Development of F-14A Wing Center Section

2008
Carlos A. Paez
Grumman Aerospace
Corporation
8/1/2008
The Development of the F-14A
Wing Center Section

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The Development of the F-14A Wing Center Section by Carlos A. Paez, August 2008
The Development of the F-14A Wing Center Section
Introduction
The F-14A was designed to perform optimally in every air combat, at all altitudes and speeds by
reconfiguring itself via a variable sweep wing in combination with other maneuvering devices. This
gave the F-14A a marked maneuvering advantage over contemporary fighters of its era. The dogfight
mission is the most demanding airframe and system design consideration. The variable sweep wing
pays off permitting automatic aircraft aerodynamic configuration adjustments during a dogfight, STOL
like take offs and landings under 1500 feet, and approach speeds of less than 120 knots. The variable
geometry concept requires that a wing center structure provide support for the movable wing outer
panel. The two wing outer panels generate large bending and torsional loads, which the center wing
structure must react as the wings are moved back and forth throughout the flight regime. While
envisioning the aerodynamic benefits was enlightening, designing and building this aerodynamicist
dream was another matter; and is the subject of this anecdotal write up.
Early development Work
For the VFX proposal Grumman had selected a variable sweep aircraft concept based largely on the F-
111A straight pivot concept. The next big decision was what material to make it out of. The F-111A
had been made of D6AC Steel, while all of Grumman prior airplanes had aluminum wings. Titanium

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emerged as the best material due to its strength to weight ratio and its immunes to corrosion. Also in
comparable fatigue test coupons, titanium also showed superior performance in terms of fatigue life,
thus reduced weight. In 1968 as part of Grumman’s VFX proposal effort, Grumman designed,
fabricated and tested a thick section electron beam welded 6 Al-4 v titanium box beam. Al Hallock, Ed
Mulcahy and Dietrich Helms were the main leaders assigned to develop preliminary methods to
design, analyze and fabricate the wing carry through box using 6 Al-4V titanium alloys. Helms for his
part developed weld joining techniques and made various types of Electron Beam (EB) welded
assemblies, while Al Hallock developed sketches and drawings for wing box concepts and Ed Mulcahy
did the stress analysis.
This resulted in the fabrication of two simple rectangular boxes EB welded to form what was meant to
be a set of simplified left and right side assemblies of the wing box. These two rectangular assemblies
were subsequently EB welded to each other on four sides. The ends of each assembly terminated in
thick plates to which the test structure was attached to simulate the wing outer panel and impart
wing loads to the carry through box. The test specimen was statically tested and provided some
further encouragement that the approach would be successful. This was the first time that a major
structure was using welds in tension as part of its basic design. This structure was later fatigue tested
using what was thought to be a severe spectrum. After surviving its intended life, the loads were
increased 12% to cut down the test time. Eventually the structure failed at a tool mark in the lower
skin. However, there were no failures in any of the welds. In a way this was good, because to resolve
fatigue issues would have been a lengthy process that no one was prepared to undertake at that time
and could have squelched this initiative.
The VFX Proposal Offering
The Grumman VFX proposal showed a wing carry through box somewhat similar to the F-111A
configuration at the pivot, except that the main structure was much more geometrically complex. The
attachments to the center fuselage were not shown and the nacelles were only shown attached rigidly
to the front and rear beam with bolts. Also, how the pylon or main landing gear strut would be
designed, or supported were not shown. This was typical of the level of detail portrayed in technical
proposals, but it obscured some critical details which had to be sorted out later on.
Traditionally, welds are never used for primary structure and they certainly were never used in
tension applications. The Navy customer, while enthused about the potential weight savings of this
approach needed to see much design, analysis, testing and quality controls before they would accept
such an innovation into one of their naval aircraft.
In October 1968 Vinnie Padden and Carlos Paez (structural designers from the Lunar Module Ascent
Stage) were assigned by Jim Brennan to the VFX proposal, reporting to Al Hallock, who was the
Structural Design Group Leader. Ed Mulcahy was the Structural Analysis Project Leader, while Bill
Stewart was the Structural Design Project Leader. Bill was known for his short temper, but also for
having previously worked very well for Lockheed. Hallock and Mulcahy went a long way back and
were very good friends, having worked together before on other Grumman projects. Sid Johnson was

