Transcript Design Procedure

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
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
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
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
Design Parameters for Composite Blade
Spar
Web
Skin
Foam
Composite Blade Cross Section
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
Composite Blade Cross Section Modeling
Spar
Web
Skin
Foam
Composite Blade Cross Section
State of Art of Modeling
• Simplified composite model
• Vlasov Theory
Smith, Chopra
Chandra, Chopra; Smith, et. al
For design and beam
aeroelasticity analysis
• VABS (Variational Asymptotical Beam Sectional Analysis )
Hodges, et. al
• 3-D shell element
Analyze and
detail designs
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 Hamilton’s Principle, based on Vlasov beam equations
Composite Blade Cross Section Modeling
– Laminated Plate Theory
y
t
x
p
T
L
n General Plate
Segment
s
z
 zs , N zs
 z , Nz
 z ,M z
 zs , M zs
 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 θp and the stiffness properties of the ply itself
 L   Q11 Q12

 
 T   Q21 Q22
   0
0
 LT  
0   L 
 
0   T 

Q66   LT 
 x   Q11 Q12
  
 y   Q21 Q22
  Q Q
26
 xy   16
Q16    x 
 
Q26    y 
Q66   xy 
 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
 Nz 
z 
 N  A B  
  zs 
 zs  


 
  
M
B
D
 z 
 z 
 M zs 
 zs 
A ,B ,D   
ij
ij
ij
tp
2
tp

2
Qij (1, z , z 2 )dz
N
Plate stress resultants

Plate strain
M
Plate moment resultants

Plate bending curvature
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

1
( N z z  N zs zs  M z z  M zs zs )ds
2 s
My
• 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
 N   k11 k 12
 M  k
k22
 x   21
 M y    k31 k 32
M  k
k42
    41
 Ts   k51 k52
k13
k14
k23
k24
k33
k34
k43
k44
k53
k54
k15   w' 
 
k25   y' 
 
k35  x' 

k45  z'' 
 

