirement for improved understanding of the fundamental issues associated with fan noise source mechanisms. At the present time, most of the prediction methods identify the dominant acoustic sources to be associated with the stator vanes or blade trailing edges which are downstream of the fan face. However recent studies have shown that acoustic waves are significantly attenuated as they propagate upstream through a rotor, and if the appropriate corrections are applied, sound radiation from the engine inlet is significantly underpredicted. The prediction models can only be applied to fans with subsonic tip speeds. In contrast, most aircraft engines have fan tip speeds which are transonic and this implies an even higher attenuation for upstream propagating acoustic waves. Consequently understanding how sound propagates upstream through the fan is an important, and not well understood phenomena. The objective of this study is to provide improved insight into the upstream propagation effects through a rotor which are relevant to full scale engines. The focus of this study is on broadband fan noise generated by boundary layer turbulence interacting with the trailing edges of the fan blades. If this source mechanism is important upstream of the fan, the sound must propagate upstream through a transonic non uniform flow which includes large gradients and non linearities. Developing acoustic propagation models in this type of flow is challenging and currently limited to low frequency applications, where the frequency is of the same order as the blade passing frequency of the fan. For trailing edge noise, much higher frequencies are relevant and so a suitable approach needs to be developed, which is not limited by an unacceptably large computational effort. In this study we are in the process of developing a computational method which applies for the high frequencies of interest, and allows for any type of flow field associated with the fan. In this progress report the approach to be used and the basic equations will be presented. Some initial results will be given, but these are preliminary and need further verification.

Glegg, Stewart↗
NASA NTRS1975

Study of noise sources in a subsonic fan using measured blade pressures and acoustic theory

Sources of noise in a 1.4 m (4.6 ft) diameter subsonic tip speed propulsive fan running statically outdoors are studied using a combination of techniques. Signals measured with pressure transducers on a rotor blade are plotted in a format showing the space-time history of inlet distortion. Study of these plots visually and with statistical correlation analysis confirms that the inlet flow contains long, thin eddies of turbulence. Turbulence generated in the boundary layer of the shroud upstream of the rotor tips was not found to be an important noise source. Fan noise is diagnosed by computing narrowband spectra of rotor and stator sound power and comparing these with measured sound power spectra. Rotor noise is computed from spectra of the measured blade pressures and stator noise is computed using the author's stator noise theory. It is concluded that the rotor and stator sources contribute about equally at frequencies in the vicinity of the first three harmonics of blade passing frequency. At higher frequencies, the stator contribution diminishes rapidly and the rotor/inlet turbulence mechanism dominates. Two parametric studies are performed by using the rotor noise calculation procedure which was correlated with test. In the first study, the effects on noise spectrum and directivity are calculated for changes in turbulence properties, rotational Mach number, number of blades, and stagger angle. In the second study the influences of design tip speed and blade number on noise are evaluated.

Hanson, D. B.↗
NASA NTRS1987

High speed propeller performance and noise predictions at takeoff/landing conditions

The performance and noise of a high speed SR-7A model propeller under takeoff/landing conditions are considered. The blade loading distributions are obtained by solving the three-dimensional Euler equations and the sound pressure levels are computed using a time domain approach. At the nominal takeoff operating point, the blade sections near the hub are lightly or negatively loaded. The chordwise loading distributions are distinctly different from those of cruise conditions. The noise of the SR-7A model propeller at takeoff is dominated by the loading noise, similar to that at cruise conditions. The waveforms of the acoustic pressure signature are nearly sinusoidal in the plane of the propeller. The computed directivity of the blade passing frequency tone agrees fairly well with the data at nominal takeoff blade angle.

Nallasamy, M.↗
NASA NTRS1988

High speed propeller performance and noise predictions at takeoff/landing conditions

The performance and noise of a high speed SR-7A model propeller under takeoff/landing conditions are considered. The blade loading distributions are obtained by solving the three-dimensional Euler equations and the sound pressure levels are computed using a time domain approach. At the nominal takeoff operating point, the blade sections near the hub are lightly or negatively loaded. The chordwise loading distributions are distinctly different from those of cruise conditions. The noise of the SR-7A model propeller at takeoff is dominated by the loading noise, similar to that at cruise conditions. The waveforms of the acoustic pressure signature are nearly sinusoidal in the plane of the propeller. The computed directivity of the blade passing frequency tone agrees fairly well with the data at nominal takeoff blade angle.

