Supersonic Fan Research Articles

Transonic Relief in Fans and Compressors

Abstract Every supersonic fan or compressor blade row has a streamtube, the “sonic streamtube,” which operates with a blade relative inlet Mach number of one. A key parameter in the design of the “sonic streamtube” is the area ratio between the blade throat area and the upstream passage area, Athroat/Ainlet. In this article, it is shown that one unique value exists for this area ratio. If the area ratio differs, even slightly, from this unique value, then the blade either chokes or has its suction surface boundary layer separated due to a strong shock. Therefore, it is surprising that in practice designers have relatively little problem designing blade sections with an inlet relative Mach number close to unity. This article shows that this occurs due to a physical mechanism known as “transonic relief.” If a designer makes a mistake and designs a blade with a “sonic streamtube,” which has the wrong area ratio, then “transonic relief” will self-adjust the spanwise streamtube height automatically moving it toward the unique optimal area ratio, correcting for the designer’s error. Furthermore, as the blade incidence changes, the spanwise streamtube height self-adjusts, moving the area ratio toward its unique optimal value, effectively controlling the blade’s incidence range. Without “transonic relief,” supersonic and transonic fan and compressor design would be impossible. This article develops a simple model that allows “transonic relief” to be decoupled from other mechanisms and to be systematically studied. The physical mechanism on which it is based is thus determined and a universal relationship for core compressor preliminary blade design is presented. Finally, its implications in relaxing manufacturing tolerances and in the design of future distortion tolerant blades are discussed.

Cyclostationary Analysis of Fan Noise Influenced by an Inflow Control Device

Acoustic data measured in a fan test facility with and without an inflow control device (ICD) over the inlet were analyzed using the cyclostationary analysis technique, which separates the total power into a rotor-locked and a fluctuating component. Measurements were taken with an inlet mode detection ring upstream of the rotor and with far-field microphones for configurations with subsonic and with supersonic fan blade tip speeds. The ICD reduces the fluctuating component and stabilizes the rotor-locked component of the frequency spectrum at the blade passing frequency and its harmonics in the way that the mode amplitudes of the Tyler–Sofrin modes increase, whereas neighboring mode orders are suppressed. The mode distribution and intermode coherence in the inlet are also changed, which changes the far-field directivity when an ICD is installed. It is demonstrated that cyclostationary analysis is a powerful tool to investigate the acoustic effect of an ICD on turbofan noise, revealing more details about the temporal and spatial structures of the generated sound field than a classical spectral analysis.

Nonlinear Propagation of Supersonic Fan Tones in Turbofan Intake Ducts

Supersonic fan tones are a critical issue for large fans; it is well known that these tones are a key noise source at high-power operating conditions for modern turbofan aeroengines. Nonlinear propagation of the rotor-locked pressure field generated in a turbofan intake duct is calculated by implementing a nonlinear weak-shock theory numerically via a combined time-/frequency-domain algorithm. The aim of the prediction method is to model the nonlinear attenuation and liner absorption of the rotor-locked pressure field within the intake duct. In this hybrid method, the time-domain approach provides a robust and accurate prediction of the nonlinear attenuation, whereas the frequency domain is required to represent the liner attenuation. The focus of this work is to compare the time-/frequency-domain method with time-domain or frequency-domain methods that have been developed previously to calculate the nonlinear propagation of the rotor-locked pressure field in either a rigid or lined intake duct, respectively. The capability of the time-/frequency-domain method to accurately calculate the nonlinear propagation of the rotor-locked pressure field, which can provide an engineering method for buzzsaw noise prediction, is demonstrated.

Near-field sound radiation of fan tones from an installed turbofan aero-engine.

The development of a distributed source model to predict fan tone noise levels of an installed turbofan aero-engine is reported. The key objective is to examine a canonical problem: how to predict the pressure field due to a distributed source located near an infinite, rigid cylinder. This canonical problem is a simple representation of an installed turbofan, where the distributed source is based on the pressure pattern generated by a spinning duct mode, and the rigid cylinder represents an aircraft fuselage. The radiation of fan tones can be modelled in terms of spinning modes. In this analysis, based on duct modes, theoretical expressions for the near-field acoustic pressures on the cylinder, or at the same locations without the cylinder, have been formulated. Simulations of the near-field acoustic pressures are compared against measurements obtained from a fan rig test. Also, the installation effect is quantified by calculating the difference in the sound pressure levels with and without the adjacent cylindrical fuselage. Results are shown for the blade passing frequency fan tone radiated at a supersonic fan operating condition.