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the senior Project leader for stress, Dick Cyphers was Project Engineer and Nat Kotlarchyk was his
assistant for Analysis, together with Ron Heitzmann who was the Materials and Process Project Lead.
These three really run the vehicle engineering airframe part of the program.
The Go-Ahead
The go ahead to Grumman for the F-14A came in mid January, 1969. Even though this was a big
accomplishment, most people realized that while this was a great win, it was also the start of a long,
intense, difficult and challenging development effort.
Larry Mead (an ex- structural analyst going back to the F-4 Wild Cat and many years later to become
one of Grumman’s VP of Engineering) was asked to oversee the overall manufacturing development
of the wing carry through box. His principal aim was to make sure the structure could be produced.
I.G. (Grant) Hedrick, VP of Engineering looked after the structural integrity of the design. Larry Mead
started his reviews with weekly meetings of all concerned. Al Hallock asked Carlos Paez to represent
the design effort and to prepare simple sketches and or drawings to illustrate what was being address
each week by the engineering team. Doug Hutching was appointed Structural Analysis group leader
and was new to the F-14A at that time. He came from the E-2A project and had worked for Nat
Kotlarchyk before. The other team members representing various disciplines were asked to provide
their inputs as well. Ed Rolko and Tom Thielemann were the machining experts brought in by the
Producibility leader Tony Ioppolo and provided guidance on machining criteria such as cutter speeds,
diameter and corner radius for a given depth of cut. Tom Tatarian was the welding tooling manager
assisted by Bob Dooley who really did most of the tooling design work. Dietrich Helms was the
welding leader together with Alan Loftgren and Frank X. Drumm. Materials and Process was
represented by Tom Main and Ron Heitzmann, Steve De May and Steve Banks. Weights
control/optimization was under Ed Tobey, John Raha and Ray Labell. Lou Veprek represented the QC
laboratory and was supported by Skip Chance and John Munyak Jr.
The Start of F-14A WCS Development
While 6 Al-4V titanium was a well known material in Grumman, it had never been used before to
fabricate such a large and complex structure. Grumman realized after winning the F-14A competition
that additional Electron Beam welding tunnels would be required to meet the fabrication rate
expected. Boeing’s Super Sonic Transport, or SST was cancelled that year and Grumman quickly
purchased the welding tunnels from Sciaky that were initially destined for Boeing in Seattle.
In 1968 and early 1969, Grumman’s EB welding area resided in a small space in the middle of Plant 3,
where the Lunar Module Ascent stage cabin was welded out of 2219 Aluminum. Dietrich Helms and a
couple of weld technicians did all of the titanium experimental work in Plant 3. After the F-14A
contract win, in the south east corner of Plant 2 an EB welding center was established and became the
focus of all wing box welding activities, see Figure 1.

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Figure 1 - The New EB Welding Facilities in Plant 2, Circa 1970
Carlos Paez became a good friend of Dietrich Helms and tried to understand what made Electron
Beam welding work. There were issues such as gun to work piece distance, part maximum gaps, gun
size, electron beam distortion, stop-start holes, welding tabs and many more detail issues to consider
in designing an EB welded joint. Carlos Paez also befriended Bob Dooley who was designing the
tooling approach for the production of F-14A WCS. Prior to that, Dietrich Helms just built prototype or
sample EB welded joints. Now a tooling approach for production was needed. Unlike Tungsten Inert
Gas welding or TIG, EB welding required only one pass to consolidate the joint, whereas TIG required
many passes and the number grew as the thickness of the joint grew, see Figure 2.
Figure 2 - Comparison of TIG and EB Welding on a thick Titanium Section

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In addition, inspection criteria were to be incorporated into the weld joints, as well as surface
machining to remove potential fatigue site initiation points. As more production like welds were
made on representative samples, a host of new welding irregularities were identified. They included
excessive under fill, lack of penetration, root side suck back and partially missed seams (more about
that in the Test 9 specimen). Therefore, an inspection technique dubbed the “witness lines” was
developed and used on all welds, see Figure 3.
Two important EB weld requirements at that time were to use constant thickness weld and to have
both parts to be welded in intimate contact with a maximum gap of no more than 0.015 inches. This
was required to provide adequate electron beam power to melt the adjoining pieces without causing
undercuts, or lack of weld penetration. The design of the upper covers, which are primarily in
compression, drove the maximum EB weld thickness to approximately 2.25 inches. The front and rear
beams were at a nominal 1.25 inch thick, including most rib intersections. During the early
development period, several undesirable weld defects emerged, thus ways to avoid them had to be
developed, see Figure 3.
Figure 3- Details of EB Welding and defects Screened by Visual Inspection
In addition, whenever an EB weld is started, the beam in essence drills a hole into the two adjacent
titanium parts, then as the beam moves across the seam, it melts both parts and solidifies the joint. A
similar thing happens when the weld is stopped. These stop and start areas form an undesirable
defect which must be removed from the structure. Thus this required that the design account for
these issues and methods to remove the defects as part of the manufacturing process needed to be
developed, see Figure 4.

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Figure 4 – The EB Welding Manufacturing Methodology Developed for the F-14A
John Grandy, an English engineer who had worked on the Lunar Module with Carlos Paez, joined the
team as a producibility engineer. He worked for George Petronio who headed up the producibility
group. Grandy was new to the EB welding technology, but he had many years of experience in aircraft
design and fabrication having worked in the aerospace industry at Vickers, back in England.
The wing box structure, which was initially called the “wing carry through box” had a name change in
mid 1969. It became known as the Wing Center Section, or WCS. This was a public relations move by
Grumman to distance itself from the F-111A wing carry through box, which at the time was in the
news receiving bad press for having failed its structural tests related to TIG (Tungsten Inert Gas)
welding of D6AC steel. In fact, Nat Kotlarchyk dropped a six inch thick Air Force report on Carlos
Paez’s desk and said: “read all of this, but don’t do any of it on the F-14A”. A comparison between the
two airplane’s structures is shown in Figure 5.

