My Airframe Metallic Design Capability Studies..pdf

Mr. G. A. Wardle MSc. MSc C.Eng. MRAeS.
MY AIRFRAME METALLIC DESIGN CAPABILITY STUDIES.
By Mr. Geoffrey Allen Wardle MSc. MSc. C.Eng. MRAeS. Current Capabilities
My Concept Advanced Variable Cycle Engine XE-137B with
vectoring LOAN Nozzle.
My Multi Materials Structural Layout in Fwd Fuselage of my Future
Deep Strike Aircraft concept design study.

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My design capabilities of metallic material structures for commercial aircraft.
The objective of this presentation is to demonstrate my metallic airframe structural capabilities and
knowledge base of the design, materials, and processes, developed through my career and
academic studies and currently applied to both my FDSA RAeS (APG), and ATDA AIAA design
studies.
Shock Strut
Assembly
Upper
Folding
Backstay
Strut Brace
Backstay Torque Tube
Trunnion
22” diameter
wheel / tyre
assembly
NLG Bay wall / frame attachments
My Nose Landing Gear for my Future Deep
Strike Aircraft concept design study.
Titanium
Steel

 Section 1:- Metallic materials in commercial aircraft and aeroengines.
 Section 2:- Designing Parts for NC Machining and the High Speed Machining Research for Al
and Ti alloys:
 Section 3:- Joining technologies for aerospace structures:
 Section 4:- Advanced Metallic Technology Additive Manufacturing Technology.
 Section 5:- My CATIA V5.R20 Design Capability Machined Part Examples Solid Modelling and
methodologies:
 Section 6:- My CATIA V5.R20 Design Examples of Sheet Metal Parts (Methodology and 2D
drawing development).
 Section 7:- My CATIA V5.R20 Assembly Design Examples (Methodology).
 Section 8:- Operation – oriented Machining Using the Catia V5.R20 Workbench based on
previous Unigraphics V.14 NC simulation work. (In Work).
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Metallic airframe structural design capabilities presentation contents.

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Section1:- Metallic materials in commercial aircraft and aeroengines .
This section covers the application of metallic material to commercial airframe structural
components design shown in figures 1(a)/(b), and aeroengines as shown in figures 2(a)/(b) and 3,
and will be expanded on below as applicable to my current ATDA AIAA design study. The
descriptive work contained herein is based Cranfield University MSc and University of Portsmouth
MSc academic studies Cranfield Aerospace design standards, my ATDA technology research
project, EASA CS 25-Book 1 SUBPART‟s C and D (formally JAR 25. ACJ 25.571) and referenced
texts.
This section also covers the respective design philosophies, with emphasis on damage tolerant
design of metallic components, the basics of designing for structural integrity, material types and
the influence of materials selection on damage tolerance.
As can be seen from the following figures metallic material still have very important roles in modern
commercial airframe and jet engine structures theses include but are not limited to:- wing leading
edge slats; control surface hinges; flap tracks; wing ribs; engine pylons and attachments; landing
gear (struts, stays, attachment pintle‟s, bogie units, torque links, etc.); in earlier aircraft:- fuselage
skin panels: wing and empennage cover skins. For HBR Turbofan aeroengines:- fan blades;
compressor blades; combustion chambers; turbine blades; stationary guide vanes; stationary
nozzle vanes; spools; bearings; and parts of the engine casing. The core body of this presentation
will look at the design for manufacturing and manufacturing processes technology for these
components relative to the ATDA airframe design development study, as well as the generic design
standards applied to their manufacture.

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Figure 1(a):- Materials utilization on current generation commercial airframes.
AL/Li Alloy
CFRP MONOLITIC
CFRP SANDWICH
TITANIUM
QUARTZ GLASS
By weight percentage.
Composites 50%
Titanium 15%
Steel 10%
Other 5%
Figure 1(a)i:- AIRBUS A350-900 XWB Airframe
(external structure application).
Figure 1(a) ii:-BOEING 787-8/-9/-10 Airframe
(external structure application).

Figure 1(b):- My ATDA Port OB Wing section multi material structural assembly model.
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PRSEUS stitched
composite stitched ribs.
Additive Manufacturing
Technology (laser disposition)
Al/Li tip rib.
Additive Manufacturing
Technology (laser disposition)
Al/Li Aileron actuator
attachment ribs.
CFC Thermoplastic
resin spars.

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Figure 1(c):- Materials utilization on current generation commercial airframes.
Figure 1(c) i:- Al/Li wing ribs (5% density reduction over Al alloy).
Figure 1(c) ii:- GLARE (Al alloy and S-2 Glass Fibre) A380 skin panels.
Figure 1(c) iii:- Al/Li Fuselage Cross Beams and
Seat rails (5% density reduction over Al alloy).
Figure 1(c) iv:- Ti Flap Drop Hinges.

Figure 2(a): - Current engine materials are considered for engine environments.
Titanium Fan LP and IP spool compressor.
Nickel HP spool compressor, and turbine blades
and combustors.
Steel used in bearings and stationary vane rows.
Aluminium used in fan case.
Composites engine casings research in
Advance Engine into CFC fan.
Front Fan compressor either
SPF/DB Ti or monolithic CFC
with Ti leading / trailing edge
blades.
Intermediate Pressure 8 stage
Compressor BLISK Ti or BLING Ti
MMC blades and Ti /Steel Stators.
High Pressure Compressor 6 stage
machined solid Ni blades BLISK.
High Pressure / Intermediate Pressure
turbine single crystal Ni blades. Low Pressure
Turbine multi pass air cooled Ni blades.
Steel Stators in the intermediate
Pressure Compressor IP are used
to eliminate the risk of Ti surfaces
friction welding or catching fire.
Low Pressure Compressor
SPF/DB or BLISK Ti or BLING
Ti MMC blades and Ti Stators.
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Figure 2(b):- RR-Trent 1700, 87,000lbs thrust, 3 shaft turbofan example.
The Forward engine
mount takes vertical
and side loads .
The Aft engine mount takes engine
thrust loads, vertical side loads,
and torque moment Mx .
The Fan 118” diameter
SPF/DB Ti or monolithic
CFC blades with kevlar or
R2 glass faces and Ti
blade edges.
Low pressure Fan stage
compressor SPF/DB Ti alloy.
Intermediate 8 stage pressure
compressor machined solid Full
3D aero Ti blades.
High 6 stage pressure compressor
machined solid Ti blades BLISK.
High 1 stage pressure turbine with
directionally solidified hollow Nickel
alloy air cooled blades soluble core.
Low 5 stage
pressure turbine
with directionally
solidified hollow
Nickel alloy air
cooled blades.
Intermediate 2 stage
pressure turbine
Nickel alloy blades.
Ti BLISK technology.

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Figure 3: - RR Advance HBR engine using Monolithic CFC Fan blades.
The Advance HBR Turbofan monolithic
CFC Fan blade which could use kevlar or
R2 glass faces and Ti blade edge and root
members.
The Advance HBR Turbofan engine layout using monolithic CFC Fan
blades and proposes more extensive BLISK, and BLING MMC technology
in compressor stages pioneered on the current Trent family.

Aircraft structures fall into 3 categories which are as follows:-
Class 1:- structural component the failure of which will result in structural collapse; loss of control;
failure of motive power, injury or fatality (to any occupant); loss of safe operation of the aircraft.
Class 2:- Stresses components but not Class1.
Class 3:- Unstressed or lightly stresses component which is neither Class 1 or 2.
Structural integrity is defined as the capability of the structure to exceed applied design loading
throughout its operational life, and the selection of a design philosophy to achieve this from the start
of the design process is extremely important as this selection impacts on:- airframe weight;
maintainability; service life; and any future role change of the airframe. The approaches available to
the designer are:- Static Strength; Safe Life; Failsafe; Damage Tolerance; and Fatigue Life, the last
four of which, are expanded below (ref:-4). See tables 3 through 5 for ATDA candidate materials
selection.
(a) Safe Life:- The important criterion in this approach is the time before a „crack or flaw‟ is initiated
and the subsequent time before it grows to critical length. It can be seen from a typical S-N
curve that low levels of stress at high frequency of application theoretically do not cause any
fatigue damage. However it is necessary to allow for them, possibly by introducing a stress
factor such that effectively damage dose not occur.
(b) Fail-safe:- In this approach the dominate factors are the crack growth rate and the provision of
structural redundancy in conjunction with appropriate structural inspection provision.
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Structural design philosophy of airframe structural components.

There are several ways of ensuring that fail safety is achieved:-
i. By introducing secondary, stand-by components which only function is in the event of a
failure of the primary load path, to carry the load. This may consist of a tongue or a stop
which is normally just clear of the mating component. A mass penalty may be implied but in
same circumstances it is possible to use the secondary items in another role, for example
the need for a double pane assembly on cabin windows for thermal insulation purposes.
ii. By dividing a given load path into a number of separate members so that in the event of the
failure of one of them the rest can react the applied load. An example of this is the use of
several span wise planks in the tension surface of metallic wing boxes. When the load path
is designed to take advantage of the material strength the use of three separate items
enables any two remaining after one has failed to carry the full limit load under ultimate
stress. In some instances the „get home‟ consideration may enable a less severe approach
to be adopted.
iii. By design for slow crack growth such that in the event of crack initiation there is no danger
of a catastrophic failure before it is detected and repaired.
c) Damage tolerant:- With this philosophy it becomes necessary to distinguish between
components that can be inspected and those that cannot. Effectively either the fail-safe or
safe-life approaches are then applied, respectively, in conjunction with design for slow crack
growth and crack stopping (e.g. panel braking web stiffeners).
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Structural design philosophy of airframe structural components.

