University of Texas VSP Structural Analysis Module Update...
Transcript of University of Texas VSP Structural Analysis Module Update...
© 2013 Armand J. Chaput
University of Texas VSP Structural
Analysis Module Update - Overview
2nd VSP Workshop, San Luis Obispo, CA
Armand J. Chaput, Principal Investigator
Hersh Amin, Undergraduate Research Assistant
Department of Aerospace Engineering and
Engineering Mechanics, University of Texas at Austin
7 August 2013
Dr. Armand Chaput, Department of Aerospace Engineering and
Engineering Mechanics (ASE/EM), University of Texas at Austin (UT)
• Director, Air System Laboratory (2008 - present)
• Senior Lecturer, Air System Engineering Design (UAS focus)
• Lockheed Martin – 30 years, advanced product development
• Air System Design and Integration - Senior Technical Fellow
including assignment as F-35 Joint Strike Fighter Weight “Czar”
• Unmanned Combat Air Vehicles - Integrated product team lead
• National Aerospace Plane (NASP) - National team Chief Engineer
• Advanced Design Department – Manager
Undergraduate Research Assistant Team
Hersh Amin, Research Ass’t
Josh Eboh, Team Lead
Natalie Maka , Research Ass’t
Patil Tabanian, Research Intern
Self Introductions
Based on work performed under NASA/NIA Task Order 6322-UTEX,
“Advanced Conceptual Design Tools and Development”
Armand J. Chaput 2013
Sarah Brown, Research Ass’t
Jose Galvan, Research Ass’t
Alex Haecker, Research Ass’t
Tejas Kulkarni, Team Lead
Current Previous
© 2012 Armand J. Chaput
(1) Expand VSP user capabilities for employing higher
order, physics based tools and methods during
conceptual design (CD)
(2) Integrate VSP FEA structures module with an open
source finite element method (FEM) structural analysis
program in a user friendly interactive environment
- Currently focused on CalculiX (available under terms of GNU
General Public License as published by the Free Software Foundation)
(3) Develop basic capabilities for open source application
of FEM-based mass property (MP) methods to CD
- Current effort develops and validates fundamental wing
and tail mass estimation methodologies
- Follow-on effort (?) will expand applications
VSP Structural Analysis Module (SAM) R&D Objectives
© 2013 Armand J. Chaput
2012 VSP Workshop: Version 0 (VSP structures module
integrated with CalculiX, posted Sept 2012)
- UT Java scripts simplify setup and run unitary VSP FEM model
- GUI inputs (loads, constraints, material prop, trim to wing box)
- CalculiX solution and display of stress, strain, displacement
- Calculate mass of input FEM model
- Initial stress/mass results (convergence stability issues)
May 2013: Version 1 (fully stressed coarse grid FEM & mass)
- Separate upper and lower skin, spar, rib FEMs connected by
rigid body nodes resolved convergence issues
- Adds: skin section trim (1.0) and solution status feedback (1.1)
Today: Version 2 (inertia loads and sizing for multiple load cases)
- Angle of attack plus fuel and discrete mass inertias
- Convergence for multiple load cases
- Initial calibration/validation results
Objective 1 - Expanded Capabilities
Codes and users guide posted
at: http://vspsam.ae.utexas.edu/
© 2013 Armand J. Chaput
Objective 2 - CalculiX Integration
UT Input Executable (Java)
Boundary Conditions and Load Cases
CalculiX Input File
Vehicle Sketch Pad
External and Internal Mesh
Generation
Parametric
External Geometry
Parametric
Internal Geometry
CalculiX
FEM Solution
FEM Input
FEM Post Process
and Graphics
Output Files
UT Convergence Executable
(Java)
Solution Files
Thickness Iteration
Stress Convergence
Thickness and Material Properties
Mass Calculation
Wing Trim
© 2013 Armand J. Chaput
Last iteration –
thickness “converged”
yellow = 28.9 ksi
2 spars,
5 ribs
Last iteration –
thickness “converged”
yellow = 28.7 ksi
2 spars,
7 ribs
Last iteration –
thickness “converged”
yellow = 28.4 ksi
5 spars,
15 ribs
Objective 3 – FEM mass for fully stressed trade study
wings with minimum gage constraints (from last year)
• Lightly loaded notional wing
• VSP defined spars and ribs
• VSP SAM defined thickness, materials
and loads (2D running load along 0.25c)
• 30 ksi fully-stressed design objective
• FEM mass calculated
• Time to generate-solve-converge for all 3
solutions from scratch < 3 hr
• Issue – solution stability
© 2012 Armand J. Chaput
So What’s New?