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Figure 5 – Comparison of F-14A and F-111A Wing Support Structures
The F-111A structure was welded in what might be thought of as a shoe box, with the top skin cover
bolted on. Interestingly enough, the producibility leader on the WCS, Tony Ioppolo, wanted the F-14A
to be done the same way. Carlos Paez was asked many times to draw up such a configuration, but he ,
having being advised not to do that by the Analysis Project Engineer Nat Kotlarcyk, refused to
accommodate the request for a long time making excuses related to the compressed schedule, more
about that later.
Configuration Development
There were some serious issues to be resolved in the configuration development and integration of
the WCS with the rest of the F-14A structure. They included the attachment to the mid fuselage, the
nacelle attachment, the main landing gear attachment, the pivot joint, the over wing fairing and the
pylon attachment. Al Hallock made a management decision to assign Phil Bourque and Vinnie Padden
to concentrate in the design of the pivot joint; and Carlos Paez to look after the configuration
development of the rest of the WCS.
The Grumman VFX proposal did not have a well developed approach for the WCS. In the proposal
plan-view drawing the rear beam was crooked (there were three kinks which made things very
difficult to analyze, integrate and build), adding great structural inefficiency and fabrication
complexity. The nacelles were attached directly with bolts to the front and rear beams of the WCS.
However, they really needed to be attached so the two structures could operate independently and
minimize induced loads to each other. The inlets were designed by hammer shock loads caused by

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potential engine stalls, while the WCS bottom skin cover adjacent to the nacelles was stretching due
to wing bending stress. These two structures wanted to somehow be independent of each other.
Also the center fuselage attachments were not even properly shown in the proposal. They needed to
be attached so that the fuselage mid section did not pick up WCS induced loads (stretching) caused by
wing bending. Carlos Paez introduced the idea to attach the fuselage close to the no-stretch axis of
the WCS. Doug Hutchings and John Sposito (group leader for mid fuselage stress) participated in its
review and opted to endorse the approach after performing some rudimentary stress analysis checks.
Other big inputs were the way he detailed configured the other attachments to the fuselage, the
nacelles and the wing actuator supports. The fuselage and nacelle supports are shown in Figure 6.
Figure 6 - WCS Attachments to Mid Fuselage and Nacelles
This includes a clever way to isolate the WCS stretch from the fuselage, as well as the interaction of
the nacelles with the WCS by the use of links reaching down to the nacelle rectangular framework.
Carlos Paez managed to straighten the rear beam by negotiating compromises with the mid fuselage
group. This resulted in a larger wing chord box and the ability to carry some additional fuel, all good
things to have. However, the weld assembly still posed some serious difficulties. Consulting with the
machine parts specialists and Bob Dooley the weld tool engineer leader, Carlos Paez developed the
welding sequence for the entire box, as well as a way to weld the upper wing skin cover utilizing a
unique weld joint concept. This is shown on Figure 7 a, b and c.

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Figure 7 a – Welding Details of Various Joints

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Figure 7 b – Weld Details of the Upper Covers
Figure 7 c - WCS Centerline Weld Sequence Concept
Producibility wanted the upper wing skin cover to be bolted. Carlos Paez did not want to spend much
time laying out a bolted approach because it would add weight and fuel tank sealing complexities, so
he took a reduced scale drawing and added an “x” for each bolt, which made the sketch look very bad

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in terms of complexity, cost and weight and presented it to the team at a regularly scheduled
meeting. The sketch was distributed to the machining, assembly drilling and weights team
representatives for their input. Larry Mead, as the then Production Manager, quickly assessed the cost
and weight benefits on an EB welded approach and the EB welding concept was selected over the
bolted one over the Producibility leader’s objections.
WCS Mock-Up
In order to make sure that the WCS could be fabricated as drawn, the project leaders decided to have
design and fabricate a wooden mock-up containing all of the structural and welding details. Shown in
Figure 8 is Carlos Paez (Design) and John Grandy (Producibility) around the wooden mock-up that was
built to prove out how the EB welding and the systems assembly and installation would be performed.
Figure 8 - Mock-Up of WCS
This was necessary since a welded box could only be reached inside to install systems via access holes.
The team wanted to make sure that all internal components (structural members bolted-in and
system components mechanically and electrically attached ones, such as fuel lines, pumps, fuel
quantity indicators, etc. could be properly installed and inspected externally. A more mature and
updated mock-up is shown in Figure 9, including all the fuel systems provisions. Note that the access
holes do not have a series of small holes drilled into the beam web to attach the fuel cover. This was
intentional, since the F-111A wing carry through box had cracks at these small holes and resorted to
interference tapered bolts to fix the problem at great cost and complexity. Grumman’s approach was
to use clamp-on covers with a built-in groove “Parker” seal to provide fuel tank integrity. This is
similar to the approach used in commercial airliners. At first glance this approach appears heavier
than the bolted access cover design, but detail analyses proved otherwise.

