Design Procedure

1
Low Weight Rotor Blade Structural Design
Ed Smith
Jianhua Zhang Professor
Research
Associate Rotorcraft Center of Excellence
Department of Aerospace Engineering The
Pennsylvania State University June 2005

2
Background
A low weight rotor system is an important goal
for helicopters and tiltrotors, and is an
enabling technology for a cost-effective large
transport rotorcraft Primary operating cost
drivers are weight and power - Rotor
system weight blade, hub and controls -
Power low disk loading and low aircraft drag
Reduced weight and lower disk loading lead to
- Larger, lighter rotors with novel hub
and control concepts - Radically altered
dynamic characteristics Computations by
government and university personnel, and industry
provide design experience needed to critique and
guide work
3
Background
Government
Stability, loads, vibration, deflection,
performance compare with design criteria
Industry
Identify new materials and other technologies
examine feasibility and develop design guidelines
University
Determine minimum weight structural design for
specific loads define blade structural and
inertia properties
4
Background
Specific technologies explored for low weight
rotor design All graphite blades (compared to
graphite/glass hybrid blades) Possible
Flexible composites for hubs and blades
Bearingless hub (unproven to date for large rotor
system) Composite Tailored blades hubs for
stability augmentation Reduced or eliminated
leading edge weight system (aft. cg
stabilization) Active controls for primary
flight control and loads management Additional
technologies, as needed, will be evaluated and
discussed by entire team
5
Design Parameters for Composite Blade
Some design parameters which will be evaluated
Ply orientation Ply thickness number of
layers Spar location designs
Leading edge mass New materials
Active control concept
6
Composite Blade Cross Section Modeling
State of Art of Modeling Simplified composite
model Vlasov Theory Smith, Chopra
Chandra, Chopra Smith, et. al VABS
(Variational Asymptotical Beam Sectional Analysis
) Hodges, et. al 3-D shell element
For design and beam aeroelasticity analysis
Analyze and detail designs
7
Composite Blade Cross Section Modeling
Current composite cross-section model is based
on Vlasov theory Translate the
two-dimensional plate equations into
one-dimensional beam equations that are
only a function of the axial coordinate
Originally derived for isotropic sections, and
was later extended to composites -
Provide a method to determine the beam properties
from the laminate properties and the
cross-section geometry Composite modeling is
based on classical laminated plate theory, where
each individual lamina are summed together to
form properties of the laminate. The stiffness
of the individual lamina depend on their
orientation within the laminate. Blade cross
section stiffness matrix and force vector terms
are derived using Hamiltons Principle, based on
Vlasov beam equations
8
Composite Blade Cross Section Modeling
Laminated Plate Theory
? Classical laminated plate theory is a method to
determine the properties of a laminate composed
of individual lamina stacked together. The
stiffness coefficients are a function of the ply
angle ?p and the stiffness properties of the ply
itself
? The properties of a composite laminate are
calculated by integrating through the thickness
of the plate. The Classical relationship between
the stress resultants and the linear laminate
strains is given by
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Composite Blade Cross Section Modeling Vlasov
Theory for Closed Sections
Using the classical laminated plate theory as a
basis, Vlasov theory is used to relate beam
displacements and rotations to the beam forces to
find the blade stiffnesses
The plate forces are related to the blade
forces through the principle of virtual work
The plate strains ( ) and first order
curvatures ( ) in the above equation can be
expressed in terms of the blade displacements and
rotation through geometric considerations