k55  z' 
y
z
M
N : Axial force
Mx
: Lag bending moment
M y:
Ts
Flap bending moment
: Torque
M :
Bimoment
Ts
N
x M x
k11  EA
(Extension stiffness)
k22  EI y
(Flap bending stiffness)
k33  EI x
(Lag bending stiffness)
k55  GJ
(Torsion stiffness)
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
Spar
Web
Skin
Foam
1.0
0.75
0.5
0.25
r/R
Composite Blade Cross Section
Design Procedure
Materials; airfoil; cross section configuration; sectional loads
Design variables: skin, spar thickness,
ply angles, web location
Adding or reducing # of plies of skin and spar
and changing ply angles
Analysis of cross section stiffness and stress
via thin-walled composite beam theory
Targeted blade cross
section properties:
Inertia, stiffness, etc
cg offset and amount of nonstructural mass needed
Satisfied ?
Strength or Strain criteria
no
yes
Output blade cross section inertia, stiffness, offsets,
etc. for the next design iteration
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
Blade section Load
● In the current stress or strain analysis, the normal shear loads and chordwise shear loads
are neglected
Max.
1/2 Peak-Peak
Average
o
Azimuth
Min.
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
Composite Blade Cross
Section Modeling
K d    f 
Blade Loads
Blade Displacement
Geometric Consideration
Plate Strain & Curvatures
Laminated Plate Theory
Stress Distribution Across
Blade Cross Section
Design Criteria - Strength
• Macromechanical failure theory
▪ Maximum stress
▪ Tsai-Hill (Deviatoric energy theory)
▪ Maximum strain
▪ Tsai-Wu (Interactive tensor theory)
• Tsai-Wu strength failure criterion is applied in the
strength analysis
F1112  F22 22  2F121 2  F66 62  F11  F2 2  1
1
2
6
normal stress in longitudinal direction
normal stress in transverse direction
shear stress
X longitudinal tensile strength
X’ longitudinal compressive strength
1
1
, F12 
, F12  0.5 F11F22 ,
XX 
YY 
1
1
1
1 1
F66  2 , F1   , F2   ,
S
X X
Y Y
F11 
Y transverse tensile strength
Y’ transverse compressive strength
Limitations:
Does Not Address Laminate Failure Modes
- Delamination
- Damage Tolerance (Holes, Notches, etc.)
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
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
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
1.771810-3
1.7718 10-3
1.7718 10-3
Slugs/cubic inch
Xt
331
553
395
ksi
Xc
208.9
349
245
ksi
Yt
8.3
16.6
16.1
ksi compared to AS4
Yc
33.1
99.3
21.8*
ksi
S
10.3
17.2
17.4
ksi
EL
ET
G LT
 LT
Density
Torsion
stiffness
Requirement is
hard to meet
Strength
Almost doubled
* 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
Blade Cross Section Design
LCTR
croot
ctip
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/R=0.25, 0.5, 0.75, 1.0)
Blade Cross Section Design
LCTR
Stiffness
requirements
1.0
Blade
Loads
0.75
Design drivers
0.5
r/R
0.25
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
Blade Cross Section Design
LCTR Cross section (the 10th iteration)
Radial
station
(r/R)
Chord
(c) (ft)
Thickness
ratio (t/c)
Skin
(trailing
edge)
0.25
3.5
0.20
[(±45)10]
[(±45)33/(0)60/(±45)2] [(±45)35]
40
0.50
3.3
0.18
[(±45)10]
[(±45)33/(0)30/(±45)2] [(±45)35]
40
0.75
3.1
0.12
[(±45)5]
[(±45)28/(0)20/(±45)2] [(±45)30]
40
1.00
2.9
0.08
[(±45)5]
[(±45)28/(0)10/(±45)2] [(±45)30]
40
Skin+Spar
Web
Web
location
(%c)
Ply
thickness
(in)
0.005
• Uniform skin lay-ups for 0-50%R and 50-100%R for trailing edge
• Slight skin taper for leading edge
• Moderate spar taper
Cross Section Strain Analysis
Normal Strain (Microstrain)
1
r/R=0.25 root
Min (-1800)
Max (2800)
0.5
2.9E-03
2.5E-03
2.0E-03
1.5E-03
1.0E-03
5.0E-04
0.0E+00
-5.0E-04
-1.0E-03
-1.5E-03
-1.8E-03
0
0
1
0.5
1
2
Max (2900)
Min (-2300)
0.5
2.2E-03
2.0E-03
1.5E-03
1.0E-03
5.0E-04
0.0E+00
-5.0E-04
-1.0E-03
-1.4E-03
r/R=0.50 mid span
Min (-1300)
Max (2200)
0
0
3
2.9E-03
2.5E-03
2.0E-03
1.5E-03
1.0E-03
5.0E-04
0.0E+00
-5.0E-04
-1.0E-03
-1.5E-03
-2.0E-03
r/R=0.75
1
1
0.5
1
2
3.0E-03
2.5E-03
2.0E-03
1.5E-03
1.0E-03
5.0E-04
0.0E+00
-5.0E-04
-1.0E-03
-1.5E-03
r/R=1.0 tip
Max (2900)
3
Min (-2000)
0
0
0
1
2
3
0
1
2
The normal strains across the section are all within 3000 micro strain
3
Cross Section Stress Analysis
Normal stress (Ib/sq.foot)
1
0.5
r/R=0.25 root
Max (9.2E+6)
Min (-5.9E+6)
0
0
1
2
8.0E+06
7.0E+06
6.0E+06
5.0E+06
4.0E+06
3.0E+06
2.0E+06
1.0E+06
0.0E+00
-1.0E+06
-2.0E+06
-3.0E+06
-4.0E+06
-5.0E+06
1
0.5
Max (6.9E+6)
0
1
8.0E+06
6.0E+06
4.0E+06
2.0E+06
0.0E+00
-2.0E+06
-4.0E+06
-6.0E+06
r/R=0.75
0.5 Max (9.2E+6)
Min (-6.7E+6)
0
Min (-4.2E+6)
0
3
1
6.0E+06
5.0E+06
4.0E+06
3.0E+06
2.0E+06
1.0E+06
0.0E+00
-1.0E+06
-2.0E+06
-3.0E+06
-4.0E+06
r/R=0.50 mid span
0.5
1
2
r/R=1.0 tip
3
8.0E+06
6.0E+06
4.0E+06
2.0E+06
0.0E+00
-2.0E+06
-4.0E+06
Max (8.7E+6)
Min (-4.9E+6)
0
0
1
2
3
0
1
2
3
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
Blade Cross Section Properties
2.0E+08
EI lag (Ib-ft^2)
EI flap (Ib-ft^2)
2.0E+07
1.5E+07
1.0E+07
5.0E+06
1.5E+08
1.0E+08
5.0E+07
0.0E+00
0.0E+00
0.25
0.5
0.75
0.25
1
Section Mass (slug/ft)
GJ (Ib-ft^2)
3.0E+07
2.0E+07
1.0E+07
0.0E+00
0.5
0.75
Radial station
0.75
1
Radial station
Radial station
0.25
0.5
1
1
0.8
0.6
0.4
0.2
0
0.25
0.5
0.75
1
Radial station
• 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
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
Section Mass (slug/ft)
Sensitivity Studies – Blade Load Reduction
1
0.8
0.6
0.4
0.2
0
0.25
Baseline
0.5
0.75
Radial station
1
50% flap m om ent 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
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
Section Mass (slug/ft)
Sensitivity Studies - Materials
1
0.8
0.6
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%
0.4
0.2
0
0.25
0.5
0.75
Radial station
1
Baseline
50% increase of Ultim ate Strength
25% increase of m odulus (longitudinal)
• 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%
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°
Blade Design - Off-axis Plies in Spar
Flap Stiffness
Torsion Stiffness
2.0E+07
2.0E+07
1.0E+07
0.0E+00
2.0E+08
EI lag (Ib-ft^2)
EI flap (Ib-ft^2)
3.0E+07
GJ (Ib-ft^2)
Lag Stiffness
1.5E+07
1.0E+07
5.0E+06
0.5
0.75
1
Radial station
Initial design
1.0E+08
5.0E+07
0.0E+00
0.0E+00
0.25
1.5E+08
0.25
0.5
0.75
Radial station
50% of ±10° plies in spar
1
0.25
0.5
0.75
1
Radial station
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%)
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
Blade Design - CG Placement via Tailoring Spar
mass
Section Mass (slug/ft)
Torsion stiffness
GJ (Ib-ft^2)
3.0E+07
2.0E+07
1.0E+07
0.0E+00
0.25
0.5
0.75
1
0.8
0.6
0.4
0.2
0
1
0.25
Radial station
0.35
0.5
0.75
Radial station
1
C.G. Location
Initial Design
0.3
Quarter of Chord
0.25
After CG placement design
0.2
0.25
0.5
0.75
1
• 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%
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
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
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.
Summary – Design Issues
• Off–Axis 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
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
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