Nallasamy, M.↗
NASA NTRS1972

The theoretical and experimental investigations on multiple pure tone noise.

An investigation on multiple pure tone (MPT) noise is described. The model fan is operated, in Freon 12, in a closed loop acoustical facility. It has been found experimentally that the rotor relative Mach number and incidence angle are parameters important to the evolution of the MPT sound and the inlet duct length has a significant influence on the MPT sound emission. The experimental results are compared with a previous analysis. From the known blade nonuniformities of the model fan, the MPT distributions are computed. The analysis correctly predicts the frequency of the dominant multiple pure tone and sound pressures of the blade passing frequency noise and the dominant MPT that are in the same range as the experimental values.

Kantola, R. A.↗
NASA NTRS1976

Noise suppression by an acoustically treated three-ring inlet on a TF-34 engine

Acoustic performance tests were conducted with a three-ring inlet noise suppressor designed for a TF-34 engine. For all tests the aft noise sources were highly suppressed. The measured inlet suppression was large, reaching levels greater than 30 db at the peak. Comparisons of the data and the performance predictions were in reasonably good agreement. The frequency of peak attenuation was well predicted; the magnitude and spectral shape were reasonably well predicted. Agreement was best when the distribution of sound energy across the inlet was taken into account in the performance predictions. Tests in which the length of treatment was varied showed an orderly progression of attenuation with length; performance predictions for the different lengths also showed an orderly progression with length. At the highest speed of the engine, multiple pure tones were present throughout the spectrum in the source noise signature. These tones were effectively suppressed by the inlet liner, even at low frequencies, although the liner was designed to work best at the blade-passing frequency.

Minner, G. L.↗
NASA NTRS1982

Vibration analysis of three guyed tower designs for intermediate size wind turbines

Three guyed tower designs were analyzed for intermediate size wind turbines. The four lowest natural frequencies of vibration of the three towers concepts were estimated. A parametric study was performed on each tower to determine the effect of varying such tower properties as the inertia and stiffness of the tower and guys, the inertia values of the nacelle and rotor, and the rotational speed of the rotor. Only the two lowest frequencies were in a range where they could be excited by the rotor blade passing frequencies. There two frequencies could be tuned by varying the guy stiffness, the guy attachment point on the tower, the tower and mass stiffness, and the nacelle/rotor/power train masses.

Christie, R. J.↗
NASA NTRS1988

Investigation of the validity of Reynolds averaged turbulence models at the frequencies that occur in turbomachinery

Turbulent flows subjected to various kinds of unsteady disturbances were simulated using a large-eddy-simulation computer code for flow in a channel. The disturbances were: a normal velocity expressed as a traveling wave on one wall of the channel; staggered blowing and suction distributions on the opposite walls of the channel; and oscillations of the mean flow through the channel. The wall boundary conditions were designed to simulate the effects of wakes of a stator stage passing through a rotor channel in a turbine. The oscillating flow simulated the effects of a pressure pulse moving over the rotor blade boundary layer. The objective of the simulations was to provide better understanding of the effects of time-dependent disturbances on the turbulence of a boundary layer and of the underlying physical phenomena regarding the basic interaction between the turbulence and external disturbances of the type found in turbomachinery. Results showed that turbulence is sensitive to certain ranges of frequencies of disturbances. However, no direct connection was found between the frequency of imposed disturbances and characteristic burst frequency of turbulence. New insight into the nature of turbulence at high frequencies was found. The viscous phenomena near solid walls was found to be the dominant influence for high frequency perturbations. At high frequencies, the turbulence was found to be undisturbed, remaining the same as for the steady mean flow. A transition range exists between the high frequency range and the low, or quasi-steady, range in which the turbulence is not predictable by either quasi-steady models or the steady flow model. The limiting lowest frequency for use of the steady flow turbulence model is that for which the viscous Stokes layer based on the blade passing frequency is thicker than the laminar sublayer.