The generalized blade force to generalized
blade displacements relation is derived as
Axial force
(Extension stiffness)
Lag bending moment
(Flap bending stiffness)
Flap bending moment
(Lag bending stiffness)
Torque
(Torsion stiffness)
Bimoment
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Design Procedure
Identify design variables selecting
materials, skin and spar thickness (No. of
plies), web location, ply orientation at 4 radial
locations (r/R 0.25, 0.50, 0.75, 1.00)
Thickness of the skin and spar are changed by
increasing or reducing number of plies. The ply
angle for the spar starts at 0? and the ply angle
for the skin starts at 45?. The ply angles of
skin and spar will be varied to meet the
stiffness and strength requirements The design
process will continue until the blade cross
section inertia and stiffness properties are
within the targeted range, and the stresses or
strains satisfy the failure criteria
Non-structural mass is assumed to be zero it
will be added in the comprehensive analysis of
the rotor system
1.0
0.75
0.5
0.25
r/R
11
Design Procedure
12
Blade Cross Section Loads
? The blade loads supplied by NASA are based on
the speed sweep (up to the total rotor power
15000 hp) and the load factor sweep (up to
1.54g). Blade Loads at different flight
conditions will be evaluated and the worse case
will be used in the blade design ? All six
section load components flap bending moment, lag
bending moment, torsion moment, chordwise force,
normal force and axial force are given at
designated radial locations in the four forms
max., min., average,1/2 Peak-Peak ? Based on
these loads provided, the worst loading
conditions are sorted out by assuming all six
components reach the largest at the same time ?
In the current stress or strain analysis, the
normal shear loads and chordwise shear loads are
neglected
13
Blade Cross Section Stress Analysis
Stress/Strain calculation is based on the worst
loading cases and are calculated at the middle of
each layer of each segment along the cross
section ? The stiffness matrix is derived from
the composite blade modeling, and the blade cross
section loads are from the rotor comprehensive
analysis. Then, the displacements can be found
using a linear solver ? Once the blade
displacements are known, the blade strains and
curvatures can be found by geometric
consideration ? Finally, using laminated plate
theory, the stress distribution across the blade
cross section can be obtained
14
Design Criteria - Strength
Macromechanical failure theory ?
Maximum stress ? Maximum strain ?
Tsai-Hill (Deviatoric energy theory) ? Tsai-Wu
(Interactive tensor theory)
Tsai-Wu strength failure criterion is applied
in the strength analysis
normal stress in longitudinal direction
normal stress in transverse direction
shear stress
X longitudinal tensile strength Y
transverse tensile strength X longitudinal
compressive strength Y transverse compressive
strength
Limitations Does Not Address Laminate Failure
Modes - Delamination - Damage Tolerance
(Holes, Notches, etc.)
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Design Criteria Laminate Strain Allowables
Industrial Design Practice Typically, an
allowable of 3000 microstrain is a laminate
allowable associated with a particular lay-up
(usually quasi-isotropic) with all kinds of knock
downs. The lamina level strength values are not
typically referred to as allowable and not used
in design Carbon fiber design strains for
aircraft structure are typically in the range of
3000-4500 microstrain range because that has been
found to provide a realistic conservative design
allowable for a damaged structural laminate under
cyclic loading Current allowables for IM7/8552
are on the order of 4500 microstrain
(compression) and 6000 microstrain (tension).
However, the factor of safety (or some may say
ignorance) reduces these to less than 3000
microstrain in design
Industry design practice of 3000 microstrain
allowable will be adopted for the current blade
structural design
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Blade Structural Design

