VSP Structural Analysis Module

Transcription

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
© 2013 Armand J. Chaput
Self Introductions
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
Current
Hersh Amin, Research Ass’t
Josh Eboh, Team Lead
Natalie Maka , Research Ass’t
Patil Tabanian, Research Intern
Previous
Sarah Brown, Research Ass’t
Jose Galvan, Research Ass’t
Alex Haecker, Research Ass’t
Tejas Kulkarni, Team Lead
Based on work performed under NASA/NIA Task Order 6322-UTEX,
“Advanced Conceptual Design Tools and Development”
 Armand J. Chaput 2013
VSP Structural Analysis Module (SAM) R&D Objectives
(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
Objective 1 - Expanded Capabilities
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 Codes and users guide posted
at: http://vspsam.ae.utexas.edu/
Objective 2 - CalculiX Integration
Vehicle Sketch Pad
Parametric
Internal Geometry
Parametric
External Geometry
External and Internal Mesh
Generation
UT Convergence Executable
(Java)
UT Input Executable (Java)
Wing Trim
Thickness and
Material Properties
Boundary Conditions
and Load Cases
CalculiX Input File
CalculiX 
FEM Input
Thickness Iteration
FEM Solution
Solution Files
Stress Convergence
FEM Post Process
and Graphics
Mass Calculation
Output Files
Objective 3 – FEM mass for fully stressed trade study
wings with minimum gage constraints (from last year)
Last iteration –
thickness “converged”
Last iteration –
2 spars,
5 ribs
yellow = 28.9 ksi
2 spars,
7 ribs
yellow = 28.7 ksi
Last iteration –
5 spars,
15 ribs
yellow = 28.4 ksi
• 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
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
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
Why we need improved CD methods
- Example from UT method development effort (A-6E)
Typical CD wing parametric estimates:
Wdg = 36526 lbm
Wto (land) = 60705 lbm
1. Raymer (fighter-attack):
= 0.0103 [(Wdgnzdu)0.5 Sref0.622 AR0.785
(1+)0.05 Scsw0.04]] / [(t/c)0.4
Cos(0.25c)]
= 4092 lbm
WING STRUCTURE - BASIC
SECONDARY STRUCTURE
TRAILING EDGE DEVICES
LEADING EDGE DEVICES
SPEED BRAKES
WING GROUP - TOTAL
Inc. wing fold unique
3443
931
593
241
145
5352
297
3443
4374
4966
5207
5352
 117”
2. Nicolai (USN fighter):
= 19.29 [(Wtonzdu)/(t/c)] {[(Tan le 2(1-) / AR(1+)] 2 +1] 10-6} 0.464
0.70S 0.58 = 7057 lbm
[(AR(1+

)]
A-6E WING GROUPref
(lbm)
Sum
MP Data from Grumman Aerospace
A-6E Weight Report
WT-128R-1S37 Aug 1988
Courtesy of Paul Kachurak, NAVAIR
FS 228.2
0.15 c
182.6
FS 283.9
0.83 c
0.70 c?
318”
much unknown
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 known
Perspective – Why airframes weigh what they weigh
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,
Iter
1
2
3
4
5
6
7
B747 Simplified Rib Model before & after Trim
Spar
Mass
(lbm)
13392
10506
8984
8157
7677
7280
7023
Skin
Mass
(lbm)
26457
21179
18572
17315
17189
17022
17238
Rib Mass
(lbm)
6249
3750
2472
1817
1481
1308
1221
Total
Mass
(lbm)
46100
35435
30030
27290
26348
25610
25482
Solution for DNS = 46.5 Ksi
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
CDN Mass Estimation Methodology Strategy
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.)