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Figure 9 - WCS Mock-Up with Fuel Systems Provisions
Helpful Inputs
During the development of the WCS structural arrangement, Carlos Paez got a visit one Saturday
morning by SVP Grant Hedrick inquiring about the status of development concepts for the BL 65
fittings used to attach the nacelles. Hedrick, still in his tennis uniform sat together for about a half
hour reviewing all of Carlos Paez’s seven different layout designs. Hedrick finally lightly sketched how
he thought the fitting should look using what he thought was one of Carlos’ best concepts. He was
careful not to show precisely how the design should look, rather he conveyed the principles to be
used, and the rest was up to the WCS team. He stated that they should keep the bolts above the
lower cover as much as possible and consider the effects of shear lag on the bolt pattern. The task of
detail designing of the attachment then went to Harry Kreckman, he was guided by Hank Aurwater, a
stress analyst well versed and regarded by the stress department. The result was the fitting
design that flew on the airplane and was commonly referred as the “Y” fittings, see Figure 10 a & 10 b.
Figure 10 a - Forward BL 65.00 “Y” Fitting Attachment

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Figure 10 b - Aft BL 65.00 “Y” Fitting Attachment
Pivot Design Concept
The design of the pivot came to Grumman from Vought via a courtesy of the US Navy. The Navy liked
the Vought concept proposed under their VFX proposal, but could not directly offer it to Grumman,
nor order Grumman to use it. Instead the Navy arranged for a meeting between Grumman technical
experts and the Vought team. After a couple of days of reviews by several Grumman experts, the F-
14A technical and program management decided it was a good way to go. Grumman offered Vought
its structural analysis methods as an exchange for the pivot design and Vought accepted the offer, see
Figure 11 a, as well as 11 b.
However, as Grumman began to analyze the pivot joint, it found that they had to change it, to
prevent the spherical balls from popping out (as had being the case with Vought initial test of the
pivot). Thus, Grumman added three posts to maintain the pivot geometry and a diagonal strut to
provide a truss support to the joint. It was SVP Grant Hedrick who suggested the change. Carlos Paez
did the diagonal strut detail design to accommodate the changes, this is shown in Figure 10. Vinny
Padden and Phil Bourque finalized the bearing and platter attachment details with much help from
Doug Hutchings and Ernie Ranalli, including finite element analysis by Larry Brown, the results are
shown in Figure 12.

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Figure 11 a - Focused Pivot Concept by Vought
Figure 11 b – Basic Focus Pivot Concept, Without Grumman Additions

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Figure 12 – Grumman Structural Additions to Stabilize the Pivot Joint
Figure 13 - F-14A Pivot Structural Details

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Structural Analysis
Late in 1969 the structural arrangement of the WCS began to mature rapidly, however much was
required in order to release drawings. The F-14 had (for its day) a large finite element model. It took
many trials before the analysis group was satisfied that it accurately represented the evolving
structure. This was a Herculean job. It was FEM model E5 that was finally run for a period of a few
days and established the first set of good internal loads for detail design; this is shown in Figure 14.
Doug Hutchings developed his own FEM for the WCS with the help of Larry Brown to a much higher
fidelity than the rest of the airplane to be able to more accurately predict the structural behavior of
the WCS; and also not be held captive to fuselage design changes not affecting the WCS. This FEM is
shown in Figure 15. Doug was a most valuable leader and member of the team. He was cool under the
most difficult times. He had learned to simplify the evolving problems such that a close form solution
could be used to approximate the state of stress in an effort to aid the design maturation process. He
had a small group of about 10 analysts that supported him. He spent a great deal of time with them as
well as with the designers trying to guide the layouts. Hutchings was well liked and a great asset to
Grumman.
Figure 14 - F-14A Fuselage and WCS FEM Idealization

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Figure 15 - F-14A WCS FEM Idealization
From these analyses typical pivot bolt load distributions and magnitude were established and used for
design purposes. These bolt load values are shown in Figure 16 below. Also a typical WCS cross
sectional stress distribution in the inboard BL 65.00, at ultimate load are shown in Figure 17.
Figure 16 - WCS Pivot Joint Bolt Load Distribution
Figure 17 – WCS Cross Sectional Stress Distribution at Ultimate Load
The biggest interface for the F-14A WCS was the wing outer panel. The structural analyst leader for
that was Ernie Ranalli, a well respected stress analyst at Grumman. He took over the pivot joint
overall analysis and established some criteria for the stiffness of the main members. There was a big
difference between Doug Hutchings and Ernie Ranalli. Ranalli could drive the team and direct the
design by his virtual presence and come back the next day (after re-thinking his position) and change
it all over. Doug Hutchings on the other hand was much slower, but sure footed. He was also very
approachable and would explain to less knowledgeable designers or analysts, how the structure really
worked and what needed to be done to fix a poor concept, or when a good concept required a bit of