A. Safe-life and Fail–safe design processes (see Chart 1):- There is a commonality in the design
process for the safe –life and fail-safe concepts. The material to be used for the structure must
be selected with consideration of the critical requirements for crack initiation or crack growth
rate, as most relevant, together with the operating environment. A vital consideration for fail-
safe design is the provision of the alternative load paths, possibly together with crack
containment or crack arresting features. When these decisions have been made it is possible to
complete the design of the individual components of the structure and to define the
environmental protection necessary.
In the case of the safe-life concept the life inclusive of appropriate life factor follows directly
from the time taken initiation of the first crack to failure. Inspection is needed to monitor crack
growth. In the fail-safe concept the life is determined by the structure possessing adequate
residual strength subsequent to the development and growth of cracks.
In both cases it essential to demonstrate by testing, where possible on a complete specimen of
the airframe, that the design assumptions and calculations are justified. Further, in fail-safe
design it is necessary to inspect the structure at regular, appropriate intervals to ensure that any
developing cracks do not reach the critical length and are repaired before they do so.
As the design process is critically dependent upon assumed fatigue loading it is desirable, if not
essential, to carry out load monitoring throughout the operational life of the airframe. This is
used either to confirm the predicted life, or where necessary, to modify the allowable
operational life.
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Structural design philosophy application processes.

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Safe-Life.
Crack Initiation time.
Fail-Safe.
Crack growth rate.
Provision of redundancies.
Crack containment.
Environment.
Material: Component Design:
Corrosion protection: Testing.
Life. Residual strength.
In service load monitoring.
Chart 1:- Application of Safe-life and Fail-safe structural design philosophies.

B. Damage Tolerant Design process (see Chart 2):- The damage tolerant approach commences
with the assumption that cracks or faults are present in the airframe as manufactured.
Experience suggests that these vary in length from 0.1mm to as much as 1.5mm. Those items
of the structure which may be readily inspected can be designed by selecting an appropriate
material and then applying essentially a fail-safe approach. The working stress level must be
selected and used in conjunction with crack stopping features to ensure that any developing
cracks grow slowly. Inspection periods must be established to give several opportunities for a
crack to be discovered before it attains a critical length.
When it is not possible to inspect a particular component it is essential to design for slow–crack
growth and ensure that the time for the initial length to reach its critical failure value is greater
than the required life of the whole structure. Since this approach is less satisfactory than that
applied to parts that can be inspected it is desirable to develop the design of the airframe such
that inspection is possible, wherever this can be arranged. As with safe-life and fail-safe
philosophies testing is needed to give confidence in the design calculations. Likewise, in-service
load monitoring is highly desirable for the same reason. This design philosophy is employed on
this project using techniques from ref:-4, CS-25 Book 1: SUBPART C, and data sheets, MSc
F&DT module notes.
C. Fatigue-life Design process (see Chart 3):- The first stage in the fatigue-life approach is the
definition of the relevant fatigue loads and the determination of the response of the aircraft
structure to these loads. The analysis for this follows that for limit load conditions, which
enables the loading on individual components of the airframe to be determined, and the
airframe structural response to be assed and the best design philosophy to be applied. 15
Structural design philosophy application processes.

Chart 2:- Application of the Damage Tolerance structural design philosophy.
Damage Tolerant.
Crack in structure as manufactured.
Is the component inspectable?
Yes. No.
Fail-safe approach.
Slow crack growth.
Crack arrest features.
Inspection periods.
Crack growth to initiate
failure to be more than
service life.
Testing.
In service load monitoring (FTI / G monitors / SHM).
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Chart 3:- Application of the Fatigue-life structural design philosophy.
Fatigue-life.
Aircraft structural response.
Fatigue load spectra.
Design philosophy selection.
Damage Tolerant.
Safe-Life. Fail-Safe.

Fatigue Design Requirements:- The emphasis of the requirements specified to ensure the integrity
of the airframe design under fatigue loading is on the methods of analysis and the means of
determination of a satisfactory fatigue life. Only in the United States military code is there a
specification of a magnitude and frequency of repeated loading and this is outlined below. Loading
conditions for all categories of aircraft are discussed below.
1) Civil transport aircraft CS-25 Book 1:- This standard outlines the basic requirements for fatigue
evaluation and damage tolerance design of transport aircraft. The paragraph outlines the
general requirements for the analysis and the extent of the calculations. Amplification of the
details is given in the associated „acceptable methods of compliance‟ given in CS-25 Book1
SUBPART C (formally JAR 25.ACJ 25.571).
2) UK Military Aircraft:- The basic requirements for fatigue analysis and life evaluation are
specified in Def Stan 00970 Chapter 201. This covers techniques for allowing for variances in
the data as well as overall requirements and the philosophy to be adopted. Detail requirements
of the frequency and magnitude of the repeated loading are given in the particular specification
for the aircraft.
3) US Military Aircraft:- The United States military aircraft stipulations are to be found in three
separate documents:- In MIL-A-8866A the emphasis is on the detail of the required magnitude
and frequency of the repeated loading rather than on analysis the data covers;- maneuver;
gust; ground and pressurization conditions for fighter, attack, trainer, bomber, patrol, utility and
transport aircraft.
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Structural design fatigue requirements for design philosophy application.

MIL-A-8867 prescribes the ground testing to be undertaken as part of the demonstration of the
life of the airframe. The final document the is MIL-8868 paragraph 3.4 and 3.5 which stipulate
the information to be provided in the form of reports outlining the analysis and testing
undertaken to substantiate the life of the airframe.
 The types of repeated airframe load data required for design against fatigue and to apply in the
selected component design philosophy are outlined below.
1) Symmetric manoeuvre case:- Extensive information is available in relation to symmetric
manoeuvres of both military and civil aircraft, e.g. Van Dijk and Jonge‟s work which outlines a
fatigue spectrum obtained from flying experience of fighter / attack aircraft which is known as
the FALSTAFF spectrum, based on the maximum value of peak stress (s) and loading
frequency (n) the peak stress selected being the Input Parameter.
2) Asymmetric manoeuvre case:- Fatigue loading data for asymmetric manoeuvre loading is
sparse, and these originate from the roll and yaw controls, the texts of Taylor derives data from
early jet fighter experience. As for civil aircraft it has been determined that atmospheric
turbulence is of much greater significance.
3) Atmospheric turbulence:- Fatigue loading due to encounters with discrete gusting or the effect
of continuous turbulence is of importance for all classes of aircraft, but especially for those
where operational role does not demand substantial manoeuvring in flight. ESDU data sheets
69023 (Average gust frequency for Subsonic Transport Aircraft (ESDU International plc. May
1989) is used in this study.
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There are two main types of turbulence which are:- (a)Symmetric Vertical Turbulence, and
(b)Lateral Turbulence.
a) Symmetric Vertical Turbulence:- where gust magnitude is a function of both flight altitude and
terrain over which the aircraft is flying, e.g. low level penetration bombing missions B-1B,
Tornado, and B-52H, where there are more up gusts than down, these are allowed for by
using correction factors.
b) Lateral Turbulence:- there is less information on the frequency and magnitude of lateral
turbulence for aircraft but it has been suggested that at altitudes below about 3km the
frequency of a given magnitude is some 10-15% greater than those of the corresponding
vertical condition.
4) Landing gear loads:- these fall into three categories;- (a) loads due to ground manoeuvring e.g.
taxiing; (b) the effects of the unevenness of the ground surface e.g. unpaved runways, rough
field poor condition runways, major consideration in troop / cargo military transports, and
forward based CAS aircraft; (c) landing impact conditions. The texts of Howe (ref:-4): Niu: and
MIL-A-8866A are employed in this project.
5) Buffeting Turbulence:- Flow over the aircraft may break down at local points and give rise to
buffeting. This induces a relatively high – frequency variation in the aerodynamic loads,
possibly resulting in the fatigue of local airframe components such as metallic skin panels.
6) Acoustic Noise Turbulence:- local high frequency vibration or flow field loading, and ESDU Data
sheets 75021 and 89041 were used in this project.
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Chart 4:- Structural damage tolerance requirements for application.
Usage Load
Spectrum
Damage
Summation Life.
Component
Stress
Analysis
Material c/a
fatigue data
Component fatigue data.
Loads and
usage
variation.
Material and
component
variation.
Errors in
models.

Damage Tolerance definition:- A structure which is designed to retain its required residual
strength for a period of use after the structure has sustained specific levels of detectable fatigue,
corrosion, or accidental damage.
The inputs required for damage tolerance analysis are as follows:-
 Service loading spectra:- must be typical for anticipated use and service environment; for crack
growth analysis, sequencing effects are important therefore the most probable ordering of
cycles should be maintained:
 Stress analysis is required to convert the strain gauge measurements to stresses at sites of
crack initiation:
 Analytical or numerical calculations are required to obtain the stress intensity as a function of
crack length from crack start length to final failure, however there are possible problems due to
lack of knowledge of what the actual crack path will be. Standard solutions for simple
geometries are available in reference works. Difficulties are within thumbnail cracks and cracks
at notches, or in multiaxial states of stress which are difficult to design for, require detailed
analysis:
 Fatigue crack growth rate data for selected material:
 Starting defect sizes (distributions):
 Non destructive inspection capability (limits to defect size detection):
 Crack growth rate predictive models:
 Structural testing.
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Structural damage tolerance requirements for design philosophy application.

