MANUFACTURING EXERCISE INVOLVED IN THE REDESIGN

OF THE HAWKER SIDDELEY TRIDENT (TRI-JET) FUSEIAGE

John Fielding

Chief Materials Engineer. Hawker Siddeley Aviation Limited

Woodford, Cheshire, England

Design Exercises

The purpose of certain design studies was to examine the application of titanium construction to replace existing aluminium alloy structure using the same design loadings, applying the same structural philosophies, and accepting the same practical con-straints on geometry. Under these design conditions weight savings result from the relative specific material properties of titanium alloys and aluminium alloy, the reduction in sizes permissible in titanium and, also from the exploitation of the weldability of titanium to produce more efficient configurations. Ti 8Al, lMo, lV was specified (Duplex Annealed). The relatively thin fuselage skin (0.022 in.) was expected to be sufficiently free from stress corrosion hazards under aqueous conditions. Three particular areas were chosen for evaluation, viz., the sheet/stringer/frame struc-ture in the keel area, the upper fuselage, and a window panel area. The usual attention was given to fatigue strength, critical crack length, and residual strength. Fusion welding was used whenever practicable, i.e., for skin to stringer joints and panel butt welds, with a little electrical resistance spot welding for the frame to fuselage skin a'ttachment. The weight savings possible with the titanium design as compared with the aluminium structure were as follows:

Fuselage keel area 26.3%

Upper fuselage area - 17.6%

Window panel area 28.0%

The overall weight saving on the complete fuselage section was 23.6%.

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46 J. FIELDING

Chemical Milling

Chemical milling (also known as chemical machining or contour etching) has been extensively used for shaping various materials for many years. Basically, the process consists of chemically etching away (using a suitable medium) the unwanted material, areas to be preserved being protected by a suitable maskant which is usually elastomeric in nature.

An initial consideration of nine etching solutions including various combinations of hydrofluoric, nitric and chromic acids and ferrous sulphate indicated two interesting possibilities. One being 25% hydrofluoric acid (by volume) and the other 3% hydro-fluoric, 30% nitric acid (by volume). The 3% HF, 30% HN03 mixture was found to be too slow in action and when heated the solution rapidly went out of balance, The 25% HF solution used at ambient temperature was very active and this feature combined with the exothermic nature of the reaction required adequate circulation and water cooling facilities. Satisfactory control was quite practicable and the process is now used for production work.

Panels to be chemically milled are hand degreased, and sprayed with "Coverlac" synthetic rubber maskant 0.008 in. thick. The shape to be etched is marked out, cut, etched in the 25% HF solu-tion, rinsed and dried. The process works well with both 6Al, 4V, and 8Al, lMo, lV titanium alloys with a normal rate of metal re-moval of 0,025 in. to 0.030 in. per; hour. Several hundred H2 analyses gave results between 35 and 90 ppm. A considerable amount of data has accumulated from the chemical machining at Hawker Siddeley Hamble and the conclusions were that the H2 due to chemi-cal milling was concentrated near to the sheet surface. This meant that the sheet thickness after milling had a marked affect on.the average H2 content determined by analysis, the H2 content not being related to surface area. It was considered advisable to restrict chemical milling on t.itanium as follows:

a) To give a mininrum thickness remaining after etching of 0.020 in. irrespective of the original thickness

b) Every effort to be made to use material with an initial hydrogen content of less than 80 ppm when etching to a final thickness below 0.025 in.

c) To restrict etching to one side only when reducing to a thickness below 0.025 in.

REDESIGN OF THE HAWKER SIDDELEY TRIDENT FUSELAGE 47

Through Welding under Tension

Before choosing the welding processes to be used consideration was given to electron beam, tungsten inert gas (T.I.G.) and plasma arc. Electron beam was not chosen for certain specific reasons. The fast welding speeds are not always possible with the relatively thin sheet involved without compromising surface contour and fatigue life. Preliminary tests also indicate that porosity could be a problem. These factors combined with difficulties of extremely ac-curate alignment and "set-up" together with the large chamber re-quired, prompted the use of T.I.G. welding. T.I.G. welding was known to give good weld shapes, excellent weld properties, reason-able speed and was a fully developed process. Plasma arc welding would probably have been rather more suitable than T.I.G. but equipment was not available at the time. Experience with T.I,G. welding was expected to "read across" to plasma arc.