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improvement. Phil Bourque and Vinny Padden supported by various materials and process experts
matured the pivot joint. Somehow all of these different personnel chemistries worked well enough
that an excellent design emerged. There was a sense of responsibility and a great technical challenge
to all involved. Everyone wanted the project to succeed in the same vein as the Lunar Module for the
Apollo program.
Pivot Detail Design and Analysis
A focus pivot configuration was selected, including a double shear connections at the upper and lower
wing covers. The upper and lower lugs sets were not parallel to each other. They converged to
intersect at an average wing center of pressure or CP somewhere outside on the wing. This is shown in
Figure 9a.
The F-14A wing pivot was designed to insure its structural integrity under either seized (i.e., locked by
friction) or normal aligned pivot bearing conditions. In order to be sure satisfactory structural and
bearing liner behavior under the most adverse bearing friction conditions, extensive analysis and
friction tests were performed. The structure and bearing system successful passed all planned tests.
In any pivot design larger stiffness discontinuities will exist in the pivot region. In the older F-111A
type of single shear pivot joint, the wing panel has continuous shear structure completely supporting
the pivot lugs. This shear type structure stiffens the outer panel skin and yields composite section
bending moments of inertias in the order of 1500 in 4.
The WCS however, has only unsupported lugs
without shear structure resulting in stiffness in the order of 5 in4
. Hence all secondary stresses
resulting from plane sections not remaining plane after bending and are felt by the WCS lugs. This
situation would occur when the bearings seize, (that is the failure to misalign due to excessive
friction). The structure then behaves as a continuous elastic structure with a small radius of
curvature, or high local bending stresses occurring in areas of low stiffness.
Note that as a point of comparison, the F-111A bearing cannot misalign (refer to Figure 5). In addition,
shear loads may be taken by the cover skins in bending due to local stiffness considerations of the
cover skins relative to the shear transfer structure. The F-14A pivot employs a different pivot concept
which inherently minimizes these effects. The joint is a double shear focused joint which carries most
of the applied wing loads as a component of a truss loaded axially (refer to Figures 9a and 9b). This
concept also allows the deletion of a shear web in the WCS pivot region and brings a better balance
across the joint. The F-14A outer panel has a bending stiffness of approximately 5 in4
,xE of 16M for
Titanium while the WCS combined clevis platters are in the order of 60 in4
xE of 16M for Titanium.
A local analysis of the pivot joint using beam type elements was done in the early design stages. These
results showed that locally a high curvature would occur in the wing lugs resulting in undesirable
peaking of bearing stresses on the lower and large weight increment required to obtain a satisfactory
fatigue life. The solution to this problem was to purposely build an eccentricity in the outer panel to
absorb the curvature and relieve the lug curvature. The outer panel shear web structure immediately
outboard of the closure rib is a flexible beaded web to permit a locally high radius of curvature to take
place.

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To insure that the pivot would not have frozen bearings; fatigue testing below ambient temperatures
were performed. The Teflon impregnated fiberglass used on the bearing, had only 3 % friction
coefficient at room temperature, but 12% at minus 40 degrees F. The concept of high wing curvature
outboard and built in eccentricities helping maintain flat lugs and good liner wear were verified on
Pivot Test 9A.
The F-14A does not have a redundant rotational path in case the spherical surfaces freeze. To
overcome this, the lugs were designed to insure that the local out of plane bending stresses due to
non-alignment would not be excessive. A full 5 % positive margin on the maximum design bending
moment was applied. The most critical point occurs at a wing sweep of 68 degrees at limit load and
then un-swept to the 25 degree position and loaded to its maximum limit load at that angle. The final
wing pivot joint details are shown in Figure 18.
Figure 18 - Wing Pivot Bearing Details
Fracture Control Plan
During this time frame, a new structural analysis requirement began to evolve, it was fracture
mechanics. Most Structural analysis group leaders in fact were not familiar with this new
methodology, thus an expert from the Analysis Home section was brought in to help. He was Al
Wolfman. Al was also learning how to use, or apply this new technology. He worked in Plant 35 in a
quieter surroundings not affected by F-14A schedule pressures. Wolfman came over frequently to see
how the design was evolving and suggested structural test coupons to validate his analytical
equations and predictions. Not only was this new, but the effect of EB welding on primary structure
added yet another bit of unknown risk to the design approach.
A premium grade of annealed Ti-6Al-4V material was used for the lower covers and all the pivot
platters. This titanium was purchased to a specification that controlled the raw material input to the

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ingot as well as melting and forging parameters which also included a guaranteed minimum fracture
toughness (KIC) of 60 ksi Early experience in producing thick welds (over 2 inches thick) indicated
that hydrogen in the titanium out-gassed during welding (melting) process and created gas bubbles or
porosity. Therefore, all detail parts underwent vacuum degassing to reduce the hydrogen level from
130 ppm to 75 ppm.
In addition, to insure that production procedures were equivalent to those used during the
development stages of the WCS, Grumman initiated a self imposed fracture control program, one of
the first in the aerospace industry. For the first production run, fracture toughness specimen were cut
from parent material of the most critical machined parts upon receipt of raw material. One set of
specimen was tested in the as received condition and one set accompanied the parent part through all
of its processing, including EB welding; and was then tested for comparison purposes. The results of
these tests provided verification of design data and proved that production processing did not have
any adverse effects on the design properties, see Figure 19.
Figure 19 – Wing Center Section Fracture Toughness Comparison
Grumman recognized that a strong fracture control program was needed to support the design. The
manufacturing processes were controlled to minimize initial flaws, but non destructive testing NDT
procedures to identify them were needed, as well as analytical models to intelligently help make
decisions, or to make dispositions on flaws, or defect size during production fabrication.
Grumman funded a program at Battelle Laboratories to determine the growth characteristics of flaws
in EB welded Ti-6Al-4v sheet and plate. The variables selected covered the expected range of stresses,
stress amplitude spectra, flaw configuration and environments related to service life. The welds
studied included both transverse and transverse with intersecting longitudinal welds. Flaws studied
included those in the weld and in the weld heat affected zone. Excellent results were obtained; all

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tests lives fell within a limited scatter band. The test data is shown in Figures 20 and Figure 21, and
provide some insight into the results of this effort.
Figure 20 – Weld’s Life as a Function of Stress and Initial Flaw Size

