The calculation of acoustic shielding of engine noise by the Silent Aircraft Airframe

The Silent Aircraft airframe has a flying wing design with a large wing planform and a propulsion system embedded in the rear of the airframe with intake on the upper surface of the wing. In the present paper, boundary element calculations are presented to evaluate acoustic shielding at low frequencies. Besides the three-dimensional geometry of the Silent Aircraft airframe, a few two-dimensional problems are considered that provide some physical insight into the shielding calculations. Mean flow refraction effects due to forward flight motion are accounted for by a simple time transformation that decouples the mean-flow and acoustic-field calculations. It is shown that significant amount of shielding can be obtained in the shadow region where there is no direct line of sight between the source and observer. The boundary element solutions are restricted to low frequencies. We have used a simple physically-based model to extend the solution to higher frequencies. Based on this model, using a monopole acoustic source, we predict at least an 18 dBA reduction in the overall sound pressure level of forward-propagating fan noise because of shielding.

Effects of rotating inlet distortion on multistage compressor stability

In multi-spool engines, rotating stall in an upstream compressor will impose a rotating distortion on the downstream compressor, thereby affecting its stability margin. In this paper experiments are described in which this effect was simulated by a rotating screen upstream of several multistage low-speed compressors. The measurements are complemented by, and compared with, a theoretical model of multistage compressor response to speed and direction of rotation of an inlet distortion. For co-rotating distortions (i.e., distortions rotating in the same direction as rotor rotation), experiments show that the compressors exhibited significant loss in stability margin and that they could be divided into two groups according to their response. The first group exhibited a single peak in stall margin degradation when the distortion speed corresponded to roughly 50% of rotor speed. The second group showed two peaks in stall margin degradation corresponding to distortion speeds of approximately 25-35% and 70-75% of rotor speed. These new results demonstrate that multistage compressors can have more than a single resonant response. Detailed measurements suggest that the two types of behavior are linked to differences between the stall inception processes observed for the two groups of compressors and that a direct connection thus exists between the observed forced response and the unsteady flow phenomena at stall onset. For counter-rotational distortions, all the compressors tested showed minimal loss of stability margin. The results imply that counter-rotation of the fan and core compressor, or LP and HP compressors, could be a worthwhile design choice. Calculations based on the two-dimensional theoretical model show excellent agreement for the compressors which had a single peak for stall margin degradation. We take this first-of-a-kind comparison as showing that the model, though simplified, captures the essential fluid dynamic features of the phenomena. Agreement is not good for compressors which had two peaks in the curve of stall margin shift versus distortion rotation speed. The discrepancy is attributed to the three-dimensional and short length scale nature of the stall inception process in these machines; this includes phenomena that have not yet been addressed in any model.

Multi-objective design and optimisation of bypass outlet-guide vanes

This paper describes the development of an automated design optimization system that makes use of a high fidelity Reynolds-Averaged CFD analysis procedure to minimize the fan forcing and fan BOGV (bypass outlet guide vane) losses simultaneously taking into the account the down-stream pylon and RDF (radial drive fairing) distortions. The design space consists of the OGV's stagger angle, trailing-edge recambering, axial and circumferential positions leading to a variable pitch optimum design. An advanced optimization system called SOFT (Smart Optimisation for Turbomachinery) was used to integrate a number of pre-processor, simulation and in-house grid generation codes and postprocessor programs. A number of multi-objective, multi-point optimiztion were carried out by SOFT on a cluster of workstations and are reported herein.

Movement of deposited water on turbomachinery rotor blade surfaces

A theoretical approach for calculating the movement of liquid water following deposition onto a turbomachine rotor blade is described. Such a situation can occur during operation of an aero-engine in rain. The equation of motion of the deposited water is developed on an arbitrarily oriented plane triangular surface facet. By dividing the blade surface into a large number of facets and calculating the water trajectory over each one crossed in turn, the overall trajectory can be constructed. Apart from the centrifugal and Coriolis inertia effects, the forces acting on the water arise from the blade surface friction, and the aerodynamic shear and pressure gradient. Non- dimensionalisation of the equations of motion provides considerable insight and a detailed study of water flow on a flat rotating plate set at different stagger angles demonstrates the paramount importance of blade surface friction. The extreme cases of low and high blade friction are examined and it is concluded that the latter (which allows considerable mathematical generalisation) is the most likely in practice. It is also shown that the aerodynamic shear force, but not the pressure force, may influence the water motion. Calculations of water movement on a low-speed compressor blade and the fan blade of a high bypass ratio aero-engine suggest that in low rotational speed situations most of the deposited water is centrifuged rapidly to the blade tip region. Copyright © 2006 by ASME.

