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- Publisher
- Emerald Publishing
- Copyright
- Copyright © Emerald Group Publishing Limited
- ISSN
- 0002-2667
- DOI
- 10.1108/eb033150
- Publisher site
- See Article on Publisher Site
Abstract
340. Automatic Curve Follower Using Polarized Light Beam Chopper. By T. Isobe, H. Hatanaka and I. Uchida. December 1958. This paper describes a new automatic curve follower, which can follow a graph given in a thin black line on a Under this heading are published regularly abstracts of all Reports and Memoranda of the Aeronautical white paper or a curve photographically recorded as an Research Council, Reports and Technical Memoranda of the United States National Advisory Com oscillogram. The detector consists of two illuminators of the line to be followed; a microscope objective lens mittee for Aeronautics and publications of other similar Research Bodies as issued. by which the image of the line is formed; a polaroid disk which is fixed at the image plane and is divided into two equal parts having their principal planes perpendicular to each other; a rotating polaroid disk driven by a synchronous motor; and a phototube. GREAT BRITAIN 3105. Aerodynamic Characteristics in the Approach, The beam of light passed successively through the two Superfine and Negative Pitch Ranges of two Four- polaroid disks flickers alternatively at the two parts of AERONAUTICAL RESEARCH COUNCIL Bladed Propellers with N.A.C.A. Series 16 Blade the field. But the total flux, which the phototube H.M. Stationery Office, London Sections. A Report of Tests Carried out Jointly by the receives, remains constant so far as the curve runs Technical Staffs of de Havilland Propellers Limited and through the centre of the field. When the curve Rotol Limited. April 1957. (8s. 6d.) CURRENT PAPERS deviates from this position, the deviation is picked up by the phototube as an alternating change of the total 427. The Flow Upstream of Finite Span Spoilers at Tests were conducted on two similar full-scale light flux, having an amplitude proportional to the propellers in the 24-ft. wind tunnel at the Royal Air Supersonic Speeds. By A. Stanbrook. September 1957. deviation and a reverse phase corresponding to its craft Establishment by representatives of Rotol Ltd. (1s. 9d.) direction. This detector constitutes a servomechanism and de Havilland Ltd. to measure the aerodynamic It is suggested that the flow upstream of swept and in conjunction with a voltage amplifier, a thyratron characteristics over a blade-angle range of —20 deg. unswept spoilers in a supersonic stream may be circuit and a split field motor, which drives the lead to +40 deg. measured at the 0·7 radius and over a explained in terms of a vortex type of flow. The pre screw of the follower. An appropriate masking of the range of advance ratio of 0·1 to 2·4. sence of this type of flow is shown to be consistent field much improved the performance of the detector. with experimental pressure distributions. 3106. Some Visual Observations of the Effects of Sweep on the Low-Speed Flow Over a Sharp-Edges 341. Experimental Study on the Supersonic Flow 428. Some Simple Conical Camber Shapes to Around Blunt-Nosed Bodies of Revolution. By H. Plate at Incidence. By N. C. Lambourne and P. S. Produce Low Lift-Dependent Drag on a Slender Delta Oguchi. March 1959. Pusey. January 1958. (14s.) Wing. By G. G. Brebner. September 1957. (4s.) Qualitative descriptions are given of the flow over a The flows around several blunt-nosed bodies of This note presents some of the theoretical aerody series of sharp-edged semi-span plates having various revolution were experimentally investigated over the namic characteristics of narrow delta wings having Mach number range from 3·76 to 4·78 at the intermit angles of sweep and at incidences as high as 30 deg. flat centre portions and two types of simple conical tent blow-down type supersonic wind tunnel of the The main interest was the type of flow resulting from camber designed to produce zero load at the leading Aeronautical Research Institute, Tokyo. In the separation at the leading edge and the changes that edge at the design CL. Slender wing theory is used in present experiment we are concerned with the flow occur when the angle sweep is varied within the range the calculations. It is shown that the drag due to lift is behaviour ahead of the bodies so that the interfero- 0 to 60 deg. both backward and forward. comparable with the low values found by Smith and metric study was primarily made. The flow was visualized in a water tunnel both by Mangier for wings with more wavy camber lines. The main results are summarized as follows: using air bubbles and also with small spherical particles of polystyrene, whilst information concern (1) The surface pressure referred to the stagnation 429. Atmospheric Turbulence encountered by Bristol ing the surface flow on the plates was obtained using pressure, shock-wave pattern, sonic-line pattern, and Freighter Aircraft in United Kingdom, West Africa and flow directions were found to be closely related to oil-film techniques in a wind tunnel. New Zealand. By J . R. Heath-Smith. October 1958. those found both experimentally and theoretically for Attention was mainly directed to the flow over the (3s. 6d.) the case of flow with a much higher Mach number upper or suction surface and in the region immediately than in the present test. downstream of the plate. The observations trace out Gust accelerations were recorded during 1,500 (2) The region with the local Mach number lower the systematic changes that take place in the behaviour hours of flying below 10,000 ft. by Freighter aircraft in than 0·5 was confirmed