The diffraction of externally generated sound in a uniformly moving flow at the trailing edge of a semi-infinite flat plate is studied. In particular, the coupling of the sound field to the hydrodynamic field by way of vortex shedding from the edge is considered in detail, both in inviscid and in viscous flow.In the inviscid model the (two-dimensional) diffracted fields of a cylindrical pulse wave, a plane harmonic wave and a plane pulse wave are calculated. The viscous proess of vortex shedding is represented by an appropriate trailing-edge condition. Two specific cases are compared, in one of which the full Kutta condition is applied, and in the other no vortex shedding is permitted. The results show good agreement with Heavens’ (1978) observations from his schlieren photographs, and confirm his conclusions. It is further demonstrated, by an explicit expression, that the sound power absorbed by the wake may be positive or negative, depending on Mach number and source position. So the process of vortex shedding does not necessarily imply an attenuation of the sound.In the viscous model a high-Reynolds-number approximation is constructed, based on a triple-deck boundary-layer structure, matching the harmonic plane wave outer solution to a known incompressible inner solution near the edge, to obtain the viscous correction to the Kutta condition.

Sound diffraction at a trailing edge

This paper determines the ray structure of a spinning acoustic mode propagating inside a semi-infinite circular cylindrical duct, and thereby determines the ray structure of the field radiated from the end of the duct. Inside the duct, but outside of a caustic cylindrical surface, the rays are piecewise linear helices; on striking the rim of the end-face of the duct, these rays produce ‘Keller cones’ of diffracted rays. The cones determine the structure of the radiated field: for example, no rays penetrate two cone-shaped far-field quiet zones centred on the duct axis; two rays pass through each point in a forward loud zone; and one ray passes through each point in a rearward loud zone. The two rays through each point in the forward loud zone interfere to produce an oscillatory directivity pattern. One quarter of the rays on each cone point back inside the duct and produce the reflected field. Thus the rim of the end-face of the duct acts as a ‘ring source’, in which the radiated and reflected fields have their origin. Every propagating duct mode determines a polar angle and an azimuthal angle; these are taken as parameters specifying the mode and are used to calculate the positions and angles of all the rays. The mathematical method on which the paper is based is Debye's approximation for the Bessel function which appears in the expression for the duct modes; the approximation shows also that the duct contains a region of smooth helical rays on which the field consists of inhomogeneous waves: this region is the inner cylinder, lying inside the annulus of piecewise linear helical rays. The results of the paper are very promising for the application of Keller's geometrical theory of diffraction to detailed calculations of the sound radiated from aeroengine ducts. An alternative description of the field, using Cargill's meridional rays, is summarized.

Sound radiation from a cylindrical duct. Part 1. Ray structure of the duct modes and of the external field

This study compares a hybrid-electric aircraft featuring a propulsive empennage and over-the- wing, distributed-propulsion to a conventional regional turboprop. The impact of multiple design parameters, mission requirements, and technology assumptions on maximum take-off mass and payload-range energy efficiency is evaluated, in order to illustrate the sensitivities of the design. A preliminary sizing method that incorporates aero-propulsive interaction effects is used to obtain rapid estimations. Results show that, for the baseline mission, the hybrid electric variant is 2.5% heavier and consumes 2.5% more energy than the reference aircraft. In this process, several key design guidelines and challenges for distributed-propulsion aircraft are identified. Firstly, when comparing a hybrid-electric configuration to a conventional one, each aircraft must be sized at its optimum cruise altitude for the same payload and range requirements. Secondly, the hypothetical advantages of distributed propulsion described in literature do not easily lead to a benefit at aircraft level, if the aero-propulsive interaction effects and associated dependencies are incorporated in the design process. Thirdly, the power-control parameters affect practically all characteristics of the aircraft, and the optimal control strategy is highly dependent on the aero-propulsive interaction. The results suggest that the proposed configuration can constitute a low-noise alternative for the regional transport market if the performance of the over-the-wing distributed-propulsion system is optimized.

