Downstream Stator Vanes Research Articles

Forced Response Analysis of an Embedded Compressor Rotor Induced by Stator Disturbances and Rotor–Stator Interactions

Accurate predictions of the blade response in a multi-row compressor is one of the most important tasks within the design process of compressor blades. Some recent studies have shown that the decoupled method considering only the stator disturbances cannot obtain accurate results for cases with strong rotor–stator interactions, especially for the interaction between the rotor and downstream stator, and the coupled method with multi-row configurations is necessary. Factors determine what computational domains to model need to be clarified to find a balance between accuracy requirements and computational costs. To this end, this study conducted full-annulus unsteady calculations with decoupled and coupled configurations to investigate the forced response of an embedded compressor rotor induced by upstream and downstream stator disturbances and rotor–stator interactions, respectively. The results show that the upstream IGV disturbances were dominated by the wake, and the IGV and S1 potential fields had little effect on the R1 response. Meanwhile, the IGV–R1 interactions and S1–R1 interactions were dominated by one cut-on mode, respectively. The comparisons of the blade vibration amplitude and the unsteady pressure field calculated by decoupled and coupled methods revealed the mechanism of the forced response, namely, for the R1 response induced by upstream aerodynamic disturbances, the dominant excitation source was the IGV wake, and the blade vibration amplitude can be predicted by the decoupled method. In terms of the response induced by downstream disturbances, the cut-on S1-R1-interaction mode was dominant and the use of the decoupled method without considering its influence will lead to an inaccurate prediction. This study concluded that the formation process of rotor–stator interactions was the key factor that determines whether the decoupled method or coupled method should be used, and analogized a process independent of the downstream stator disturbance. The results can provide a preliminary configuration for accurate and efficient blade response predictions and explain the reason why including downstream stator vanes is very important.

Influence of Different Flow Solvers and Off-Design Conditions On the Determination of Fan-Rotor Wakes for Broadband Noise Prediction

Abstract The acoustic interaction of fan-rotor wakes with the downstream stator vanes is considered as an important noise source of an aircraft engine. The turbulence induced by the rotor generates a stochastic acoustic source that appears as broadband noise in the acoustic spectrum. During the preliminary design phase of an engine, established meanline and throughflow solvers usually do not resolve turbulence and associated unsteady flow parameters. But such solvers provide rotor pressure losses that can be used to estimate the mean and turbulent rotor wakes. A crucial step is the deduction of turbulence parameters from the mean wakes. A semi-empirical model for rotor-wake turbulence estimation is presented in this paper. The meanline method and the throughflow solver are compared to three-dimensional computational flow simulations investigating the capabilities of the different solvers to provide flow data for broadband wake interaction noise prediction. The methods are applied to a representative modern fan stage at a comprehensive number of operating points, comprising several speed lines from surge to choking conditions. Microphone measurements are consulted to assess the noise predictions. The evaluation confirms the applicability of the meanline and throughflow method in combination with the turbulence model for broadband noise estimation during the preliminary design phase. The underestimated turbulence in the tip region of the fan is found to be negligible even during off-design conditions.

Large Eddy Simulation of a scale-model turbofan for fan noise source diagnostic

A wall-modeled statistically converged Large Eddy Simulation (LES) of the turbulent flow in the NASA Source Diagnostic Test turbofan has been successfully performed for the first time. A good agreement with aerodynamic measurements is observed for both Reynolds Averaged Navier-Stokes and LES results, although the LES provides better results in the tip regions where large coherent structures appear and no flow separation on the stator vanes is observed. In the LES the boundary layer naturally transition to turbulence on the blade suction side but remains quasi laminar over most of its pressure side. The rotor-wake turbulence yielding the stage broadband noise is then seen to be quasi isotropic. Transition on the downstream stator vanes is not triggered by the wake impingement but rather occurs at mid-chord. Finally, acoustics are investigated using both Ffowcs Williams & Hawkings' and Goldstein's analogies from the recorded LES noise source on the stator vanes. The latter analogy provides levels closer to the measurements especially at high frequencies, although the results are most likely still influenced by too coherent rotor tip secondary flow at low frequencies.