Proposed methods are open source and publically available
to encourage collaborative development approach
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
UT Methodology Development/Calibration Status
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
Advanced Composite Technology (ACT) Static Test
Article (as reconstructed and analyzed)
Wing carrythrough not
included in
ACT test box
ACT thickness
not linear
Mass analysis based on data from NASA
CR-2001-210650-AST Composite Wing
Program-Executive Summary
Sources: multiple NASA and Boeing ACT Documents
Reconstructed midchord thickness
Stringer runout
LE  30
Stringer
runout
ACT Test
Article
lbm
1656
1238
350
3244
Fraction
0.41
0.31
0.09
0.81
Aero ribs and intercostals
MLG rib blkhd
MLG pad up
Bolts and nuts
SOB Pad up
Access panel pad-up
Total
520
75
31
80
25
45
4020
0.13
1
2
3
4
5
6
7
8
ACT Transport
Fiber – IM7 and AS4
Process - GrEP VARTM
Mass Categories
Upper and lower skins
Spar caps and stringers
Spar webs
Stress based (subtotal)
Load actuators
Actuator
load (lbf)
40500
99750
-3000
21000
15000
-45000
45000
6000
Spar web CDN = 27.0 ksi
Ribs CDN = 37.9 ksi
Skin CDN = 36.5 ksi
Min gage (given)
= 0.22 in
0.03
0.02
0.02
1.00
2y/b
0.947
1.000
0.634
0.704
0.384
0.382
0.202
0.296
Generic Boeing Transport Wings – B727, B737, B747
Parametric geometry from Analytical Fuselage and Wing Weight Estimation of Transport Aircraft, NASA TM 110392, May 1996
Half
Half
span
S
ref Sweep
Aircraft or
Half
span
DUL
2
article
[ft
]
[deg]
span
[ft]
AR
WDG [lbs] Nz
747-100
737-200
727-300
ACT Test
A-6E
A-6E
A-6E
713,000
100,800
160,000
unk
43,077
43,077
43,077
B747-100 VSP
Simplified to 15 ribs
3.75
3.75
3.75
3.75
9.75
9.75
-4.5
2790
502.5
793.5
264.5
264.5
264.5
37.5
25
32
98.5
45.4
55.2
6.96
8.21
7.67
25.5
26.5ksi 5.31
CDN = 46.5
25.5
25.5
26.5
26.5
5.31
5.31
TR
(t/c)r
0.265 0.1794
0.220 0.126
0.265 0.154
0.390
0.390
0.390
B747, Min gage =
0.1” exc. ribs = 0.36”
(t/c)t
Dihed
[deg]
0.078
0.112
0.09
7
6
3
0.9
0.59
NASA 110392
0.9
0.59
Wing
0.9 Weights
0.59
B-727
B-737
B-747
B-720
DC-8
MD-11
MD-83
L-1011
0
PDCYL
0
(lbm)
0
8688
5717
52950
13962
22080
33617
6953
25034
CDN = 27.4 ksi
B727-100 VSP
Full 26 ribs
B727, Min
gage = 0.1”
Engine
pylon or Engine or
Wing store (per store
Fuel
side)
mass (ea)
Y
2
8608
Y
1
Y
0
N LoadNcarrying 0 PrimaryN
Total
Ystructure 2 structure
2296 structure
(lbm)
N (lbm) 0 (lbm) N
8791
5414
50395
11747
19130
35157
8720
28355
12388
7671
68761
18914
27924
47614
11553
36101
CDN = 29.8 ksi
B737-100 VSP
Simplified to 14 ribs
B737, Min gage =
0.1” exc ribs = 0.15”
Wing Boxes Notional and Extrapolated to Centerline
17860
10687
88202
23528
35330
62985
15839
46233
Reconstructed A-6E Geometry
 BL 66
BL 38.9
BL 0
FS 0
28.5
•
•
•
•
•
•
•
•
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)
A-6 graphic and raw data from A-6
Post Design Analysis Report (CORL
G003), Dec 1993
 117”
All locations and linear
dimensions in inches
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
 BL 78
Wing Box (theo)
b/2 = 305”
Cr (theo) = 117”
Ct = 33”
Theo taper ratio = 0.282
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
BL 144
BL 318
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
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 305
FS 228.2
0.15 c
 0.05 c
56”
 0.15 c
 28
182.6
FS 283.9
 0.65 c
33”
 0.70 c
0.83 c
0.70 c?
57”
318”
A-6E MP Summary Data and VSP SAM Results
PRIMARY STRUCTURE
Wing box
mat'l
Wing primary - lbm
lbm
CENTER SECTION
INTERMEDIATE PANEL 1077.80
OUTER PANEL 814.70
LE 142.70
TE
68.60
Wing and intergration
WING TOTAL 2964.90
Less center section 2103.80
Skins
lbm
718.00
681.40
488.30
86.30
37.40
2011.40
1293.40
Skin
Spars and
stiffners stiffeners
lbm
lbm
0.00
39.50
0.00
200.00
1.30
239.50
13.30
15.40
0.00
19.20
14.60
14.60
513.60
474.10
Ribs
lbm
0.00
0.00
44.20
27.70
12.00
Rib Bkhd
lbm
103.60
196.40
41.40
Js&F
lbm
35.40
106.10
58.90
6.40
0.00
83.90
83.90
341.40
237.80
206.80
171.40
A-6E @ Wfdg = 36526 lbm
no ext tanks, no wing fuel,
Min gage = 0.040”
SECONDARY STRUCTURE
TipFaFen
lbm
0.00
0.00
0.00
0.00
15.20
35.8
51.00
Other
lbm
3.40
0.00
0.00
0.00
0.00
3.40
Pri - sum TipFaFen Doors Acs Panel Fin & walk Misc
Sec - sum CV Unique Total
lbm
lbm
lbm
lbm
lbm
lbm
lbm
lbm
lbm
899.90
1799.80
Data from: MODEL A-6E ACTUAL DETAIL WEIGHT AND
1183.90
2367.80
BALANCE REPORT, NO. WT-128R-1S37, August 1988