• A composite rotor blade modeling program has
  been developed and adapted for blade cross
  section design. Detailed cross section
  stress/strain analysis, have been formulated and
  applied in the design process. Design criteria
  have been established
• The baseline blade properties were from the
  scaled XV-15 blade, from which the initial blade
  loads were calculated based on the loads and
  other design requirements, such as stiffness,
  C.G. location, etc., a new blade design will be
  conducted and the blade properties will be fed
  back for comprehensive analysis until the design
  process converges
• Ten design iterations of LCTR blade have been
  accomplished the newly designed blade properties
  satisfy the requirements from rotor comprehensive
  analysis, and the overall weight reaches the
  targeted 50 reduction
• Two iterative design process for LABC and one
  for LCTC have also been completed

17
Composite Materials Properties
For the preliminary studies, AS4/3501 composite
materials were used then a new advanced
composite material was postulated with high
modulus and high strength (1.67 times higher than
AS4/3501-6) to study the influence of material
properties on low weight blade design IM7/8552
is chosen in the final design because of its
higher modulus and higher allowables
AS4/3501-6 AS4/3501-6 (with factor of 1.67) IM7/8552 Unit
20.6 34.4 23.8 msi
1.49 2.49 1.7 msi
1.04 1.74 0.754 msi
0.27 0.27 0.32
Density 1.7718?10-3 1.7718 ?10-3 1.7718 ?10-3 Slugs/cubic inch
Strength
Xt 331 553 395 ksi
Xc 208.9 349 245 ksi
Yt 8.3 16.6 16.1 ksi
Yc 33.1 99.3 21.8 ksi
S 10.3 17.2 17.4 ksi
Torsion stiffness Requirement is hard to meet
Almost doubled compared to AS4
Generic IM6/Expoxy UD prepreg Sources 1.
Engineering Mechanics of Composite MaterialsIsaac
M. Daniel and Ori Ishai, Oxford University Press,
1994. 2. Hexply 8552 from Hexcel
Composites 3. Industries and
Government
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Blade Cross Section Design
LCTR
0.75 r/R
r/R 0.25 0.5 0.75
1.0
t/c 0.20 0.18 0.12
0.08
Blade with different thickness ratios and taper
ratios were investigated for their effects on the
blade weight. Tip to root taper ratio of 0.8 was
chosen for LCTR The blade was designed at four
radial locations (r/R0.25, 0.5, 0.75, 1.0)
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Blade Cross Section Design
LCTR
Stiffness requirements
1.0
Blade Loads
0.75
Design drivers
0.5
0.25
r/R
r/R 0.25 0.5 0.75 1.0
Design loads loads level60 loads level60 loads turn80sls rotor2 loads turn80sls rotor2
Design drivers are different at each cross
section. In the current design, the design
drivers are either load limits or the stiffness
requirements The worst loading condition is
chosen for each cross section by evaluating the
max. loads among different flight conditions
20
Blade Cross Section Design
LCTR Cross section (the 10th iteration)
Radial station (r/R) Chord (c) (ft) Thickness ratio (t/c) Skin (trailing edge) SkinSpar Web Web location (c) Ply thickness (in)
0.25 3.5 0.20 (45)10 (45)33/(0)60/(45)2 (45)35 40 0.005
0.50 3.3 0.18 (45)10 (45)33/(0)30/(45)2 (45)35 40 0.005
0.75 3.1 0.12 (45)5 (45)28/(0)20/(45)2 (45)30 40 0.005
1.00 2.9 0.08 (45)5 (45)28/(0)10/(45)2 (45)30 40 0.005
Uniform skin lay-ups for 0-50R and 50-100R
for trailing edge Slight skin taper for
leading edge Moderate spar taper
21
Cross Section Strain Analysis
Normal Strain (Microstrain)
r/R0.50 mid span
r/R0.25 root
Min (-1800)
Max (2200)
Min (-1300)
Max (2800)
r/R1.0 tip
r/R0.75
Max (2900)
Max (2900)
Min (-2000)
Min (-2300)
The normal strains across the section are all
within 3000 micro strain
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Cross Section Stress Analysis
Normal stress (Ib/sq.foot)
r/R0.50 mid span
r/R0.25 root
Max (6.9E6)
Max (9.2E6)
Min (-4.2E6)
Min (-5.9E6)
r/R1.0 tip
r/R0.75
Max (9.2E6)
Max (8.7E6)
Min (-6.7E6)
Min (-4.9E6)
Tapered airfoil thickness
Stresses shown are calculated based on the
loading condition when the blade cross section
loads axial force, flap and lag bending moments,
and torsion moment are all the largest
23
Blade Cross Section Properties
Moderate taper of spar and tapering of blade
chord and thickness save blade weight, but the
stiffnesses also drop quickly, especially the
blade flap stiffness The blade has thick torque
box to meet the requirement of high torsion
stiffness
24
Sensitivity Studies Blade Load Reduction
Recent studies show that blade loads can be
significantly reduced by the active blade
management, such as the concept of dual active
trailing edge flaps In the current sensitivity
study, the flap bending moment is reduced by 50
Deformed blade w/o control
Opposite action of dual flap
lift due to outboard flap
Opposite lift due to inboard flap
Straightened blade

• Dual flap concepts
• - Generate additional moments
• Results in reducing blade load
• Reduce blade stresses and increase blade life
• - Effect to trim by dual flap could be minimized
  (net lift is nearly zero)
• Control inputs include 1/rev and higher
  harmonic components