Broadband noise prediction of fan outlet guide vane using a cascade response function

An analytical model of the broadband noise produced by both the interaction of ingested turbulence with a fan rotor blades and the rotor-wake impingement on downstream stator vanes is proposed and detailed. The noise prediction methodology is a strip-theory approach based on a previously published formulation of the three-dimensional unsteady blade loading for a rectilinear cascade. This three-dimensional cascade response applied in each strip combined with an acoustic analogy in an annular duct have been chosen to account for the main three-dimensional effects. To further improve some of the identified limitations of this approach, a correction is added to mitigate the effects of the non-coincidence of the cut-on frequencies of the annular duct modes and of the modes of the rectilinear cascade. A correction of the unsteady blade loading formulation, previously developed in a tonal configuration, is also introduced to account for the dispersion relation of annular duct modes in the rectilinear-cascade model. The model is compared with experimental results of the 22-in source diagnostic test (SDT) fan rig of the NASA Glenn Research Center. A numerical assessment of the simplifications proposed in the model and of the convergence of the truncated sums in spanwise wavenumbers and azimuthal orders of the incident perturbation is carried out. The subcritical gusts are shown to have a crucial effect at low frequencies, whereas they become negligible at higher frequencies. Furthermore, alternative high-frequency formulations lead to a satisfactory accuracy above a Helmholtz number based on the duct radius of 20. The strong reduction in computational time associated with these formulations could justify their use for parametric studies in industrial context. The effect of the turbulence model is also investigated showing the relevance of Liepmann's isotropic model in the SDT case, and a possible strong effect of anisotropy in static tests. Finally, the model is compared with NASA's experimental results for two outlet guide vanes at approach condition, showing a very good agreement upstream, whereas an underestimate of 3–5 dB is observed downstream in the middle frequency range.

On the Phenomenon of Pressure Pulses Reflecting Between Blades of Adjacent Blade Rows of Turbomachines

The recently revived interest in “acoustic resonances,” whose details are still not well defined or understood, points to a realization that a new look at some previously unrecognized findings is needed to explain problems encountered in operation of compressors and turbines. The purpose of this paper is to call the attention of the turbomachinery community to an important physical phenomenon of pressure waves in form of pulses, which reflect between blades of adjacent blade rows of turbomachines discovered more than 40 years ago, about whose existence and consequences there is little awareness today. The turbine test results which led the author in 1957 to hypothesize the existence of the phenomenon of reflecting pressure pulses are described. Subsequently, his 1966 ASME paper is discussed. In it, the author reported on the photographed observations of pressure pulses reflecting between stationary nozzles and moving blades of a water-table turbine at Lehigh University, on the description of the various types of such waves, and on an explanation of some of the resonant blade excitation frequencies observed by National Advisory Committee for Aeronautics (NACA) in a turbine of turbojet engine. This is followed by a description of his 1984 ASME paper, in which more general formulae were derived for the blade excitation frequencies caused by the reflections of pressure pulses between the rotor blades, and both upstream and downstream stator vanes. These equations were subsequently used to explain the blade excitation frequencies measured in an axial compressor stage. Finally, his 1992 AIAA paper is discussed, in which additional formulae relating to the reflecting pressure pulses were derived, and the process of formation of a pressure pulse was explained. To put this work in perspective, the author provided, in mostly chronological order, excerpts from reports on operational problems encountered with turbomachines in service and brief descriptions, from selected publications, of pertinent research work.

Comparison between Measurements and Predictions of Fan Broadband Noise on a Low Speed Fan Rig

This paper presents a comparison between measurements and predictions of the fan broadband noise in a low speed fan rig. It is assumed here that the broadband noise is due to the interaction between turbulence generated in the wake of the rotor and the downstream stator vanes. The contribution due to the background turbulence level is also included in the analysis. The stator vanes are assumed to be unloaded flat plates and the wakes are assumed to satisfy two-dimensional plane wake theory. The wake turbulence velocity correlation function is assumed to be periodic in the blade pitch direction (i.e., cyclo-stationary). Aerodynamic input data to the model, both steady and unsteady, is taken directly from hot wire velocity measurements. Comparisons between predicted and measured sound power spectra are in reasonable agreement with spectral shape being particularly well captured by the model. We first review the details of the broadband noise prediction model. Hot wire velocity data will then presented by way of validating the model simplifying assumptions, and to provide the input data to the model. Finally, comparisons between the measured and predicted sound power spectrum will be presented. The work reported here was undertaken as part of the European project, PROBAND.

Flow Field Computations of Combustor-Turbine Interactions Relevant to a Gas Turbine Engine

The current demands for high-performance gas turbine engines can be reached by raising combustion temperatures to increase power output. High combustion temperatures create a harsh environment that leads to the consideration of the durability of the combustor and turbine sections. This paper presents a computational study of a flow field that is representative of what occurs in a combustor and how that flow field convects through the first downstream stator vane. The results of this study indicate that the development of the secondary flow field in the turbine is highly dependent on the incoming total pressure profile. The endwall heat transfer is also found to depend strongly on the secondary flow field.