/pdf/preliminary-sizing-of-a-hybrid-electric-passenger-aircraft-4k64y2lmyo.pdf

Preliminary Sizing of a Hybrid-Electric Passenger Aircraft Featuring Over-the-Wing Distributed-Propulsion

Ase ightspeedincreasesinto thehypersonicregime, stagnation pressureand temperatureinsidean airbreathing engine become so great that, for practical structures of acceptable mass, the e ow must pass through the engine at supersonic speeds so as to limit the static pressures, hence the term scramjet (supersonic combustion ramjet ). The classicapplication forthescramjet isto along-rangeairlinerusing hydrogenfueland offeringextended hypersonic e ight. The scramjet is also widely accepted as propulsion for cruise missiles and as a possible complement to the rockets conventionally used for space launchers. The topic of how scramjets may best be used, especially in the near term, is explored. In particular, the advantages are demonstrated of using hydrocarbon fuel such as kerosene in a scramjet-powered second stage of a two-stage-to-orbit aerospace plane. Nomenclature Ai = intake capture area CL = lift coefe cient (referred to planform area ) D = drag f = acceleration g = acceleration due to gravity h = specie c enthalpy L = lift L=D = lift-to-drag ratio .L=D/p = L=D calculated from pressure forces .L=D/v = L=D calculated from pressure lift and the sum of pressure and friction drags M = mach number n = number of shock waves in scramjet intake (Fig. 3) p1 = ambient pressure T = thrust or temperature t = time 1f = gain in acceleration

Domain of the Scramjet

This paper addresses the aerodynamic performance and numerical modeling of over-the-wing propellers. Installing the propeller above a wing has the potential to increase wing lift-to-drag ratio, high-lift capabilities, and to reduce flyover noise. However, the prediction of its performance is difficult, since research on the aerodynamic interaction effects of over-the-wing propellers has been limited so far. For this reason, an exploratory wind tunnel campaign was performed with a wing featuring a fowler flap. A single propeller was installed above the wing at different chordwise locations and inclination angles. Wing surface-pressure and wake-pressure measurements showed strong, bilateral aerodynamic coupling between the propeller and wing. A configuration with the propeller attached to the flap showed wing lift increases of 8% and 3% in cruise and high-lift conditions, respectively. The key findings of the wind tunnel campaign were used to validate a low-fidelity numerical tool, which combines a non-uniform inflow blade-element model for the propeller, a panel method for the wing, and a vortex lattice model for the propeller slipstream. The numerical model was used to assess the effect of propeller axial location and diameter. Results indicated that the optimal axial propeller position is near the trailing edge of the wing, and that reducing the propeller diameter at constant thrust coefficient at this location is beneficial for distributed propulsion applications. The tool allows a rapid computation of over-the-wing propeller and wing performance in cruise conditions. This enables an efficient design space exploration during the conceptual design process of such configurations.

/pdf/aerodynamic-investigation-of-an-over-the-wing-propeller-for-544cc3k0rx.pdf

Aerodynamic investigation of an over-the-wing propeller for distributed propulsion

Noise shielding for aircraft

Flows with heat addition

Certain shapes of high speed aircraft, such as caret wings, make use of twodimensional flow fields and in principle they can be propelled by purely external heat addition without any ducted flow. Suitable under-surfaces can be designed by an exact numerical method [1]. Various examples of the method are given, and comparisons are made of propulsive efficiency, temperature distribution and other variables.

External heat addition for supersonic propulsion

A method of calculation is presented for determining the inviscid flow field with heat addition, that surrounds the recirculating bubble, in an axisymmetrie experiment with burning behind a blunt base. As in previous work, the streamlines are chosen and the mode of heat addition determined, although in the present paper it is the streamtube area that is prescribed rather than the streamlines themselves. A discussion is given of the relation to experiment, and of the efficiency of the process.

E.G. Broadbent

Papers

Noise shielding for aircraft

Flows with heat addition

External heat addition for supersonic propulsion

Axisymmetric flow with heat addition to simulate base-burning

Some shockless axisymmetric flows with heat addition