The coupling between the bleed system and the flowfield of a downstream compressor stage is studied using two approaches. In the first, three-dimensional, full annulus, unsteady computations simulate the flow in a low speed research compressor with non-uniform bleed extraction. Comparisons with experimental data show that the flow prediction in the main annulus is accurate to within 0.005 of flow coefficient and 0.5◦ of flow angle. The CFD is then used to provide a description of flow within the bleed system itself. In the second approach, a two-dimensional mean radius model, similar to that adopted by Hynes and Greitzer in previous work on compressor stability, is used to simulate the response of the compressor to non-uniform bleed. This model is validated against experimental data for a single stage compressor and despite the inherent assumptions (two dimensional flow and simplified compressor response) provides a satisfactory prediction of the flow for preliminary design purposes with orders of magnitude less computational cost than full 3D CFD. The model is then used to investigate the effect of different levels of bleed non-uniformity and of varying the axial distance between the bleed and the downstream stage. Reducing bleed non-uniformity and moving the stage away from the bleed slot are predicted to reduce the circumferential non-uniformity of the flow entering the stage. INTRODUCTION The bleed extraction from an axial compressor is typically circumferentially non-uniform. The principal cause for this is that, although the casing slot may be axisymmetric, the bleed flow is often distributed to other parts of the gas turbine via a finite number of off-take ducts. The result of this non-axisymmetric geometry is the imposition of a circumferential 1 TURBO-16-1024 (GRIMSHAW) static pressure distortion on the main annulus flow. Studies of uniform bleed have been performed using computations of compressors [1–3] and linear cascade experiments [4–6], as well as by the present authors using a rotating rig [7]. These studies show, that for typical design operating point bleed rates of 1% to 5%, there is limited spanwise redistribution of the flow due to bleed and a passage-averaged one-dimensional treatment is sufficient to model the compressor response. Non-uniform bleed has been investigated with annular cascade experiments [8, 9] and by the present authors [7]. Our tests demonstrated that the distribution of non-uniform bleed is linked to a reduction in compressor operating range and that this can be quantified using an approach similar to the DCθ criterion that was developed to characterise inlet total pressure distortion [10]. The purpose of this paper is to address the question, “How should one analyse a compressor geometry with non-uniform bleed?” The answer to this question depends on the fidelity required at a given stage in the design process. We propose two strategies: for the final design, a high fidelity approach based on a full annulus unsteady CFD analysis of the compressor stage and bleed system; for preliminary design, a low fidelity approach based on a mean radius model that assumes the flow upstream of the compressor, including where the bleed is extracted, is linear and the compressor behaviour is input via a measured pressure-rise characteristic. The paper is organised as follows: we first present the test compressor used to provide data to validate the modelling approaches. The CFD code and the meshing strategy used are then described and the results from the CFD calculations are compared with measurements and discussed. A qualitative description of the flow in the bleed system is developed based on the computations. Next we describe the approach, assumptions and specific elements that make up the low fidelity mean radius model. Finally, we show comparisons of the mean radius model predictions against test data, discuss the sensitivity of the model to assumptions, and investigate the effect of changing the bleed distribution and moving the downstream stage further away from the bleed extraction. LOW SPEED COMPRESSOR RIG WITH BLEED Experiments were performed with a one-stage, axial-flow, low-speed compressor with inlet guide vanes (IGVs) and a hub-to-tip radius ratio of 0.75, Grimshaw et al. [7]. Bleed air is extracted from the compressor through a bleed system which is located upstream of the stage. The compressor and bleed system design is typical of the rear stages of a land-based gas turbine and details are given in Table 1. Figure 1 shows the layout of the rig and the measurement stations referred to in this paper. The bleed system has a small, axisymmetric plenum chamber, one off-take duct and bleed rate of 4.1%. This produces non-uniform extraction through the bleed slot which is measured using 16 pairs of static pressure tappings and stagnation pressure probes as described in [7]. The resulting distortion into the downstream stage causes a 3.0% increase in the stage inlet stalling flow coefficient as compared to the uniform bleed rate case. Five-hole probe area traverses are performed with an automated traversing system. In all cases, each traverse is over one stator pitch and contains 21 equi-spaced pitchwise points and 21 radial points clustered towards the end walls. The relative position of the off-take duct and the traverse location is changed by moving the off-take duct to different circumferential positions. 2 TURBO-16-1024 (GRIMSHAW) Hub-to-tip radius ratio 0.75

