Wr Graham

Active flow control using a reduced order model and optimum control

An approach to developing active control strategies for separated ows is presented. The methodology proposed is applied to the incompressible unsteady wake ow behind a circular cylinder at Re = 100. Control action is achieved via cylinder rotation. Low order models which are amenable to control and which incorporate the full non-linear dynamics are developed using the Proper Orthogonal Decomposition (POD) technique. These models are generated from data provided by numerical simulation. The model predictions are found to be capable of describing control-induced departures from stationary conditions, making it feasible to envisage employing them for model-based control of the vortex-shedding ow. The use of these models and the resulting optimal control results are discussed.

Implementation of a phased microphone array in a closed-section wind tunnel

This paper describes the design and implementation of a phased microphone array system for aeroacoustic measurements in a closed-section wind tunnel at the Department of Engineering at the University of Cambridge. The tunnel has a test-section area of 1.67 x 1.22 m and can generate wind flows up to 60 m/s. The research started by conducting a feasibility study to find out whether an array could be used in the presence of the tunnel background noise. The boundary-layer pressure fluctuations were found to be correlated according to the Corcos model (Corcos, G. M., Journal of Fluid Mechanics, Vol. 18, No. 3, 1964, pp. 353-378.). Their level can be reduced by recessing microphones away from the flow, but the benefits of such an approach in this facility were found to be insufficient for it to be followed. Two different size arrays, for frequency ranges 650 ∼ 6500 Hz and 5 ∼ 50 kHz, were designed and installed. A range of characterization and shakedown experiments are described; in particular, it is shown that averaging can increase signal-to-noise ratio and that the array can detect sources up to 6 dB below the mean cross-spectral level of the background noise.

Trailing Vortices from a Wing with a Notched Lift Distribution

5Hsu, C.-H., Hartwich, P.-M., and Liu, C. H.,“Incompressible Navier– Stokes Computations for a Rounded-Edged Double-Delta Wing,” Journal of Aircraft, Vol. 25, No. 8, 1988, pp. 675, 676. 6Fujii,K., and Schiff, L. B., “Numerical SimulationofVortical Flows over Strake-DeltaWing,”JournalofAircraft,Vol. 27,No. 9, 1989,pp.1153–1162. 7Gordnier, R. E., and Visbal, M. R., “Unsteady Vortex Structure over a Delta Wing,” Journal of Aircraft, Vol. 31, No. 2, 1994, pp. 243–248. 8Ishizaka, K., Ikohagi, T., and Daiguji, H., “A High-Resolution Finite Difference Scheme for SupersonicWet-Stream Flows,” Transactions of the Japan Society of Mechanical Engineers, Series B, Vol. 60, No. 579, 1994, pp. 3887–3892 (in Japanese). 9Frenkel, J., Kinetic Theory of Liquids, Dover, New York, 1955, Chap. 7. 10Yamamoto, S., and Daiguji, H., “Higher-Order-Accurate Upwind Schemes forSolving the CompressibleEuler andNavier–StokesEquations,” Computers and Fluids, Vol. 22, No. 2/3, 1993, pp. 259–270. 11Yuan,X., Yamamoto, S., andDaiguji, H., “A Higher-ResolutionShockCapturing Scheme for Simulating Unsteady Three-Dimensional Transonic Flows in Turbomachinery,” AIAA Paper 94-3199, July 1994.

Aeroacoustic study of airfoil leading edge high-lift devices

The aim of this research is to attempt to mitigate the sources of noise in the leading-edge slat region without compromising performance or safety. Noise from the slat originates due to unsteady flow within the slat cove and in the tr ailing edge wake of the slat. In order to further understand and treat these sources, an aero acoustic study was conducted in the 5Ωby-4-foot Cambridge University Markham wind tunnel. The study was carried out using a two-part wing and slat model that spanned the tunnel horizontally. Separate low and high frequency phased microphone arrays consisting of 48 channels each were used to determine the source strengths associated with the noise with in the slat cove. The aerodynamic forces on the model were determined from separately-conducted overhead balance measurements. Modifications to the slat geometry were made in order to analyse their effectiveness in reducing the slat noise source str engths. The alterations reported here are taping over the upper surface gap and filling in th e slat cove region. The models were tested at angles of attack from six to sixteen degrees, an d at flow speeds of 20, 30, and 40 m/s. Boundary layer trips were used to simulate the full -scale, high Reynolds number flow. The tests confirmed that the slat is a significant sour ce of noise. A sealed slat gap is very effective at reducing the noise, but limits the aerodynamic p erformance. A filled cove on the other hand reduces the sources of noise by a few decibels . However, the concept is less effective than hoped due to the enhancement of other sources of noise. Also, the force balance results show that, after filling in the slat cove, the mode l stalls prematurely. Thus, for modifications reported here, any acoustic benefits are likely to be accompanied by aerodynamic performance penalties.

