CoA. Reports

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Open Access
Aerofoil theory for swallow tail wings of small aspect ratio
(College of Aeronautics, Cranfield., 1950-10) Robinson, A.
A method is developed for the calculation of the aerodynamic forces acting on a ‘swallow tail’ wing of small aspect ratio. Lift, induced drag, and aerodynamic centre position of simple swallow tail wings (Fig. 1 (b)) are computed as an application. For a given incidence, lift and induced drag are, within the limits of the theory, proportional to aspect ratio and independent of speed. The chordwise life distribution rises linearly from zero at the apex, drops rapidly in the region of the root chord trailing edge, and then decreases gently to zero.
Open Access
The application of memo-motion to industrial operations
(College of Aeronautics, Cranfield., 1954-12) Norbury, C. J.
Memo-Notion, or Spaced Shot Photography, was introduced as a tool of Work Study Dr. Mundel, then of Purdue University, in 1946, as a means of reducing the cost of film analysis on long operations by using a camera driven by a geared down electric motor, giving exposures every second. Since its introduction Memo-Motion has been applied to a limited extent in the U.S.A., but it is not known to be in use at all in England.
Open Access
The combustion characteristics of a cylindrical-rod burner system
(College of Aeronautics, Cranfield., 1955-06) Goodger, E. M.
Combustion systems constructed of a multiplicity of small elements appear attractive from the aspects of low pressure loss and short chamber length. A combustion chamber has been proposed, by Spalding of Cambridge University, in which fuel is fed under gravity down the surfaces of vertical cylinders located normally to the air stream. Preliminary tests were made at Cambridge, and the work has been continued at Cranfield under a Ministry of Supply Contract.
Open Access
The economics of aircraft production : a study of the control of overhead costs in aircraft manufacture
(College of Aeronautics, Cranfield., 1954-03) Twiss, B. C.
This paper represents an attempt to appraise the different factors which must be considered in applying modern methods of cost control to aircraft production, and to suggest those positive measures best calculated to effect such control. The emphasis throughout is on control of overhead costs as these are by far the largest element in the total cost make-up.
Open Access
The effect of the sweepback of delta wings on the performance of an aircraft at supersonic speeds
(College of Aeronautics, Cranfield., 1947-03) Robinson, A.; Davies, F.T.
The variation with sweepback of total drag of an aircraft in level flight at supersonic speeds is calculated. It is shown that sweepback is not uniformly beneficial, but that in general the optimum amount of sweepback depends in the design speed and altitude.
Open Access
The effect of vibrations (sonic and subsonic frequencies) during the period of solidification on the mechanical properties of castings of gas turbine materials with special reference to H. R. Crown Max and Nimonic C. 75
(College of Aeronautics, Cranfield., 1955-09) Jagaciak, Jerzy
Test castings were designed to produce data which would help in the production of gas turbine blades as castings. For castings in H. R. Crown Max the investigation is limited to the effect of subsonic frequencies and range of amplitudes; but in the experiments conducted on Nimonic C. 75, both sub- sonic and sonic ranges of frequency were investigated.
Open Access
The equations of motion and energy and the velocity profile of a turbulent boundary layer in a compressible fluid
(College of Aeronautics, Cranfield., 1951-01) Young, A. D.
As far as the author is aware the derivation of the equations of motion and energy for a turbulent boundary layer in a compressible fluid have not yet been given in detail in any publication. To meet a possible need in this connection this paper puts on record the analysis underlying the equations quoted by the author in Chapter x of the forthcoming Vol. 111 of Modern Developments in Fluid Dynamics.
Open Access
The evaluation of matrix elements for the analysis of swept-back wing structures by the method of oblique coordinates
(College of Aeronautics, Cranfield., 1951-04) Lewis, S. R.
This report is an addition to the college of Aeronautic Report No. 31. The purpose of the report is to enable one to obtain the matrix elements used in the analysis of swept-back wing structures by oblique coordinated in a very rapid manner.
Open Access
An experimental investigation into some of the problems associated with stress diffusion in the vicinity of chord-wise cut-outs in the wing, and a comparison with existing theories
(College of Aeronautics, Cranfield., 1954-09) Brown, L. W.
Chord-wise openings in the skin between the spars of the wing are designed in some aircraft for undercarriage doors, bomb bay doors, and the wing fold joints of naval Aircraft. Stress concentrations exist in the region of these cut-outs where the load is transferred from the stringers and skin into the concentrated load carrying members. Two theories have evolved to predict the resulting behaviour of the structure. The stringer sheet' theory predicts an infinite shear stress in the corners of the sheet; the 'finite stringer' theory predicts a high, finite shear stress in the corners, the magnitude of which increases with the number of stringers.
Open Access
The flexure-torsion flutter of cambered aerofoils in cascade
(College of Aeronautics, Cranfield., 1955-12) Craven, A. H.
This report contains the results of a series of tests on the flexure-torsion flutter of cascades of aerofoils of 30° and 45° camber. The critical flutter speeds and frequencies in cascade are expressed as ratios of the values for the aerofoil in the isolated condition, The tests cover stagger angles between -30° and +30° and gap chord ratios up to 1.5 at a Reynolds number of 1 x 10 to the power of 5 based on aerofoil chord.
Open Access
Flutter and divergence of sweptback and sweptforward wings
(College of Aeronautics, Cranfield., 1950-06) Babister, A. W.
