An overview of the Aeroelastic Design Optimization Program (ADOP) at the Douglas Aircraft Company is given. A pilot test program involving the animation of mode shapes with solid rendering as well as wire frame displays, a complete aircraft model of a high-altitude hypersonic aircraft to test ADOP procedures, a flap model, and an aero-mesh modeler for doublet lattice aerodynamics are discussed

Rommel, Bruce A.

English

NASA Technical Reports Server

N89- 25222 - 
a *  - -  
MEETING THE CHALLENGES WITH THE DOUGLAS AIRCRAFT COMPANY 
AEROELASTIC DESIGN OPTIMIZATION PROGRAM (ADOP) 
BRUCE A. ROMMEL 
Long Beach, California 
Douglas Aircraft Company, McDonnell Douglas Corporation 
1369 
https://ntrs.nasa.gov/search.jsp?R=19890015851 2020-03-20T01:49:06+00:00Z
THE CHALLENGES OF AEROELASTIC DESIGN OPTIMIZATION 
Aeroelastic design optimization for commercial aircraft must consider the full range 
of design loads and aeroelastic constraints. If significant constraints are omitted 
from the design optimization process, the result may be far from satisfactory and 
perhaps worse than a nonoptimal design obtained from conventional procedures. Fatigue 
is a major consideration: As shown in Figure 1, a design which is close to optimum 
for only a few load conditions may result in a severely short service life. Nearly 
5,000 load conditions resulting from variations in speed, weight, maneuvers, and gust 
loads must be considered when designing a typical commercial aircraft. Typically, 
service lift goals are on the order of 60,000 flight hours, with 30,000 flights and 
landings over a service life of 20 years, To achieve these goals the fatigue life is 
normally set at about twice the number of flight hours and landings of the design 
service l i f e .  Recent experience indicates that these goals may double or even triple 
with the next generation of commercial aircraft. 
To achieve these goals the analysis and design optimization program must be able to 
handle the large-scale, finite-element models required for detailed stress analysis, 
On the other hand, to automate the design process, common structural models for static 
strength, structural dynamics, flutter, and aeroelastic constraints are required. 
Critical flutter and aeroelastic constraints often arise near the region of transonic 
flight. A high-speed commercial transport to service the Pacific Basin will cruise 
at supersonic or even hypersonic speed, resulting in high heat loading conditions. 
Buckled skin finite elements used for high load conditions In static strength analysis 
are inappropriate for modal analysis resulting in a model dependency. Future aircraft 
will undoubtedly use more advanced composite materials to reduce weight and achieve 
desirable aeroelastic characteristics. It is with these considerations that we are 
developing an Aeroelastic Design Optimization Program at Douglas Aircraft Company. 
CRITICAL LOAD CONDITIONS FOR STATIC STRENGTH (5,000 LOAD CASES) 
FATIGUE DESIGN FOR A 20-YEAR SERVICE LIFE 
0 60,000 FLIGHT HOURS/FATIGUE LIFE 120,000 HOURS 
0 30,000 FLIGHTS/FATIGUE LIFE 60,000 FLIGHTS 
LARGE FINITE-ELEMENT MODELS (10,000 - 60,000 DOF) 
COMMON MODELS FOR STRENGTH AND FLUTTER DESIGN 
DESIGN FOR TRANSONIC/SUPERSONIC TRANSPORTS 
Figure 1 
1370 
ORIGINAL PAGE IS 
OF POOR QUALITY 
THE PILOT TEST PROBLEM (FIN OR SMALL WING) 
Since t h e  e a r l i e s t  days of ADOP development w e  have used the  f i n  o r  small  wing model 
t o  develop our da t a  flows and procedures.  This  model i s  small  enough so t h a t  much of 
t he  requi red  da ta  can be ca l cu la t ed  by hand f o r  v e r i f i c a t i o n .  The amount of da ta  i s  
small  enough t h a t  t he  procedure may, i n  many cases ,  be s i g h t  v e r i f i e d .  Figure 2 dep ic t s  
one of t he  simple wing modes. The mesh shown superimposed over t he  w i r e  frame model 
i s  the  ae ro  c o n t r o l  p o i n t  mesh. It  i s  a pseudo mesh generated t o  show how w e l l  the  
s p l i n e  between the  s t r u c t u r a l  mesh and the  aero  mesh has  f i t  the  da t a .  A Harder s p l i n e  
was used, and the  f i t  w a s  q u i t e  good even though the  s p l i n e  had t o  ex t r apo la t e  beyond 
