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pressure distribution more especially in the subsonic flow region. This could be explained by assuming that the flow in this region separated from the duct wall. The results of tests on models of two forms of annular entry (the Q1 and E24/43 entries respectively) Under this heading are published regularly abstracts of all Reports and Memoranda of the Aeronautical showed that two types of flow regime were possible Research Council, Reports and Technical Memoranda of the United States National Advisory Com depending on the outlet pressure. For low outlet pressures a shock-wave was formed at the lip of the mittee for Aeronautics and publications of other similar Research Bodies as issued entry and the flow passed into the entry but for higher outlet diffuser pressures there was no distinct shock- wave from the annular lip and the flow through the annulus was reversed. The outlet diffuser pressure at which the flow changed direction for each entry tested was about 0·5 of the free-stream pitot pressure GREAT BRITAIN The influence of the following parameters is in and was thus considerably lower than for the un vestigated : AERONAUTICAL RESEARCH COUNCIL obstructed type of duct. (a) The degree of sweep. REPORTS AND MEMORANDA With the airflow passing into the annulus the flow (b) Wing torsional and flcxural stiffness. through the Q1 entry was independent of the outlet H.M. Stationery Office, London (c) Wing plan form. pressure over the range of Mach numbers tested and (d) Aileron plan form. R. & M. No. 2802. Two-dimensional Wind-tunnel was about 0·88 of the flow through an area equal to Families of curves are given for extended variation Interference. By L. G. Whitehead. June, 1950. (4s. 6d.) the intake area in the free stream. For the E24/43 of these parameters which may be used for the direct Exact solutions are given for the inviscid flow past entry the airflow decreased as the outlet diffuser estimation of the reversal speed of a given wing by two cylindrical profiles in the centre of a stream of pressure increased probably due to changes in the interpolation. limited depth. The first of these relates to a nearly boundary-layer thickness at the annulus. A solution is given for the wing divergence speed circular cylinder and the second to a thin section of a swept wing. giving a constant pressure drop over the greater part SWEDEN The general results have been obtained using simple of its surface. The stream has cither parallel walls, modes of wing deformation but equations are quoted KUNGL TEKN1SKA HOGSKOLAN constant pressure walls, or the boundaries may be for any given modes of deformation and the adopted (ROYAL INSTITUTE OF TECHNOLOGY) partly parallel and partly of the constant pressure modes are compared with the actual deformations INSTITUTIONEN FOR FLYGTEKNIK type. For the thin profiles the changes of thickness produced by the aerodynamic loading for an extreme (DIVISION OF AERONAUTICS) ratio required to give the same pressure distribution case. A suggestion is put forward for improving the Stockholm 70 as in an unlimited flow are found. accuracy of the semi-rigid approach by an iterative TECHNICAL NOTES method of solution and the flexural mode of distortion K.T.H. Aero T.N. 41. On Calculating Incompres is investigated for a particular case. R. & M. No. 2810. Downwash Measurements sible Turbulent Boundary Layers with Arbitrary Pres behind a 12-ft. diameter Helicopter Rotor in the sure Distribution. By H. Schuh. 24-ft. Wind Tunnel. By R. A. Fail and R. C. W. R. & M. No. 2821. Critical Mach Numbers for The turbulent boundary layer can be characterized Eyre. September, 1949. (3s. 6d.) Thin Untanered Swept Wings at Zero Incidence. By with sufficient accuracy by two quantities: The S. Neumark. November, 1949. (20s.) Some measurements of downwash have been made momentum thickness and a profile parameter. More in a plane behind a 12-ft. diameter helicopter rotor In this paper, which is a continuation of two earlier recent papers show agreement between various authors ones (R. & M.'s 2713 and 2717), the subsonic flow over a range of shaft inclination and tip speed ratio. on the method of calculating the momentum thick past untapered swept wings, at zero incidence, is In the various operating conditions, the tunnel tests ness, but for the profile parameter different methods further investigated using linear theory. Methods for are in reasonable agreement with the theoretical have been put forward. For calculating the latter calculating 'lower' and 'upper' critical Mach numbers results for the appropriate type of loading. quantity an integral form of the boundary layer equa are given, the solution of the main problem being tion is used in the present paper in a form which preceded by a short analysis of critical Mach numbers allows the coefficients in this equation to be opened R. & M. No. 2812. Design and Use of Counting for the simpler cases of infinite wings (straight, sheared directly from experiments. For flow with adverse pres Accelerometers. By J . Taylor. June, 1950. (3s. 6d.) and yawed). sure gradients, solutions of the equation for the profile The fundamental principles underlying acceleration The determination of critical Mach numbers de parameter are obtained in a particularly simple form, recording by means of a counting accelerometer are pends on the knowledge of velocity distribution over from which a diagram is derived for predicting separa analysed. The essential design requirements for a the wing surface, the problem dealt with in the pre tion. However, the new method for calculating the counting accelerometer are presented. A design that vious reports mostly for the case of the simple bicon profile parameter is only applicable to fully developed has been specially made to meet these requirements vex parabolic profile. These earlier results have been turbulent flow and it does not cover the transitional is described. Both mechanical and