• FEM structural methods have been available for decades
- FEM analysis requires a well-defined representation of the
airframe structure; design details are not available during CD
- CD design and analysis cycles are typically incompatible with
time required for FEM model development and turn around
- By the time a FEM model is developed the CD team has
usually moved on to another concept
- CD budgets are often incompatible with specialized FEM
analyst staffing requirements
• VSP SAM lets structural design and analysis keep up with
other CD participants
- Traditional FEM model definition, solution and analysis time
and skill requirements limit wide scale application
VSP SAM enables requirement-based CD mass estimation
© 2013 Armand J. Chaput
Background – Airframe Mass Property (MP) Estimation
Airframe mass is driven by multiple requirements; many of
which are not captured by traditional CD analysis methods
- Current state-of-the-art MP methods still rely on parametric
(statistical or regression analysis of historical data) methods
- A problem when trying to predict mass for new vehicles, new
materials, new processes or new design requirements
Primary loads drive 60% of load carrying airframe mass
- Calibrated FEM analyses should be able to predict primary
structural mass with better accuracy than parametrics
Secondary structure mass is driven by non-primary loads
- Many of which could be captured by FEM-based methods
System installation and integration effects are problematic
- Not defined until much later in the design process
Bottom line: FEM-based methods can improve the quality
of at least CD and PD primary structure mass estimates
- It doesn’t cost any more or take any more time
© 2013 Armand J. Chaput
FS 228.2
FS 283.9
0.70 c?
182.6
0.83 c
0.15 c
318”
117
”
MP Data from Grumman Aerospace
A-6E Weight Report
WT-128R-1S37 Aug 1988
Courtesy of Paul Kachurak, NAVAIR
Why we need improved CD methods - Example from UT method development effort (A-6E)
1. Raymer (fighter-attack):
= 0.0103 [(Wdgnzdu)0.5 Sref
0.622 AR0.785
(1+)0.05 Scsw0.04]] / [(t/c)0.4
Cos(0.25c)]
= 4092 lbm
2. Nicolai (USN fighter):
= 19.29 [(Wtonzdu)/(t/c)] {[(Tan le -
2(1-) / AR(1+)] 2 +1] 10-6} 0.464
[(AR(1+ )] 0.70Sref0.58 = 7057 lbm
A-6E WING GROUP (lbm) Sum
WING STRUCTURE - BASIC 3443 3443
SECONDARY STRUCTURE 931 4374
TRAILING EDGE DEVICES 593 4966
LEADING EDGE DEVICES 241 5207
SPEED BRAKES 145 5352
WING GROUP - TOTAL 5352
Inc. wing fold unique 297
Typical CD wing parametric estimates:
Wdg = 36526 lbm
Wto (land) = 60705 lbm
© 2013 Armand J. Chaput
Perspective – Why airframes weigh what they weigh
For a good design, the driver is structural requirements
- Operating environment (speed, altitude and temperature)
- Almost always known and available up front
- Failure modes, Durability and Damage Tolerance (DaDT)
- Loads inc. primary air loads, secondary loads and accidents
- All are quantifiable but often missed in CD (inexperience)
- Systems integration (loads, penetration, installation access)
- Predictable but only when design teams are integrated
- In-flight moving parts (control surfaces, doors, gaps & locks)
- Ground handling and maintenance access