The coupling between the bleed system and the flowfield of a downstream compressor stage is studied using two approaches. In the first approach, three-dimensional, full annulus, unsteady computations simulate the flow in a low-speed research compressor with nonuniform bleed extraction. Comparisons with experimental data show that the flow prediction in the main annulus is accurate to within 0.005 of flow coefficient and 0.5deg of flow angle. The computational fluid dynamics (CFD) is then used to provide a description of flow within the bleed system itself. In the second approach, a two-dimensional mean radius model, similar to that adopted by Hynes and Greitzer in the previous work on compressor stability, is used to simulate the response of the compressor to nonuniform bleed. This model is validated against experimental data for a single-stage compressor, and despite the inherent assumptions (two-dimensional flow and simplified compressor response), provides a satisfactory prediction of the flow for preliminary design purposes with orders of magnitude less computational cost than full 3D CFD. The model is then used to investigate the effect of different levels of bleed nonuniformity and of varying the axial distance between the bleed and the downstream stage. Reducing bleed nonuniformity and moving the stage away from the bleed slot are predicted to reduce the circumferential nonuniformity of the flow entering the stage.

This paper investigates numerically the acoustic sources and far-field noise of chevron and round jets. The acoustic sources are described by the fourth-order space-time velocity cross correlations, which are calculated based on a large-eddy simulation flowfield. Gaussian functions are found to fit the axial, radial, and azimuthal cross correlations reasonably well. The axial length scales are three to four times the radial and azimuthal length scales. For the chevron jet, the cross-correlation scales vary with azimuthal angle up to six jet diameters downstream; beyond that, they become axisymmetric like those for a round jet. The fourth-order space-time cross correlation of the axial velocity R1111 is the dominant source component, and there are considerable contributions from other source components such as R2222, R3333, R1212, R1313, and R2323 cross correlations where 1, 2, and 3 represent axial, radial, and azimuthal directions, respectively. For the chevron jet, these cross correlations decay rapidly with axial distance whereas for the round jet, they remain roughly constant over the first 10 jet diameters. The chevron jet intensifies both the R2222 and R3333 cross correlations within two jet diameters of the jet exit. The amplitude, length, and time scales of the cross-correlations of a large-eddy simulation velocity field are investigated as functions of position and are found to be proportional to the turbulence amplitude, length, and time scales that are determined from a Reynolds-averaged Navier-Stokes calculation. The constants of proportionality are found to be independent of position within the jet, and they are quite close for chevron and round jets. The scales derived from Reynolds-averaged Navier-Stokes are used for source description, and an acoustic analogy is used for sound propagation. There is an excellent agreement between the far-field noise predictions and measurements. At low frequencies, the chevron nozzle significantly reduces the far-field noise by 5-6dB at 30deg and 2-3dB at 90deg to the jet axis. However, the chevron nozzle slightly increases high-frequency noise. It was found that R1212 and R1313 cross correlations have the largest contribution to the jet noise at 30deg to the jet axis, whereas the R2323 cross correlation has the largest contribution to the jet noise at 90deg to the jet axis. The Reynolds-averaged Navier-Stokes calculations are repeated with different turbulence models, and the noise prediction is found to be almost insensitive to the turbulence model. The results indicate that the modeling approach is capable of assessing advanced noise-reduction concepts.

An optimization process has been used to design an ultra-low count fan outlet guide vane with an unconventional leading edge profile to reduce the interaction noise. Computational fluid dynamics has been used to predict the aerodynamic and acoustic performance of the stator vane. The final stator design has been built and tested in a representative fan stage rig to determine its tone noise characteristics. The stator vane is found to give significant tone noise reduction at the fundamental blade passing frequency at cut-back in line with design expectations. Detailed comparisons of predicted circumferential and radial modes levels against measured mode detection data are also presented. A good agreement was found between numerical predictions and experimental data.

This paper provides a physical interpretation of the mechanism of stagnation enthalpy and stagnation pressure changes in turbomachines due to unsteady flow, the agency for all work transfer between a turbomachine and an inviscid fluid. Examples are first given to illustrate the direct link between the time variation of static pressure seen by a given fluid particle and the rate of change of stagnation enthalpy for that particle. These include absolute stagnation temperature rises in turbine rotor tip leakage flow, wake transport through downstream blade rows, the influence on mixing losses of turbine wake behavior in downstream blade rows, and effects of wake phasing on compressor work input. Fluid dynamic situations are then constructed to explain the effect of unsteadiness, including a physical interpretation of how stagnation pressure variations are created by temporal variations in static pressure; in this it is shown that the unsteady static pressure plays the role of a time-dependent body force potential. It is further shown that when the unsteadiness is due to a spatial nonuniformity translating at constant speed, as in a turbomachine, the unsteady pressure variation can be viewed as a local power input per unit mass from this body force to the fluid particle at that point.