electrical counting extended here to cover a wide class of profiles. Hence region from the laminar to the turbulent flow regime. are considered, but mechanical counting is found to it has been possible to determine critical Mach num be superior. bers for wings with four different profiles, showing U.S.A. the effect of thickness ratio and of angle of sweep-back NATIONAL ADVISORY COMMITTEE FOR (or sweep-forward) in each case. The method applies R. & M. No. 2814. Velocity Distribution on Thin AERONAUTICS strictly to wings of large aspect ratio, but no significant Bodies of Revolution at Zero Incidence in Incompres TECHNICAL REPORTS corrections are necessary except for very low aspect sible Flow. By S. Neumark. July, 1950. (12s. 6d.) ratios. Government Printing Office, Washington, D.C. A new method of determining velocity distribution The results and examples, illustrated by a number (Foreign Annual Subscription Rate: 11-0 dollars) on slender bodies of revolution in axial flow is ex of tables and graphs, provide a basis for more general pounded, analogous to the linear perturbation method 1137. Initial Results of Instrument-Flying Trials discussion. Several conclusions concerning the practi widely used for slender symmetrical profiles in two Conducted in a Single-Rotor Helicopter. By Aimer D. cal use of swept-wing designed are presented. dimensions. The proposed method leads to simple Crim, John P. Reeder and James B. Whitten. approximate formulae for velocity distribution on a Instrument-flying trials have been conducted in a R. & M . No. 2827. Report on the Flow Phenomena body, once the equation of the meridian line is given, single-rotor helicopter to determine the adequacy of at Supersonic Spe?d in the Neighbourhood of the Entry either in the form of a polynomial, or a square root of existing longitudinal flying-qualities requirements of a Propulsive Duct. By G. H . Lean. (4s. 6d.) one. The new method avoids many inconveniences under instrument conditions. In addition, lateral- The present work continued that reported in an of the older procedures, and is much more rapid. directional characteristics were examined. The suit N.P.L. Engineering Division Report of 1944 and Although theoretically applicable to bodies of small ability, for helicopter use, of standard aeroplane extended some of the results there described to lower thickness only, it works with satisfactory accuracy instruments was also investigated. entry Mach numbers (1·3 to 1·9). It was found, as up to quite considerable thickness ratios. It has been in the previous Report, that with a parallel entry further improved by taking into account not only axial 1138. Study of Inadvertent Speed Increases in duct followed by a straight divergent diffuser of but also radial velocity components, following a sug Transport Operation. By Henry A. Pearson. 10 deg. total angle the flow inside the parallel tube gestion of Lighthill's supersonic theory. It may be Some factors relating to the speed and Mach num was supersonic provided the outlet pressure of the easily applied to compressible subsonic flow. ber margins required in operation of transport aero diffuser was less than a certain critical value (about The method has been used for computing velocity planes are discussed for the purpose of indicating how 0·93 of the upstream pitot pressure). In this case the distributions on twelve different bodies, of seven these margins should vary with the aeroplane qualities. mass flow of air through the tube was equal to that different thickness ratios (0·04-0·28) each, so as to Most of the cases investigated show that on a percent calculated, assuming that all the air incident on the exhibit the most characteristic features in typical age basis smaller margins should be required of the internal section of the tube entry passed through it. cases, and especially to show some unexpected effects faster aeroplane than the slower aeroplane. Equations For pressures higher than the critical value the flow of thickness changes. Several practical conclusions are suggested which allow these margins to be esti became subsonic at the duct entry, a shock-wave have been derived from the examination and com mated. was formed at the entrance lip and the rate of airflow parison of these results. through the tube decreased. The method may find useful applications in the 1139. Charts and Approximate Formulas for the Similar results were obtained for a uniformly design of fuselages, nacelles and wing junctions, and Estimation of Aeroclastic Effects on the Lateral Con divergent tube of 7 deg. total angle; in this case, especially in determining critical Mach numbers for trol of Swept and Unswept Wings. By Kenneth A. Foss however, an outlet pressure equal to 0·97 of the up such bodies. and Franklin W. Diederich. stream pitot pressure was attained before the shock- Charts and approximate formulas are presented for wave left the lip. R. & M. No. 2817. Aileron Reversal and Wing the estimation of static aeroclastic effects on the span- For outlet pressures less than the critical the flow Divergence of Swept Wings. By E. G. Broadbent and wise lift distribution, rolling-moment coefficient, and was supersonic for a distance inside the duct entry Ola Mansfield. September, 1947. (7s. 6d.) rate of roll due to aileron deflexions on swept and un depending on the outlet pressure, the flow becoming swept wings at subsonic and supersonic speeds. Some A method of solution for the aileron reversal speed subsonic further along the duct. The assumption of design considerations brought out by the results of this of a swept wing (with emphasis on sweepback) is unidimensional flow in the duct led to results which paper are discussed. developed on the lines of strip and semi-rigid theories. showed considerable disagreement with the observed 130 Aircraft Engineering
Aircraft Engineering and Aerospace Technology – Emerald Publishing
Published: Dec 1, 1954
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