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Figure 21 – Environmental Flaw Growth Rate for Ti 6AL-4V
In static and fatigue tests EB welded samples which had been stressed relieved and machined
exhibited nearly 100 % base metal properties, see Figure 22.
Figure 22 – Fatigue Properties of EB Welds

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Figure 23- Fracture Toughness of EB Welded Joints
However, the fracture toughness in the vicinity of the EB weld was found to be reduced to 56 % of the
parent or base metal allowables using symmetrical welds lands. A sample of these test results is
shown in Figure 23.
As a further safety consideration that all processes, analyses and tests were valid, the first 36 EB
welded WCS boxes were statically proof tested to 115 percent of limit load in both bending and
torsion. No anomalies were ever detected and the proof testing was deemed successful and
terminated at the 36th
production unit of the WCS.
Stress Relieve & De-scaling
The Materials and Process engineers had a great input into the design as well. It was established by a
series of test samples that the EB welding, if left in the as-welded state would have very high residual
stresses. Therefore a means of stress relieving the welds needed to be developed. Steve Banks of
M&P did a literature search and consulted with various other industry sources and came up with a
production approach. Test coupons were made to verify the assertions and eventually it was agreed
that the WCS had to be stressed relieved at 1200 degrees F for four (4) hours to eliminate any
detrimental residual stresses.
The stress-relieving process fixed one problem and created another. When titanium is exposed to
1200 degrees F in air, it forms a layer of alpha case, or titanium oxide. This layer is only a few
thousands of an inch thick, but it has very fine surface cracks, that if left intact would become fatigue

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cracks initiation sites in service. Therefore, this alpha case had to be removed. Steve Banks, being a
good M&P engineer did another literature search, consulted others in industry and found that if we
coated the wing box with a substance known as “Turco” and descaled, or etched the stress relieved
assembly, the alpha case could be easily removed. This then became the standard production
process. Large chemical treatment tanks to accommodate these processes were installed in Plant 3.
Other Critical Design Issues
There were many design issues that kept coming out as the design matured. For instance, the
propulsion group wanted to make sure that all of the fuel in the WCS could be fed to the engines at a
great rate. This resulted in a requirement for a set of rather large holes in the bottom skin that would
allow as much as 6 lbs per second fuel flow during maximum acceleration to Mach 2.2. The fuel
cutouts were detailed analyzed and photo elastic models made to make sure the stress concentration
due to the open hole were kept low enough (around a value of K t=2.20) for an adequate fatigue life,
this is shown in Figure 24 below.
Figure 24 – WCS Lower Skin Cut Out for Fuel Transfer
Another propulsion requirement that emerged was to be able to get all the trapped fuel out of the
WCS. This required that we design what became to be known as “weeping” holes in the bottom waffle
skin machined part. The team consulted Ed Rolko and Tom Thielemann who brought in a machining
expert. He in turn evaluated the situation and proposed to used “Electric Discharge Machining” or
EDM to accomplish the task. A series of samples were made and the approach was adopted by the
team, since no drilling or reaming method could accomplish this task due to the limited available
room to insert such tools.
The transfer of fuel from the outer panel to the WCS also required some ingenious design. Propulsion
worked very hard to develop a “trombone-like design that accommodated rotation in two axes and

26
translation of the fuel joint. The design met its intended requirements and worked very well in
service.
The attachment of the main landing gear up-lock was another critical detail. The concern was that the
use of a bolted fitting to the lower skin would introduce a fatigue stress concentration and a potential
fuel leak source. After many iterations a set of tabs were provided outside the fuel tank, but integral
to the lower skin, to which a separate fitting was fastened which in turn provided the main landing
gear support in its stowed and locked position.
Another critical detail was the attachment of the wing outer panel actuator. This actuator had a
250,000 lbs maximum load and needed to be able to be severed without damaging the WCS if the
F-14A had a net arrestment on the carrier. Working with Bob Vogel of Mechanical Design, several
concepts were tried until the final design emerged. A rather large wish bone like titanium fitting was
the winning configuration.
Full depth ribs were designed at the centerline, BL 65 joint, BL 87 rib, and the closure rib. The WCS
upper skin needed to be stabilized for compression column loads. Use of full depth ribs would have
added about 65 lbs of weight and increased weld and manufacturing complexity. An effort to find a
better way resulted in the use of a Warren truss approach. The problem then was how match to drill
the attachment holes inside the WCS, after it was welded and stress relieved. Manufacturing
engineers came up with a way of measuring the distance from hole to hole in the WCS with a tool and
transferring that dimension to the struts that had only one hole in it. The idea worked well, and is
shown in Figure 25.
Figure 25 – Warren Type Truss for WCS Inboard Covers
The pylon attachment was another critical joint. It was designed by Vinny Padden and supported by
others. It is a robust concept that was elegantly executed and incorporated into the WCS structure. A
three point support system was developed with a lug each on the upper and lower covers, plus a so-
called “Lollypop” fitting extending from the lower platters of the pivot joint.
Technical Reviews