- Manufacturing and assembly (including workforce skill level)
- LCC cost, schedule, risk and growth considerations
During early phases, many designers use rule of thumb or
program defined knock-downs to cover unknowns
- Generally expressed in terms of % design stress (or strain)
- Similar to our “Conceptual Design Nominal Stress (CDN )”
much k
now
n
much u
nkno
wn
© 2013 Armand J. Chaput
FEM Approach to Nominal Stress (CDN) for CD
Step 1 - Develop CD-Level FEM Models of Existing Designs
- Capture representative geometry, material and primary loads
- Focus on primary structure: Spars, Ribs and Skins,
Step 2 - Back out CDN to correlate calculated FEM mass with
actual mass consistent with min gage
1. Repeat Process for Multiple Vehicles in Given Class
2. Use Correlated CDN for CD and Early PD Designs
B747 Simplified Rib Model before & after Trim
Iter
Spar Mass (lbm)
Skin Mass (lbm)
Rib Mass (lbm)
Total Mass (lbm)
1 13392 26457 6249 46100
2 10506 21179 3750 35435
3 8984 18572 2472 30030
4 8157 17315 1817 27290
5 7677 17189 1481 26348
6 7280 17022 1308 25610
7 7023 17238 1221 25482
Solution for DNS = 46.5 Ksi
© 2013 Armand J. Chaput
1. Solve simple problems first (current effort)
• Geometry - trapezoidal wing box, equivalent skin thickness
• Loads - symmetrical pull-up, push-over, 2-D distributed loads
(inc. Schrenk approx.) with user defined spar load fraction
• In-plane isotropic properties – assumed for simplicity
• No fasteners or other non-optimums - i.e. “knocked-down”
static stress sized structural representation of a real wing
2. Address buckling as separate issue using CalculiX
buckling factor
3. Next (?) - apply design “rules of thumb” to estimate
basic non-optimums (fasteners and spacing, fuel tanks and
sealant, load introduction fittings, hinges, tracks, etc.)
CDN Mass Estimation Methodology Strategy
Proposed methods are open source and publically available
to encourage collaborative development approach
© 2013 Armand J. Chaput
Good/Bad News - CD Nominal Stress (CDN ) Method
Pro
• Simple and straight forward, anybody can do it
• Based on physics, geometry and CD design requirements
• Based on familiar structural tools and methods
• Reduces reliance on Fudge Factors for MP estimates
• Can be tailored for internal capabilities and skills
• Good risk tracking metric for customer engineers
Con
1. Little CD history
2. Limited publically available correlation data
3. Methodology for Non-Primary Loads, Fittings and Fasteners
and System Integration needs development
© 2013 Armand J. Chaput
Transport Wings
• Advanced Composite Technology (ACT) Test Wing
- PDF sketch level geometry, acceptable MP detail
- Materials defined, no allowables (nominal GrEP assumed)
• B707, B727, B737 (early models)
- Based on 1990s NASA supported PDCyl published data
- PDF sketch level geometry, no detail below wing box level
- Materials assumed (nominal 2024 T3 )
A-6E Wings (metal and composite replacement)
• Biz-jet type planform (exc. for folding wing)
- Good MP data quality, good for Version 2 validation
- PDF sketch level geometry plus handbook data
- Materials assumed (nominal 2024 or GrEP)
UAV Wings (in progress)
• X-56A Wing (1 of 4)