Effect of Nonuniform Inlet Conditions on Endwall Secondary Flows

Exit combustor flow and thermal fields entering downstream stator vane passages in a gas turbine engine are highly nonuniform. These flow and thermal fields can significantly affect the development of the secondary flows in the turbine passages contributing to high platform heat transfer and large aerodynamic losses. The flow and thermal fields combine to give nonuniform total pressure profiles entering the turbine passage which, along with the airfoil geometry, dictate the secondary flow field. This paper presents an analysis of the effects of varying total pressure profiles in both the radial and combined radial and circumferential directions on the secondary flowfields in a first-stage stator vane. These inlet conditions used for the first vane simulations are based on the exit conditions predicted for a combustor. Prior to using the predictions, these CFD simulations were benchmarked against flowfield data measured in a large-scale, linear, turbine vane cascade. Good agreement occurred between the computational predictions and experimentally measured secondary flows. Analyses of the results for several different cases indicate variations in the secondary flow pattern from pitch to pitch, which attributes to the rationale as to why some airfoils quickly degrade while others remain intact over time.

Effect of Inlet Conditions on Endwall Secondary Flows

Exit combustor e ow and thermal e elds entering downstream stator vane passages in a gas turbine engine are highly nonuniform. These e ow and thermal e elds can signie cantly affect the development of the secondary e ows in the turbine passages attributing to high platform heat transfer and large aerodynamic losses. An analysis is presented of the effects of both the temperature and velocity proe les on the secondary e ows in the endwall region of a e rst-stage stator vane geometry. These effects were assessed using the predicted e owe eld results from computational e uid dynamics (CFD) simulations. Prior to using the predictions, these CFD simulations were benchmarked against e owe eld data measured in a large-scale, linear, turbine vane cascade. Good agreement occurred between the computational predictions and experimentally measured secondary e ows. Analyses of the results for several different cases indicate the stagnation pressure gradient is a key parameter in determining the character of the secondary e ows.

Experimental Validation of Lifting Surface Theory for Rotor-Stator Interaction Noise Generation

An important element in the National Aerospace Laboratory NLR prediction model for rotor-stator interaction noise is the unsteady lifting surface theory for a rotating blade row. To validate this lifting surface theory, a wind-tunnel experiment has been carried out using a through-flow nacelle with a hub, a 16-bladed rotor, and an 18-vane stator. Two mechanisms of sound generation were recognized. As usual in rotor-stator configurations, sound was produced by interaction of rotor viscous wakes with the downstream stator vanes. In addition, due to the small gap between rotor and stator, sound was generated by interaction of the stator vanes, with flow distortions caused by the thickness of the rotor blades and vice versa. Trip wires were attached to blades and vanes to simulate high-Reynolds-number flow. Unsteady pressures were measured on a stator vane, on a rotor blade, and in the inlet. Experiment and theory are briefly described, and results from both are compared. For most of the measured cases, the agreement is good.

Experimental validation of lifting surface theory for rotor-stator interaction noise generation

An important element in the National Aerospace Laboratory NLR prediction model for rotor-stator interaction noise is the unsteady lifting surface theory for a rotating blade row. To validate this lifting surface theory, a wind-tunnel experiment has been carried out using a through-flow nacelle with a hub, a 16-bladed rotor, and an 18-vane stator. Two mechanisms of sound generation were recognized. As usual in rotor-stator configurations, sound was produced by interaction of rotor viscous wakes with the downstream stator vanes. In addition, due to the small gap between rotor and stator, sound was generated by interaction of the stator vanes, with flow distortions caused by the thickness of the rotor blades and vice versa. Trip wires were attached to blades and vanes to simulate high-Reynolds-number flow. Unsteady pressures were measured on a stator vane, on a rotor blade, and in the inlet. Experiment and theory are briefly described, and results from both are compared. For most of the measured cases, the agreement is good.

Compressor Blade Forced Response Due to Downstream Vane-Strut Potential Interaction

A blade forced response prediction system has been developed using an implicit two-dimensional CFD solver to model the rotor blade forced response due to the static pressure distortion (potential disturbance) from the downstream stator vanes and struts. The CFD solver predicts the static pressure distortion upstream of the stator vanes and struts, which is used to calculate the induced velocity perturbation at the rotor inlet. Using the velocity perturbation and the blade’s natural frequencies and mode shapes from a finite element model, the unsteady aerodynamic modal forces and the aerodynamic damping are calculated. A modal response solution is then performed. The results show that the stator vanes cause a significant amplification of the potential disturbances due to the struts. Effects of strut and vane modifications are examined using the analysis. A vane modification with an “optimized” flow angle distribution shows that the disturbance can be greatly reduced. Recent testing of the strut modification shows exceptional correlation with the prediction.