“Buzz-saw” noise: Prediction of the rotor-alone pressure field

Public expectations of lower environmental noise levels, and increasingly stringent legislative limits on aircraft noise, result in noise being a critical technical issue in the development of jet engines. Noise at take-off, when the engines are at high-power operating conditions, is a key reference level for engine noise certification. “Buzz-saw” noise is the dominant fan tone noise from modern high-bypass-ratio turbofan aircraft engines during take-off. Rotor-alone tones are the key component of buzz-saw noise. The rotor-alone pressure field is cut-off at subsonic fan tip speeds; buzz-saw noise is associated with supersonic fan tip speeds, or equivalently, high power engine operating conditions. A recent series of papers has described new work concerning the prediction of buzz-saw noise. The prediction method is based on modelling the nonlinear propagation of one-dimensional sawtooth waveforms. A sawtooth waveform is a simplified representation of the rotor-alone pressure field. Previous validation of the prediction method focussed entirely on reproducing the spectral characteristics of buzz-saw noise; this was dictated at that time by the availability of spectral data only for comparison between measurement and prediction. In this paper, full validation of the method by comparing measurement and prediction of the rotor-alone pressure field is published for the first time. It is shown that results from the modelling based on a one-dimensional sawtooth waveform capture the essential features of the rotor-alone pressure field as it propagates upstream inside a hard-walled inlet duct. This verifies that predictions of the buzz-saw noise spectrum, which are in good agreement with the measured data, are based on a model which reproduces the key physics of the noise generation process. Validation results for the rotor-alone pressure field in an acoustically lined inlet duct are also shown. Comparisons of the measured and predicted rotor-alone pressure field are more difficult to interpret because the acoustic lining significantly modifies the sawtooth waveform, but there remains good agreement with the measured spectral data. The buzz-saw noise prediction code used to generate the simulations in this paper has been used by the Rolls–Royce Noise Department since 2004.

Shock wave structures ahead of nonuniform fan cascades

We study shock wave structures (SWS), consisting of shock waves and expansion waves between them, that occur in supersonic flow past nonuniform fan cascades when the velocity component normal to their front (“axial” component) is subsonic. The cascade nonuniformity is due to the scatter in the setting angles of identical blades, either sharp or blunt. A result of the uniformity is the generation of combined noise, whose frequencies are much smaller than the fundamental frequency of the uniform cascade, and slower nonlinear SWS attenuation. The accurate and fast “simple wave method” and “nonlinear acoustics approximation”, together with numerical algorithms for integrating Euler equations on overlapping grids (in calculating flow past blunt edges) and on SWS-adapted grids, are applied to determine the “guiding” action of nonuniform cascades and to describe the SWS evolution. The application of the Fourier analysis gives the sound field spectrum. The use of blades with rectilinear initial regions of the “backs” for reducing supersonic fan blade noise is efficient only at small (less than 0.25°) scatter in the setting angles. The shock wave structures attenuate more rapidly ahead of nonuniform cascades composed of blunt blades than ahead of those with sharp blades. For uniform cascades the blade bluntness effect is not large.

“Buzz-saw” noise: A comparison of modal measurements with an improved prediction method

“Buzz-saw” noise is radiated from a turbofan inlet duct when the fan tip speed is supersonic. In a recent article the effect of an acoustic liner on buzz-saw noise has been examined. Spectral measurements in a rigid and an acoustically lined inlet duct have been compared. Also these measurements have been utilized to assess a buzz-saw noise prediction method. The prediction method is based on a one-dimensional nonlinear propagation model. Sound absorption by an acoustic lining can be included in the model. In this article, the buzz-saw noise prediction method is improved by the inclusion in the modelling of the effect of a boundary layer on absorption of sound in a lined duct. Also, modal measurements from a circumferential microphone array have been examined. These show that the principal source of buzz-saw noise is not always the rotor-alone pressure field. Non-rotor-alone scattered tones can be a significant source of buzz-saw noise at low supersonic fan speeds. The numerical simulations, which only predict the rotor-alone tones, have been re-evaluated in light of these new modal measurements.