The problem of "through welding" a typical fuselage panel (see Figs. 1 and 2) is not usually machine power but heat b~lance and distortion - the larger the weld pool the greater the distor-tion due to shrinkage. Excessive distortion led to the dev~lopment of a tension draw welding process (see Figs. 3 and 4) Patent Spec. No, 37125/69 UK. 52835 USA, 7026963 France, and P20,37 ,3493 Germany. The component parts were assembled together and a tension load ap-plied, thus holding the parts in line over their entire length and applying the correct stress in the components to overcome the weld shrinkage. The die holds the component parts together locally whilst traversing the length of the component to weld the stringer to the skin. Alternatively the die and welding equipment may be stationary and the tension frame complete with the component parts under tension traversed,

A small machine was constructed to apply 80 tons end load producing average stresses up to the yield strength of the titanium. Both the stringers and the skin were gripped in specially designed clamps using a strip spot welded to the transverse edges of the skin and around the stringer. Means were provided to apply tension to the stringer sections quite independently from the skin so that any small discrepancies in the clamps or length of stringer were compensated while maintaining even tension across the panel. The aim was to load the stringer and skin to the same stress when the hydraulic load was applied and welding conunenced. The twin welding head was water cooled with argon backing, current to the head being from a conunon power source manually controlled.

48 J. FIELDING

72"

I

~ Fig. 1. Typical fuselage panel with window

apertures 8Al-1Mo-1V titanium.

Fig. 2. T.I.G. through welded joint for stringer -skin fuselage panels.

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REDESIGN OF THE HAWKER SIDDELEY TRIDENT FUSELAGE

Fig. 3, Draw welding machine. Top view showing twin welding heads.

Fig. 4. Draw welding machine. Underside of panel showing bottom die assembly.

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50 J. FIELDING

Welding on the Tension Draw Welding Machine

The aim was to produce skin to stringer welds in 8Al, lMo, lV with adequate penetration, no porosity, no undercutting, and with acceptable surface finish, which would not require any further machining process. For the fuselage panels a slight "weld bead" could be tolerated if this did not reduce the fatigue life of the joint.

The skin and stringers were pre-stressed to 80,000 psi, which was believed to be the stress required to eliminate quilting. How-ever owing to the local changes of skin neutral axis at the trans-verse lands, there was a natural tendency for the panel to produce slight humps when the load was applied. This was unacceptable for welding and so the stress had to be reduced to 62,000 psi, at which load the effect was negligible.

Using this preload a number of panels were welded. These panels were very good from the overall flatness viewpoint, but slight quilting was present which would be removed during hot sizing. The surface appearance of the welds was very good with consistent penetration to form a slight fillet against the stringer. The process was developed satisfactorily and the re-quired number of fuselage panels were satisfactorily welded. The surface appearance of the T.I.G. welds was satisfactory giving in the main a slight protuberance on the surface. Although under-cutting of the "weld bead" did occur it was only in isolated local areas where the stringer legs were not exactly normal to and in line with the torch.

The following important points emerge from the welding exercise:

1. All components T.I.G. welded must either be stress relieved or hot sized after welding.

2. Porosity in titanium welds is still a problem. Even with careful preparation, cleaning, and adjustments to welding speed and conditions, occasional porosity was noted.

3. The surfaces to be welded must be pickled or degreased, they should only be handled with clean white gloves. The atmosphere must be reasonably free from dust.

4. To avoid excessive distortion the skin panels should be held under tension during welding.

REDESIGN OF THE HAWKER SIDDELEY TRIDENT FUSELAGE 51

Vacuum Hot Sizing

The stringer-skin panels for the fuselage were welded up as flat panels. These were subsequently to be stress relieved and contoured to the fuselage shape, i.e., 72.75 in. curvature radius. The stress relieving - contouring operations were combined into a single hot sizing operation whereby the panels were clamped to the required shape and heat-treated for 30 min at 1320°F. The first three panels were hot sized successfully in steel male and female dies made from 3/8-in.-thick plates suitably reinforced with an "egg-box" structure. However the fixture showed increasing dis-tortion after each heating and had to be discarded after the third time.

A vacuum clamping system was devised (Fig. 5). The fixture was fitted with a series of hoop frames so that steel bars could be clamped by a double wedge action around the periphery of the panel to seal the skin to the base of the fixture. Accepting the fact that the sealing would not be very efficient a high capacity vacuum pump was used with cooling between the fixture and the pump. The vacuum pressure at temperature was 20 in. Hg and the panels were sized exactly to contour. The process was considered quite suitable for panels of this type and dimensions.