873.60
1747.20
149.10
298.20
JS&F = Joints, splices and fasteners
83.80
167.60
TipFaFen = tips, fairings and fences
35.80
3190.30
231.80
165.90
2.00
125.60
3.80
529.10
295.70
4015.10
A-6E @ Wfdg = 36526 lbm with
2x2300 lbm ext tanks + 3462 lbm
wing fuel, Min gage = 0.040”
Note - Rib definition includes bulkheads and store stations
- Spar web fuel pressure, cat. and arrest loads not included
VSP SAM CDN Preliminary Results
CD Nominal Sress Correlation
50
A-6 spweb
45
A-6 rib
Sigma CDN (ksi)
40
A-6 skin
ACT spweb
35
ACT rib
ACT skin
30
B727-300
25
B737-200
B747-100
20
A-6E weighted
15
ACT weighted
10
Overall wing box mass weighted results
5
0
0
200
400
600
800
WL x Nz (psf)
Overall results correlate with expectations
- Additional load cases and refined geometry expected to improve results
VSP SAM CDN Issue – VSP Geometry Constraint
A-6E Wing Chord Distribution
A-6E Wing Thickness Profile
Wing
Wing box
1.20
Wing
Theo box
Theo box less ctr
0.15
1.00
0.13
t/c
t/c
0.80
0.60
0.10
Actual t/c distribution
0.40
0.08
0.20
Assumed t/c distribution
0.00
0.05
0
50
100
150
200
250
300
350
0
50
100
BL (in)
150
200
250
300
350
BL (in)
VSP structural module grid limited to single
trapezoidal planform and stream-wise chord
- Even CD level structures need more
flexible multi-panel capability
CDNS (psi)
- Effect on ACT
ACT Wing Box t/c
45.00
0.4
40.00
0.35
Max to min t/c
35.00
0.3
30.00
ht (in)
t/c
0.25
0.2
0.15
25.00
20.00
Skin
Ribs
Spars
Least Squares Fit
15.00
0.1
10.00
0.05
Case 1
36.46
37.85
27.02
Case 2
47.31
31.94
34.04
Diff
30%
-16%
26%
- In order of priority we need (1) non-linear
t/c, (2) non-linear chord and (3) nonstream-wise cut capabilities
5.00
0
0
100
200
300
BL (in)
400
500
0.00
0
100
200
300
400
500
600
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
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
Version 2 Overview - New Features
Vehicle Sketch Pad
Parametric
Internal Geometry
Parametric
External Geometry
External and Internal Mesh
Generation
UT Convergence Executable
(Java)
Wing Trim
Thickness and
Material Properties
Boundary Conditions
and Load Cases
CalculiX Input File
CalculiX 
FEM Input
Thickness Iteration
FEM Solution
Solution Files
Stress Convergence
FEM Post Process
and Graphics
Mass Calculation
Output Files
Wing Trim – Structural Model Geometry
• 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
Wing Trim
Thickness and
Material Properties
Boundary Conditions
and Load Cases
CalculiX Input File
Inertia Loads (relief)
• 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
Fuel mass
External store
Engine and pylon
Front view – Notional Wing
Wing Trim
Thickness and
Material Properties
Boundary Conditions
and Load Cases
CalculiX Input File
Multiple Load Cases
1 – Max +
1
2 - Min +
2
Wing Trim
Thickness and
Material Properties
Boundary Conditions
and Load Cases
CalculiX Input File
4
4 - Max -
3
3 - Min -
Adapted from http://upload.wikimedia.org/wikipedia/commons/1/13/PerformanceEnvelope.gif
Multiple load case methodology sizes structure (and
calculates mass) for most demanding of multiple cases
Running CalculiX
Plus user feedback on solution status
CalculiX Buckling Factor (BLF)
CalculiX Linear Buckling Analysis
CalculiX 
FEM Input
1.20
Buckling Factor (BLF)
FEM Solution
1.10
FEM Post Process
and Graphics
1.00
Output Files
0.90
0.80
0.70
0.60
4
8
12
16
Number of Ribs
CalculiX BLF output can provide design guidance on
spacing for stability but at cost of 2x solution time
Questions
ord
TE sta.
ction BL from Max span Chord
Rib # fraction
LocationLoadfm
CL
m TE Actuator
SOB (in) # fraction
(lbf)
1
18
543.5
0
410.7
0.942
0 TE
27000
2
18
543.6
1
435.9
1.000
1 LE
66500
3
13
396.5
0
277.5
0.637
0 TE
-2000
4
13
396.5
1
308.3
0.707
1 LE
14000
5
279.5
0
171.5
0.3939
0 TE
10000
6
7.50
236.75
1
169.7
0.389
1 LE
-30000
7
194
0
94.0
0.2166
0 TE
30000
8
194
1
132.6
0.3046
1 LE
4000
bte
453.2
453.3
306.2
306.2
189.2
146.45
103.7
103.7
Chord
C perp to fraction BL from Max span Chord
TESources:
(in)
from TE
SOB (in)NASA
fraction
Load
(lbf)
multiple
andfraction
Boeing
ACT
59.42
0
410.7
0.942
0
27000
59.42
1
435.9
1.000
1
66500
72.86
0
277.5
0.637
0
-2000
72.86
1
308.3
0.707
1
14000
83.55
0
171.5
0.393
0
10000
87.46
1
169.7
0.389
1
-30000
91.37
0
94.0
0.216
0
30000
91.37
1
132.6
0.304
1
4000
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