the  domain of t he  s t r u c t u r a l  box. This  r e s u l t e d  i n  a s l i g h t  cusping of t he  po in t s  
f a r t h e s t  from the  s t r u c t u r a l  box. I n  genera l ,  the  Harder s p l i n e  w i l l  work b e s t  when 
no ex t r apo la t ion  i s  requi red  and the  sp l ined  p o i n t s  are near  t he  da ta  po in t s .  Even the  
h igher  frequency modes were f i t t e d  adequately wi th  t h i s  s p l i n e ;  however, o the r  s p l i n e  
procedures a r e  being developed f o r  use i n  ADOP. With t h i s  model we developed the  
animation of  mode shapes wi th  s o l i d  rendering a s  w e l l  as w i r e  frame d isp lays .  The use r  
may r o t a t e  during modal animation t o  b e t t e r  v i s u a l i z e  t h e  modes. 
This  same model was used t o  develop a s t r e s s  contours  post-processor  d isp lay .  
_-- 
Figure 2 
THE BIG TEST PROBLEM 
The next test problem developed to test the ADOP program procedures was a complete 
aircraft model of a high-altitude hypersonic aircraft. Most advanced design study 
models use 5,000-10,000 DOF. Modal analysis of these models can be completed in a few 
minutes during a *TSO interactive session on the IBM 3090 computer without resorting 
to batch-mode processing. Such models can be rotated during animation, even with mirror 
imaging and solid rendering. 
The model shown in Figure 3 was developed by Alan Dodd for an advanced design study. 
This configuration, which has many interesting design features such as the V-tail, was 
abandoned as impractical. It makes an ideal study model because no military or com- 
mercial trade secrets are revealed in its configuration. It was the first model with 
which we encountered significant modeling discrepancies with modal display. Some sin- 
gular modes were developed that showed incorrect element nodal connectivity, even 
though an interactive graphics system was used to generate the model. These "marble 
modes'' consisted of large deflections of a single node while the rest of the model was 
relatively stationary. Once these nodes and elements were defined, the modeling prob- 
lems could be corrected. Subsequently, this condition was discovered in several other 
models developed for static strength analysis. The ability to compute and display the 
modes of large-scale, finite-element models has thus become a very valuable diagnostic 
tool for finite-element structural analysis. This was an unanticipated benefit of ADOP 
development. 
*Time-sharing Option 
Figure 3 
1372 ORIGINAL PAGE 
BLACK AND WHITE P H O T o c i R W  
LARGE SCALE VERIFICATION MODEL (MD-80 SERIES) 
When we started ADOP development, models like the one in Figure 4 - originally developed 
during the detailed design development of the MD-80 - were considered too large for 
use in any aeroelastic design optimization procedure. However, after we had developed 
procedures for weight and balance and large-scale modal analysis, we were challenged 
to test our procedures against real results from real aircraft. Since we don't do ground 
vibration tests on paper airplanes, we found ourselves translating this model, com- 
puting its weights, and calculating its modes. This model uses about 30,000 DOF and 
about 10 modes can be computed in about 1 .5  hours on the IBM 3090 computer. Modal 
animation displays can be achieved, but view selection is much slower than for the 
smaller models. We divided the model into bays for weights analysis. When we translate 
the finite-element model we have only the finite-element weight. Other weights are 
determined by our semi-empirical weights program. This program computes the total bay 
weights and moment of inertia. From this data we compute the tear weight difference 
between the finite-element model and the bay weight. This is accomplished by dividing 
the model into bays with our IMAGES interactive graphics program. The tear weight is 
then distributed to the surface elements in each bay as a nonstructural mass. Major 
masses may be excluded from this distribution and lumped directly to one or more nodes 
using the IMAGES system. Passenger floor loading is assigned to the model using the 
IMAGES system. 