Kuhn, Gary D.↗
NASA NTRS2008

A Numerical Investigation of Turbine Noise Source Hierarchy and Its Acoustic Transmission Characteristics

Understanding the relative importance of the various turbine noise generation mechanisms and the characteristics of the turbine acoustic transmission loss are essential ingredients in developing robust reduced-order models for predicting the turbine noise signature. A computationally based investigation has been undertaken to help guide the development of a turbine noise prediction capability that does not rely on empiricism. The investigation relies on highly detailed numerical simulations of the unsteady flowfield inside a modern high-pressure turbine (HPT). The simulations are developed using TURBO, which is an unsteady Reynolds-averaged Navier-Stokes (URANS) code capable of multi-stage simulations. The purpose of this study is twofold. First, to determine an estimate of the relative importance of the contributions to the coherent part of the acoustic signature of a turbine from the three potential sources of turbine noise generation, namely, blade-row viscous interaction, potential field interaction, and entropic source associated with the interaction of the blade rows with the temperature nonuniformities caused by the incomplete mixing of the hot fluid and the cooling flow. Second, to develop an understanding of the turbine acoustic transmission characteristics and to assess the applicability of existing empirical and analytical transmission loss models to realistic geometries and flow conditions for modern turbine designs. The investigation so far has concentrated on two simulations: (1) a single-stage HPT and (2) a two-stage HPT and the associated inter-turbine duct/strut segment. The simulations are designed to resolve up to the second harmonic of the blade passing frequency tone in accordance with accepted rules for second order solvers like TURBO. The calculations include blade and vane cooling flows and a radial profile of pressure and temperature at the turbine inlet. The calculation can be modified later to include the combustor pattern factor at the turbine inlet to include that contribution to turbine noise. We shall present preliminary analysis of the results obtained so far in order to assess the validity of such an approach and to seek feedback on improving the approach. This work addresses both Area 1 (Turbine Tone Noise) and Area 5 (Influence of the Turbine on Combustor Noise) topics.

VanZante, Dale↗
NASA NTRS1999

Acoustic Reflection and Transmission of 2-Dimensional Rotors and Stators, Including Mode and Frequency Scattering Effects

A reduced order modeling scheme has been developed for the unsteady acoustic and vortical coupling between blade rows of a turbomachine. The essential behavior of the system is governed by modal scattering coefficients (i.e., reflection and transmission coefficients) of the rotor, stator, inlet and nozzle, which are calculated as if they were connected to non-reflecting ducts. The objective of this report is to identify fundamental behavior of these scattering coefficients for a better understanding of the role of blade row reflection and transmission in noise generation. A 2D flat plate unsteady cascade model is used for the analysis with the expectation that the general behavior presented herein will carry over to models that include more realistic flow and geometry. It is shown that stators scatter input waves into many modes at the same frequency whereas rotors scatter on frequency, or harmonic order. Important cases are shown here the rotor reflection coefficient is greater than unity; a mode at blade passing frequency (BPF) traveling from the stator with unit sound power is reflected by the rotor with more than unit power at 2xBPF and 3xBPE Analysis is presented to explain this unexpected phenomenon. Scattering curves are presented in a format chosen for design use and for physical interpretation. To aid in interpretation of the curves, formulas are derived for special condition where waveforms are parallel to perpendicular to the rotor.

Hanson, Donald B.↗
NASA NTRS1990

Potential reduction of en route noise from an advanced turboprop aircraft

When the en route noise of a representative aircraft powered by an eight-blade SR-7 propeller was previously calculated, the noise level was cited as a possible concern associated with the acceptance of advanced turboprop aircraft. Some potential methods for reducing the en route noise were then investigated and are reported. Source noise reductions from increasing the blade number and from operating at higher rotative speed to reach a local minimum noise point were investigated. Greater atmospheric attenuations for higher blade passing frequencies were also indicated. Potential en route noise reductions from these methods were calculated as 9.5 dB (6.5 dB(A)) for a 10-blade redesigned propeller and 15.5 dB (11 dB(A)) for a 12-blade redesigned propeller.