In general the hot sizing operation was a very time consuming operation requiring considerable care in the preparation of the panels. The cleaning and Turco spray, followed by Turco spraying of all metal parts of the fixture which were to contact the tita-nium panel, coupled with the actual lay-up of the wire mesh fixture frame and clamps, required detail attention to achieve the desired result. The achieving of a satisfactory seal sometimes called for re-assembly of sealing bars, clamps, etc. before a 20 in. Hg pres-sure was obtained.

Fig. 5. Vacuum hot sizing fixture.

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The Manufacture of Fuselage Frames

The "Z" section members required 0.043-in.-thick 8Al, lMo, lV material with l~t bend radii, the overall curvature of the section (outer radius) being 72.75 in. The maximum length (measured chord-wise) of the required section was 56 in. Hot forming techniques were necessary in order to achieve the l~t bend radii and to stretch the section to the required curvature. As hot rolling dies were not available the frames were fabricated from a hot bent angle section ~ x ~ in. with a small bend radius, welded to a cold bent channel to form the "Z" section. These straight sections were then cold stretch wrapped in a Hufford machine using 1% pre-strain before wrapping to a strain of 2~%. They were then hot sized for ~hour at 12500F producing accurate contours to fuselage.

This method (i.e., cold stretch, clamp and hot size) whilst enabling the manufacturing exercise to be completed (8 components) proved to be a very slow and time consuming method. Hot stretch forming or hot rolling would be preferred for large scale produc-tion. Scaling of the steel tools created difficulties and care was necessary to prevent contamination of the titanium. Cadmium plated bolts were inadvertantly used in the fixture for one hot forming operation which caused severe cracking of the component on heating.

Value Engineering Study

The exercise described was used as a basis to estimate costs for a representative Trident type fuselage 12 ft diameter x 27 ft long. It assumed that the manufacture of continuous skin-stringer assemblies up to 30 ft long would be possible. Manufacture of the fuselage shell in titanium alloy gives a weight saving of 23.6% over the aluminium alloy. To achieve this there is a cost increase of 1.94 to 1. The estimate assumes the manufacture of 100 fuselages.

From the results of the practical work and the resulting cost estimate, it would seem that the three high cost areas require further investigation. Draw welding needs development to make it less labour consuming and hence a more viable production process. Chemical milling needs careful control, and since it is a high labour cost low material utilization process serious thought should be given to its ut:ilization in the design. If chemcial milling were not used weight saving would obviously not be as high but with the resultant cost saving on optimized cost-weight strength solution may be obtainable.

REDESIGN OF THE HAWKER SIDDELEY TRIDENT FUSELAGE

The Use of Ultra High Strength Titanium Alloys

in a V/STOL·Military Aircraft Application.

The H. s. A. Harrier

This part of the paper was contributed by Mr. H.J. Sayer, Chief Materials Engineer, Hawker Siddeley Aviation Limited, Kingston upon Thames.

Weight saving in V/STOL aircraft structure is of great significance and a prime considerationo Maximum use must be made of lightweight materials that provide satisfactory strength with-out compromising structural integrity or incurring too great an increase in expense. It is literally true that in the Harrier aircraft the pound (weight) ·added to the airframe, engine or equipment is the pound (weight) that can keep the aircraft on the ground during VTO - until the equivalent fuel is burnt off.

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This weight saving applies to the whole aircraft and includes metallic and non-metallic materials, but this paper deals mainly with ultra high strength titanium alloy, although some other titanium alloys are used. ·

Titanium alloys have been used extensively on the engines for the Harrier aircraft accounting for some 20% of the dry engine weight. A range of titanium alloys are used from the well known 6Al 4V alloy to newer stronger alloys. The high strength alloys were developed between the late 1950's and early 1960's in response to demand from both airframe and engine manufacturers. A range of alloys was described by R. M. Duncan and C. Minton (Ref.·2) in a paper which included IMI 318, IMI 550,.Ti 679, and Ti 7Al,4Mo, all four having a UoT.S. of 157,000 psi. A range of higher strength alloys (180,000 psi), were also available, i.e., Ti 6Al, 6V, 2Sn, Ti 680, I.M.I. 551, and Ti 13V, llCr, 3Al.