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Sensitivity Studies Blade Load Reduction
r/R 0.25 0.50 0.75 1.0 Overall blade weight reductions
Weight reduction 21 15 13 12 18
Reducing the flap bending moment by 50 can save
18 of weight. It is most effective at the blade
root, where the flap bending moment is the largest
26
Sensitivity Studies - Materials
The new advanced composite materials with high
modulus and high allowables are expected in the
future, therefore some sensitivity studies of
materials are essential for the success of low
weight rotor design Increase the ultimate
strengths by 50 via improved materials and
better damage design and detection Increase
elastic modulus in fiber direction by 25 via
using stiffer fibers
27
Sensitivity Studies - Materials
r/R 0.15 0.55 0.75 0.98 Overall reduction
50 increase of ultimate strength 23 14 12 10 20
25 increase of El (longitudinal) 0 16 18 13 8
Increasing the ultimate strengths by 50
achieves about 20 weight reduction It is most
effective at the blade root, where the flap
bending moment is the largest, and also the
driving factor for blade section design
Increase the material elastic modulus by 25 is
not effective at those blade section, where
driving factor is the loads (0.15 R), but they
are as effective at those blade sections, where
the stiffness is the driving factor (0.50, 0.75
and 1.0 R for example). The Overall weight
reduction is about 8
28
Blade Design - Off-axis Plies in Spar
The structural design iterations on the blade
have included spars with 0 plies and skins with
45 plies. According to industries comments,
there needs to be some off-axis plies in the
spar, in order for it to provide torsional
resistance to the applied loads The current
LCTR blade design requires high torsion
stiffness, the off-axis plies in the spar will be
investigated for their influences on the blade
stiffness, especially how they contribute to
torsion stiffness The spar with 50 of
off-axis plies will be investigated for three
cases with different ply angles 10, 30
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Blade Design - Off-axis Plies in Spar
Flap Stiffness
Torsion Stiffness
Lag Stiffness
Initial design
50 of 10 plies in spar
50 of 30 plies in spar
The largest torsion stiffness increase is
achieved at root section about 15 for the case
of 50 of 30 plies in spar The flap and lag
stiffness decrease according (28 and 11)
30
Blade Design - CG Placement via Tailoring Spar
Tailoring spar topology to move CG forward such
that less dead weight needed for stability Add
more 45 plies at leading edge such that the CG
location can be pushed forward and the blade
torsion stiffness can be increased
31
Blade Design - CG Placement via Tailoring Spar
mass
Torsion stiffness
By adding more 45 plies at leading edge, CG
locations at all four cross section are ahead of
quarter of chord Torsion stiffness has been
significantly increased by 35 while the
sectional mass has also been increase by 35
32
Summary Blade Design
A composite rotor blade modeling program has
been developed and adapted for blade cross
section design. The program is capable of
calculating blade inertia and stiffness
properties, all offsets, and the stress and
strain distribution across the blade cross
section based on the blade loads provided
After discussion with industries and government,
the blade design criteria have been established.
The Industry design practice of 3000 microstrain
allowable is used for the current blade
structural design studies The baseline blade
properties were from the scaled XV-15 blade, from
which the initial blade loads were calculated
based on the loads and other design requirements,
such as stiffness, C.G. location, etc., a new
blade design will be conducted and the blade
properties will be fed back for comprehensive
analysis until the design process converges
33
Summary Blade Design

• For preliminary design, AS4/3501 composite
  material was used IM7 was used in final design
  for its high modulus and high strength. In order
  to investigate the material in the future
• Blade with different thickness ratios and taper
  ratios were investigated for their effects on the
  blade weight. Tip to root taper ratio of 0.8 was
  chosen for LCTR
• Ten design iterations of LCTR blade have been
  accomplished the newly designed blade properties
  satisfy the requirements from rotor comprehensive
  analysis, and the overall weight reaches the
  targeted 50 reduction
• Two iterative design process for LABC and one
  for LCTC have also been completed

34
Summary Sensitivity Studies
Sensitivity studies of material properties on
blade weight Increasing the ultimate
strengths by 50 achieves about 20 weight
reduction It is most effective at the blade
root, where the flap bending moment is the
largest, and also the driving load for blade
section design Increase the material
elastic modulus by 25 is not effective at those
blade section, where driving factor is the loads
(0.15 R), but they are as effective at those
blade sections, where the stiffness is the
driving factor (0.55, 0.75 and 0.98 R for
example). The Overall weight reduction is about
8 Study of active loads control (active
trailing edge flaps) on blade weight
Reducing the flap bending moment by 50 can save
18 of weight. It is most effective at the blade
root, where the flap bending moment is the
largest.
35
Summary Design Issues
OffAxis Plies in Spar Using off-axis plies in
the spar can increase the torsion stiffness and
decrease the flap and lag stiffness. This method
can be used for tailoring blade frequencies
C.G. Placement by Tailoring Spar CG placement
by adding more 45 plies at leading edge can
push the CG ahead of quarter of chord and also
increase the torsion stiffness, however, the
sectional mass also increases. Therefore, it may
help for CG placement, but it may not change the
torsion frequencies
36
Summary Research Topics
Materials The new advanced composite materials
with high modulus and high allowables are
essential for the success of low weight rotor
design. It is a challenging task to predict the
composite material development in the next 15
years The current blade design uses all
Graphite material. As an opposite, all glass
blades could be designed to explore the design
boundary and a hybrid blade (Gr. glass) could
be investigated for the tradeoff and benefits of
using multiple composite materials for blade
design Blade design Details The leading edge
protection cap, the trailing edge block, the
blade dead weight, as well as anti-icing blanket
should be taken into consideration for design in
order to more accurately estimate the blade
weight
37
Summary Research Topics
Modeling Enhancement Validation An automated
blade cross section optimization programming will
be developed. It is expected that by using this
optimization program, the topology of the spar
and skin can be further optimized to reduce the
blade weight and the design process can be more
efficient Validation studies with other more
sophisticated composite blade modeling algorithms
such as VABS (In progress) All researches will
be documented in the final report to NASA as well
as the following three conferences AHS
October 2005 Rotorcraft Structures and
Survivability Specialist Meeting AHS November
2005 2nd International Basic Research Conference
AHS January 2006 Design Conference