Dittmar, James H.↗
NASA NTRS2013

Limitations of Phased Array Beamforming in Open Rotor Noise Source Imaging

Phased array beamforming results of the F31/A31 historical baseline counter-rotating open rotor blade set were investigated for measurement data taken on the NASA Counter-Rotating Open Rotor Propulsion Rig in the 9- by 15-Foot Low-Speed Wind Tunnel of NASA Glenn Research Center as well as data produced using the LINPROP open rotor tone noise code. The planar microphone array was positioned broadside and parallel to the axis of the open rotor, roughly 2.3 rotor diameters away. The results provide insight as to why the apparent noise sources of the blade passing frequency tones and interaction tones appear at their nominal Mach radii instead of at the actual noise sources, even if those locations are not on the blades. Contour maps corresponding to the sound fields produced by the radiating sound waves, taken from the simulations, are used to illustrate how the interaction patterns of circumferential spinning modes of rotating coherent noise sources interact with the phased array, often giving misleading results, as the apparent sources do not always show where the actual noise sources are located. This suggests that a more sophisticated source model would be required to accurately locate the sources of each tone. The results of this study also have implications with regard to the shielding of open rotor sources by airframe empennages.

Horvath, Csaba↗
NASA NTRS1976

Effect of loading and rotor wake characteristics on the acoustic field of stator blades

The generation of blade passing frequency noise due to rotor wakes interaction with the downstream stator is examined using a recent model for the unsteady aerodynamic forces developed by Goldstein and Atassi. The effects of blade geometry and loading are investigated. It is shown that at low reduced frequencies loading significantly reduces the noise level while at high frequencies loading substantially increases the sound level radiated. It is also shown that for the same loading a higher angle of attack generates more noise particularly at higher reduced frequencies. The noise level shows a quasi-periodic pattern with pronounced minima and maxima as the frequency increases. These conclusions are valid for different wake models.

Atassi, H.↗
NASA NTRS1987

Advanced turboprop wing installation effects measured by unsteady blade pressure and noise

A single rotation model propeller (SR-7A) was tested at simulated takeoff/approach conditions (Mach 0.2), in the NASA Lewis 9- by 15-Ft Anechoic Wind Tunnel. Both unsteady blade surface pressures and noise measurements were made for a tractor configuration with propeller/straight wing and propeller alone configurations. The angle between the wing chord and propeller axis (droop angle) was varied along with the wing angle of attack to determine the effects on noise and unsteady loading. A method was developed that uses unsteady blade pressure measurements to provide a quantitative indication of propeller inflow conditions, at least for a uniform (across the propeller disk) inflow angle. The wing installation caused a nearly uniform upwash at the propeller inlet as evidenced by the domination of the pressure spectra by the first shaft order. This inflow angle increased at a rate of almost 150 percent of that of the wing angle-of-attack for a propeller-wing spacing of 0.54 wing chords at a constant droop angle. The flyover noise, as measured by the maximum blade passing frequency level, correlates closely with the propeller inflow angle (approx. 0.6 dB per degree of inflow angle) for all droop angles and wing angles of attack tested, including the propeller alone data. Large changes in the unsteady pressure responses on the suction surface of the blade were observed as the advance ratio was varied. The presence of a leading edge vortex may explain this behavior since changes in the location of this vortex would change with loading (advance ratio).

Heidelberg, Laurence J.↗
NASA NTRS1987

Advanced turboprop wing installation effects measured by unsteady blade pressure and noise

A single rotation model propeller (SR-7A) was tested at simulated takeoff/approach conditions (Mach 0.2), in the NASA Lewis 9- by 15-Ft Anechoic Wind Tunnel. Both unsteady blade surface pressures and noise measurements were made for a tractor configuration with propeller/straight wing and propeller alone configurations. The angle between the wing chord and propeller axis (droop angle) was varied along with the wing angle of attack to determine the effects on noise and unsteady loading. A method was developed that uses unsteady blade pressure measurements to provide a quantitative indication of propeller inflow conditions, at least for a uniform (across the propeller disk) inflow angle. The wing installation caused a nearly uniform upwash at the propeller inlet as evidenced by the domination of the pressure spectra by the first shaft order. This inflow angle increased at a rate of almost 150 percent of that of the wing angle-of-attack for a propeller-wing spacing of 0.54 wing chords at a constant droop angle. The flyover noise, as measured by the maximum blade passing frequency level, correlates closely with the propeller inflow angle (approx. 0.6 dB per degree of inflow angle) for all droop angles and wing angles of attack tested, including the propeller alone data. Large changes in the unsteady pressure responses on the suction surface of the blade were observed as the advance ratio was varied. The presence of a leading edge vortex may explain this behavior since changes in the location of this vortex would change with loading (advance ratio).

Heidelberg, Laurence J.↗