Challenges in the Silent Aircraft engine design

The Silent Aircraft Initiative goal is to design an aircraft that is imperceptible above background noise outside the airport boundary. The aircraft that fulfils this objective must also be economically competitive with conventional aircraft of the future and therefore fuel consumption and mechanical reliability are key considerations for the design. To meet these ambitious targets, a multi-fan embedded turbofan engine with boundary layer ingestion has been proposed. This configuration includes several new technologies including a variable area nozzle, a complex high-power transmission system, a Low Pressure turbine designed for low-noise, an axial-radial HP compressor, advanced acoustic liners and a low-speed fan optimized for both cruise and off-design operation. These technologies, in combination, enable a low-noise and fuel efficient propulsion system but they also introduce significant challenges into the design. These challenges include difficulties in predicting the noise and performance of the new components but there are also challenges in reducing the design risks and proving that the new concepts are realizable. This paper presents the details of the engine configuration that has been developed for the Silent Aircraft application. It describes the design approach used for the critical components and discusses the benefits of the new technologies. The new technologies are expected to offer significant benefits in noise reduction without compromising fuel burn. However, more detailed design and further research are required to fully control the additional risks generated by the system complexity.

Performance of a Boundary Layer Ingesting (BLI) propulsion system

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.

Active wavelength control of tunable lasers for use in Un cooled WDM systems

It is shown for the first time that uncooled tunable DBR-laser diodes can be used as athermal WDM sources. Using novel bias current control, absolute wavelength control to within 6Å has been achieved for temperatures up to 70°C. © 2000 Optical Society of America.

Low on-resistance SOI dual-trench-gate MOSFET

A low specific on-resistance (R-{{\rm on}, {\rm sp}}) integrable silicon-on-insulator (SOI) MOSFET is proposed, and its mechanism is investigated by simulation. The SOI MOSFET features double trenches and dual gates (DTDG SOI): an oxide trench in the drift region, a buried gate inset in the oxide trench, and another trench gate (TG) extended to a buried oxide layer. First, the dual gates form dual conduction channels, and the extended gate widens the vertical conduction area; both of which sharply reduce R-{{\rm on}, {\rm sp}}. Second, the oxide trench folds the drift region in the vertical direction, resulting in a reduced device pitch and R-{{\rm on}, {\rm sp}}. Third, the oxide trench causes multidirectional depletion. This not only enhances the reduced surface field effect and thus reshapes the electric field distribution but also increases the drift doping concentration, leading to a reduced R-{{\rm on}, {\rm sp}} and an improved breakdown voltage (BV). Compared with a conventional SOI lateral Double-diffused metal oxide semiconductor (LDMOS), the DTDG MOSFET increases BV from 39 to 92 V at the same cell pitch or decreases R-{{\rm on}, { \rm sp}} by 77% at the same BV by simulation. Finally, the TG extended synchronously acts as an isolation trench between the high/low-voltage regions in a high-voltage integrated circuit, saving the chip area and simplifying the isolation process. © 2006 IEEE.

Influence of film cooling hole angles and geometries on aerodynamic loss and net heat flux reduction

