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Interfacing External, High Order Aerodynamicsinto MSC/NASTRAN for Aeroelastic Analyses

Brent WhitingLead Engineer, Loads Integrated Product Team

Boeing Defense and Space Group

Douglas J. NeillSr. Staff Engineer

Aeroelasticity and Design OptimizationThe MacNeal-Schwendler Corporation

Abstract

In the design of aircraft, it is important to have an accurate simulation ofboth the structural characteristics and the aerodynamic characteristics of thevehicle. For static aerodynamic loads, MSC/NASTRAN uses unsteadyaerodynamics at zero reduced frequency. To utilize the accurate structuralrepresentation of MSC/NASTRAN in the computation of aeroelastic loads, itis desirable to incorporate aerodynamic and aeroelastic data into theMSC/NASTRAN solution sequence that better represents the actualgeometry of the vehicle. This entails importing data that replaces oraugments the unsteady data. The Boeing Company, in conjunction with MSCpersonnel, has prototyped an interface procedure that allows the staticaeroelastic loads computations in MSC/NASTRAN to use the rigidaerodynamic forces and aeroelastic corrections from the A502 High OrderPanel Method.

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Introduction

In the design of aircraft, it is important to have an accurate simulation ofboth the structural characteristics and the aerodynamic characteristics of thevehicle to produce accurate trimmed loads. While MSC/NASTRAN has longhad a static aeroelastic analysis capability (Ref. 1), it utilizes the embeddedunsteady aerodynamic methods (e.g., Doublet-Lattice) at zero reducedfrequency. While these methods are very useful for aeroelastic analyses atthe conceptual and preliminary design stage, the geometric simplificationsnecessary for the unsteady methods can make it difficult to obtain highfidelity aerodynamic forces for static trim, either because the actual geometrycannot be modeled accurately or because viscous effects dominate the flowfield. In both cases, model correction factors need to be used; however, thegeometric assumptions of the model can make it difficult to define and createthe correction data.

In particular, the unsteady model in MSC/NASTRAN is comprised of planarlifting surfaces and slender bodies. The lifting surface edges must be alignedalong the far field streamlines (not inclined to the onset flow) and thesesurfaces produce only normal forces and moments about one axis. The axis ofthe slender bodies must also be aligned in the freestream direction and thesebodies produce only forces and moments in one or two axes. For manyaircraft, however, there are important aerodynamic forces arising fromcambered wings, bodies and nacelles that need more geometric fidelity onwhich to map high order potential method data or even nonlinearaerodynamic data to obtain good trimmed forces. This need encompasses alarge speed range in that low speed configurations will typically require largecontrol surface deformations to trim the vehicle while cruise flight will causetransonic flows to occur that can affect the trimmed load distributions. Toprovide improved aerodynamics for static aeroelastic analysis, it wasdesirable to use geometric and pressure data from Boeingís A502 panel code(Ref. 2) for the aerodynamics.

The purpose of the work documented in this paper is to create a system foraeroelastic analysis that provides for higher fidelity aeroelastic loads. Thesystem is automated to reduce the level of expertise required to obtain goodaeroelastic loads and can be extended to include unsteady aerodynamic data.The approach is to remove the geometric restrictions imposed by theembedded unsteady methods so as to allow import of high fidelity geometrieswith high fidelity aerodynamic data into the static aeroelastic analysissolution sequence of MSC/NASTRAN. To that end, a system has been builtusing MSC/PATRAN V6, MSC/NASTRAN V69.1, the A502 panel method anda small translator program to allow the use and display of A502 data in theStatic Aeroelastic (Sol 144) solution sequence.

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Several important issues have been addressed in this development. First isthe removal of the geometric restrictions/assumptions from the aerodynamicgeometry. This leads to the most critical area that has been investigated: themeans to apply the existing splining (aero/structural coupling) methods tothe enhanced aerodynamic geometry. The existing splining methods areinherently "two dimensional" in that they all assume that the aerodynamicforces are normal to one or two orthogonal planes and that the aerodynamicgeometry consists of collections of points on a plane or along an axis. Finally,the problem of data visualization has been addressed to allow the analyst tobetter understand the behavior of the aeroelastic system and, hopefully, findand correct modeling problems more quickly.

Aerodynamic Modeling Approaches

The purpose of this effort is to calculate the aeroelastic loads on complexblended wing/body fighter configurations. The configurations that thecurrent MSC/NASTRAN doublet lattice aero package can represent consist offlat lifting surfaces and interference bodies as shown in Figure 1. Toaccommodate the blended wing/body configurations it was necessary tocouple high order panel method aerodynamics with MSC/NASTRAN to obtainpressure distributions for these complex configurations. The A502aerodynamic method models the aircraft configuration as a outer-mold-linesurface mesh divided into rectangular or triangular ìnetworksî of points. Ithas the capacity to apply a number of boundary conditions to the networksincluding flow tangency and mass flux to model important flow phenomenonwithin the realm of potential aerodynamics.

In using a code like A502 that can model a complex surface geometry, it wasnecessary to accommodate geometric generality within MSC/NASTRAN’saeroelastic solutions. And, having done so, there exists the potential tointroduce aerodynamics from other sources such as Euler or Navier-Stokessolutions or even wind tunnel data. All that is required is to provide theappropriate surface mesh and data through the interfaces defined in thisdevelopment.

