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DESIGN An Article Based on a Paper Read by the Author at the 0 1 0 0 VIIth International Congress of Applied Mechanics 2·096 1·62 1·37 London, Sept. 5—11, 1948 1·97 5·07 0 0 By J. Lockwood Taylor, D.Sc. VI. Honeycomb Cores There appears to be some difference of opinion as to whether isotropic-core theory, as here con I . Wide Compression Panels, Isotropic Core Twisting moment: sidered, can be applied to pa ls having cores of HIS is the type of panel which has received built-up honeycomb construction. There is no most attention from investigators and the question that the rigidity modulus G is simply critical load for local wrinkling of a single that for the core material, reduced in the ratio also skin of thickness t is most simply expressed as of the wall cross-sectional area to the total panel area, and further modified by a geometrical P=G2/3(E")1/3t factor which is easily derived and is equal to per unit width of panel, in terms of G the shear ¾ for regular hexagonal construction. This factor modulus of the core material, E" the plate where D is the bending rigidity and q the applied allows for the fact that part of the core is inclined modulus of the skin (=E'÷(1— σ'2)) and t the normal load, if any. skin thickness. Wrinkling occurs with a wave to the direction of loading and is therefore not Introducing shear deflexion v by the relations length 2π/k, where fully effective in shear. The main difference of opinion is over the kt=2(G/E")1/ 3 longitudinal E of the core; for the honeycomb The above solution is subject to the core thick alone this is very small, due to bending of the ness exceeding a minimum value 2d=4/k, say. where C is the shear rigidity, having the value walls, and if this were the true effective modulus This will always be the case for panels of optimum 2Gd for a sandwich, 2Ghx4/5 approximately of the core in situ the buckling load of the panel design, i.e. minimum weight for given strength, for a plate. would be appreciably reduced. However, the for which the local wrinkling load is equal to The total deflexion w=(u+v) satisfies effect of the skin in preventing transverse spread the load for Euler failure over the whole panel ing and hence wall bending of the core must be span. It is implied that a core of minimum allowed for and when this is done the longitudinal density is used having a shear modulus only just E is found to be comparable with the full E of sufficient to stabilize the skin, viz.: I n the special case of buckling under end load P, the material, again subject to a geometrical factor. This appears to be borne out by test results. where e, the skin strain, is defined as (P/E"t). VII. Balsa Core III . Compression Panel of Finite Width The complete expression for wrinkling load Balsa, besides being non-isotropic, differs from covering all core thicknesses and wavelengths Using the equations just developed, a rectangu the cores previously considered in that it is is rather cumbersome but can be simplified to lar plate buckling in the form w=sin mx sin ny more rigid than is necessary to stabilize the skin give the Euler load (modified for core shear has a critical load against local wrinkling. This applies to any skin, deflexion) applicable to all normal panel lengths P=(m2+n2)2÷{m2/D+m2(m2+n2)/C} even steel, and disposes of the argument in favour 2P=( 1 +t/d)÷(½Gd+½E'td2k2) The assumed buckling form shows that this of transverse grain. The Euler load, including where 2P now refers of course to both skins of applies to a supported edge panel of dimensions shear effect, is the same for either direction of the panel. (π/m) X (Π/n). As a special case a square panel grain (apart from the contribution of the longi The limiting value P=Gd(1+t/d) for short buckling in one half-wave, m=n, gives tudinal grain to the bending rigidity), but as this wavelength (k large) can also be shown to be a P=4m 2 ÷(1/D+2m 2 /C ) critical load is shared by skin and core in one safe value for the local wrinkling load, when the A long panel of width (π/n) has a minimum case and is practically all taken by the skin in core thickness is less than the minimum given critical load the other, the skin stress is obviously less in the above. For a skin material such as Al alloy the P=4n 2 /D÷(1/D+n 2 /C) 2 former case, at a given load; which definitely optimum panel will be found to have a working occurring when favours the longitudinal grain, as in British compressive stress well below the proof, due to m 2 =(n 2 /D+n 4 /C)÷(1/D-n 2 /C ) practice. falling-off in (tangent) modulus E'. This may give a lower buckling load than for m=n. I L Modified Plate Equations Local wrinkling as considered in Section I has still to be checked as a separate case. Fo r problems involving bending of a panel AIR REGISTRATION BOARD NOTICES in two directions at right angles as, for instance, British Civil Airworthiness Requirements IV. Shear Panel compression of panels of finite width, shear Sub-section R.2, Issue 4, of British Civil Airworthi stability and transverse loading, it is convenient Buckling of a long panel in shear can be ness Requirements, was published on April 1. to have available modified flat-plate equations examined using the energy method in conjunc This Sub-section replaces the existing R.2, Issue 3, including shear deflexion. Love's classical treatise tion with the relation between bend and shear together with its Appendix, Issue 1, both of which treats a kindred problem, that of a moderately deflexions as given under II. Results can be should be removed from the folder and filed. thick homogeneous plate, but his solution is un summarized as follows: S=critical shear, l= Notices to Licensed Aircraft Engineers and to Owners necessarily complicated due to his working in panel width of Civil Aircraft terms of total deflexion, instead of splitting the deflexion into bend and shear components. In The Council of the Air Registration Board announces 0 0·2 1 large terms of bend deflexion u the usual equations the issue of the undermentioned: hold : Notice No. 3, Issue 3. April 1, 1949. Duties of 5·65D 3·3D C 1·01D Aircraft Engineers licensed in Categories 'A' and 'C'. Notice No. 8, Issue 2. April 1, 1949. Radio Appara As before, local wrinkling under compression tus installed in Civil Aircraft. S (at 45 deg. t o the panel edges) has to be checked and in doing so the perpendicular tension S can Notice No. 9, Issue 2. April 14, 1949. Aircraft be ignored. Maintenance Schedules and Certificates of Safety Shear components: for Flight. V. Transverse Loading Notice No. 17, Issue 4. April 14, 1949. Certificates Supported-edge panels yield the same solution of Safety. as ordinary plates, apart from additional shear Notice No. 30, Issue 2. January 24, 1949. Blade- deflexion, but clamped edges introduce a correc Retaining Nut and Barrel Failures—De Havilland 4-Bladed Propellers for Hercules Engines. tion, since the clamping only applies to the bend component of the deflexion. The result is a relief Notice No. 31, Issue 1. April 14, 1949. De Havilland of maximum bending moment as indicated below 4-Bladcd Propeller Types CD and PD27/445/1 or Bending moments: for the particular case of a square panel: 2 and CD and PD 54/445/1 or 2. Aircraft Engineering
Aircraft Engineering and Aerospace Technology – Emerald Publishing
Published: Jun 1, 1949
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