Turbine design engineers have to ensure that film cooling can provide sufficient protection to turbine blades from the hot mainstream gas, while keeping the losses low. Film cooling hole design parameters include inclination angle (α), compound angle (β ), hole inlet geometry and hole exit geometry. The influence of these parameters on aerodynamic loss and net heat flux reduction is investigated, with loss being the primary focus. Low-speed flat plate experiments have been conducted at momentum flux ratios of IR = 0.16, 0.64 and 1.44. The film cooling aerodynamic mixing loss, generated by the mixing of mainstream and coolant, can be quantified using a three-dimensional analytical model that has been previously reported by the authors. The model suggests that for the same flow conditions, the aerodynamic mixing loss is the same for holes with different α and β but with the same angle between the mainstream and coolant flow directions (angle κ). This relationship is assessed through experiments by testing two sets of cylindrical holes with different α and β : one set with κ = 35°, another set with κ = 60°. The data confirm the stated relationship between α, β, κ and the aerodynamic mixing loss. The results show that the designer should minimise κ to obtain the lowest loss, but maximise β to achieve the best heat transfer performance. A suggestion on improving the loss model is also given. Five different hole geometries (α =35.0°, β =0°) were also tested: cylindrical hole, trenched hole, fan-shaped hole, D-Fan and SD-Fan. The D-Fan and the SD-Fan have similar hole exits to the fan-shaped hole but their hole inlets are laterally expanded. The external mixing loss and the loss generated inside the hole are compared. It was found that the D-Fan and the SD-Fan have the lowest loss. This is attributed to their laterally expanded hole inlets, which lead to significant reduction in the loss generated inside the holes. As a result, the loss of these geometries is ≈ 50 % of the loss of the fan-shaped hole at IR = 0.64 and 1.44. Copyright © 2011 by ASME.

Shock Propagation and MPT Noise From a Transonic Rotor in Nonuniform Flow

One of the major challenges in high-speed fan stages used in compact, embedded propulsion systems is inlet distortion noise. A body-force-based approach for the prediction of multiple-pure-tone (MPT) noise was previously introduced and validated. In this paper, it is employed with the objective of quantifying the effects of nonuniform flow on the generation and propagation of MPT noise. First-of-their-kind back-to-back coupled aero-acoustic computations were carried out using the new approach for conventional and serpentine inlets. Both inlets delivered flow to the same NASA/GE R4 fan rotor at equal corrected mass flow rates. Although the source strength at the fan is increased by 38 dB in sound power level due to the nonuniform inflow, far-field noise for the serpentine inlet duct is increased on average by only 3.1 dBA overall sound pressure level in the forward arc. This is due to the redistribution of acoustic energy to frequencies below 11 times the shaft frequency and the apparent cut-off of tones at higher frequencies including blade-passing tones. The circumferential extent of the inlet swirl distortion at the fan was found to be two blade pitches, or 1/11th of the circumference, suggesting a relationship between the circumferential extent of the inlet distortion and the apparent cut-off frequency perceived in the far field. A first-principles-based model of the generation of shock waves from a transonic rotor in nonuniform flow showed that the effects of nonuniform flow on acoustic wave propagation, which cannot be captured by the simplified model, are more dominant than those of inlet flow distortion on source noise. It demonstrated that nonlinear, coupled aerodynamic and aero-acoustic computations, such as those presented in this paper, are necessary to assess the propagation through nonuniform mean flow. A parametric study of serpentine inlet designs is underway to quantify these propagation effects. © 2013 American Society of Mechanical Engineers.

Shock propagation and MPT noise from a transonic rotor in non-uniform flow

One of the major challenges in hig4h-speed fan stages used in compact, embedded propulsion systems is inlet distortion noise. A body-force-based approach for the prediction of multiple-pure-tone (MPT) noise was previously introduced and validated. In this paper, it is employed with the objective of quantifying the effects of non-uniform flow on the generation and propagation of MPT noise. First-of-their-kind back-to-back coupled aero-acoustic computations were carried out using the new approach for conventional and serpentine inlets. Both inlets delivered flow to the same NASA/GE R4 fan rotor at equal corrected mass flow rates. Although the source strength at the fan is increased by 45 dB in sound power level due to the non-uniform inflow, farfield noise for the serpentine inlet duct is increased on average by only 3.1 dBA overall sound pressure level in the forward arc. This is due to the redistribution of acoustic energy to frequencies below 11 times the shaft frequency and the apparent cut-off of tones at higher frequencies including blade-passing tones. The circumferential extent of the inlet swirl distortion at the fan was found to be 2 blade pitches, or 1/11th of the circumference, suggesting a relationship between the circumferential extent of the inlet distortion and the apparent cut-off frequency perceived in the far field. A first-principles-based model of the generation of shock waves from a transonic rotor in non-uniform flow showed that the effects of non-uniform flow on acoustic wave propagation, which cannot be captured by the simplified model, are more dominant than those of inlet flow distortion on source noise. It demonstrated that non-linear, coupled aerodynamic and aeroacoustic computations, such as those presented in this paper, are necessary to assess the propagation through non-uniform mean flow. A parametric study of serpentine inlet designs is underway to quantify these propagation effects. Copyright © 2011 by ASME.