- PDF based geometry, good MP, nominal GrEP assumed
UT Methodology Development/Calibration Status
© 2013 Armand J. Chaput
Advanced Composite Technology (ACT) Static Test
Article (as reconstructed and analyzed) Mass Categories lbm Fraction
Upper and lower skins 1656 0.41
Spar caps and stringers 1238 0.31
Spar webs 350 0.09
Stress based (subtotal) 3244 0.81
Aero ribs and intercostals 520 0.13
MLG rib blkhd 75
MLG pad up 31 0.03
Bolts and nuts 80 0.02
SOB Pad up 25
Access panel pad-up 45 0.02
Total 4020 1.00
Mass analysis based on data from NASA
CR-2001-210650-AST Composite Wing
Program-Executive Summary
Reconstructed mid-
chord thickness
Load actuatorsLoad actuators
LE 30
Stringer runout
Stringer runout
Fiber – IM7 and AS4
Process - GrEP VARTM
Sources: multiple NASA and Boeing ACT Documents
ACT Transport
ACT Test Article
ACT thickness
not linear
Actuator
load (lbf) 2y/b
1 40500 0.947
2 99750 1.000
3 -3000 0.634
4 21000 0.704
5 15000 0.384
6 -45000 0.382
7 45000 0.202
8 6000 0.296
Wing carry-
through not
included in
ACT test box
Spar web CDN = 27.0 ksi
Ribs CDN = 37.9 ksi
Skin CDN = 36.5 ksi
Min gage (given)
= 0.22 in
© 2013 Armand J. Chaput
Generic Boeing Transport Wings – B727, B737, B747
Aircraft or
article WDG [lbs]
DUL
Nz
Half
span Sref
[ft2]
Sweep
[deg]
Half
span [ft]
Half
span
AR TR (t/c)r (t/c)t
Dihed
[deg]
Wing
Fuel
Engine
pylon or
store (per
side)
Engine or
store
mass (ea)
747-100 713,000 3.75 2790 37.5 98.5 6.96 0.265 0.1794 0.078 7 Y 2 8608
737-200 100,800 3.75 502.5 25 45.4 8.21 0.220 0.126 0.112 6 Y 1
727-300 160,000 3.75 793.5 32 55.2 7.67 0.265 0.154 0.09 3 Y 0
ACT Test unk 3.75 N
A-6E 43,077 9.75 264.5 25.5 26.5 5.31 0.390 0.9 0.59 0 N 0 N
A-6E 43,077 9.75 264.5 25.5 26.5 5.31 0.390 0.9 0.59 0 Y 2 2296
A-6E 43,077 -4.5 264.5 25.5 26.5 5.31 0.390 0.9 0.59 0 N 0 N
NASA 110392
Wing Weights
PDCYL
(lbm)
Load-
carrying
structure
(lbm)
Primary
structure
(lbm)
Total
structure
(lbm)
B-727 8688 8791 12388 17860
B-737 5717 5414 7671 10687
B-747 52950 50395 68761 88202
B-720 13962 11747 18914 23528
DC-8 22080 19130 27924 35330
MD-11 33617 35157 47614 62985
MD-83 6953 8720 11553 15839
L-1011 25034 28355 36101 46233
B747-100 VSP
Simplified to 15 ribs
B747, Min gage =
0.1” exc. ribs = 0.36”
CDN = 46.5 ksi
B727-100 VSP
Full 26 ribs B727, Min
gage = 0.1”
CDN = 27.4 ksi
B737-100 VSP
Simplified to 14 ribs B737, Min gage =
0.1” exc ribs = 0.15”
CDN = 29.8 ksi
Parametric geometry from Analytical Fuselage and Wing Weight Estimation of Transport Aircraft, NASA TM 110392, May 1996
Wing Boxes Notional and Extrapolated to Centerline
© 2013 Armand J. Chaput
Reconstructed A-6E Geometry
All locations and linear
dimensions in inches
Wing Box (theo)
b/2 = 305”
Cr (theo) = 117”
Ct = 33”
Theo taper ratio = 0.282
Taper (fold) = 0.661
Sref = 317.7 sqft
Aspect ratio = 8.13
t/c (BL0) = 0.143
t/c (BL144) = 0.137
t/c (BL 305) = 0.115
Kc = 1.0
Theoretical Wing
b/2 = 318”
Cr (theo) = 182.6”
Ct (theo) = 57”
Taper ratio = 0.312
Sref = 528.9 sqft
Aspect ratio = 5.31
Taper ratio = 0.312
Taper (fold) = 0.689
t/c (BL33) = .09
t/c (BL144) = 0.084
t/c (tip) = 0.059
LE flap = 15% c
Flaperon LE at 65% c
Est. box Kc = 0.5
BL
30
5
BL
31
8
FS 228.2
FS 283.9
33”
BL
0
0.70 c?