The importance of weight saving on V/STOL aircraft prompted serious consideration of the use of the high strength titanium alloys to replace 1,240 steel components on the H. s. Pll27, under development in 1964/650 Two of the higher strength alloys were chosen for further evaluation: I.M.I. 680, an alloy originally developed for gas turbines, and Jessop-Saville Hylite 51 (now I.M.I. 551), a complex titanium aluminium alloy of 180,000 psi. U.T.S. Messrs. High Duty Alloys Ltd., and Dowty Rotol Ltd,, collaborated in this programme and as a result of the investiga-tion I.M.I. 551 was chosen as the high strength titanium alloy for the Harrier components,

I.M,I. 551 is an alpha/beta alloy which followed the lower strength I.M.I. 550 (Ref, 2). It is a Ti, 4Al, 4Mo, 4Sn, 0,5 Si

54 J. FIELDING

alloy giving a min. 0,2% proof stress of 155,000 psi and giving a min. U.T.S. of 180,000 psi obtained by solution treatment at 1650°F and ageing at 930°F, both treatments being followed by air cooling. The alloy has a density of 4.62 g/cm3 (0.166 lb/in3) and a modulus of elasticity of 16 .4 x 106 psi. Typical "cut up properties" obtained from forgings are

0.2% proof stress 164,000 psi

Ultimate tensile stress 186,000 psi

% Elongation 4 VSo 15%

The fatigue properties are satisfactory - a fatigue to tensile strength ratio of 0.50 for plain specimens in rotating bending with the reduction due to a notch (Kt 3,2) less than the predicted value.

Stress Corrosion Cracking. There has been very little stress corrosion trouble with titanium forgings and even with the higher strength alloys now in use the problems appear less severe than with the high strength steels. The experience gained so far with Harrier aircraft suggests that I.M.I. 551 is no more susceptible than other titanium forging alloys.

Fracture Toughness. The rather limited range of tests to date indicate a range of Klc from 30 to 40 ksi Vin: The feature of a reduction in toughness with increasing strength is a problem with the very strong titanium alloys, and the approach has been to fix a minimum acceptable critical crack length which will meet these re-quirements.

Details of I.M.I. 551 Components replacing steel components

Leading Edge Ribs No. 18 Flap Lever Undercarriage Brackets Nose Undercarriage Pivot Nose Undercarriage Pivot Bracket Spar Reinforcing Booms Outboard and Inboard Spigots Fin Fittings

See Figs. 6, 7, and 8.

saving 0.90 kb (2 lb) saving 0.90 kg (2 lb) saving 2.16 kg (4.8 lb) saving 0.81 kg (1.8 lb) saving 7.65 kg (17 lb) saving 1.80 kg (4 lb) saving 7.10 kg (15.62 lb) saving 1.80 kg (4.0 lb)

Use of Lower Strength Alloys. There are 16,000 bolts in 6Al-4V titanium, also brackets, mountings, skins and duct clamps, also in 6Al-4V on each aircraft. A conunercially pure titanium welded water tank has also proved very successful,

REDESIGN OF THE HAWKER SIDDELEY TRIDENT FUSELAGE 55

Fig. 6. Nose undercarriage pivot bracket.

Fig. 7. Centre spigot - pylon.

Fig. 8. Engine mounting bracket.

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Weight and Cost. The use of ultra high strength I.M.I. 551 for 35 components originally in steel of 180,000 psi U.T.S., saved 23 kg (50 lb) weight. The 16,000 bolts also gave savings of 23 kg (50 lb) at a cost of fl3 per lb weight saved. The plate, skins, and other components made from bar and forgings bring the total weight saved to 120 kg (280 lb). The total weight of titanium alloy used is 182 kg (400 lb) which is 8% of the structure weight.

With the weight of titanium and its alloys in the engine which is nearly 320 kg (700 lb) the sum total in the Harrier is between 9-10% of the operational weight. This is achieved with an overall cost per pound weight saving of from fl3 for bolts to i45 to the most expensive components. This is an entirely acceptable price for such a significant contribution to the success of the only operational strike V/STOL aircraft in the world.

Acknowledgements

The authors wish to thank the directors of Hawker Siddeley Aviation Limited, for permission to publish the paper, also the Procurement Executive Ministry of Defence who sponsored some of the work. The views expressed are not necessarily the views of the company.

References

1. Duncan, R. M. and Minton, C. D. T., "The Role of Depth Hardenability in the Selection of High Strength Titanium Alloys for Aircraft Applications•" Proceedings of the First International Conference on Titanium, London 1968, Paper VII (b) 5.

2. Duncan, R. M. and Hubbard, R., "The Application of the High Strength Alloy Ti 550 in European Airframe Projects." To be presented at the Second International Conference on Titanium, Cambridge, Massachusetts, May 1972.