Damage tolerance design recommendations of CS 25.571 which are applied to the ATDA
project:-
 Design features are to be considered to achieve damage tolerance:- Multiple load path
construction and the use of crack stoppers: Materials and stress levels that after crack initiation
provide controlled, slow rates of crack growth: Arrangement of design details to ensure a high
probability of crack detection before strength is reduced below the limit load capability:
Provision should be made in the design to preclude the possibility of multiple site damage.
 Full scale fatigue test of two or more times the design life to demonstrate damage tolerance,
plus inspection to assess damage growth:
 Probabilistic analysis may be used particularly with fail safe structures:
 Examples of critical features to be considered:
 As far as is possible all structural parts are to be inspectable:
 Inspection stipulations very general: – Inspection is ultimate control and guidance information is
provided by the manufacture to assist operators in the frequency, extent and methods of
inspection of the critical structure:
 Thresholds for inspection – where it can be shown that a load path failure in fail safe, or a
partial failure in a crack arrest structure, can be detected then thresholds can be established via
fatigue analysis or slow crack growth analysis:
 For single load path structures, thresholds should be based on crack growth analysis, assuming
maximum manufacturing defect size.
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CS 25.571 Damage tolerance and fatigue evaluation:-
 Show that catastrophic failure due to fatigue will be avoided throughout the life of the aeroplane:
 Each evaluation must include;- typical loading spectra; identification of critical points the failure
of which will cause catastrophic failure of the aeroplane; and an analysis of these points:
 Inspection or other procedures must be established as necessary to prevent catastrophic
failure:
 The evaluation must include a determination of the probable location and modes of damage
due to fatigue, corrosion or accidental damage.
 At any time in the service life the residual strength must withstand all limit load conditions:
 The aeroplane must successfully complete a flight in which likely structural damage occurs as a
result of bird impact or uncontained engine burst.
USAF and Civilian (FAA / JAA) Damage Tolerance:-
 Same basic approach as above:
 USAF is prescriptive on the following;- Initial defect size; Inspectability; Design lives:
 But does permit uninspectable details:
 The FAA / JAA has a more general approach;- that is a manufacturer must demonstrate that
damage can be detected and that an inspection regime can be defined:
 The USAF has a fail safe and damage tolerance by slow crack growth as separate categories,
where as the FAA / JAA prefers to combine them, but will accept slow crack growth without fail-
safe; but fail-safe without slow crack growth is not permitted. 24

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Figure 4:- Multiple Load Path Structure, external Inspectable.
External Inspection
Broken stiffener cannot
be seen from outside
Critical at limit load
Safe period for inspection
Failure of primary
member (stiffener)
Crack
Size.
Flights.
Detectable skin crack

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Figure 5:- Multiple Load Path Structure, Inspectable for less than load path failure.
aρ
FAIL – SAFE FOR PLANK COMPLETELY FAILD.
Inspectable crack in plank.
Safe period for inspection.
Critical at limit load.
Life in secondary
member subsequent
to primary member
failure.
Crack
Size.
Flights.
Secondary member.
Detectable crack in
primary member.
 This approach may be permitted but the
crack in the primary member must be
readily inspectable.

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Figure 6:- Multiple Load Path Structure, damage tolerant.
SINGLE LOAD PATH
INTEGRALLY STIFFENED.
Obvious partial failure.
LOWER WING COVER SKIN.
Safe period for
inspection.
Critical at limit load.
Detectable.
Crack
Size.
Flights.
 This type of structure is not
recommended but allowed.
 If used it must be shown that
damage will be readily found
before it becomes critical.

Key assumptions and required capabilities for damage tolerant design:-
 Significant proportion of the structures life to failure is occupied by crack growth:
 Predictive capability for damage growth rates and damage extent:
 Damage detection and monitoring techniques with accuracy compatible with rates of damage
growth and damage influence on residual strength of the structure:
 Accurate calculation of damaged residual strength of the structure:
 The ability to design, select materials, and fabricate structures that are resistant to discrete
damage. The design practices to produce a resistant design based on experience and analysis
are collected in the Reference Structural Design Principles Documents (RSDPD) or Company
Design Standards. Figures 7 through 10 give examples of design for structural integrity best
practice from the ATDA RSDPD the layout of which is based on the Airbus titles and the
substance is derived from references 1 &2. Therefore for the ATDA project the RSDPD is split
into two books with seven specific volumes:- Book (1) Generic Design Principles consisting of:-
Volume 1;- General Design Principles; Volume 2;- Composite Design Principles; Volume 3;-
Metallic Design Principles: Book (2) Component Specific Design Principles consisting of :-
Volume 4;- Fuselage Design; Volume 5;- Wing Design; Volume 6;- Propulsion Integration
Design; Volume 7;- Empennage Design.
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T
Max 0.3 T
Min 2.0 rad
Counterbores
T
0.3 „T‟ min
Countersink
1) The depth of counterbores should be no
greater than 0.3 times the thickness of the
material.
2) Countersink should be no more than
70%
Fastener Skin OML intersection points.
Distance from
flange edge 2 x
fastener diameter.
Fastener
Vector.
Fastener
Vector.
Fastener Rib IML intersection points.
For metallic parts the minimum distance
to flange edge is 2 x fastener diameter.
Example taken from Rib 12 of my ATDA
Airframe Design Project.
Figure 7(a):- Design for structural integrity metallic fastener examples.

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Figure 7(b):- Design integrity stress raisers at changes in section and fillet radii.
Designing for thickness changes and fillet radii.
t
r
h
Radius „r‟ : The lesser of
r = 0.5 t
r = 2 h
t 1
r
h
r = t 1
r = 2 h
t 1 r
t 2
r = t 1
r = t 2
t 2
t 1
r = t 1
r = 0.5 t 2
r
r = 0.5 t 1
r = 0.5 t 2
t 2
t 1
r
 Note : where the rule results in a radius of less than minimum Reference Structural Design Principle minimum then
the RSDPD minimum will be used.
Examples taken from my ATDA Airframe Design Project Al/Li Wing Root Rib.

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Figure 8(a):- Design for integrity avoiding KT stress raisers in machined parts.
KT
Corner
radius
Flange
radius
X To be avoided
RSDPD for minimum
separation
Corner
radius
Flange
radius
 Recommended
Where corner radius and flange radius meet.
Fillet
Corner
rad
KT
X To be avoided
Fillet
Corner
radius
RSDPD for minimum
separation
 Recommended
Example of a 5-Axis landing
 Basically:- avoid design stress concentrations in
to the part which could act as damage sites.

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Figure 8(b):- Design for integrity avoiding KT stress raisers in machined parts.
RSDPD for minimum
separation
KT
External
radius
Fillet
X To be avoided
External
radius
Fillet
 Recommended
Where an external radius and fillet meet.
Stiffener
radius
KT
Flange
radius
X To be avoided
Stiffener
radius
RSDPD for
minimum
separation
Flange
radius
 Recommended
Example of a stiffener radius and flange radius meeting.
 Basically:- avoid design stress concentrations in
to the part which could act as damage sites.

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Figure 9:- Web scalloping stress raisers in an acoustic fatigue area.
Crack failure
at fillet / radii.
Figure 9(a).
Make flange top flat.
Figure 9(b).
Figures 9(a) and 9(b) Scalloping of stiffeners to reduce structural weight should be
avoided in areas of acoustic loading, cyclic loading and vibration, as cracks can
start at fillets / radii, so where possible keep stiffener tops flat.

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• Details parts that are made out of extruded plate, extrusions, and forgings have „Grain Direction‟
identified see figure 10:
• Grain direction is determined by the structures group and is to be shown in 3d models and on
2d drawings:
• If removal of Dead Zones is critical, then a note in the model or on the drawing is required:
• If grain direction is not critical then a note is required in the 3d model or on the 2d drawing
stating that either the Grain direction immaterial or Grain direction control not required for
structural purposes.
• Max Material sizes : Material ThicknessLength Width
Aluminium RSDP (Metallic design principles).
(Thickness in 10mm increments)
Titanium RSDP (Metallic design principles).
(Thickness in 10mm increments)
• Any components outside these sizes would require a forged billet or forging
• Material has different thickness bands which are defined as „Ruling Section‟ or „Ruling
Dimension‟.
• Each band has different properties, Structural engineers will determine the „Maximum Ruling
Dimension‟ for each detail.
Design for integrity effect of grain direction on machined parts.

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Figure 10:- Design for integrity grain direction definition.
ST
LT
L
L
LT
LT
ST
ST
ST
Parting Plane
ST Across parting plane.
Figure 10(a):- Plate, Strip,
and Sheet.
Figure 10(b):- Extrusion.
Figure 10(c):- Forging.

The role of strength level and toughness.
 Effect of strength on smooth specimen fatigue properties:- It is found experimentally that fatigue
initiation lives as measured on smooth specimens are sensitive to both the static ultimate
tensile strength and also the ductility of the material. High cycle fatigue strength is dependent
almost exclusively on the static ultimate tensile strength. For quenched and tempered steels,
the fatigue strength at 106
to 107
cycles is ≈ 0.5 of the static UTS, although for certain types of
steel – notably ferrite – perlite microstructures the ratio can be much lower – 0.35 to 0.4, for
aluminium alloys, the high cycle fatigue strength is ≈ 0.3 of the UTS (Ultimate Tensile Strength).
This lesser influence of strength level is also found in comparison of the fatigue performance of
the 2XXX series alloys (e.g. 2024) with the higher strength 7XXX series alloys (e.g. 7075). It
has been found experimentally that the high strength alloys have little or no improvement in
smooth specimen fatigue properties.
Ductility is an important property at low cycles (102
to 104
cycles) and high applied stress where
plasticity is occurring in the sample. Under these conditions the greater the static ductility, the
better the fatigue life. At low cycles plastic strain range is a better predictor of life than the
applied stress range. Two materials – one a high strength low ductility, and the other a lower
strength with higher ductility, will have a cross over in their S-N behaviour.
Variable amplitude loading, relevant to service applications will contain cycles of ranges that
causing both high and low cycle fatigue, and both will contribute to damage. The best material
for a particular load spectrum will depend on the match of the S-N curve to the spectrum in
question.
36
Materials Selection for fatigue and damage tolerance design.

The material with the longest life under variable amplitude loading can be determined by a life
calculation in which the damage contributions of both low and high cycle portions of the
spectrum are evaluated using Miners rule.
 Effect of material strength on crack growth rates:- Increases in material strength level dose not
improve resistance to fatigue crack growth rates at cracks greater than 1mm to 2mm. In fact
high strength low ductility materials tend to have marginally faster growth rates for the same
value of ∆K than do lower strength variants of the same alloy system.
The material property with most effect on fatigue crack growth rates is the stiffness or elastic
modulus E. Crack growth rates are inversely proportional to this parameter, and a plot of ∆K / E
for common aerospace materials e.g. Steel, Aluminium, and Titanium alloys reveals a common
line. Generally steels have the slowest crack growth rates followed by titanium and with
aluminium having the fastest crack growth rates.
Static fracture toughness will influence fatigue lives in fatigue crack dominated regimes in two
ways:-
1) Firstly, as the crack grows and stress intensity factor K increases, Kmax , the maximum in
the stress intensity cycle approaches K1c or Kc the material fracture toughness. In this
region crack acceleration occurs, as static fracture modes such as cleavage; void
coalescence; and intergranular failure add to the normal cyclic fatigue crack growth
increment. This process will occur to a greater extent and at smaller values of Kmax for low
K1c materials.
37
Materials Selection for fatigue and damage tolerance design (continued).

