If the fundamental frequency of this species of pendulum vibration is numerically equal to nN where n is the number of blades and N is the frequency of rotation of the rotor then serious resonant forced vibration may ensue and it would appear that this is Under this heading are published regularly abstracts of all Reports and Memoranda of the Aeronautical quite likely to occur in practical cases with the blades Research Council, Reports and Technical Memoranda of the United States National Advisory Com in vibration in the plane of rotation of the rotor. mittee for Aeronautics and publications of other similar Research Bodies as issued. U.S.A. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ing the size of the wing and tunnel. These functions FRANCE TECHNICAL REPORTS have been tabulated and used to estimate the effect Superintendent of Documents, Government Printing PUBLICATIONS SCIENT1FIQUES ET TECH on C and C , for wings of a variety of sizes and Office, Washington, D.C. L M NIQUES DU MINISTERS DE L'AIR shapes. The variation of mean induced incidence (Foreign subscription rate: 11.0 dollars) Magazin C.T.O., 2 Avenue dc la Porte d'Issy, Paris, with sweep and taper was found to be small. A formula (15e.) is given for computing the residual correction to C RAPPORTS for each special case. In the cases computed the cor 1102. The Linearized Characteristics Method and rections to aerodynamic centre were found to be No. 288. Essai sur la Convection Naturelle. (An its Application to Practical Non-Linear Supersonic negligible for most purposes. Essay on Natural Convection.) By P. Vernotte. Problems. By A. Ferri. The tunnel sections considered are square, duplex, (600 fr.) The method of characteristics has been linearized 18×7 and 9×7. A system of equations involving five unknowns is by assuming that the flow field can be represented as formulated to describe natural convection in any a basic flow field determined by nonlinearized methods R. & M . No. 2787. Experiments on the Flow past a fluid. The approximate relation 0 =Ah3+Bh6 is and a linearized superposed flow field that considers Porous Circular Cylinder fitted with a Thwaites Flap. By derived for the velocity distribution created by a hot small changes in boundary conditions. The method R. C. Pankhurst and B. Thwaites, with an Appendix by body of temperature 0 above the surroundings, has been applied to two-dimensional rotational flow, W. S. Walker. October 1950. (7s. 6d.) where h is the 'convective disturbance' to pure con to calculations of axially symmetric flow, to slender This paper describes wind-tunnel experiments on duction. It is shown that the coefficient of natural bodies without symmetry, and to wing problems. a porous circular cylinder of 3 in. diameter fitted with convection can be expressed, over a large range of 0 , a Thwaites Flap. Measurements were made of the by ∞=A+B0 1/6(1+C0 ), which reduces to pure 1103. Generalized Theory for Seaplane Impact. O 0 pressure distribution at mid-span, together with a conduction as a limiting case. In practice, the vari By B. Milwitzky. number of wake traverses, over a range of suction ability of ∞ with 0 is not so serious as might be The motions, hydrodynamic loads, and pitching quantity, flap size, wind speed and flap setting. expected, and can often be assumed constant so as moments experienced in impacts of V-bottom sea The distributed suction effectively prevented to make theoretical calculations possible. Some ex planes are analysed and compared with experiment. boundary-layer separation and enabled a close ap perimental results, and methods of studying con The analysis is presented in terms of generalized proximation to potential flow to be achieved. The vection, under stationary conditions, are discussed. variables which are related through a single para flap was essential to the attainment of steady flow meter, called the approach parameter K. For use in conditions with suction; without a flap the pressure the design of seaplanes, charts are presented showing GREAT BRITAIN recovery at the rear of the cylinder was incomplete the generalized relationships which apply throughout and the pressure distribution fluctuated. In view of this the impact; charts are also presented which show the AERONAUTICAL RESEARCH COUNCIL unsteadiness in the flow without a flap, the circulation variations with K of the generalized variables at the H.M. Stationery Office, London could scarcely be expected to remain, as had previously instants of maximum acceleration, maximum pitching been conjectured, when the flap was withdrawn. moment about the step, maximum penetration, and R. & M . No. 2769. Some Additional Notes on the Over the limited Reynolds number range of the exit during rebound. The effects of chine immersion Derivation of Airworthiness Performance Climb tests, the minimum suction quantity needed to prevent on the maximum load are also determined. Extensive Standards. By A. K. Weaver. February 1950. (3s. 6d.) separation (C ) appeared to be proportional to experimental data are presented to permit evaluation After the publication of a report (R. & M. 2631) Qmin R— n with n rather greater than the theoretical value of the theoretical results. on the derivation of airworthiness performance climb of ½ for a laminar boundary-layer. For a given Rey standards, various subsidiary points raised in the nolds number, C decreased with increasing flap 1104. Preliminary Investigation of a New Type of course of discussions were examined. Some of these Qmin size. Supersonic Inlet. By A. Ferri and L. M. Nucci. have been collected together in the present note. They At small flap deflexions the wake could be com A supersonic inlet with supersonic deceleration of are in the nature of elaborations of the original method pletely suppressed. The suction quantities required the flow entirely outside of the inlet is considered. and include a refined method of deriving the take-off increased with flap deflexion, owing to the increased A particular arrangement with fixed geometry having climb standard, a method of treating