Figure 4 
1373 
FLAP MODEL 
In one of our first major ADOP applications we computed the modes of a flap model for 
an aircraft that we currently have under design and development (Figure 5). The 
widespread availability of computer graphics for finite element modeling has resulted 
in the development of very large structural models. This inboard flap model uses 35,523 
DOF which is larger than the entire MD-83 model. With our first attempt at modal 
analysis of this model, we discovered numerous singularities. Many were traceable to 
an incompatible programming assumption: that all models should be able to take a 
general inertia loading at each node. Many of the bar elements in this model had section 
properties defined in only one direction. Furthermore, many dummy bar elements had been 
introduced as a modeling convenience that did not represent any real structure. 
I These fictitious structural elements and numerous boundary condition errors were the source of most of our problems. ADOP detects joint instability by considering the 
internal load paths resulting from all elements connected to each joint. Two approaches 
to automatic joint instability correction (NASTRAN Auto SPC) have been used in other 
programs. In one approach light stability springs are attached to the unstable nodes. 
In the other approach the unstable nodes are treated as skew nodes and the global DOF 
for these nodes are rotated so that the singular directions can be grounded out. In 
ADOP we had taken the the spring approach. This model showed us that this was a bad 
approach. A s  a result, ADOP was significantly reprogrammed to use the skew node 
approach. This corrected most of the problems. A modal analysis with ADOP was then 
performed and the modes reviewed with the IMAGES system. AS a result, this fly-away 
vane deflection was discovered in some of the modes, as was an error in the model’s 
boundary conditions. These errors were corrected and the modes recomputed. 
. 
Figure 5 
FLAP MODEL TIMING RESULTS 
The modal analysis extracted eight modes. One of these is shown in Figure 6 .  ADOP’s 
front-end processes include model translation from NASTRAN or CASD model files, joint 
instability identification and correction, equation count, node and element rese- 
quencing to minimize the wavefront and envelope size, and the generation of mass and 
stiffness matrix derivatives (scaled element arrays). These expensive front-end proc- 
esses result in a well-structured data base f o r  the numerically intensive modal anal- 
ysis. ADOP uses an adaptive shift subspace iteration procedure for modal analysis. The 
actual modal analysis of this model was accomplished in 79.26 CPU minutes on the IBM 
3090 computer, and 9.70 CPU minutes on the CRAY XMP-18. This speed-up factor of 8.2 
is close to the maximum expected computational ratio for these single-CPU machines. 
Front processes were much lower, resulting in a speed-up factor of only about 2.3 
between the CRAY XMP-18 and the IBM 3090. The modes extracted with ADOP were compared 
with the modes extracted with NASTRAN, resulting in errors of about one percent for 
all modes,. The NASTRAN results were computed with both the Guyan reduction and 
Houeeholder Givens in superelement procedure6 and with the NASTRAN block Lanczos pro- 
cedure. The NASTRAN block Lanczos required 127.14 CPU minutes on the IBM 3090. Direct 
comparisons between the NASTRAN and ADOP run times are difficult to make because 
NASTRAN performed more shifts and computed more modes than did ADOP. Block Lanczos 
should be faster than subspace iteration, but the very efficient implementation of 
adaptive subspace in ADOP makes it very competitive with NASTRAN Block Lanczos. 
Select node 
Shou Undefomed OFF 
Uire frame ON 
Display Shading OFF 
Menus Ooun j=i 
Figure 6 
1375 
AERO-MESH MODELER FOR DOUBLET LATTICE AERODYNAMICS 
In a recent development, work has been initiated to add an aero-mesh modeler to the 
IMAGES system for aeroelastic analysis and design. Neither PATRAN nor Douglas’ Computer 
Graphics Structural Analysis (CGSA) program provides this kind of modeling capability. 