2) Secondly, final failure will occur at smaller values of stress intensity and shorter crack
lengths, when Kmax becomes approximately equal to Kc or K1c. This effect has little
influence on the life to failure, as the cycles occupied in growing the crack in the fast growth
rate regime are a small fraction of the total. However, it has great implications for damage
tolerance, as long cracks prior to catastrophic failure have much greater detectability than
short cracks.
 Influence of material properties on safe life and damage tolerant design:- Safe life design will be
influenced primarily by material strength. High design stresses will require high strength
materials to resist them, and these will in turn have good high cycle fatigue strength. If the
variable amplitude loading to which the component is subjected is very irregular and contains
occasional high stress cycles then good ductility will also be required. Frequently there is a
trade – off between high strength and good ductility and toughness and resistance to fatigue
crack growth. Fatigue design with high strength materials is relatively easy to achieve for safe
life designs. Damage tolerant and fail safe design is difficult to achieve with high strength
materials.
For fail safe and damage tolerant design a high static toughness is required so that the
component or structure can withstand the longest cracks possible without catastrophic fracture.
Equally fatigue crack growth rates must be as slow as possible for a given value of ∆K.
Designing for damage tolerance with high strength materials is difficult for the following
reasons:-
38

1) Failure crack length:-
a) There is an inverse relationship between fracture toughness and strength level; the grater
the strength level the lower the fracture toughness, This will reduce the crack length at a
failure for a given value of design stress.
b) As a high strength material is being used, the design stress level will be increased (there is
no point in employing a strength material otherwise). This will have the effect of reducing
the crack length at failure still further according to the expression:-
2
ɑ = 1 K1c
𝝅 σ
Where ɑ = the crack length: K1c = the fracture toughness: and σ = the applied stress. As K1c
declines and σ increased : ɑ will shrink rapidly.
In addition crack growth rates at increased stresses will be similarly driven faster by the
higher stresses, and life will be decreased according to the factor 1/∆𝛔𝐦
where ∆σ is the
applied stress range and m the exponent in the Paris law.
𝑁𝑓 = 1 (𝛼𝑓
1−𝑚/2
- 𝛼𝑖
1−𝑚/2
)
f (ɑ / w) C∆ 𝛔𝐦
π𝐦/𝟐
Where 𝑁𝑓= cycles to failure: ɑi and ɑf are initial and final crack lengths respectively: f(ɑ / w) is
compliance correction : C is the constant in the Paris law. 39

40
Figure 11:- Materials Selection the influence of strength (Design stress).
STRESS.
LIFE. DESIGN LIFE.
Peak stress in spectrum.
0.2% Proof Strength for material.
 Strength level influences crack growth behaviour
through the design stress. Greater design stresses
will increase the crack growth rates and reduce
crack lengths at failure. The K1c and da/dN
behaviour must be proportionately greater in high
strength materials to compensate for this effect.
 In fact the reverse is frequently found; high
strength materials have K1c and da/dN values no
better than lower strength materials, and frequently
they are worse. To maintain the same life, design
stress levels must be reduced, hence a smaller
fraction of the static strength of high strength alloys
can be used.

41
Figure 12:- Materials Selection the Role of K1c.
CRACK
LENGTH.
CYCLES. N1 N2
Crack length at failure low K1c.
Crack length at failure high K1c.
 K1c influences life in two ways:-
1) Firstly it determines the crack length at
failure; when Kmax = K1c sudden failure
occurs. Low K1c will produce failure at
smaller crack lengths than high K1c. It has a
relatively unimportant effect on life as most
of the life has been consumed by this stage.
2) Secondly, K1C can affect the constants C
and m and thus influence the pre-failure
crack growth rates.
Generally, high K1c is associated with ductile
materials which tend to have reduced da/dN
for a given value of ∆K.

2024-T3 Aluminium Ti-318 Titanium 4340 Steel
UTS = 480 MPa:
Smooth specimen:
Fatigue limit, R = 0:
200 MPa range.
UTS = 1,000 MPa:
Smooth specimen:
460 MPa range.
UTS = 1600 MPa:
Smooth specimen:
750 MPa range.
∆Kt = 4 MPa 𝑚1/2
:
Fatigue limit with 0.1mm crack
= 200 MPa.
∆Kt = 5 MPa 𝑚1/2
:
= 263 MPa.
∆Kt = 5 MPa 𝑚1/2
:
= 263 MPa.
= 67 MPa.
= 83 MPa.
= 83 MPa.
42
Table 1:- Initiation and crack growth behaviour of aerospace metal alloys .
*Reference 3:- Cranfield University experimental data.

Constraint
and m. 2024-T3 7075-T6
Crack Growth
Rate
m/cycles.
𝟏𝟎−𝟖
- 𝟏𝟎−𝟕
𝟏𝟎−𝟕
- 𝟏𝟎−𝟔
𝟏𝟎−𝟔
- 𝟏𝟎−𝟓 𝟏𝟎−𝟖
- 𝟏𝟎−𝟕
𝟏𝟎−𝟕
- 𝟏𝟎−𝟔
𝟏𝟎−𝟔
- 𝟏𝟎−𝟓
Ct 5.2 x 10−13
1.9 x 10−11
7.3 x 10−12 2.0 x 10−10
1.1 x 10−10
m 5.2 3.7 4.7 3.0 3.2
43
Table 2:- Crack growth constants C and m for 2024-T3 and 7075-T6.
*Reference 3:- Cranfield University experimental data.

The influence of strength level on Critical Crack Length at failure.
𝐾 = 𝜎√𝜋𝑎𝑓(
𝑎
𝑤
)
Therefore:-
𝑎 =
1
𝜋𝑓(
𝑎
𝑤
)
(
𝐾
𝜎
)2
If the value of K is reduced by 2 and σ is increased by 2 a factor of 16 reduction results.
Hence:- Reduction in K1c : Increased service stress : Reduces failure crack length.
44

In this subsection the challenging environment and materials selection will be discussed, leading
into the metallic components manufacturing and processing methods, in subsequent sections of
this presentation. Almost all jet engines in currently manufactured and near term project commercial
jet aircraft are Turbofans as shown in figure 13, twin spool for medium / short haul such as the
Airbus A-320, and Boeing B-737 families, and three spool for long haul aircraft such as the Airbus
A-350 family and Boeing B-787 family. The major reasons for the High Bypass Ratio Turbofans
dominance are the much greater fuel efficiency at subsonic speeds than turbojets, (the HBP
Turbofan produces about twice as much thrust for the same fuel consumption as a turbojet of the
same core size), and much lower acoustic footprint, only HBP Turbofans will be considered here
because of their direct relevance to my ATDA airframe design development research project,
although low to medium bypass turbofans are used in military applications.
For a modern large commercial HBR Turbofan the fan (see figure 2) passes over one tonne of
airflow per second, which produces around 75% of the engines thrust: and the overall compression
system pressure ratios are now approaching 50:1, and compressor exit temperatures can be more
than 700ºC.
In normal operating conditions, air at ambient pressure and temperature (which could range from
sea level to 35-45,000ft), is drawn into the engine through the fan at a velocity of 150m/s, and then
is subsequently compressed through the compressor stages up to 10 atm, before reaching
temperatures of 1300ºC to 1500ºC in the combustion chambers. The resulting excited gas is
expanded through the turbine stages, finally exiting the exhaust nozzle at 500ºC and recovering the
initial pressure with a velocity of 500m/s, a generic illustration of this is shown in figures 14 and 15.
45
Materials Selection and design of commercial aeroengines.

46
Requires high:-
 Overall pressure ratio:
 Turbine entry temperature:
 Bypass ratio.
Range
Fuel consumption.
Long / Medium-Haul
(40,000-100,000lbs thrust):
Three-Shaft Configuration.
Short / Medium-Haul
(8,000 - 40,000lbs thrust):
Two-Shaft Configuration.
Acquisition Cost
Maintenance
Simpler engine, hence moderate:-
 Overall pressure ratio
 Turbine entry temperature
 Bypass ratio
Figure 13: - Turbofan Engine type application for long and medium / short haul aircraft.

47
0
40
0
1500
Figure 14: - The High Bypass Turbofan Engine generic operating environment.
Pressure
(atmospheres)
Temperature
(degrees ºC)

48
Figure 15(a)/(b): - The selection of engine materials and engine operating environment.
Figure 15(a): Component Temperatures.
Figure 15(b): Component Materials Selection.
50ºC
400ºC
800ºC
1000ºC
1300ºC
Composite
Aluminum alloy
Titanium alloy
Nickel alloy
Stainless steel
C/L

Because of this extreme working environmental conditions of temperature and pressure outlined
above and shown in figures 14 and 15(a), coupled with corrosion, stress, fatigue, and weight
constraints, a current state of the art turbofan engine components are made from the following
materials as shown in figure 15(b).
 In general terms, titanium alloys are used for their high strength to weight ratio, excellent heat
and corrosion and density properties (see figure 16) in:- fan blades: tanks: and low and
intermediate pressure compressor stages: and also in exhaust nozzles.
 At high temperatures titanium alloys are replaced by nickel based alloys for example:- in the
high pressure compressor: combustion chamber: and high and low pressure turbines.
 Stainless steels like jethete are used in static parts of the compressor and bearings among
other applications.
 Aluminium alloys can be used in compressor casings: inlet ring and cone applications:
 Composites are currently used for fan casings: fan blades: and cowls.
Figure 15(b) shows the current application of material throughout the engine structure and currently
Nickel and Titanium alloys represent 70% of the weight of the engine, and the use of aluminium and
steel have diminished and composites have increased in the last few years. Materials selection, is
not only dependent on engine performance in terms of weight and fuel burn, and normal operating
conditions, but most critically on the material and manufacturing processes ability to withstand the
structural loading of impact events such as;- Bird strike; Ice ingestion; Blade out; Water ingestion;
Fatigue, and flutter loads.
49
Materials Selection and design of commercial aeroengines (continued).