Acoustic scattering by an axially-segmented turbofan inlet duct liner at supersonic fan speeds

Fan noise is one of the principal noise sources in turbofan aero-engines. At supersonic fan speeds, fan tones are generated by the “rotor-alone” pressure field. In general, these tones can be well absorbed by an inlet duct acoustic liner, except at high supersonic fan speeds when the rotor-alone pressure field is well cut-on. In this article an axially segmented liner is proposed, which is predicted to improve the attenuation of tones at high supersonic fan speeds. The analysis is based on locally reacting cavity liners. The axially segmented liner is axisymmetric and consists of two circular sections of different linings joined together. The optimum design consists of two linings with the same face-sheet resistance, but with different cavity depths. The depth of the liner adjacent to the fan is very thin. This means that where the two liners are joined there is a wall impedance discontinuity that can cause acoustic scattering. Fan tones can be modelled in terms of spinning modes in a uniform circular-section duct. The liner is axisymmetric, so modal scattering will be only between different radial modes. The optimum design minimizes the acoustic energy scattered into the first radial mode. This improves the attenuation of fan tones at high supersonic fan speeds, because acoustic energy is scattered into high radial mode orders, which are better absorbed by the lining.

Acoustic scattering by a spliced turbofan inlet duct liner at supersonic fan speeds

Fan noise is one of the principal noise sources generated by a turbofan aero-engine. At supersonic fan speeds, fan tones are generated by the “rotor-alone” pressure field. In general, these tones can be well absorbed by an inlet duct acoustic liner, apart from at high supersonic fan speeds. However, in practice inlet duct liners contain acoustically hard longitudinal splices which cause scattering. This leads to acoustic energy being scattered into a range of different azimuthal mode orders, similar to the modal content resulting from rotor–stator interactions. The effectiveness of an inlet duct lining is reduced because acoustic energy is scattered into modes that are poorly absorbed by the liner. In this article, the effect of this acoustic scattering is examined by three-dimensional finite-element simulations of sound transmission in a turbofan inlet duct. Results include predictions of the sound power transmission loss with different splice widths, and at different supersonic fan speeds. These results demonstrate how acoustic scattering by liner splices can adversely affect fan tone noise levels at low supersonic fan speeds, but have little adverse affect on noise levels at high supersonic fan speeds. The potential noise benefit that could be achieved by manufacturing thinner splices is also examined.

On the prediction of “buzz-saw” noise in acoustically lined aero-engine inlet ducts

Aero-engines operating with supersonic fan tip speeds generate an acoustic signature containing energy spread over a range of harmonics of the engine shaft rotation frequency. These harmonics are commonly known as the “buzz-saw” tones. The pressure signature attached to a supersonic ducted fan will be a sawtooth waveform. The non-linear propagation of a high-amplitude irregular sawtooth upstream inside the inlet duct redistributes the energy amongst the buzz-saw tones. In most modern aero-engines the inlet duct contains an acoustic lining, whose properties will be dependent on the mode number and frequency of the sound, and the speed of the oncoming flow. Such effects may not easily be incorporated into a time-domain approach; hence the non-linear propagation of an irregular sawtooth is calculated in the frequency domain, which enables liner damping to be included in the numerical model. Results are presented comparing noise predictions in hard-walled and acoustically lined inlet ducts. These show the effect of an acoustic liner on the buzz-saw tones. These predictions compare favourably with previous experimental measurements of liner insertion loss (at blade passing frequency), and provide a plausible explanation for the observed reduction in this insertion loss at high fan operating speeds.

An in-flight study of cabin buzz-saw noise

This paper examines the characteristics of multiple-pure-tone noise generated by high-speed turbofans under conditions of supersonic fan tip speeds, especially as it is observed in an aircraft passenger cabin. This phenomenon, also known as buzz-saw noise, is an important noise source in commercial airplane passenger cabins and has proved to be difficult to treat with sound-absorbing materials. Recent flight test experiments by The Boeing Company have demonstrated extraordinary success in suppressing cabin buzz-saw noise by strategic placement of engine inlet acoustic linings. The observed behavior is explained by a propagation and radiation model, which is validated by in-flight measurements made with a phased microphone array mounted on the fuselage skin of a Boeing 777. A structural acoustic model is also offered to explain the different transmission characteristics of buzz-saw noise and turbulent boundary layer excitation. Correlation length scales measured on the fuselage surface for these two noise sources are key inputs to the structural model.