Currently, we are developing models from an existing data base of components of generic 
aircraft parts, as well as from three-view drawings. This new capability has been under 
development since July 15, 1988. Figure 7 shows the aero-mesh model being developed 
with this new Douglas graphics system. 
Y 
..4 
I .m 
I 
i I 
Figure 7 
I 1376 ORIGINAL PAGE 
BLACK APJD WHITE PHOTOGRAPH 
LOOKING BACK 
Figure 8 shows the work completed and the work just initiated on the ADOP system. 
We have completed a large-order static and modal analysis system. We have demonstrated 
a fully stressed design capability for static strength. We have demonstrated a flutter 
analysis capability and computed design sensitivities for flutter and static strength. 
We have in place a time-domain integration system for modal models and work has been 
initiated on a direct integration package for large structural models. Work has been 
initiated on a new composite finite element library for analysis and design. We have 
in place a very powerful direct matrix abstraction language (ACL-DMAP), which in some 
respects is more powerful than the ASTROS-MAPOL or NASTRAN-DMAP languages. We have 
initiated work on the new static aeroelastic package, which will be coded largely in 
ACL-DMAP. 
WORK COMPLETED 
0 LARGE ORDER STATIC ANALYSIS AND FSD DESIGN 
0 INTERFACE WITH SEMI-EMPIRICAL WEIGHTS PROGRAM 
0 LARGE-ORDER MODAL ANALYSIS 
0 FLUTTER ANALYSIS 
' 0 DIRECT MATRIX ABSTRACTION LANGUAGE (ACL-DMAP) 
0 TIME DOMAIN INTEGRATORS FOR MODAL MODELS 
WORK JUST INITIATED 
0 INTERACTIVE GRAPHIC MODELER FOR AERO-MESH 
0 STATIC AEROELASTIC PACKAGE 
0 FLUTTER OPTIMIZER 
0 COMPOSITE FINITE ELEMENTS 
0 DIRECT INTEGRATORS FOR LARGE FINITE-ELEMENT MODELS 
Figure 8 
1377 
THE ROAD AHEAD 
While much has been done, much remains to be done. We must complete work on our flutter 
optimization system. We must complete the static aeroelastic analysis and design 
package. We must develop a substructure capability for even larger finite-element 
models than we have currently analyzed. We must develop new procedures for global-local 
stress analysis to be used in analysis and design. We must initiate new work on air- 
craft structural loads. We must coordinate with new work being done in our Dynamics 
and Loads Research group on transonic flutter and aeroelastics. We must develop new 
procedures for fatigue life prediction, analysis, and design (Figure 9). 
COMPLETE WORK JUST INITIATED 
0 INTERACTIVE GRAPHIC MODELER FOR AERO-MESH 
0 STATIC AEROELASTIC PACKAGE 
0 FLUTTER OPTIMIZER 
0 COMPOSITE FINITE ELEMENTS 
0 DIRECT INTEGRATORS FOR LARGE FINITE-ELEMENT MODELS 
AREAS OF NEW WORK 
0 SUBSTRUCTURES, SUPERELEMENTS AND COMPONENT MODE SYNTHESIS 
0 GLOBAL/LOCAL STRESS ANALYSIS 
AIRCRAFT STRUCTURAL LOADS 
0 GEOMETRIC NONLINEARITY AND BUCKLING OPTIMIZATION 
0 
0 FATIGUE LIFE, PREDICTION, ANALYSIS AND DESIGN 
TRANSONIC/SUPERSONIC FLUTTER AND AEROELASTIC ANALYSIS AND DESIGN 
Figure 9 
1 1378 


Meeting the challenges with the Douglas Aircraft Company Aeroelastic Design Optimization Program (ADOP)

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