FlightGlobal.com
Home
Premium
Archive
Video
Images
Forum
Atlas
Blogs
Jobs
Shop
RSS
Email Newsletters
You are in:
Home
Aviation History
1927
1927 - 0834.PDF
SUPPLEMENT TO FLIGHT OCTOBER 27, 1927 THE AIRCRAFT ENGINEER Other quantities to be determined are : the most suitable stiffness for the connecting shaft, the necessary shaft strength and the horse-power to be dissipated by the damper under the conditions that make this a maximum. The present investigation was initiated to throw light on the subject generally. In the mathematical working the damper is assumed to be of the Lanchester viscous-friction type, but the investigation can be used in relation to the solid-friction type. A damper of known polar moment of inertia is assumed to be driven by a shaft of known stiffness. Damping is assumed to occur within the engine itself, but not in the damper driving shaft or bearings. The engine is assumed to be either of single-crank type or of multi-crank type. in which the crank portion of the shaft is very stiff relative to the shafting between the airscrew (or large flywheel) and the first throw. Expressions are derived for the horse-power absorbed in the damper under any set of conditions, and for the magnitude of torque variation in the damper drive. The effect of the damper is virtually to lower the speed of synchronism to an extent depending upon the moment of inertia of the damper fly-wheel and the value of the damper coefficient D. To produce effects of practical utility it appears that the damper flywheel needs to have a polar moment of inertia at least equal to about one-half that of the total crank mass when the journals and cranks are relatively rigid. A second case is treated where a viscous friction damper or a flywheel is connected electrically to the crank mass. Such a system can vibrate freely in two different ways, and there are thus two peaks to the amplitude speed curves for forced vibrations. Experiments are to be made of the direct variations in a multi-crank engine in addition to experiments on a single- crank engine, so as to obtain data illustrative of the theory developed in this report. LATERAL STABILITY AT LOW SPEED. PART I. : MEASUREMENT OF ROLLING MOMENTS FOR THREE WINGS AT LOW RATES OF ROLL. PART II: PRES- SURE MEASUREMENTS ON A WING WHILST ROTAT- ING AT LOW SPEEDS. By S. SCOTT HALL, B.SC, D.I.C. Edward Busk Memorial Studentship, 1925-26. R. & M. No. 1075. (Ae. 257.) (13 pages and 45 diagrams.) January. 1927. Price l.s. net. In the general report of the Stability and Control Panel upon the control of stalled aeroplanes. R. & M. No. 1000* (see page 74), it is pointed out that experimental data have not been obtained for wings other than the conventional thin-wing biplane cell in general use in Britain :— " For other arrangements of wings or for wings of thick section, or for monoplanes, the relative magnitude of these various factors (governing the motion in stalled flight) may differ somewhat from those which have been ascertained." Part I of the present report gives the results of tests of a thick section monoplane wing upon a continuous rotation balance at low rates of roll. The results obtained when this wing is tapered and twisted are also given, and for comparison with these two wings, corresponding results for a wing of R.A.F. 15 section and rectangular plan form are included. The results include autorotation curves for all three wings, and cover a range of air speeds from 40 to 70 ft. per second The greatest scale effect is shown for R.A.F. 30 unmodified, and the least for R.A.F. 15. The effect of tapering the twisting the R.A.F. 30 aerofoil is to shorten the range over which autorotation occurs, but also to increase the " violence '" of the motion. There is an indication (from comparison with other tests) that the shape of wing tips may have an appreciable effect on their characteristics in stalled flight. Part II deals with the pressure plotting of the same R.A.F. 15 wing whilst rotating over the same range of speeds as on the rotation balance, for an airspeed of 40 ft. per second only. The method employs the use of a long oil-sealed • The lateral control of stalled aeroplanes. General Report by theStability and Control Panel. bearing between rotating and stationary parts of the pressure system. The pressure distributions for different rates of roll arc given, and curves of rolling moment coefficients have been prepared for comparison with those obtained by the previous method. It is considered that the results show the method employed to be satisfactory, and it in regretted that more time was not available to extend the experiments. Higher rotational speeds might be employed to give data on spinning. THE PRESSURES ROUND A CYLINDER ROTATING IN AN AIR CURRENT. By A. THOU, B.SC.. Ph.D. R. & M. No. 1082 (Ae. 201). (12 pages and 9 diagrams.) November. 1926. These experiments were undertaken with the idea of obtaining information which might help to further the theory of the lift of a rotating cylinder. The pressure distribution has been determined for four sections along the length of the cylinder 8 cm. diameter spanning the 2-ft. wind channel at r V — 2. and for the centre section at r V = 0. 1. 2. 3 and 4. where V is the wind speed and v is the peripheral velocity. In some respects the results agree with what would be expected from the known lift, but the curves obtained show some interesting features. These Reports are published by His Majesty's Stationery Office, London, and may be purchased directly from H.M. Stationery Office at the following addresses : Adastral House, Kingsway, W.C. 2; 28, Abingdon Street, London, S.W.I; York Street, Manchester ; 1, St. Andrew's Crescent, Cardiff ; or 120, George Street, Edinburgh ; or through any book- seller. MR. DRIGGS' ARTICLE ON PERFORMANCE In THE AIRCRAFT ENGINEER of September 22, 1927. a small section of Mr. Driggs' article on Aircraft Performance was inadvertently left out. This occurred on page 62, in the last column. The section omitted is given herewith, and should be inserted 18 lines from the bottom of the last column, between "Absolute Ceiling = 22.100 ft. (flight test)" and the next paragraph which commences "A large portion," &c. Case II :— A,,. = KWSA,,"- 1-0 ,: 27-793-6K (/max. 93-6 X 0-00397A*. = 206 sq. feet. Actual A,r = 108 sq. feet. Therefore Case II should give the best results. PK' ¥ . 1•00000872 F,VM,.-« -f —i- I > MI',' h-e VM,=11 VInin. - V \v — 59-1 m.p.h. ,4 _ 0-01113-f-0-(X)91 K ~ 01106 X 0-784 x 0-716 = 0-326. From Fig. 4, Oiling ^22,450 feet. Ceiling = 22,100 feet (flight test). AN OMISSION In Mr. Mettam's article on '"Handicapping by Formula," published in THE AIRCRAFT ENGINEER of July 28, 1927, the first sentence at the top of the left-hand column on p. 47 should read : " lch] and &t2 must vary in a manner. &c." Will readers wishing to have their copies correct kindly make this addition. 750/
Sign up to
Flight Digital Magazine
Flight Print Magazine
Airline Business Magazine
E-newsletters
RSS
Events
