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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.
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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.
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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.
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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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52 J. FIELDING
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.
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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.
53
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
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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,
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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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56 J. FIELDING
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.