Simon K. Richards

Computation of Fan Noise Radiation through an Engine Exhaust Geometry with Flow

This paper outlines a computational model of noise radiation from a realistic engine exhaust geometry with flow. The computational model described allows acoustic waves, propagating inside the bypass duct of a generic aircraft engine, to be admitted into a computational domain that includes the aft duct section, the exit plane of the duct, and the jet flow immediately downstream. The method has three parts: a matching process to admit acoustic waves into the induct propagation region; near field propagation inside the duct and diffraction at the lip of the exhaust duct; and an integral surface for far field directivity. In this model the near field propagation is determined by a numerical solution of a 2.5D form of the linearised Euler equations. The mean flow about which the equations are linearised is assumed to be axisymmetric. The proposed method is illustrated through a case study on the radiation of a typical fan assembly generated acoustic wave from a generic engine bypass duct. Inside the duct, an acoustic wave of circumferential order m = −13 and comprising five radial modes (n = 1 – 5) is admitted into the model as inputs on the boundary of the computation domain. The radiation of the acoustic wave through the exhaust geometry and mean flow is determined, with the effect of acoustic treatment through the inclusion of lined duct sections also examined.

Aximo: Automated Axiomatic Reasoning for Information Update

We present an algorithm for proving epistemic properties of dynamic scenarios in multi-agent systems and an implementation of it as the C^+^+ program Aximo. The program consists of a rewrite system and a recursive reasoner, and we prove that its decision procedure is sound with regard to the algebraic axiomatics of dynamic epistemic logic. We study the termination and complexity of the program and show its applicability, by proving properties of honest and also newer dishonest versions of the the muddy children puzzle as well as a coin toss scenario.

Attenuation of slat trailing edge noise using acoustic liners

Noise generated by high-lift devices such as the slats on the leading edge of a wing is a major contributor to the overall airframe noise during the landing approach of a commercial aircraft. In this work the concept of attenuating noise sources located at the trailing edge of a slat using absorptive acoustic liners in the slat gap is explored using a time-domain computational aeroacoustic (CAA) scheme. The aim of the computational modelling is to demonstrate the feasibility of controlling slat noise using acoustic liners, and to optimize the design for a future experiment. A model scale three-element high-lift airfoil is modelled; the slat and flap deflection angles are 23 deg and 32 deg respectively and the slat has a blunt trailing edge. The freestream Mach number is 0.2 and the main element angle of attack is 5 deg, corresponding to a typical approach condition. The Reynolds number is 3.6 million, based on the airfoil chord with the high-lift devices retracted. The computational results are divided into three parts: (1) an unsteady Reynolds-Averaged Navier-Stokes (URANS) simulation using high-order spatial and temporal schemes is conducted for a wing without acoustic liners; the computed flow shows the presence of vortex shedding and acoustic sources behind the slat trailing edge; (2) a calculation of the radiated sound field is made for a range of liner impedance values by solving the linearised Euler equations (LEE) for a modelled acoustic source located at the trailing edge of the slat; (3) URANS computations for the wing with liner treatment are conducted. The results show that acoustic liners on the slat cove and on the main element can provide useful attenuation of slat trailing edge noise.

Parallel computation of 3D acoustic radiation from an engine intake

Tonal noise from an engine intake is generally dominated by only a few cut-on modes. For the case of an axisymmetric intake geometry, the three-dimensional (3D) problem can be reduced to a more convenient set of 2D problems through a modal decomposition. However, for the more general case, the 3D formulation must be maintained. In this paper the radiation of 3D acoustic modes from a realistic engine intake duct with background mean flow is studied numerically. The method is based upon a parallel implementation of a computational scheme which allows acoustic modes, propagating inside the intake duct of a generic aircraft engine, to be admitted into a computational domain that includes the duct section, the exit plane of the duct, and the surrounding flow. The method comprises three elements: a matching process to admit acoustic waves into the in-duct propagation region; near-field propagation inside the duct and diffraction at the lip of the duct; and an integral surface for far-field directivity. The wave admission is realised through an absorbing non-reflecting boundary treatment which admits incoming waves and damps spurious waves generated by the numerical solutions. The wave propagation and diffraction are calculated by solving the linearised Euler equations, using high-order compact schemes. Far-field directivity is estimated via an integral surface solution of the Ffowcs Williams - Hawkings equation. The 3D parallel solver is used to determine multi-mode propagation from a realistic engine intake geometry with background mean flow. The solver can account for the effect of a swirling flow. Validation of the solver is performed by comparing 3D propagation results with a previously developed 2.5D formulation. The simulations include the effect on an acoustic lining modeled using a time-domain impedance boundary condition. The 3D method is extended to the case of multi-mode calculations.