A ray tracing approach to calculate acoustic shielding by the Silent Aircraft airframe

The Silent Aircraft is in the form of a flying wing with a large wing planform and a propulsion system that is embedded in the rear of the airframe with intakes on the upper surface of the wing. Th ...

The effect of mean flow on the radiation efficiency of rectangular plates

The acoustic power radiated by a rectangular plate bounded by a uniform mean flow is investigated. Asymptotic expressions for the modal radiation efficiency of a simply supported plate are derived, and confirmed by comparison with numerical results. The asymptotic expressions show that the frequency at which a mode becomes an efficient radiator is reduced in the presence of mean flow, and hence that the critical frequency for a vibrating elastic plate is similarly lowered. The effects of these changes on the radiation damping of an aircraft panel in cruise, and on its radiation efficiency on landing, are investigated. The asymptotic results are then extended to more general plate edge conditions via analysis of appropriate infinite and semi–infinite problems. Hence the ‘3 dB rule’, that a plate with clamped edges, vibrating well below the critical frequency, radiates twice the power of a plate with simple supports, is shown to apply also in the presence of mean flow.

An engine demonstration of active surge control

Epstein, Ffowcs Williams and Greitzer presented in 1986 arguments that, provided small perturbation stability models were relevant to engine surge and rotating stall, then both could be avoided by active control. The idea was tested and validated on turbocharger compression systems at MIT and Cambridge University. Engine experiments are more difficult, the surge more violent and the signals usually contaminated by noise. This paper describes the early results obtained on a gas turbine engine subjected to the same control strategy. The control system can be switched on when the engine operates in violent surge and smooth operation is restored. This proof that the technique has practical relevance is interesting enough to report in advance of detailed investigations of its scope. This paper reports the experimental details, describes the relevant theoretical model and gives the first positive test results.Copyright

Analytical approximations for the modal acoustic impedances of simply supported, rectangular plates.

Coupling of the in vacuo modes of a fluid-loaded, vibrating structure by the resulting acoustic field, while known to be negligible for sufficiently light fluids, is still only partially understood. A particularly useful structural geometry for the study of this problem is the simply supported, rectangular flat plate, since it exhibits all the relevant physical features while still admitting an analytical description of the modes. Here the influence of the fluid can be expressed in terms of a set of doubly infinite integrals over wave number: the modal acoustic impedances. Closed-form solutions for these impedances do not exist and, while their numerical evaluation is possible, it greatly increases the computational cost of solving the coupled system of modal equations. There is thus a need for accurate analytical approximations. In this work, such approximations are sought in the limit where the modal wavelength is small in comparison with the acoustic wavelength and the plate dimensions. It is shown that contour integration techniques can be used to derive analytical formulas for this regime and that these formulas agree closely with the results of numerical evaluations. Previous approximations [Davies, J. Sound Vib. 15(1), 107-126 (1971)] are assessed in the light of the new results and are shown to give a satisfactory description of real impedance components, but (in general) erroneous expressions for imaginary parts.

The influence of curvature on the sound radiated by vibrating panels

The influence of curvature on sound radiation by a rectangular panel is investigated by considering the situation where the panel is set in an infinite cylinder. As in the flat panel case, acoustic effects may be characterized by a set of modal impedances, whose relationship to the corresponding flat panel terms is calculated for large cylinder radius. It is found that the impedances arising from the external acoustic field tend to their flat panel counterparts, but those associated with the internal field only do so under suitable conditions on the cylinder internal acoustic damping. Specifically, the interior wall must be sufficiently hard that it effectively appears rigid around the panel, but sufficiently dissipative that reflected ray and near‐field line‐of‐sight contributions to the impedance are negligible. The conditions are investigated in detail for parameters appropriate to a wide‐body civil aircraft, and curvature is found to be potentially important for the mode classes of interest.

Stability of Hybrid-Wing-Body-Type Aircraft with Centerbody Leading-Edge Carving

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.