182.6
BL
14
4
FS 0
56”
0.83 c
0.15 c
57”
318”
B
L 7
8
B
L 6
6
0.05 c
0.70 c
28.5
28
117
”
• Wing fuel – L/R full 6923 lbm
• 2 Ext fuel inbd – L/R 4010 lbm (2)
• 2 Tank w/adapter inbd L/R 398 lbm
• 2 Pylon inbd L/R : 192.6 lbm (WS 95)
• 2 Ext fuel outbd – L/R 4010 lbm
• 2 Tank w/adapter outbd L/R 398 lbm
• 2 Pylon outbd L/R : 183.4 lbm (WS 141)
• Pylon (WS 187) – replacement wing only
• Max GW landplane: 60705 lbm
• Max t/o landplane: 60400 lbm
• Flt des GW (landplane): 36526 DUL = 9.75
• Max GW zero fuel, zero stores = 39781 (body
fuel =9016 lbm)
BL
38
.9
Wing Box (less ctr sect)
b/2 = 239”
Cr = 98.8”
Ct = 33”
Theo taper ratio = 0.334
Taper (fold) = 0.783
Sref = 218.8 sqft
Aspect ratio = 7.25
t/c (BL66) = 0.140
t/c (BL144) = 0.137
t/c (BL 305) = 0.115
Kc = .5/.65 = 0.77
0.15 c
0.65 c
A-6 graphic and raw data from A-6
Post Design Analysis Report (CORL
G003), Dec 1993
© 2013 Armand J. Chaput
A-6E MP Summary Data and VSP SAM Results
JS&F = Joints, splices and fasteners
TipFaFen = tips, fairings and fences
Wing box
mat'l Skins
Skin
stiffners
Spars and
stiffeners Ribs Rib Bkhd Js&F TipFaFen Other Pri - sum TipFaFen Doors Acs Panel Fin & walk Misc Sec - sum CV Unique Total
Wing primary - lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm lbm
CENTER SECTION 718.00 0.00 39.50 0.00 103.60 35.40 0.00 3.40 899.90 1799.80
INTERMEDIATE PANEL 1077.80 681.40 0.00 200.00 0.00 196.40 106.10 0.00 0.00 1183.90 2367.80
OUTER PANEL 814.70 488.30 1.30 239.50 44.20 41.40 58.90 0.00 0.00 873.60 1747.20
LE 142.70 86.30 13.30 15.40 27.70 6.40 0.00 0.00 149.10 298.20
TE 68.60 37.40 0.00 19.20 12.00 0.00 15.20 0.00 83.80 167.60
Wing and intergration 35.8 35.80
WING TOTAL 2964.90 2011.40 14.60 513.60 83.90 341.40 206.80 51.00 3.40 3190.30 231.80 165.90 2.00 125.60 3.80 529.10 295.70 4015.10
Less center section 2103.80 1293.40 14.60 474.10 83.90 237.80 171.40
SECONDARY STRUCTUREPRIMARY STRUCTURE
Data from: MODEL A-6E ACTUAL DETAIL WEIGHT AND
BALANCE REPORT, NO. WT-128R-1S37, August 1988
Spar web CDN = 19.9 ksi
Ribs CDN = 12.4 ksi
Skin CDN = 46.2 ksi
A-6E @ Wfdg = 36526 lbm
no ext tanks, no wing fuel,
Min gage = 0.040”
A-6E @ Wfdg = 36526 lbm with
2x2300 lbm ext tanks + 3462 lbm
wing fuel, Min gage = 0.040”
Spar web CDN = 15.0 ksi
Ribs CDN = 9.0 ksi
Skin CDN = 42.6 ksi
Note - Rib definition includes bulkheads and store stations
- Spar web fuel pressure, cat. and arrest loads not included
Overall results correlate with expectations
- Additional load cases and refined geometry expected to improve results
© 2013 Armand J. Chaput
VSP SAM CDN Preliminary Results
CD Nominal Sress Correlation
0
5
10
15
20
25
30
35
40
45
50
0 200 400 600 800
WL x Nz (psf)
Sig
ma
CD
N (
ks
i)
A-6 spweb
A-6 rib
A-6 skin
ACT spweb
ACT rib
ACT skin