Slat noise attenuation using acoustic liner

Noise generated by high-lift devices such as slats on a wing is a major contributor to the overall airframe noise during the landing approach of a commercial aircraft. In this work the concept of attenuating slat noise using absorptive acoustic liners in the slat gap is explored using a time-domain computational aeroacoustic (CAA) scheme. The aim of the computational modeling is to demonstrate the feasibility of controlling slat noise using acoustic liners, and to optimize the design for a future experiment. A model scale threeelement high-lift airfoil is modeled; the slat and flap deflection angles are 23 deg and 32 deg respectively and the slat has a blunt trailing edge. The freestream Mach number is 0.2 and the main element angle of attack is 5 deg, corresponding to a typical approach condition. Reynolds number is 3.6 million, based on the airfoil chord with the high-lift devices retracted. The computational results are divided into three parts: (1) an unsteady Reynolds-Averaged Navier-Stokes (URANS) simulation using high-order spatial and temporal schemes is conducted for a wing without acoustic liners; the computed flow shows the presence of vortex shedding and acoustic source behind the slat trailing edge; (2) an exercise is conducted on a range of liner impedance values by solving the linearised Euler equations (LEE) for a modelled acoustic source located at the trailing edge of the slat; (3) URANS computations for the wing with liner treatment are conducted. The results show that acoustic liners on the slat cove and on the main element can provide useful attenuation of slat trailing edge noise.

Multi-Blade Row Interactions in a Low Pressure Ratio Centrifugal Compressor Stage With a Vaned Diffuser

Previous experimental and CFD investigation of a GE Oil and Gas centrifugal compressor stage with a vaneless diffuser revealed a complex excitation mechanism caused by an aero-acoustic interaction between three blade rows. In stages with vaned diffusers, additional sources of aeromechanical excitation on the impeller can be expected. This unsteady CFD investigation is a follow-up from the previous vaneless diffuser study to identify any additional sources of excitation that arise in the presence of a vaned diffuser in preparation for aeromechanic tests to be conducted later. The study confirms that excitation from impeller-diffuser interaction generated acoustic modes can dominate the potential field excitation from the diffuser vanes. In addition, a significant aero-acoustic excitation to the impeller at a vane pass frequency corresponding to the sum of the vane counts in the two downstream vane rows is observed, and its origination is discussed. The latter excitation is different from that observed in the vaneless diffuser stage where the vane pass frequency observed by the impeller corresponds to the sum of the vane counts in the upstream and downstream vane rows.Copyright

Numerical Prediction of Exhaust Fan Tone Noise from High Bypass Aircraft Engines

The ability to accurately predict fan noise is important in designing and optimizing aircraft engine turbofans for low noise emissions. In this paper, a prediction methodology for exhaust fan tone noise analysis is described and validated against various canonical test cases and NASA Source Diagnostic Test (SDT) data. The prediction process consists of solving Reynolds-Averaged Navier-Stokes (RANS) equations to compute the fan wake, and calculating the acoustic response of the outlet guide vanes (OGV) to the fan wake using linearized Euler equations. Very good agreement is observed between numerical predictions and semi-analytical results for canonical cases. Detailed comparisons against SDT data are presented for unsteady vane pressure and integrated in-duct exhaust noise power levels. Geometric trends for different OGV configurations at various operating conditions are also analyzed.

Unsteady Acoustic Forcing on an Impeller Due to Coupled Blade Row Interactions

This article contains an investigation of the unsteady acoustic forcing on a centrifugal impeller due to coupled blade row interactions. Selected results from an aeromechanical test campaign on a GE Oil and Gas centrifugal compressor stage with a vaneless diffuser are presented. The most commonly encountered sources of impeller excitation due to upstream wake interaction were identified and observed in the testing campaign. A 30/rev excitation corresponding to the sum of upstream and downstream vane counts caused significant trailing edge vibratory stress amplitudes. Due to the large spacing between the impeller and the return channel vanes, this 30/rev excitation was suspected to be caused by an aero-acoustic excitation rather than a potential disturbance. The origin of this aero-acoustic excitation was deduced from an acoustic analysis of the unsteady compressor flow derived from CFD. The analysis revealed a complex excitation mechanism caused by impeller interaction with the upstream vane row wakes and subsequent acoustic wave reflection from the downstream return channel vanes. The findings show it is important to account for aero-acoustic forcing in the aeromechanical design of low pressure ratio centrifugal compressor stages.