 Low Pressure Spool:- Material selection Titanium Alloy Compressor fwd of the
combustor, because of its higher specific strength over the temperature range of 0ºC up to
150ºC as Titanium has a higher specific strength over this temperature range than either Nickel
alloys or Steel alloys and is considerably lighter for the equivalent strength. For the main
forward swept fan either SPF/DB Titanium alloy or Monolithic CFC materials can be used with a
metallic leading edge. Nickel Alloy Turbine aft of the combustor, would be used for the low
pressure 6 turbine stages with blades being multi pass machined air cooled blades as these
alloys are better suited to long exposure to the highest part of the this temperature range (see
figures 14,15,&16).
 Intermediate Pressure Spool:- Again the material selection would be Titanium Alloy
Compressor fwd of the combustor, because of its higher specific strength over the
temperature range of 150ºC up to 450ºC. Using BLISK Ti blades for compressor stages or
BLING Ti MMC to reduce weight for the 8 IP compressor stages, as Titanium has a higher
specific strength over this temperature range than either Nickel alloys or Steel alloys and is
considerably lighter for the equivalent strength. Nickel Alloy Turbine aft of the combustor, for
the single stage Intermediate turbine stage, the blades could be single crystal thermally coated
blades to withstand long exposure to the higher operating temperature (see figures 14,15,& 16).
 High Pressure Spool:- For this spool Nickel alloys All , alone would be used because at the
higher temperature range of up to 750ºC Titanium alloys begin to loose their Specific strength
advantage over Nickel alloys and therefore the 6 high pressure compressor stages, and the
single high pressure turbine single crystal blades stage would be Nickel alloy (see figures
14,15,& 16). 50
Summary of HBP Turbofan 3 spool engine material applications.

51
Specific
Strength
Nickel Alloy
Steel
Aluminium Alloy
Titanium Alloy
Temperature
Figure 16:- Metal alloy strength variation with temperature.
*Reference 4:- Rolls Royce Published data.

The front fan blade disc is the first component of the low pressure compressor of the turbofan
engine and a very large mass of air is drawn through the front fan and is accelerated and split into
two flows, the first flow passes into the turbofan‟s core, and the second flow “bypasses” the core
altogether, passing through an annular bypass duct that surrounds the core, and eventually both
flows pass through separate or integrated propelling nozzles at the rear of the engine to generate
thrust. This cooler air flow also provides a degree of noise suppression whilst adding thrust. The
ratio of mass – flow air bypassing the engine compared to the amount of air passing through the
core is termed the “bypass ratio” and these engines derive their thrust by accelerating a large
volume of air to a velocity just above the aircraft‟s flight velocity, and as thrust is proportional to Vjet
but fuel consumption goes with V²jet, the turbofan gives about twice as much thrust for the same
fuel consumption as a turbojet of the same core size.
The fan system has two primary functions:- (1) Compress the bypass air: (2) Feed supercharged air
into the core. The commercial High Bypass Ratio Turbofan figure 17(a), has a pressure ratio in the
order of 2:1, and this bypass air expands through the exhaust nozzle and contributes approximately
75% of the engine thrust. A military Low Bypass Ratio Turbofan figure 17(b), has a pressure ratio in
the range 3:1 to 4:1, and this air passes down a bypass duct and is then mixed with the core airflow
from the turbines, and expanded through the exhaust nozzle. In this case the bypass air can also
be used for afterburning and to cool the reheat and nozzle system. The fan system functionality is
achieved at a high level of aerodynamic efficiency, at a low life-cycle cost, weight, and diameter,
and low noise level, the system must an have adequate stability margin and be able to cope with
harsh operating environments.
52
SPF / DB technology applied to commercial aeroengine front fan blades.

53
Figure 17: - High and low bypass ratio turbofan engines and their applications.
Figure 17(a):- Commercial
high bypass ratio turbofan
RR- Trent family.
Figure 17(b):- Military
turbofan - EJ200.

The system has to pass rigorous certification tests which include;- rain; hail; icing; operability; bird
strike; requirements; fan - blade – off; any distortion of inlet airflow resulting from aircraft
manoeuvring or cross – wind; altitude; and compatibility with the intake and thrust reversers; and
achievement of noise targets which are of critical importance to commercial turbofan aeroengines.
Focussing on bird strike, the fan system must be designed to cope with impact from a range of bird
sizes at various portions of the fan face; the size of the bird is a function of the intake diameter, i.e.
the larger the intake diameter, the larger the weight of bird impact that the fan must be able to
withstand. The system must be able to demonstrate integrity for all types of bird specified in the
certification requirements at all probable impact locations on the fan blades.
The major components of the commercial fan system are:- the blades: fan disc: containment
casing: and the front bearing housing structure containing the bypass vanes and engine section
stators, (in this sub - section I will be considering the fan blades). In order to reduce the fan
diameter and hence weight and drag, the inlet hub – tip ratio is minimised subject to meeting the
mechanical criteria for the hub design.
The Fan Blade:- The fan blade comprises of an aerofoil with a root attachment that secures the
blade into the fan disc. The rotor is attached to the fan shaft, which is connected to and driven by
the Low Pressure Turbine. The whole fan rotor assembly is supported by the front bearing housing.
The flow leaving the Fan Outlet Guide Vane (FOGV) ring is axial, but the flow leaving the engine
section Stators axial or swirling, depending on the engine configuration.
The hollow wide-core fan blades of current generation engines allows higher efficiency, and is
quieter than its predecessor, the snubbed blade, see figure 18.
54

55
Figure 18: - Fan Blade performance improvement and weight reduction technologies.
Figure 18(a):- Snubbered blade. Figure 18(b):- Hollow Wide-chord fan blade.
+ 4% efficiency
Snubber.
Root attachment.
Root attachment.
Aerofoil.
Aerofoil.

The snubbered blade shown in figure 18(a), consists of a solid aerofoil, which has two appendages
or snubbers attached at right angles to the aerofoil span at about three quarters of the blade height,
(these snubbers are also known as clappers). When the blade is mounted in the blade disc these
snubbers form a structural support which resists the twisting of the aerofoil when subjected to cyclic
loading caused by aerodynamic distortion and wakes. The also function as a source of damping by
raising the natural frequency of the blade.
In order to reduce blade structural weight and increase efficiency modern fan blades are
snubberless as shown in figure 18(b), but simply removing the snubbers resulted in a design that
was too flexible (its natural frequencies were too low), and also removed the mechanism for
damping any aerofoil vibration. In order to overcome this modern fan blade designs have increased
chord‟s thereby stiffening the blade and allowing a reduction in the number of aerofoils.
One of the main reasons for adoption of the wide – chord fan blade design is greater aerodynamic
efficiency over the older design. This is due to the fact that the snubbers introduced a significant
amount of aerodynamic loss, resulting in a very inefficient design, and they also, present a
blockage to the airflow, requiring the frontal area to be increased with the resulting drag and weight
increase. In order to reduce the fan module weight, these wide – chord fan blade aerofoils are
hollow, which not only reduces the weight of the individual fan blade, but also the whole system (i.e.
disc, front structure, containment casing, and bearings and housing). These hollow blades have a
cavity within the aerofoil and are formed from three sheets of titanium alloy (Ti - 6Al – 4V) (grade 5):
the two outer sheets form the aerofoil OML surface skins and the one inner sheet forms a Warren
Girder supporting core as shown in figures 19(b) / (c) / (d) which also shows the critical parameters
for modelling these structures.
56

Figure 19(b):- Warren- Girder Structure 3 sheet SPF/DB.
Figure 19(a):- Current Ti alloy SPF/DB Swept Fan Blades used on Trent family.
Figure 19(c):- Important Parameters in SPF/DB fab blade construction.
Figure 19(d):- Section of actual Fan blade and DB joint close up.
Figure 19: - Rolls Royce Wide Chord Swept Fan Blade SPF/DB manufacture.
57

These blades are produced using the processes of diffusion bonding and super – plastic forming,
which are used together for these components, but can be used individually to form other engine
components, as illustrated in figure 20. This process has the ability to manufacture large complex
structures eliminating subassemblies and fasteners, and is applied to many critical aerospace
components.
The Ti – 6Al – 4V alloy (grade 5) is considered the workhorse amongst all other grades of titanium
alloy because of its following properties:-
 Fully heat treatable for sections up to 15mm for temperatures in the order of 400ºC:
 Corrosion resistance:
 Weldability:
 Ease of fabrication.
The mechanical properties are as follows:-
 Young's modulus = 110 GPa:
 Density = 4420kg/m³:
 Tensile strength = 1000MPa:
 Poisson‟s ratio = 0.35 – 0.37
Outline of the three sheet Warren Girder core fan blade DB / SPF process:- The production
process begins with three sheets of the Ti alloy describes above, and a barrier material called Stop-
off (Yttrium boron nitride), which is screen printed on to the internal surfaces of the two sheets
which will form the aerofoils external skins to ensure that boning only occurs at the designed areas.
58

59
Figure 20: - Possible SPF Engine Components in addition to Wide Chord Fan Blade.
Dovetail seal
Inlet ring
Fan Blades and
Fan duct Outlet
guide vanes
Nacelle Panels
Inlet cone
Compressor Blades
Oil tanks
Power plant casings
Compressor Ducts Piping Components Pylon Panels
Top Core Vanes
Exhaust Nozzle Components
 Exhaust cone:
 Fairing flaps and heat shields:
 Exhaust ducts.
Drive shaft fairing (helicopter turboshaft engines)

The core sheet is then pre-drilled with gas transfer holes for the post diffusion bonding, super
plastic forming process.
Next the sheets are sandwiched together to form a 3 sheet bonding stack which is TIG welded
around the periphery, and then placed into a tool in a furnace and heated to 950ºC in order for
Diffusion Bonding by atomic diffusion to take place (see figure 21(a). This atomic self – diffusion
is a property of Ti – 6Al – 4V alloys and the kinetic occurs slowly and is obtained by a microscopic
viscoplastic deformation required to expel all porosities from the mating surfaces of the sheet stack,
and this process requires typically a constant hydrostatic pressure at temperature of 3MPa for at
least 2 hours to create the viscoplastic deformation. The procedure of diffusion transports
molecules through the cross section structure of a crystalline solid using vacancy transition and
filling, therefore the measure of contact between the surfaces is of prime importance, and this can
be improved by, mechanical machining and polishing, etching, cleaning, covering and inching the
material under high temperature and pressure.
The next stage is Super Plastic Forming of the blade OML aerofoil shape, and the supporting IML
Warren Girder core (see figure 21(b)). The ability to undergo superplasticity is one of the most
famous mechanical properties of titanium alloys, and this was the first application of Ti – 6Al – 4V
alloy to aerospace in the 1970‟s. This alloy is the perfect material for Super Plastic Forming
because of its natural mechanical ability of grain boundary sliding which occurs at temperatures in
the order of 925ºC when stretched below 10−3
s−1
. In rheology, superplasticity corresponds to the
maximum strain rate sensitivity which is approximately 0.5 for Ti – 6Al – 4V alloy. Unlike aluminium
or nickel – based alloys, Ti – 6Al – 4V does not exhibit cavitation and can be stretched to 700% of
elongation with aerospace qualified sheets. 60