B727-300
B737-200
B747-100
A-6E weighted
ACT weighted
Overall wing box mass weighted results
© 2013 Armand J. Chaput
VSP SAM CDN Issue – VSP Geometry Constraint
A-6E Wing Thickness Profile
0.05
0.08
0.10
0.13
0.15
0 50 100 150 200 250 300 350
BL (in)
t/c
WingTheo boxTheo box less ctr
A-6E Wing Chord Distribution
0.00
0.20
0.40
0.60
0.80
1.00
1.20
0 50 100 150 200 250 300 350
BL (in)
t/c
Wing
Wing box
Actual t/c distribution
Assumed t/c distribution
0.00
5.00
10.00
15.00
20.00
25.00
30.00
35.00
40.00
45.00
0 100 200 300 400 500 600
ht
(in
)
ACT Wing Box t/c
Least Squares Fit
0
0.05
0.1
0.15
0.2
0.25
0.3
0.35
0.4
0 100 200 300 400 500BL (in)
t/c
Max to min t/c
VSP structural module grid limited to single trapezoidal planform and stream-wise chord
- Even CD level structures need more flexible multi-panel capability - Effect on ACT
- In order of priority we need (1) non-linear t/c, (2) non-linear chord and (3) non-stream-wise cut capabilities
Case 1 Case 2 Diff
Skin 36.46 47.31 30%
Ribs 37.85 31.94 -16%
Spars 27.02 34.04 26%
CDNS (psi)
© 2013 Armand J. Chaput
Concluding remarks - VSP Structural Design and Analysis
1. FEM design and analysis can be accommodated during
conceptual design without adding onerous requirements
for higher levels of design detail
- Serious structural design issues can be assessed, identified
and resolved without slowing down the pace of design
2. Design and analysis methods can be applied intelligently
by designers who are not structural specialists
- Tool specific details can be pushed into the background
3. FEM model results correlate with actual mass property
data through use of data validated CDN
4. VSP SAM based methods are now ready to replace
parametric mass estimate methods for primary load
carrying wing structure
© 2013 Armand J. Chaput
Future Plans
1. Contract nearing completion – tasks to complete in
order of priority
a. Document results, present at 2014 AIAA ASM
b. Complete Version 2 validation
c. Post Version 2 Software and Users Guide
d. Additional methodology calibration
© 2013 Armand J. Chaput
Version 2 Overview - New Features
UT Input Executable (Java)
Boundary Conditions and Load Cases
CalculiX Input File
Vehicle Sketch Pad
External and Internal Mesh
Generation
Parametric
External Geometry
Parametric
Internal Geometry
CalculiX
FEM Solution
FEM Input
FEM Post Process
and Graphics
Output Files
UT Convergence Executable
(Java)
Solution Files
Thickness Iteration
Stress Convergence
Thickness and Material Properties
Mass Calculation
Wing Trim
© 2013 Armand J. Chaput
Wing Trim – Structural Model Geometry
UT Input Executable (Java)
Boundary Conditions and Load Cases
CalculiX Input File
Thickness and Material Properties
Wing Trim • Deletes non-primary load carrying structure