In this note, the equations of the flexural-torsional flutter of a swept wing are established, assuming the wing to be semi-rigid and fixed at the root. The general effect of sweepback, wing stiffness and position of the inertia axis are determined. The critical speeds for flutter and for wing divergence are determined (i) for incompressible flow (ii) for compressible flow, assuming a modified Glauert correction. The critical flutter speed is in general higher for a sweptback wing having the same wing stiffness as the upswept wing; foe a swept forward wing, divergence will occur before flutter.
Open Access
An investigation into the effect of the application of sub sonic vibrations during the period of solidification of castings with particular reference to a material for gas turbine blades - 'H.R. Crown Max'
(College of Aeronautics, Cranfield., 1955-04) Hinchcliff, S.
The report considers the theoretical relations between microstructures of castings and their mechanical properties and the effects and advantages of vibration during solidification; the design of a melting furnace and a mechanical vibrator to be used together, and the use of sillimanite bonded with Ethyl Silicate as a material for moulds to withstand vibration.
Open Access
An investigation into the machining of titanium 150A in the Forged State
(College of Aeronautics, Cranfield., 1954-08) Holt, J. T. D.
Tests to investigate various aspects of machining have been carried out at The College of Aeronautics on a billet of Titanium - 150A,hot forged by High Duty Alloys. Measurements were made of the drilling forces, and an investigation of the cutting force when turning the blank was made. Tests were also made to determine the effect on tool life of tool shape, cutting speeds, of various rates of feed and of cutting lubricants and coolants.
Open Access
An investigation of the noise field from a small jet and methods for its reduction.
(College of Aeronautics, Cranfield., 1952-01) Westley, R.; Lilley, G. M.
Sound measurements have been made on the noise from the jet of a one inch diameter convergent nozzle at atmospheric temperature and at speeds above and below choking. The noise level and spectrum have been investigated in both the near and distant fields. The results agree in some measure with the predictions of the Lighthill theory, that the elementary sound radiator is an acoustic quadrupole. The agreement is more marked if attention is confined to the higher frequencies. Simple empirical formulae are derived giving the overall sound intensity and frequency spectrum in terms of the position relative to the jet, the stagnation pressure excess over the atmospheric pressure, and the frequency. The results of tests on various noise reduction devices are discussed. These tests indicate promising lines of investigation. The maximum reduction in total noise level was about 10 db.
Open Access
Measurement of the derivative zw for oscillating sweptback wings
(College of Aeronautics, Cranfield., 1955-07) Whitmarsh, G. E.
Measurements have been made of the derivative Zw for rigid sweptback wings mounted at zero incidence and oscillated with simple harmonic motion. The Reynolds number was in the range 1.2 x 105 to 4.1 x 105.
Open Access
Measurement of the derivative zw for oscillating wings in cascade
(College of Aeronautics, Cranfield., 1955-07) Milne, R. D.
Experimental results are reported of the damping derivative zw. for rigid rectangular wings of various aspect ratios in cascades having gap-chord ratios of 2, 1,1/2, 1/3, 1/4. The results show fair agreement with two-dimensional theory. The ranges of Reynolds numbers and frequency parameters were 0.8 to 2.5 x 10) and 0.1 to 0.45 respectively. The results show a strong dependence on Reynolds number which increases with decrease in gap-chord ratio. This effect was eliminated by transition fixation by wires placed at suitable positions down-stream of the wing leading edge.
Open Access
Measurements of the pressure distribution on swept back wings with trailing edge split flaps: summary of wind tunnel work at the College of Aeronautics 1948 - 50
(College of Aeronautics, Cranfield., 1951-03) Babister, A. W.
This is an interim report giving measurement s that have so far been made of pressure distribution on two untapered wings swept back 45◦, aspect ratio 2 and 4, fitted with both full span and part span trailing edge split flaps. The Reynolds number in these tests was about 0.5x106.
Open Access
Nose controls on delta wings at supersonic speeds
(College of Aeronautics, Cranfield., 1950-05) Bolton Shaw, B. W.
Expressions are derived for lξ and a2 of nose ailerons and nose elevators on a delta wing, as depicted in Fig. 1, in supersonic flight. Nose and trailing edge controls on delta wings in supersonic flight are compared.
Open Access
Note on the efficiency of adiabatic shock
(College of Aeronautics, Cranfield., 1948-09) Morley, A. W.
This note records some calculations of the efficiency of adiabatic shock in air (x=1.4); where the efficiency is defined as the ratio of the work used to compress the air in the shock wave, from the inlet static to the outlet total head pressure, to the work required if the compression were isentropic.
Open Access
Note on the results of some profile drag calculations for a particular body of revolution at supersonic speeds
(College of Aeronautics, Cranfield., 1954-07) Wedderspoon, J. R.
Details additional to those discussed in College of Aeronautics Report No. 73 are given of the method developed for the calculation of the profile drag of bodies of revolution at supersonic speeds and zero incidence. The method has been applied to a particular body of fineness ratio 7.5 (see Fig. 1) for Mach numbers ranging from 1.5 to 5.0, Reynolds numbers ranging from 106 to 108 and transition positions ranging from the nose to the tail end of the body. The calculations assume zero heat transfer. The results indicate that the overall difference in profile drag between fully laminar and fully turbulent flow decreases rapidly with mainstream Mach number and rather more rapidly than does the corresponding difference for a flat plate, and at Mach numbers greater than about 2 the profile drag of the body with fully turbulent flow is less than that of a flat plate (Fig. 14).