Prévision du bruit de raie émis par un rotor : application à l'aéronautique

Prediction of rotor tone noise is a major issue both in industry andin aeronautics for reducing sound radiation. Two frequency methods are described, the first applies to open rotors and the second solves the wave equation inside a cylindrical duct. It is shown that the main terms are the same in both cases. This is particularly useful for studying supersonic fans of turbofans because computations of isolated propellers are then valid. Emphasis is put on the differences with the directivities radiated by conventional subsonic propellers

Interrelationship of a rotating stall and vibrations of a blade ring

As a result of summarization of experimental data obtained in investigation of the parameters of a rotating stall and also simultaneously of the parameters of the vibration condition of blade rings of single-stage supersonic fans and of multistage axial compressors, the interrelationship of the two phenomena, a rotating stall and the vibrations of the blade ring accompanying it, has been observed. The derivation of the relationships between the parameters of these phenomena is presented.

Annular Cascade Study of Low Back-Pressure Supersonic Fan Blade Flutter

Low back-pressure supersonic fan blade flutter in the torsional mode was examined using a controlled-oscillating annular cascade test facility. Precise data of unsteady aerodynamic forces generated by shock wave movement, due to blade oscillation, and the previously measured data of chordwise distributions of unsteady aerodynamic forces acting on an oscillating blade, were joined and, then, the nature of cascade flutter was evaluated. These unsteady aerodynamic forces were measured by direct and indirect pressure measuring methods. Our experiments covered a range of reduced frequencies based on a semichord from 0.0375 to 0.547, six interblade phase angles, and inlet flow velocities from subsonic to supersonic flow. The occurrence of unstalled cascade flutter in relation to reduced frequency, interblade phase angle, and inlet flow velocity was clarified, including the role of unsteady aerodynamic blade surface forces on flutter. Reduced frequency of the flutter boundary increased greatly when the blade suction surface flow became transonic flow. Interblade phase angles that caused flutter were in the range from 40 to 160 deg for flow fields ranging from high subsonic to supersonic. Shock wave movement due to blade oscillation generated markedly large unsteady aerodynamic forces which stimulated blade oscillation.

Flutter of a Fan Blade in Supersonic Axial Flow

An application of a simple aeroelastic model to an advanced supersonic axial flow fan is presented. Lane’s cascade theory is used to determine the unsteady aerodynamic loads. Parametric studies are performed to determine the effects of mode coupling, Mach number, damping, pitching axis location, solidity, stagger angle, and mistuning. The results show that supersonic axial flow fan and compressor blades are susceptible to a strong torsional mode flutter having critical reduced velocities that can be less than one.

Mass Balancing of Hollow Fan Blades

This paper uses a typical section model to investigate analytically the effect of mass balancing as applied to hollow, supersonic fan blades. A procedure to determine the best configuration of an internal balancing mass to provide flutter alleviation is developed. This procedure is applied to a typical supersonic shroudless fan blade which is unstable both in the solid configuration and when it is hollow with no balancing mass. The addition of an optimized balancing mass is shown to stabilize the blade at the design condition.

Propagation of Supersonic Fan Noise in a Nonuniform Medium

Propagation of Supersonic Fan Noise in a Nonuniform Medium

Characteristic surfaces in one-dimensional unsteady and two-dimensional steady expansion fans

The paper deals with the construction of general characteristic surfaces in one-dimensional unsteady and two-dimensional steady supersonic expansion fans. Of special interest are the properties of domains of dependence and influence for arbitrary points within such an expansion. It is shown that the effect of the rotation of the wave-normals along bicharacteristics is of importance for their calculation.

Wave forms for a supersonic rotor

The geometry of the wave forms of a supersonic rotor is described, based on a ray theory. The waves formed by a non-translating supersonic rotor are found to be limited by a hyperboloid, causing a cone of silence centred on the axis where only radiation from conventional noise sources may be observed. The wave region is limited by cusps which imply local focusing of the radiated energy. These fundamental features apply also to cases with hub motion but their geometry becomes complex. The approach is one in which non-linear effects necessarily are ignored, but the results appear to have value in delineating potential regions of shock formation in the free field. Applications include the shock radiation from supersonic fans in jet engines and the advancing blade slap problem in helicopters.