61
Warren Girder Blade Core
Blade Skin
Blade Skin
Mould Tool
Mould Tool
Diffusion Bond Blade
Core to Cover Sheets
Argon gas inflation pressure Fig 21(b):- Super Plastic Forming.
Figure 21: - Overview of DB/SPF process for Wide Chord Swept Fan Blades.
OML Face Sheet
OML Face Sheet
Stop off locations
Core Sheet
TIG Seal Weld
TIG Seal Weld
Mould Tool
Mould Tool Fig 21(a):- Diffusion Bonding.
Argon gas bonding pressure

These Ti - 6Al - 4V sheets have an alpha – beta equiaxed microstructure with a typical grain size of
10 - 20 μm (micrometers). To effect the Super Plastic Forming of the Fan Blade, the Diffusion
Bonded three sheet stack mounted in the tool is heated to 950ºC and Argon gas is blown into the
cavities crated in the stack where bonding was prevented by the Stop-off applied prior to diffusion
bonding of the stack, at a typical pressure of 2MPa and over a cycle time on approximately one
hour, this gas inflates the stack and expands the OML Skin sheets into the OML Mould tool to crate
the aerofoil surface and the pressure is evenly distributed through the stack by the core gas
transfer holes. The core sheet extends under the expansion of the skin sheets to which it is
bonded, to form the Warren Girder core which will support the shape of the aerofoil skin under
operational loading. Post processing the blade is removed from the tool and the root attachment is
machined into the blade, which is subsequently cleaned to make the surface aerodynamically
smooth to ensure there is no drag producing blemishes.
The application of the SPF / DB process to complex rotating parts such as fan blades as outlined
above, or even simple static fabrications (see figure 20), comes with the mandatory requirements to
ensure an extremely high process control in order to avoid catastrophic problems. To meet these
requirements X-ray for core bonding, and ultrasonic for skin bonding, along with metallographic and
tensile of almost 100% of engine components manufactured by the SPF / DB process has enabled
it to become a common certifiable practice. FEA idealisations are being used to research for current
and future designs and the key parameters in this research are shown in figure 19(c) and are as
follows:-
1) Thickness of the OML Skin sheet T1 :
2) Thickness of the original core sheet T2 : 62

3) The thickness of the core sheet after the super plastic forming process T3 :
4) The Skew angle of the Warren Girder α :
5) Length of the hollow region bond D, The length of the leading edge bond D1, and the length of
the trailing edge bond D2 ideally (D1 = D2).
The majority of this research is looking at modeling impact effects from hail, stones, birds etc.
For the Fan Blade this SPF / DB process produces a net shape product, and has no effect on the
sheet metallurgy properties in volume but usually an oxygen rich layer appears on the surface
which must be removed by Chemical Milling. Some small machine operations can be
accommodated post processing, such as contour reworking by mechanical machining or laser
cutting.
Advantages of the SPF / DB process:-
 Applicable to complicated monolithic components:
 Reduces the number of fasteners reducing weight and assembly complexity and time:
 Eliminates geometry elimination in fabrication:
 More precise than fabrication with very good mechanical properties:
 Very good process repeatability.
Disadvantages of the current SPF /DB process:-
 High energy consumption:
 Tool can attain thermal damage:
 Expensive tooling and equipment.
63

These Titanium alloy SPF / DB hollow blades behave in a very similar manner to solid blades and
there is no determent in stiffness or bird strike resistance capability. The larger the blade the
greater the benefit from hollow blade technology as more weight can be saved. As blades reduce in
size they can no longer be hollow because the panel thickness would be too thin.
The Fan Disc:- The fan disc is one of the most critical components in the engine and has four main
functions:-
 React to the centrifugal loads from the fan blades – both during normal running and in the event
of fan blade off:
 Provide attachment from LP (Low Pressure) shaft to drive the fan and retain the fan blades:
 Absorb impact loads:
 Provide attachment for the nose cone and other peripheral components.
As a fan disc failure would endanger the aircraft this component is classified as a critical part, (CS-
25 airworthiness). The disc contains a number of slots (see figure 22(a)), into which the fan blades
are mounted and there is a front drive arm, which provides attachment to the nose cone assembly.
The disc is usually produced as a near net forging in Titanium.
The mechanical design of the disc is one of the key design areas because it is a safety critical
component, and is an extremely heavy part of the fan system. The role of the disc is to ensure that
the blades continue to travel in a circular path, and to resist their high centrifugal loads which are in
the order of 100 tons equal to 10 double decker buses hanging from each fan blade.
64
HBR Front Fan technology for commercial aeroengines.

65
Figure 22: - A Trent fan disc with blades mounted on slider assemblies.
Figure 22(a):- Fan
fixing arrangement of
typical wide – chord
commercial fan.
Fan Blade
Annulus
filler
fixings
Fan disc
Slider
assembly
Figure 22(b):- Fan disc with blades mounted and installed.

The total disc stress is a combination of inertia stresses of the disc itself and stresses imposed by
centrifugal force on the blades. Two key issues govern the amount of stress the disc is designed to
withstand:-
1) The burst criteria state that if the assembly overspeeds, the disc will not burst and compromise
the integrity of the engine, this provides the minimum cross-sectional area of the disc.
2) Disc life which sets the maximum stress in the disc. If it is unable to meet the life criteria various
strategies can be employed:-
 Increase the size of the disc, so that stress in the disc reduces to an acceptable levels. Any
extra material added is put at the bore, as the most weight efficient location:
 Decrease or eliminate any stress concentrations in the disc, such as small holes or tight radii:
 Increase the capability of the material by changing the specification:
 If the material properties exceed the life requirements, the disc can be reduced until it
reaches the minimum size and weight, as specified in the burst overspeed margin.
The Fan Case:- The primary functions of the fan case are to form the outer gas path, and contain
a fan blade should it disintegrate during flight. Therefore the casing must be capable of absorbing
the energy of a complete fan blade without releasing blade or case fragments and maintaining the
integrity of the engine. The energy of a released fan blade is equal to a family saloon car at
100kmph (60mph). The casing needs high strength and high ductility. In some engines the fan case
is part of the engine mounting system, and therefore transmits thrust from the core engine to the
aircraft.
66
HBR Front Fan technology for commercial aeroengines (continued).

It interfaces with the an Outlet Guide Vane ring in a high bypass ratio commercial engine, as shown
in figure 1. The fan case also provides mounts for the gearbox, ground support equipment and
other accessories mounted on the accessory flange. The casing assembly also contains acoustic
liners to attenuate noise generated by the fan. The panels are made of honeycomb structure of
composite construction. The fan case inner profile when fully assembled with in – fill panels, fan
track liner, acoustic panels and ice impact panel forms the outer annulus line. Containment system
weight is a function of the fan diameter cubed so high bypass ratio engines with large fan blade tip -
to - tip diameters have much heavier containment systems.
Core compressors:- The core compressor system has three main functions:
1) Raise the pressure of the air supplied to the combustor and deliver it at a suitable Mach number
with acceptable radial flow properties:
2) Supply bleed air for engine sealing anti – icing, cooling and aircraft environmental control:
3) Provide for any off-take requirements.
Like the fan system the core compressor system has to demonstrate a high level of aerodynamic
efficiency with adequate stability margin for all fan exit conditions, and at low life-cycle cost and
weight. It must also meet similar certification requirements.
A core compressor system can consist of one or two compressor modules each one driven by its
own turbine. Core compressor module pressure ratios are typically in the order of 5 to 16. The core
compressor configuration is dependent upon the engine application, and is derived from a series of
trade-off studies looking at performance, weight, cost, stability and life.
67
HBR Turbofan Component technology for commercial aeroengines.

For large commercial engines, the use of two core modules (the three-shaft layout) is usually
preferred (see figure 19), and provides for a very flexible robust and efficient system, allowing each
module to run at its optional rotational speed, it also has the benefit of minimising the number of
variable vane stages.
The major components of the core compressor module are :- the rotor drum: the casing and other
statics: Outlet Guide Vane assembly: combustor pre-diffuser: and one or more support structures.
For the purposes of this study I will consider the following only:-rotors, blades, and discs.
Rotors:- The core rotor configuration has typically consisted of 3 to 12 discs each with a set of
blades of aerofoil cross-section. The disc can be bolted or welded together to form an integral
drum, although recent developments have see the introduction of Blisk and Bling compressor discs
into commercial aeroengines, see figure 11 in which the blades are integral to the disc and this
technology was originally developed for military engines to reduce weight. Blades, Discs, and
Blisk‟s, are made from a range of materials, in modern engines forward Low Pressure (LP), and
Intermediate Pressure (IP) compressor stages are made from titanium alloys due to their high
strength to weight ratio over a range of temperatures (see figures 3(a) through (c)). The rear stage
High Pressure compressor stage uses nickel alloys because of their high strength at high
temperatures. Bling compressor discs are made from Metal Matrix Composites and there is current
research into Ceramic Matrix Composites for greater performance at higher temperatures, (see
figure 12). Conventional blades attached to the disc in commercial engines offer the advantage of
easy maintenance, damaged blades can be replaced relatively easily, although there is the penalty
of adding parasitic mass which increases the centrifugal load on the disc.
68