Jet noise has been a major community concern since the 1940s. Among noise-suppression devices, chevron nozzles are most effective as they significantly reduce low-frequency noise without appreciable thrust loss. The objective is to understand the noise-suppression mechanisms of chevron nozzles by identifying the noise sources and predicting the far-field noise for both chevron and round jets. To identify noise sources, the fourth-order space-time velocity cross-correlations are calculated based on an LES flow field. The Gaussian form fits axial, radial and azimuthal cross-correlations reasonably well. The axial length scales of the cross-correlations are 3-4 times the radial or azimuthal length scales for both chevron and round jets. For chevron jets, the cross-correlation lengths vary with azimuthal angle at axial positions within 6 jet diameters. Further downstream, both chevron and round jets behave similarly. Although for round and chevron nozzles R1111 the fourth-order cross-correlation of the axial velocity is the dominant component of the source, for chevron jets there is a significant contribution from other components such as R2222 and R3333. The cross-correlations decay rapidly with the axial distance for a chevron jet, whereas they remain constant for a round jet. Chevrons intensify R2222 and R3333 within 2 jet diameters downstream of the nozzle exit. The amplitude, length and time scales of the cross-correlations of the LES velocity field are investigated as functions of position and are found to be proportional to the turbulence amplitude, length and time scales determined from a RANS calculation. These proportionality constants are universal i.e. independent of nozzle geometry and position within the jet. The scales derived from RANS are used for source description and an acoustic analogy is used for sound propagation. At low-frequencies, chevrons drastically reduce noise by 5-6 dB at 30° to the jet axis and 2-3 dB at 90°. There is excellent agreement between far-field noise predictions and NASA measurements. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

A novel approach to the development of a hybrid prediction methodology for jet noise is described. Modeling details and numerical techniques are optimized for each of the three components of the model. Far-field propagation is modeled by solution of a system of adjoint linear Euler equations, capturing convective and refraction effects using a spatially developing jet mean flow provided by a Reynolds-averaged Navier―Stokes computational fluid dynamics solution. Sound generation is modeled following Goldstein's acoustic analogy, including a Gaussian function model for the two-point cross correlation of the fourth-order velocity fluctuations in the acoustic source. Parameters in this model describing turbulent length and time scales are assumed to be proportional to turbulence information also taken from the Reynolds-averaged Navier―Stokes computational fluid dynamics prediction. The constants of proportionality are, however, not determined empirically, but extracted by comparison with turbulence length and time scales obtained from a large eddy simulation prediction. The large eddy simulation results are shown to be in good agreement with experimental data for the fourth-order two-point cross-correlation functions. The large eddy simulation solution is then used to determine the amplitude parameter and also to examine which components of the cross correlation are largest, enabling inclusion of all identified dominant terms in the Gaussian source model. The acoustic source description in the present approach is therefore determined with no direct input from experimental data. This model is applied to the prediction of sound to the experimental configuration of the European Union JEAN project, and gives encouraging agreement with experimental data across a wide spectral range and for both sideline and peak noise angles. This paper also examines the accuracy of various commonly made simplifications, for example: a locally parallel mean flow approximation rather than consideration of the spatially evolving mean jet flow and scattering from the nozzle; the assumption of small radial variation in Green function over the turbulence correlation length; the application of the far-field approximation in the Green function; and the impact of isotropic assumptions made in previous acoustic source models.

The silent-aircraft experimental aircraft are balanced by generating lift near the aircraft nose through leading-edge carving of the centerbody. The use of leading-edge carving over the centerbody is novel, in that previous blended-wing body aircraft have balanced the aircraft by downloading the centerbody (via reflex camber) to achieve the effect of a tail. This paper decomposes the aerodynamic forces into contributions from spanwise sections to explain how three-dimensional flow effects are beneficial in allowing the silent-aircraft experimental aircraft to be both statically stable and to have an elliptical lift distribution over a large range of angles of attack. By analyzing the results in this manner, rationale is also given as to why, unlike other blended-wing body-type configurations, the silent-aircraft-experimental design can use supercritical unstable-outer-wing airfoil profiles to generate a balanced and stable aircraft. The results are then used to develop a methodology to aid the aircraft designer in determining the amount of leading-edge carving that is necessary to achieve static stability for blended-wing body-type aircraft.

A novel acoustic analogy approach is considered for jet noise modeling within the generalized acoustic analogy framework. The approach is based on large eddy simulation and is empirical-parameter free. Numerical predictions are provided for a subsonic isothermal axisymmetric jet case (JEAN ex- periment).