interdependence severity of the adverse pressure gradient over the of engine failure and a method for including the effect a central body with a circular annular intake is rear of the cylinder. With the available pump the of sideslip in the margin allowed for pilotage errors. analysed, and it is shown theoretically that this ar maximum lift coefficient attained was about 9, but The main principles set forth in the earlier report rangement gives high pressure recovery for a large there is no reason to doubt that still higher values remain unaffected. range of Mach number and mass flow and, therefore, is would have been reached with greater suction. practical for use on supersonic aeroplanes and missiles. The wake traverse measurements indicated only R. & M . No. 2774. The Effect of Sweepback on the Experimental results confirming the theoretical slight hysteresis according as the suction was increas Fundamental Derivative Coefficient for Flexural analysis give pressure recoveries which vary from ing or decreasing. Motion. By J . B. Bratt and K. C. Wight. October 95 per cent for Mach number 1·33 to 86 per cent for 1950. (6s.) Mach number 2·00. These results were originally R. & M. No. 2797. The Performance after Power presented in a classified document of the N.A.C.A. Measurements have been made with new equip Failure of a Helicopter with Blade Pitch Control. By in 1946. ment, designed for derivative tests in a 9 x 7 ft. tunnel, F. O'Hara and H. A. Mather. September 1951. to determine the effect of sweepback on the derivatives (4s. 6d.) λφ and λφ for a rectangular aerofoil of aspect ratio 6. 1105. Chordwise and Compressibility Corrections In Part 1 a review is made of helicopter performance A numerical reduction was observed in each case, to Slender-Wing Theory. By H. Lomax and L. Sluder. after engine failure. The transition from powered amounting to 15 per cent for λφ and 20 to 30 percent Corrections to slender-wing theory are obtained by operation to autorotation is discussed and a theoreti for λφ over the rangeω=1· 0 to 1·5 with a sweepback assuming a spanwise distribution of loading and de cal analysis of the motion is given for a single-rotor angle or 41·3 deg. termining the chordwise variation which satisfies the helicopter with blade-pitch control. The technique appropriate integral equation. Such integral equations Values of λφ and λφ for the swept model were ob of landing from a steady autorotative glide is dealt tained from measurements relating to oscillation are set up in terms of the given vertical induced velo with briefly; the possibility is indicated of making about an axis perpendicular to the leading edge. city on the centre line or, depending on the type of a safe landing before the transition to steady auto A comparison of λφ and λφ with available theoretical wing plan form, its average value across the span at rotation has been completed. Reference is also made results for finite aspect ratio is made and good agree a given chord station. The chordwise distribution is to the case of engine failure so near the ground that ment observed in the case of the former. The less then obtained by solving these integral equations. a safe landing may be made by increasing the blade satisfactory agreement with A,,, is thought to be due to Results are shown for flat-plate, rectangular, and pitch to make immediate use of the rotor energy. the lower accuracy of the theoretical values. triangular wings. In Part II, tests made to investigate the performance Some difficulty was experienced in the interpreta of a Hoverfly I in the transition to autorotation fol tion of the measurements on the swept aerofoil due 1106. The Langley Annular Transonic Tunnel. lowing power cut in level flight are described; parti to distortion of the model during oscillation. The effect By L. W. Habel, J . H. Henderson and M . F. Miller. cular attention is given to the minimum rotor speed is examined in detail in an appendix and a method of The development of the Langley annular transonic attained and to the height lost during the transition. correction devised. tunnel, a facility in which test Mach numbers from Tests were made to investigate the performance for A comparison between measurements of λφ. for the 0·6 to slightly over 1·0 are achieved by rotating the immediate reduction of pitch only; the need for quick nswept aerofoil and earlier measurements made with test model in an annular passage between two con pitch reduction however is stressed because of the the same model by the method of decaying oscillations centric cylinders, is described. Data obtained for two- rapid fall-off in rotor speed following power-cut at ves satisfactory agreement. dimensional aerofoil models in the Langley annular high pitch. transonic tunnel at subsonic and sonic speeds are R. & M. No. 2777. Corrections for Symmetrical shown to be in reasonable agreement with experi ept and Tapered Wings in Rectangular Wind Tun- R. & M. No. 2801. Resonant Vibration of Helicopter mental data from other sources and with theory when Is. By W. E. A. Acum. April 1950. (9s.) Rotor Blades. By J . Morris. June 1950. (1s. 6d.) comparisons are made for nonlifting conditions or In this report it is shown that in the case of wings The blades of the operative rotors of helicopters for equal normal-force coefficients rather than for nth straight leading and trailing edges the inter- are usually hinged both in the lift and rotational equal angles of attack. The trends of pressure distri vrence upwash due to the images of the wing in the planes and it is because of this articulation that the butions obtained from measurements in the Langley wind-tunnel walls may be determined in terms of blades in the course of rotation are akin dynamically annular transonic tunnel are consistent with distri three functions, I , I and I , of the parameters defin to 'pendulum vibration dampers.' butions calculated for Prandtl-Meyer flow. 1 2 3 July 1954
Aircraft Engineering and Aerospace Technology – Emerald Publishing
Published: Jul 1, 1954
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