Compressor blades:- Blades, disks, and blisks are made from a range of materials in modern
engines, forward LP and IP compressor stages are usually made from titanium alloys due to their
high strength to weight ratio. The rear HP stages are usually Nickel alloys due to their high strength
at high temperatures and pressures. The conventional bladed disc typical of current commercial
core compressor designs, where compressor blades are normally attached to the disc using a
mechanical feature known as the root fixing,(shown in figure 23(a)i) in general, the aim is to design
a securing feature that imposes the lightest possible load on the supporting disc thus minimising
disc weight. There are two principle fixing methods in common use namely:- (1) Axial fixing and (2)
Circumferential fixing.
1) Axial fixing:- where a series of slots are machined out of the disc to accept the dovetail or fir-
tree shaped rotor blade root fixing. Axial fixings are the more complex and costly complex
option, however they are more robust for handling foreign object damage, and better facilitate
the use of variable vanes. For these reasons the front LP stages of a compressor employ axial
fixings.
2) Circumferential fixing:- (figure 23(a)i) are simpler and cheaper than axial fixing and are common
in the rear IP / HP stages of the compressor. It is relatively easy to manufacture an annular
groove at the head of the disc. Blades are assembled on to the disc through a loading slot. The
ring then being closed with a locking device.
Blade fixings have the advantage of easy maintenance where damaged blades can be removed
and replaced easily, but they also carry the penalty, that using root fixings adds parasitic mass
which increases the centrifugal load applied to the compressor disc.
69

70
Figure 23(a):- Compressor disk design development to reduce weight.
*Reference 4:- Rolls Royce Published data.
Integral blades
Figure 23(a)i:- Conventional Ti alloy
blades mounted in the IP Compressor
disc.
Figure 23(a)ii:- Ti alloy blisk with
blades integral to the IP Compressor
disc.
Figure 23(a)iii:- Ti bling with integral
blades and MMC reinforcement ring.
Cob
Rim
Circumferential fixing blades
MMC Reinforcement ring

71
The modern 5 - axes machining centres enable machining of turbine compressor blade leading
edges, trailing edges and tips in one set-up operation. The centres of STARRAG are equipped with
CNC control systems with software for NC simulation of the blade aerofoil and blade root machining
process, enabling operators to check milling operations and tools on the computer display screen
and optimise the NC program. Until recently the finish grinding of the blade aerofoil surface has not
been possible on the STARRAG machine tools and the blades have required extensive hand –
polishing to generate the required surface finish. The machine vendors have now developed other
types of machining centres, connecting in one machine the features of the milling and grinding
machines to avoid those disadvantages and to reduce the number of required machine tools, and
further details of blade machining are given in section two.
Compressor discs:- As with the fan, the mechanical design of the compressor disc is another of
the key component design tasks, as failure of a compressor disc would seriously compromise the
integrity of the engine. Additionally the disc assembly forms a significant fraction of the modules
weight. The total disc stress comprises a combination of the stresses imposed by the blades and
spacers, the inertia stresses within the disc itself, and the thermal stresses imposed by bore to rim
thermal gradients. These thermal stresses are more significant on modern engines with increased
core temperatures, and occur when the rim heats up quicker than the cob (centre thickened ring)
during acceleration, and also when from steady state running, the speed is reduced and the cob
cools more slowly than the rim. Generally the greater the size of the disc the less thermally
responsive it is and hence the higher the thermal stress levels.

72
Titanium Metal Matrix Composite
Nickel Superalloy
Figure 23(b): - Bling Compressor weight reduction using MMC technologies.
Specific
Strength.
Temperature (degrees C)
Titanium Alloy
Bling Ti MMC Compressor disc.
Bling Ti Metal Matrix Composite microsection .
Mono filaments
Metal Matrix

The mechanical design of the disc in the rear stages of the compressor is very challenging with
three major factors to be considered:-
1) High rim speed:
2) High temperatures:
3) Additional loads from the drive arm from the shaft to the turbine.
Further significant reductions in compressor rotor weight are being achieved in the Trent WXB
through the application of blisk and bling technologies (figures 23(a)ii and 23(a)iii, and figure 23(b)).
The Blisk hybrid is an integral unit with the blades and disc combined as a single component and is
used in the IP compressor stage. The Bling replaces the current bladed disc and blisk
configurations with a high strength reinforced MMC ring with blades integrated into a single
component.
Turbine stage:- The most server temperatures and pressures are encountered in the first row of
turbine blades, where the gas entry temperature is in the order of 1600ºC, although temperatures
are kept lower at the turbine blade surface as a result of the cooling system and / or thermal
coatings, detailed below, resulting in a blade surface metal temperature in the order of 950ºC. Many
requirements need to be met when designing a new turbine, the three main aerodynamic objectives
are:- (1) to produce sufficient turbine power: (2) to pass the correct amount of gas flow, and (3) to
achieve (or exceed) the target stage efficiency. Complex 3-D aerodynamic designs are used to
accurately tailor the aerodynamic shape of the turbine blade and nozzle guide vane (NGV see
FATA support Engine study) aerofoils and platforms to suit the required stage characteristics.
73

The flow characteristics of the turbine must be carefully matched with those of the compressor to
achieve efficiency and performance targets. If the turbine components allowed too low a maximum
flow then a back pressure would build up in the engine causing the compressor to surge.
Conversely too high a flow would cause the compressor to choke, where the total gas flow entering
the compressor is greater than its working capacity due to the imbalance between the two systems.
Either condition would induce a loss in engine efficiency and performance. Modern design tools and
CFD analysis have enabled current designs to meet these criterions and incorporate features to
minimise both boundary layer flow losses and also the NGV wake forcing effects on rotors. Every
effort is made to minimise the effects of consumption and reintroduction of cooling air into the gas
path.
Every design is a compromise and the design methodologies used often require a lengthy iteration
process to achieve the best possible overall solution. Such a process is required because of each
components inter-relationship with its neighbouring component, for example a modification to a
turbine blade design may require a redesign of the shroud or a change in the hub, also any change
in the blade design may lead to modification of the disc design, and such a disc iteration may then
effect the containment requirements, possibly affecting the casing design criteria etc.
A new turbine component will be reviewed by the following disciplines before engine development
and testing begin:- (1) Aerodynamic design: (2) Cooling or thermal design and analysis: (3) Stress
analysis: (4) Mechanical design: (5) Manufacturing. The component‟s operation is then fully proven
and validated before certification is received from the relevant authority and the product rolled out
to customers.
74

Turbine blades:- At operating temperatures in the order of 1600°C the HP turbine components are
in the hottest parts of the gas flow, and are designed to operate at temperatures far greater than
the melting point of leading nickel based super alloys from which they are cast.
In order to withstand these temperatures and accomplish their prime function, the High Pressure
turbine blades, NGV‟s and seal segments are cooled internally and externally using air from the exit
of the HP compressor as shown in figure 24(a), this air itself at temperatures over 700°C (obtained
from compression alone) and feed at pressures of 3,800kPa. The gas stream pressure at the
turbine inlet being 3,600kPa, therefore the cooling feed pressure margin is only small and
maintaining this margin is critical to component operation.
Considerations in determining whether a blade is to be cooled or uncooled include;- the choice of
blade materials; the application of a thermal barrier coating (TBC); the performance requirements;
and the engine cost target. Not cooling a blade or vane gives more freedom in terms aerofoil
design, both size and shape, as no internal cooling system has to be cast within it. However the
consequences of not cooling include;- limitations on the blade or vane operating temperature,
adversely affecting performance, and limiting scope for further engine growth. TBC‟s alone provide
no benefit in reducing metal temperatures on uncooled turbine components. An uncooled
component may also have to be manufactured from an improved material with impacts on
manufacturability and cost.
Advances in metallurgy and casting technology have enabled the development of single crystal
nickel alloy components see figure 24(b)(iii). The resulting improvements in material properties
allow the components to run at increased turbine operating temperatures.
75

76
HP/IP/LP Turbine Stages
IP Compressor Stage
C/L
Combustor.
HP Compressor Stage
HP Compressor Stage air is bleed-off for turbine blade cooling through
this passage following the route indicated by yellow arrows.
Steel Stators
Figure 24(a): - Turbine blade cooling air feed technology development.

77
Equiaxed
Crystal Structure
Directionally
Solidified Structure
Single Crystal
Figure 24(b): - Turbine blade materials and processing technology development.
Increasing:- Creep resistance; Performance
and; Cost of component.
(i):- Equiaxed Grain
structure Ni alloy
(ii):- Directionally Solidified
Grain structure Ni alloy
(iii):- Single Crystal Grain
structure Ni alloy

The use of these advanced alloys cast as single crystals improves the operational life limits by
enabling the most efficient use of cooling air and by giving the designer a better understanding of
the material properties. Nickel alloys are almost the universal choice for high temperature turbine
blades, (there is also research into Ceramic Matrix Composite blades) and NGV‟s, due to their high
temperature creep resistance and strength retention and three methods of processing can be
employed each have property and cost factors and the selection is a trade between cost and
performance.
 Single crystal components have superior metallurgical properties in all directions, but come at
far greater manufacturing cost (illustrated in figure 24(b)iii).
 Similar alloys can be cast utilising directional solidification this gives the micro-structure shown
in figure 24(b)ii, and offers a cheaper solution than single crystal processing but also reduced
properties especially creep resistance, compared with single crystal blades.
 Conventional processing results in an equiaxed grain as shown in figure 24(b)i, which further
reduces cost but also reduces further the blades creep resistance.
Cooling geometry design has also greatly improved, with patented laser drilling of cooling hole
designs, and soluble ceramic core technologies enabling enhanced cooling methods (as shown in
figure 24(c)) with high levels of cooling effectiveness on blades and vanes. These methods enable
the reduction of cooling airflow, as does the controlled application of ceramic Thermal Barrier
Coatings (TBC‟s), allowing higher turbine operating temperatures, resulting in increased thrust
levels.
The metallics of combustion chambers are covered in the FATA HBT engine study presentation.
78

79
Figure 24(c): - Turbine blade cooling technology development.
Single pass
Cooling air
Multi-pass
Thermal Barrier Coating,
applied to a cooled blade

80
SECTION 2:- DESIGNING PARTS FOR NC MACHINING .
X+
Z+
Y+
A
B
X+
Z+
Y+
A
B
FATA Project Port Wing Torsion Box Metallic components.
5 Axis Machining.
FATA Wing Carry Trough Box assembly.
*See references (1) , (2), (3) and (4) for all material in this section.