This paper presents an assessment of the performance of an embedded propulsion system in the presence of distortion associated with boundary layer ingestion. For fan pressure ratios of interest for civil transports, the benefits of boundary layer ingestion are shown to be very sensitive to the magnitude of fan and duct losses. The distortion transfer across the fan, basically the comparison of the stagnation pressure non-uniformity downstream of the fan to that upstream of the fan, has a major role in determining the impact of boundary layer ingestion on overall fuel burn. This, in turn, puts requirements on the fidelity with which one needs to assess the distortion transfer, and thus the type of models that need to be used in such assessment. For the three-dimensional distortions associated with fuselage boundary layers ingested into a subsonic diffusing inlet, it is found that boundary layer ingestion can provide decreases in fuel burn of several per cent. It is also shown that a promising avenue for mitigating the risks (aerodynamic as well as aeromechanical) in boundary layer ingestion is to mix out the flow before it reaches the engine face.

We set a target for a 'Silent' aircraft to be imperceptible outside the airfield perimeter in an urban environment, and then address conceptual designs to meet this requirement. Avoiding some traditional aircraft noise sources requires a radical rethink about the configuration. An all-lifting design has many benefits, enabling a closer integration of airframe and engine than the traditional 'tube and wing. Low-noise design includes taking advantage of shielding of engine noise by the airframe; low-noise engines with large, low speed jets; an order of magnitude increase in absorption by liners; and operations for low-noise informing the design. Progress to date on the Silent Aircraft Initiative is presented, along with some conceptual aircraft and engine designs. The further work needed to develop these into viable future aircraft is discussed.

The paper presents a method to increase the computational accuracy by preserving additional constraints on the numerical solution. The new technique can noticeably increase the precision of a standard finite-volume flow solver by a modification to the flux computation procedure without changing its essential features. The efficiency of this method is demonstrated by application to the prediction of sound from high-speed helicopter blades. Several open domain boundary conditions for this application are also developed and compared for a model problem of a two-dimensional transonic aerofoil in an unsteady free stream.

To understand how natural wind close to the ground interacts with a jet engine, tests have been completed with a 1/20-scale model fan and intake rig operating within a simulation of the lower atmospheric boundary layer. Unsteady measurements are used to confirm that turbulence in the ambient flow is attenuated as the intake is approached, and analysis of unsteady pressure data reveals that the nacelle surface pressure field is most sensitive to large length-scale gusts. A new statistical approach to analyze intermittent intake separation is used to show that there is hysteresis in the unsteady separation and reattachment of the inlet flow and that there are similarities between the unsteady and steady intake performance. These findings suggest that the intake response to natural wind can be described as quasi steady, and a novel probabilistic model of nacelle behavior is devised based on this hypothesis. The model shows that by combining some statistical characteristics of the wind with the steady intake performance it is possible to synthesize the observed unsteady performance. This confirms that the intake response can be treated in a quasi-steady manner, and this has significant practical implications for engine intake testing and design.

Flow separations in the corner regions of blade passages are common. The separations are three dimensional and have quite different properties from the two-dimensional separations that are considered in elementary courses of fluid mechanics. In particular the consequences for the flow may be less severe than the two-dimensional separation. This paper describes the nature of three-dimensional (3D) separation and addresses the way in which topological rules, based on a linear treatment of the Navier-Stokes equations, can predict properties of the limiting streamlines, including the singularities which form. The paper shows measurements of the flow field in a linear cascade of compressor blades and compares these to the results of 3D computational fluid dynamics (CFD). For corners without tip clearance, the presence of three-dimensional separation appears to be universal, and the challenge for the designer is to limit the loss and blockage produced. The CFD appears capable of predicting this.

Helicopter noise is an increasingly important issue, and at large forward-flight speeds transonic rotor noise is a major contributor. A method for predicting transonic rotor noise, which is more computationally efficient than previous methods and which furthermore offers physical insight into the noise generation, is developed. These benefits combine to make it of potential use to helicopter rotor designers. The permeable surface form of the Ffowcs Williams-Hawkings (FW-H) equation is used to express the sound field in terms of a distribution of monopole and dipole sources over a permeable control surface and a distribution of quadrupole sources over the volume outside of this surface. By choosing the control surface to enclose the transonic flow regions, the noise from the quadrupole distribution becomes negligible. Only the more straightforward surface sources then need be considered, making the acoustic approach computationally efficient. By locating the control surface close to the blade subject to enclosing the transonic flow regions, efficiency in the computational-fluid-dynamics (CFD) approach is also attained. To perform noise predictions, an Euler CFD method to calculate the flowfield was combined with an acoustic method incorporating the retarded time formulation of the FW-H equation. Several rotor blades in hover and steady forward flight were considered, all of which involved transonic flows but for which shock delocalization did not occur. The predictions showed very good agreement with experimental data and with predictions obtained using more computationally intensive methods.