27
The WCS received a great deal of attention in Grumman, thus regular design reviews were called.
However, SVP Grant Hedrick had his own, held at on an irregular schedule. Several of the WCS team
members, particularly the stress group, were called to come in and provide a status on some aspect of
the evolving design. Most people felt this was the equivalent of a high school principal visit and went
to each one with a great deal of apprehension. Grant Hedrick for his part would do his homework
ahead of time. He knew in advance where the soft spots were and waited to see if the presenters
would acknowledge the issues. If one left the meeting not bleeding too badly, it was considered a
successful one. Hedrick had the ability to seek out the problems that others could not see; and later
offer some guidelines on how to go about and solve them. He did refrain on giving specific directions;
rather he described how the solution should be pursued. Of course a return visit to show the final
selected solution was inevitable to get his symbolic nod of approval. A slight smile from his face was
considered major win.
One such topic that was discussed at these meetings was how to restore the titanium fatigue life of
the external WCS surface after chemical etching for alpha case. Many approaches were discussed;
however the winning one was to apply saturation shot peening the exterior surfaces in a semi
automatic fashion in a controlled atmosphere. To be sure sample coupons were designed, built and
tested to restore the fatigue properties of the processed WCS before the approached was authorized.
Plasma Arc Welding
I remember very well my first conversation regarding “how do we repair a bad Electron Beam weld
joint" with Henry Beck, a savvy liaison engineer. My reply was to simply re-weld the joint. We found
out rather quickly, if the first E B welds was not acceptable, re-welding in most cases made the
situation even worse.
I recall that on one F-14A early unit the weld at BL. 65 lower cover had gotten so bad from all the re
welding, we really didn't know what do next . The weld was full of large pores; it looked like a zipper
on the x-ray. V P Grand Hedrick send Brandeis Wehle (the thinker) to Plant 2 to help us with this
problem, however after several weeks of trying different repair approaches, we surrendered and
scrapped a full open Wing Center Section sub assembly. On some subsequent units, we drilled out the
internal weld voids and installed interference fit pins. This approach was only acceptable in low stress
areas like between stringers; some early units went out with 4 or 5 interference pins in the weld
joints.
Henry Beck, a liaison engineer, had realized very early in this program that we must get really
smart and improve our E B weld technique, or find other ways to repair defective weld joints to
support the production flow. Henry spent many hours with my friend Dietrich Helm to search for
a solution. Helms was an ex-German U-boat captain who after the WWII moved to the United States
and began a career as a weld engineer, eventually specializing in electron beam welding. EB welding in
1968-1969 period was very new and everything about it was experimental. Therefore, the idea of
developing yet another welding technique ran against common sense. In order to understand the
reasons for bad EB welds we made many welds with various weld parameters, we even deliberately

28
contaminated the joint surfaces, and usually the specimen came out perfect. However, in production
we didn't do as well, we never found out the real cause of large internal voids.
Secretly, Henry Beck went to the Plant 35 library and searched all available Weld journals. He soon he
discovered that plasma arc welding in an argon gas environment produced a clean weld puddle much
better than the popular TIG welding. Dietrich Helm and Frank X. Drumm (materials expert) didn’t
want to hear about this. Frank X. Drumm was offended and told Henry Beck to concern himself with
his MRB (liaison) business. However, Henry was able to convince Ken Payne (a brilliant
welding engineer) to make a few Plasma Arc weld specimen on a G-Job basis (G-jobs were
Government jobs or unofficial, or not officially authorized tasks), the results went through NDT with
flying colors. The next challenge was to make some (dog bones) structural specimen for static and
fatigue testing without creating too much publicity. Henry Beck convinced his MRB Department to pay
for some testing.
Like the saying goes, “the rest is history”, Plasma Arc welding became a standard repair procedure.
Later on we managed to get the Navy to pay for a full scale test program in case we had to repair
battle damaged aircrafts. Several years after Grumman incorporated the use of Plasma Arc Welding as
a repair procedure. Sometime later, machine parts suppliers working for both Grumman and
McDonnell Douglas asked McDonnell Douglas why they did not allow titanium parts to be fixed as
Grumman did. This led to a technology transfer across both companies and now the technique is used
all over the aerospace industry. I remember that John Grandy (the engineer from Vickers) called this
the "Silver Nugget” fix.
Pivot Test 9
Pivot Test 9 was initially scheduled to be a pre-prototype of the wing pivot structure to prove out the
design, analysis, fabrication, inspection and testing procedures for this complex structure. A replica of
the pivot joint was designed and built. The ends of it were made thick enough to attach test adaptors
that would have a much longer fatigue life than the test specimen. The first Test 9 in its completed
form is shown in Figure 26.