The metallic structural components designed for AIAA design project by myself include the wing ribs
which are to be produced as double sided machining's from Aluminium Lithium alloy by 5 axis high
speed machining the machining methods, standards, and design practices (all Catia V5 parts
shown are designed by myself), which are applied in all machined component design undertaken to
date. The following sections contain my examples of machined part design, sheet metal design,
and metallic assembly, and Machining Simulation worked examples for proficiency practice more
examples will be added.
The one of the most effective weight reduction features for the all metallic aircraft wings has been
the adoption of large scale five axis high speed machining of many structural components
previously made by the sheet metal fabrication route, and the use of ruled surfaces, and minimum
fillet radii, and if essential scalloping. This includes integrally machined wing cover skin stringers,
machined spars (with web crack stoppers), and ribs, thus enabling a reduction in fastener weight,
less scope for fatigue cracking propagating from fastener holes, reduced parts count and assembly
costs. Also joining high speed machined components can be achieved with bath tub joints or
integral end tabs without the need for separate cleats and additional fasteners. Other weight
savings have been gained from the application of titanium alloy in place of steels for highly loaded
or high temperature components produced as near net shape forgings, or even in the case of
Super Plastically Formed titanium alloy structures employed as lower wing access port panel
covers, replacing the formally sheet fabricated covers. Titanium is also more compatible than
aluminum when used with composites in that it is not susceptible to galvanic corrosion and has a
compatible coefficient of thermal expansion. Also the adoption of Aluminium Lithium alloys in such
applications as wing ribs with a density saving of 5% over conventional aluminium alloy structures.
81
Section 2:- Design of Machined metallic components for FATA studies.

1 piece machining
5 piece welded assembly
Why machine? Machining verses Fabrication
Consideration should be given to integrating smaller details into 1 piece machining to reduce
weight parts count and assembly operations as shown below.
Benefits of machining detail :- Only 1 item required to manufacture, hence inventory
reduced: No sub-assembly / welding time: Weight reduction: Better quality: Better
accuracy.
82

83
Figure 25:- Example Machined components:- Ribs, and Control Surface Hinges.
Figure 25(a):- B-787 wing rib double sided 5 axis machining
Figure 25(b):- A-350 WXB ADH Flap hinge
a double sided 5 axis machining

LR & SA Pintle Pin
84
Fig 25(c):- Commercial Wing metallic components produced by GKN Aerospace.

The designing and drawing parts is an important part of any company process. However, just
designing the part dose not make the aircraft or engine or any other product leave the assembly
line. The parts for the assemblies must be manufactured. This section will look at manufacturing
machined metallic components using Catia V5, the initial example will be three axis machining
simulation to demonstrate toolset experience.
Types of NC (Numerically Controlled) Machines:-
There are many different types of CNC machines used in metallic material machining, the Prismatic
machining toolset of Catia V5 concentrates on a few types but is open to expansion, and those
types will be highlighted here, although not all machine types will be used due to the limitations of
this presentation.
Three Axis CNC Machines:- Thee axis machines are most commonly used for simple parts. Three
axis machines come in two sub-types which are:- (1) Vertical machining centres; and (2) Horizontal
machining centres.
(1) Vertical three axis machines have the tool axis locked along the Z - axis. The X - axis generally
points the length of the table, while the Y – axis runs forward and aft on the table. Several tools
are usually carried in a carousel near the head of the machine (see figure 26).
(2) Horizontal machines work in a similar fashion. The Z – axis of a horizontal machine still runs
along the tool axis, while the Y – axis points along the machine arm, and the X – axis runs
along the table. It is very common to find an additional axis on a horizontal machine, namely the
Rotational axis which is commonly found on the table (see figure 27).
85
Prismatic Machining Methods ATDA design studies.

86
Fig 26:- NC Machines, Three Axis CNC Vertical Machine, for simulation studies.
 3 Axis Machining:-
During machining the cutter can move simultaneously
along the X,Y & Z axes. The tool axis orientation is fixed
during machining. Usually used for simple geometries
where missed material is not a major issue.
(This example shows the spiral milling of a shallow pocket
feature on a compound surface).
Tool Carousel
Spindle
Control
Y
Z
X+
Z+
Y+

87
Fig 27:- NC Machines, Three Axis CNC Horizontal Machine, for simulation studies.
Spindle
Machine Tool
Control
Tool Chain
Table
Z
X

Five axis CNC Machines:- There are three rotational axis associated with the three Cartesian axes
(X, Y, Z). The three rotational axes are A,B, and C, all respectively associated with X, Y, and Z. It is
not uncommon to find CNC machines with one, two, or even all three rotation axes. Machines with
more than one rotation axis are commonly considered multi axis machines. The most common multi
axis machine is a five axis machine that has the three X, Y, and Z directions, as well as A and B
rotational components see figure 28. This capability enables the Fanning and Tilting of the tool
during machining for complex deep pockets where excess material is an issue. Although multi axis
machines are generally more expensive to operate, and keep operational, and therefore are mostly
used by major commercial aerospace manufactures e.g. Airbus, Boeing, GKN Aerospace, and
Spirit AeroSystems, and Rolls Royce, and when weight reduction is critical as in commercial
airframes. The machining principles for the machining of flanges and landings from the FATA
RSDPD Volume 3 are shown in figures 29 and 30 respectively.
Lathes:- Horizontal and Vertical lathes are other types of cutting machine which can be programed
for in Catia V5.R20. Lathes are most generally used for making round, or round shaped, parts. This
is due to the nature of the lathe. The stock material is held in a set of grips at each end, and then
the material is spun as a tool cuts. In Catia V5.R20 Lathe operation simulations have six stages as
follows:- Stage 1;- Read V5 Product and Define Part Operation; Stage 2;- Define Lathe Operation;
Stage 3;- Define Pocketing Operation; Stage 4;- Import and Apply Drilling Process; Stage 5;-
Replay Machining Operations and Video Simulation; Stage 6;- Generate HTML Documentation and
Generate APT File. This will be covered in Section 7 of this presentation.
88
Prismatic Machining Methods ATDA design studies (continued).

 5 Axis Machining:-
During machining the cutter can move along the X, Y & Z axes and rotate around e.g. the X & Y axes
(designated A & B axes motion) during the machining cycle. This capability enables the Fanning and
Tilting of the tool during machining for complex deep pockets where excess material is an issue.
Fig 28:- NC Machines, Five Axis CNC Machine for simulation studies.
89
X+
Z+
Y+
A
B
Figure 28(b):- Tool path.
Figure 28(a):- 5-axis NC machine.

 Design for Manufacture:-
 To machine an External Flange surface produced as a
result of splitting the model with a „complex‟ surface is both
time consuming and costly.
 Therefore to aid manufacturing, the „complex‟ surface can
be replaced by a „ruled‟ surface provided the Chord Height
Error (CHE) is within the values specified in Design
Standards. (see Figure 29(a))
 Where the CHE value exceeds the specified maximum, the
flange is produced by splitting the model with a „faceted‟
surface. (see Figure 29(b)).
A bespoke „Flange‟ application will be available in the near
future to automate the creation of the „Faceted Ruled
Surface‟. As this was not available at the time of writing, the
exercise accompanying the course requires manual
generation of this geometry
 External Flanges produced by complex surfaces are
permissible, but should only be used in extreme cases and
in agreement with manufacturing due excessive machining
costs
Fig 29(a)/(b):- Machined Metallics:- Chord Height Error applied for design studies.
Figure 29(a) Figure 29(b)
CHE
Preferred Non-Preferred
90

 Design for Manufacture:-
 In Figure 30(a) the area shaded in Black indicates the 5
Axis Landing, and is the remaining material following
machining of the internal face of the closed angle
flange, and represents the difference between the „as
designed‟ and „as manufactured‟ part.
 In such cases, it is a mandatory requirement for
allowances to be made for the loss of fastener seating
area.
 The remaining material can be further reduced by
additional machining.
 The area shown in Black in Figure 30(b) represents the
preferred condition of 5 axis landings following
machining.
Figure 30(a)
Figure 30(b)
Preferred
Fig 30(a)/(b):- Machined Metallics :- 5 axis landings applied in the design studies.
91

Machining Modes:- There are two different types of machining modes which are;- (1) Axial
Modes, which include drilling, reaming, and tapping, basically making the machine behave as a
drill; (2) Milling Modes, which include pocketing, facing, and contouring motions. For each of these
mode types, specific tools are used and these will be covered in some detail below.
1) Axial Modes:-
 Drilling;- This is the most basic of the axial modes. Drilling makes the machine act as
though it were a large, automatic drill press. Drilling is used for holes that vary from very
small fastener holes through to a moderate size. If a large hole ( several cm in diameter is
desired a circular motion or pocket operation is used instead.
 Spot Drilling;- Spot drilling is usually used before a drilling operation is performed, for pilot
holes, this keeps the tool from “walking” away from the centre of the hole.
 Drilling Dwell Delay;- Drilling dwell delay will drill a hole in the same fashion as a standard
drilling operation but will delay or stop when it is inside the hole. This allows the tool time to
completely finish a hole, before retracting and starting a new one. A delay at the bottom of
the hole generally results in a smoother hole cut than a standard drilling motion.
92

 Drilling Deep Hole;- Drilling deep hole is used when a large, deep hole is required. The
tool is drilled into the material a set distance, a dwell time can be added, then the drill is
completely retracted. The drill is then re-inserted into the hole, drilled a bit further. The
process is repeated until the hole is drilled to the bottom or drilled clear through.
 Drilling Break Chips;- During a drilling break chips operation, the drill bit is drilled partially
into the material, then it is reversed and then drilled further. This allows the chips bound in
the drill bit to be removed, thus breaking away any excess chips. This keeps the drill from
overheating and keeps the chips from binding around the drill bit.
 Tapping;- Tapping is the process where threads are cut into a hole. Generally a tapping
motion is for holes that are not too excessive in size. Large holes have a different method of
creating threads. A tapped hole allows for bolts or pips to be screwed into the part.
 Reverse Threading;- Reverse threading is the same as a tapping motion, with the
exception that the threads are cut by the opposite handed cutter.
93

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