- Typically leading and trailing edge devices
• Deletes non-load carrying skin panels
- To represent typical fabric or film skin sections
© 2013 Armand J. Chaput
Inertia Loads (relief)
UT Input Executable (Java)
Boundary Conditions and Load Cases
CalculiX Input File
Thickness and Material Properties
Wing Trim
Front view – Notional Wing
Fuel mass
Engine and pylon
External store
• Discrete loads applied to rib and/or spar centroid at defined rib and/or spar
• Fuel inertia applied as internal pressure load along bottom (+nz) or top (-nz)
of defined rib/spar tank boundaries
nz
© 2013 Armand J. Chaput
Multiple Load Cases
UT Input Executable (Java)
Boundary Conditions and Load Cases
CalculiX Input File
Thickness and Material Properties
Wing Trim
Adapted from http://upload.wikimedia.org/wikipedia/commons/1/13/PerformanceEnvelope.gif
1 2
4 3
1 – Max + 2 - Min +
4 - Max - 3 - Min -
Multiple load case methodology sizes structure (and
calculates mass) for most demanding of multiple cases
© 2013 Armand J. Chaput
Running CalculiX
Plus user feedback on solution status
© 2013 Armand J. Chaput
CalculiX Buckling Factor (BLF)
CalculiX Linear Buckling Analysis
0.60
0.70
0.80
0.90
1.00
1.10
1.20
4 8 12 16
Number of Ribs
Bu
cklin
g Fa
cto
r (B
LF)
CalculiX BLF output can provide design guidance on
spacing for stability but at cost of 2x solution time
CalculiX
FEM Solution
FEM Input
FEM Post Process
and Graphics
Output Files
© 2013 Armand J. Chaput
Questions
© 2013 Armand J. Chaput
Reconstructed mid-
chord thickness
Actuator # Rib # Location
TE sta.
fm CL bte
C perp to
TE (in)
Chord
fraction
from TE
BL from
SOB (in)
Max span
fraction
Chord
fraction Load (lbf)
1 18 TE 543.5 453.2 59.42 0 410.7 0.942 0 27000
2 18 LE 543.6 453.3 59.42 1 435.9 1.000 1 66500
3 13 TE 396.5 306.2 72.86 0 277.5 0.637 0 -2000
4 13 LE 396.5 306.2 72.86 1 308.3 0.707 1 14000
5 9 TE 279.5 189.2 83.55 0 171.5 0.393 0 10000
6 7.50 LE 236.75 146.45 87.46 1 169.7 0.389 1 -30000
7 6 TE 194 103.7 91.37 0 94.0 0.216 0 30000
8 6 LE 194 103.7 91.37 1 132.6 0.304 1 4000
Actuator # Rib # Location
TE sta.
fm CL bte
C perp to
TE (in)
Chord
fraction
from TE
BL from
SOB (in)
Max span
fraction
Chord
fraction Load (lbf)
1 18 TE 543.5 453.2 59.42 0 410.7 0.942 0 27000
2 18 LE 543.6 453.3 59.42 1 435.9 1.000 1 66500
3 13 TE 396.5 306.2 72.86 0 277.5 0.637 0 -2000
4 13 LE 396.5 306.2 72.86 1 308.3 0.707 1 14000
5 9 TE 279.5 189.2 83.55 0 171.5 0.393 0 10000
6 7.50 LE 236.75 146.45 87.46 1 169.7 0.389 1 -30000
7 6 TE 194 103.7 91.37 0 94.0 0.216 0 30000
8 6 LE 194 103.7 91.37 1 132.6 0.304 1 4000
Load actuators
LE 30
Stringer runout
Stringer runout
Fiber – IM7 and AS4
Process - GrEP VARTM
Sources: multiple NASA and Boeing ACT Documents
ACT Transport
ACT Test Article