Turbofan noise research at NASA

Results of recent NASA research to reduce aircraft turbofan noise are described. For very high bypass ratio turbofan engines, the dominant source of engine noise is the fan. A primary mechanism of tone noise generation is the rotor blade wakes interacting with downstream stator vanes. Methods of analyzing rotor–stator tone noise generation are described and sample results are given. The analysis includes the unsteady aerodynamic response of the stators to gusts, coupling to duct modes, and finite element calculation of the far-field radiation accounting of for flow in the fan ducts. Wind tunnel tests of model fans and nacelles are described including comparisons between measured and predicted tone directivities. A novel rotating microphone technique is used to measure the acoustic modes in the fan inlet, and the results indicate that a mechanism associated with unsteady fan tip loading can be important in addition to rotor–stator interaction. Finally, concepts for active fan noise control which emphasize control at the source are addressed by an unsteady aerodynamic analysis of compliant stator vanes.

Fan noise reduction achieved by removing tip flow irregularities behind the rotor

One of the major noise producing mechanisms in a fan stage for a turbofan engine is the interaction of the rotor wakes with the downstream stator vanes. Previous work has indicated that the rotor tip flow irregularities (vortices and velocity defects) may be as large or larger a source of blade passage tone noise as the usually considered mean rotor wake. Flow was removed between the rotor and stator of an existing fan stage by an outer wall slot and noise measurements were made to investigate the rotor tip flow irregularity‐stator interaction as a noise source. Farfield noise data taken in an inlet arc showed little tone noise reduction with flow removal, probably the result of rotor‐stator interaction noise not propagating upstream through the rotor. In‐duct measurements behind the fan stator did show noise reduction with flow removal and proportionally more reduction occurred with the first increment of flow removal than with subsequent amounts. Since the tip flow irregularities were assumedly removed with the first increment of flow, the duct measurements indicate that the tip flow irregularity‐stator interaction is probably as large a noise producing mechanism as the normally considered rotor wake‐stator interaction.

Analysis of an Axial Compressor Blade Vibration Based on Wave Reflection Theory

The paper describes application of the theory of wave reflection in turbomachines to rotor blade vibrations measured in an axial compressor stage. The blade vibrations analyzed could not be explained using various flutter prediction techniques. The wave reflection theory, first advanced in 1966, is expanded, and more general equations for the rotor blade excitation frequencies are derived. The results of the analysis indicate that all examined rotor blade vibrations can be explained by forced excitations caused by reflecting waves (pressure pulses). Wave reflections between the rotor blades and both the upstream and downstream stator vanes had to be considered.

Discrete frequency noise reduction modeling for application to fanjet engines

A model for the generation of discrete frequency noise due to rotor–stator interactions in a fanjet engine is examined to ascertain its relevance to noise reduction studies. Predictions are obtained for the noise reduction between a base fan stage and one redesigned to achieve reduced discrete frequency noise utilizing response functions corresponding to incompressible flow isolated airfoil transverse and longitudinal gust analyses as well as incompressible and compressible flat plate airfoil cascade transverse gust analyses. The correlation of the various predictions with the actual fan test results demonstrated that the viscous wake-generated noise model utilizing the compressible flow cascade response function was the one most appropriate for noise reduction studies of fanjet engines. The significant parameters and their relative importance were then identified by means of a parametric study of the base fan based on the compressible flow cascade response analysis. This study showed that the dominant parameter relevant to discrete frequency noise generation is the interblade phase angle, equal to the ratio of the number of upstream rotor blades to downstream stator vanes. Other parameters of importance included the Mach number, reduced frequency, and vane solidity.

Reducing Blade Tonal Noise Generated by Vane-Axial Fans

An experimental investigation was made to determine the effectiveness of unequally spaced blades and blade/vane interaction techniques in reducing the pure-tone aerodynamic noise generated by vane-axial fans. Balanced mirror-image and other modulated blade patterns were investigated for 13, 19, 20, and 21 blades in combination with nine equally spaced downstream stator vanes. In-duct discrete-frequency noise analyses were obtained. Blade-passing frequency and its harmonics were modulated and reduced 10–20 dB, using unequally spaced blades. Large spacing deviation had the effect of spreading the sound energy to more harmonics of rotational frequency. In some instances, significant downstream stator vane and upstream strut frequencies were observed. Predictable pure-tone multiples of rotational frequency were also detected. Elimination of inlet strut frequency and large reductions at the harmonics of blade-passing frequency were obtained upon replacement of the inlet bell and stationary nose cone and eight support struts with an inlet bell and spinning nose cone. Unequally spacing the blades of the fan had little effect on the ability of the fan to move air.

Experimental Noise Studies of Inlet Guide Vane-Rotor-Stator Interactions for a Single-Stage Axial-Flow Compressor

Farfield noise-radiation patterns for the rotor alone, the inlet guide vane and rotor, the rotor and downstream stator vanes, and the inlet guide vane-rotor-stator combination are compared for a constant rotor speed and mass flow. For the configurations studied, the radiated-noise levels due to the interactions of rotor and downstream stator vanes are appreciably less than those clue to inlet guide vane and rotor interactions. It has been possible to correlate variations in the noise levels, with a knowledge of local conditions in the vicinity of the blades.