to spread extensively ahead of the free vortex layer from the leading edge and the United Kingdom, West Africa and New Zealand. The of the nose. shape of the regions of reduced total head as the gust frequencies in the three regions are compared and sweepback is varied. (3) The shock-wave detachment distance was found the influence of the flight plan on the values recorded to be nearly independent of the nose shape when is discussed. referred to the radius of curvature of the shock wave. The average distance between gusts greater than (4) For bodies with the same nose radius, the shock- 10 ft./sec. E.A.S. within 1,000 ft. of the ground, wave detachment distances were found to be nearly varies from two miles in New Zealand to twelve miles ITALY independent of the body shoulders. in West Africa. Over the sea the distance is greater by MINISTERO DIFESA-AERONAUTICA a factor of 2 or 3. Centro Consultivo Studi e Ricerche SWEDEN Rome 431. The Effect of Axial Spacing on the Surge ROYAL INSTITUTE OF TECHNOLOGY Characteristics of Two Mismatched Axial Compressor MONOGRAPHS Department of Aeronautical Engineering Stages. By R. C. Turner. November 1956. (4s. 6d.) 4. An Exact Solution for Free Rectangular Plates Stockholm 70 This describes an investigation into two of the Laterally Loaded. By L. Broglio. December 1958. major factors governing the surge of multistage axial TECHNICAL NOTES compressors; the mismatching of the stages and the The general problem of a rectangular free plate of 47. Investigation of the Response to Random Wind spacing between them. The purpose is to contribute to constant thickness, under random load distribution is Gusts of a Typical Subsonic Fighter Aircraft in Level a fundamental understanding of surging. considered in this paper. It is shown that any given Flight With Mach Number 0 9 at Sea Level. By K. load distribution can be reduced to boundary loads Two highly mismatched stages were tested at five Flodin and M. Sundström. December 1957. and that a complete solution is reached when six interstage spacings over a wide range of flow coefficient. Frequency spectra and root mean square deviations The results were compared with those of seperate particular problems (according as to whether sym are computed for all velocity components, attitude tests on each stage. It was found that there was con metric or antisymmetric moments or reactions be angles, components of displacement, and load factor siderable mutual interference as regards the surge prescribed by the boundary conditions) are solved. for an aircraft in level flight during normal turbulence behaviour; and the aerodynamic characteristics gene conditions. For some of the outputs, the influence of rally were also affected. In particular, at the lower different modifications as to the wind gust spectra has spacings, the second stage considerably delayed the been investigated. surge of the first. JAPAN Calculations have been carried out on two assump AERONAUTICAL RESEARCH INSTITUTE tions: firstly that the wind disturbance is the same all REPORTS AND MEMORANDA UNIVERSITY OF TOKYO over the aircraft, and secondly that the disturbance Komaba, Meguro-ku, Tokyo 2982. Thermodynamic Charts for the Combustion has a velocity gradient the spectrum of which is that Products of Nitric Acid and Kerosene. By W. H. given by Decaulne and Summers. REPORTS Williams. July 1950. (7s.) 339. Fundamentals of Acoustical Silencers,!: Theory 49. A Special Method for Determining the Power Thermodynamic charts giving enthalpy, entropy, and Experiment of Acoustic Low-Pass Filters. By Spectral Densities of the Solution of Linear Differential temperature, pressure and specific volume, have been J . Igarashi and M. Toyama. December 1958. Equations with Stochastic Right Member and Coef constructed for the products of combustion of a ficients Consisting of Polynomials in the Independent hydrocarbon fuel (85 per cent C, 15 per cent H by Attenuation characteristics of several low-pass Variable. By M. Sundström. October 1958. weight) with nitric acid (98 per cent by weight). The acoustic filters have been studied theoretically and charts have been drawn for the following mixture experimentally. A method of electrical equivalent net A method is presented for transforming linear ratios of the propellants: (a) stoichiometric propor work was applied to the acoustic elements, such as differential equations with stochastic right member and tion; (b) 10 per cent by weight excess of fuel; (c) 50 per cavities, resonators, internal tube type cavities, and coefficients consisting of polynomials in the indepen cavities with absorbing material. Each element was cent by weight excess of fuel; (d) 10 per cent by weight dent variable into linear differential equations between represented by four terminal matrices, and attenuation excess of oxidant. Fourier transforms for limited intervals. From the of the system was calculated from their products. Graphs of combustion temperature and specific solutions of these equations, the corresponding power impulse against mixture ratio are also shown. The By an automatic recording system, frequency spectral densities and standard deviations of the un procedure for calculating the initial enthalpy of the characteristics of various acoustic elements and their known functions are derived. The method is developed separate propellants (before combustion) and that of combinations were measured. They showed good for only one single equation but can be extended to a system of equations. the propellant mixture is described. agreement with calculation. 258 Aircraft Engineering
Journal
Aircraft Engineering and Aerospace Technology – Emerald Publishing
Published: Aug 1, 1959