29
Figure 26 – WCS Test 9 EB Welded Assembly
The first of these tests was aimed at investigating the bearing, lug platters and pivot functions. The
structure was representative only of the pivot region; the rest was transitions to attach the test
fixtures. The assembled pivot joint is shown in Figure 27.
Figure 27 - Test 9 Pivot Joint Assembly
Test 9 was conducted in Dallas Texas at the Vought structural test lab facility. After approximately 80
percent of its equivalent flight hours and unexpectedly the pivot joint failed. The primary failure had
initiated at a bolt hole in the lower platter. This was caused by an uneven bolt load distribution due to

30
large tolerances specified in an attempt to obtain interchangeable bearings. A second failure initiated
at one of the lower platter welds.
A detail investigation identified that the platter weld had a partially missed seam. In other words the
EB weld had not melted the exit side of the joint, leaving an un-welded portion in intimate contact,
but not joined. The defect was not picked up by x-ray, penetrant, or ultrasonic inspection due to the
faying surfaces being in very intimate contact, but not welded. This realization brought about the use
of “witness lines”, a visual means of confirming that the EB weld had melted the joint in its entirety. In
addition a detail examination of the spherical ball joints at the pivot identify severe galling between
the ball and the platters was taking place. While the galling did not cause any failure at that point, it
was estimated that it would, had the testing continued.
Pivot Test 9A
The fix to the galling problem was to insert a thin (0.062 inch thick) aluminum disc made from 6061-T6
alloy covered with dry film lubricant (Molybdenum disulfide). A new test specimen Test 9A was
redesigned and built incorporating the improvement features mentioned above; plus line reaming of
the pivot bolt holes and thickening of the platters in the bolt region. The specimen was also enlarged
to add the outboard nacelle attachment and internal rib structure to further test the welding
approach selected for the WCS. This new specimen completed a total of six fatigue lives without
failure. However, cracks were found inside the upper cover welds, which at that time were left un-
machined and not stressed relieved, the test specimen is shown in Figure 28.
Figure 28 – Test 9A Showing the Final Wing Pivot Configuration and Nacelle Attachments
Prior to Test 9A it had been considered un-necessary to either machine, or stress relieve those welds
because they were primarily in compression under positive gs. These cracks were attributed to high
tensile residual stresses in the as welded condition. To preclude failures in any of the WCS production
upper covers, the assembly procedures were modified to include a final machining of all internal
welds and stress relieving the entire WCS box welded assembly.

31
Full Scale Testing
Full scale testing was a major aspect of the comprehensive structural development program. Thus a
total of five major development tests were conducted to validate the design approach selected. This
had included the pivot bearing test, two Pivot Test 9, a fatigue test article and the static and Fatigue
aircraft. The fatigue test article for the WCS was successfully completed on December 12, 1971 after
surviving 12,000 equivalent flight hours of testing.
The static test aircraft was also successfully tested to ultimate load. The Test aircraft is shown in
Figure 29 being prepared at the test laboratory in Plant 5.
Figure 29 – F-14A Static Test Article Being Assembled in Plant 5
A major feature of the WCS development was the fatigue test of a complete box subjected to a service
life 6,000 flight equivalent hours multiplied by a factor of 2.0 or 12,000 equivalent flight hours of
extreme-use. No problems were encountered in the WCS welded structure. However, the aft nacelle
attachment or “Y’ fitting had failed due to galling of the attachment pins. The fittings were re-
designed increasing its lug size, using silver plated attachment pins and adding Aluminum-Bronze
bushings to prevent fretting.
This author still has his design notebooks outlining all the issues that the Grumman team went
through and the detail calculations for preliminarily sizing the structural members and manufacturing
approaches. Dick Cyphers at one point asked Carlos Paez to go see the SVP of engineering Dick Hutton
and show him his notebook. He was congratulated for documenting such vital engineering records on
how a great deal of the WCS was being designed and analyzed. It was a very nice thing for Carlos. He
felt he had finally arrived and was properly recognized.
In 1973 Carlos Paez and Tom Taglarine were asked by SVP Grant Hedrick to write an AIAA paper on
how Grumman went about the design and analysis of the structure, it was finally published in 1976
after many months of AIAA reviews. The paper can be found under "Developing the Backbone of the
F-14, AIAA Conference, at Las Vegas, Nevada. June 1976", page 419 to 424. Some of the data from that
paper is included in this write-up.

32
Carlos Paez was inducted into the Masters Fellowship program in the fall of 1970. The program
allowed up and coming engineers to earn their MS degrees while working part time at Grumman and
attending a University during working days. The program also allowed for rotation exposure to other
engineering disciplines other than the original candidate’s home section. For his first two rotations
Carlos was assigned to the F-14A structural analysis under Doug Hutchings. Thus he made a complete
circle from design to stress on a single project. Other later assignments involved dynamic analysis and
advanced concept or preliminary design work. In May of 1972 Carlos Paez got his MS degree in
Structures from New York University.
The F-14A went on to become the mainstay of the US NAVY for over 30 years as its front line fighter.
Its unique configuration and robust structure have been a hallmark of the Grumman Aerospace
Corporation in Bethpage, Long Island, New York. It is a tribute to many who worked to make this
aircraft a success. Figure 31 shows the total structural arrangement of this great airplane.
Figure 31 – Structural Arrangement of the Entire F-14A
The development of the wing center section was a true first for the aerospace industry, an all-welded,
unitized titanium structure weighing 2065 lbs, 19 lbs under its target weight. This structure surpassed
all its structural tests (proof tests, static test and fatigue tests), and no airplanes have ever been
recalled due to any wing center section structural issues in service.
This state of the art advancement in the design and construction of high performance swing wing
aircraft was the result of an extra ordinary effort by a team of Grumman engineers. It was a detail
design and analysis effort, it involved use of untried materials and new manufacturing processes. The
success of this endeavor was the result of work by many individuals from several disciplines. It was
also the result of the talent, dedication and ingenuity of its leaders.

33

Development of F-14A Wing Center Section

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