The most critical issue in allowing complex aerodynamic geometries inMSC/NASTRAN is that the aerodynamic/structural coupling (splining)methods in MSC/NASTRAN have taken advantage of the simplifiedaerodynamic geometries of the unsteady methods. To make use of thesemethods, the aerodynamic geometry had to be made to appear as a series ofplanar facets to which Harderís Infinite Surface Spline (SPLINE1) or theBeam Spline (SPLINE2) could be separately applied. To accomplish this, theA502 code was modified to accept user-defined components that map theregular networks into fictitious “components” that approximately satisfy the

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SPLINE1 geometric restrictions. The A502 code then transforms the surfaceforces along the local (curved surface) normal to the “component” normal.These data then nearly satisfy the existing MSC/NASTRAN geometricrestrictions. The only remaining restriction that MSC/NASTRANaerodynamics methods have is that the components cannot be inclinedrelative to the onset flow. This restriction, however, comes from the unsteadyaerodynamics itself and its violation does not impact this application sincethose portions of MSC/NASTRAN are completely bypassed.

The A502 program was further modified to calculate the aerodynamic extrapoint (control surface deflections, onset angles and onset rates) forceincrements and the aeroelastic correction terms that MSC/NASTRAN needsfor aeroelastic analysis and to produce a file containing these data along withthe “component” geometry. The base force distribution (force due to onsetflow over the body) is computed in the standard A502 manner. Theincremental force distributions for pitch, yaw, pitch rate, roll rate, yaw rateand control surface deflections were calculated by A502 by applying theincrements as transpiration boundary conditions.

The aerodynamic forces arising due to structural deformations are calculatedin MSC/NASTRAN by using an AIC matrix. The AIC matrix is initiallycomputed in the aerodynamic domain and transformed to the structuraldomain using the same splining transformations that are used for theaerodynamic forces. The AIC that A502 computes is the incremental force forall points on the airplane due to a unit translational perturbation of anaerodynamic grid point. To be consistent with the splining geometry, it wasnecessary to define the direction of the deflection as being normal to theìcomponentî plane defined previously. As with the base force increments, theAICís output forces are also aligned along the MSC/NASTRAN “component”normals. By virtual work principles, the deflections of the structure can thenbe transferred to the aerodynamic mesh using the transpose of the forcespline. By premultiplying by the AIC matrix, one obtains incremental forceson the aerodynamic components due to structural deformation. These forcesare splined to the structure in the same manner as the base forces and otherforce increments.

Splining

The MSC/NASTRAN SPLINE1 (Harder Desmarais infinite plate spline) andSPLINE2 (beam spline) were used to transfer the aerodynamic forces fromthe aerodynamic grid to the structural model and to map the structuraldeflections back to the aerodynamic grid. Each of the fictitious A502 “planarcomponents” was mapped to a set of structural nodes using a SPLINE1 orSPLINE2 entry and a SET entry.

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For accurate aerodynamic loads on blended wing/bodies, it is desired to get afull three-dimensional mapping between the aerodynamic forces and thestructural deflections. However, the infinite plate spline allows the mappingof a scalar component of force normal to an infinite, planar, plate. Thisnecessitated dividing the A502 model’s collection of networks into smallcomponents (patches) that are nearly planar in their local region. The planefor the SPLINE1 is then defined as parallel to the plane of the component. Inusing this approach it was hoped that the majority of the load will betransferred and little load will be “lost”. Figure 2 shows an example of theaerodynamic grid points of a A502 model. Figure 3 shows a representation ofthe planes that are defined for the SPLINE1 card to map the forces anddeflections.

Clearly, this approximation introduces a modeling requirement on the user ofthis procedure. In particular, for each component (patch) the modeler mustdecide which direction contains the dominant effects for the analyses that areto follow. Clearly, the activity of model development is complicated in thatthe number of component patches as well as their orientation must beselected based on the nature of the dominant flow field effects AND theexpected aeroelastic effects. In the procedure that is being described, thisremains the most critical modelling issue for aeroelastic analysis.

The Bypass Procedure

To bypass the MSC/NASTRAN computation of aerodynamic matrices andgeometries, a number of datablocks had to be created using an external tooland brought into the MSC/NASTRAN database during the run. A simple userinterface was developed to allow selection of a multiplicity of externalaerodynamic models (e.g., multiple Mach numbers); however, this interactionneeded to be automated to a large degree to allow the bypass procedure to bedisseminated into a loads organization that is not familiar with DMAP or theparticulars of the aerodynamics code. Figure 4 shows the flow of theprocedure. The basic idea of the bypass procedure is to call an externalprogram to compute or provide the normal MSC/NASTRAN aerodynamicdata with the structural geometry supplied from the MSC/NASTRAN inputstream and the aerodynamic geometry, force and AIC data from the A502program.

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Legend

Procedures

Data Files

expnast.DAT

mna502_alter.dat

MSC/PATRAN

expa502.DAT

A502

expa.a502expa.aic.op4

exp_geom.op2

ISHELL forkMNASTA502

MNA502

MSC/NASTRANSOL 144 or 200

Solutionbgpa_exp.op2expa.aic.op4aicfrc_exp.op4

expnast2.F06

expa.a502 MSC/PATRAN aero.pib

Figure 4. External Loads Interface Data Flow. Items within the Dotted Lineare Run Using the Shell Procedure

The A502 code was modified as described in the preceding sections to producethe files denoted in Figure 4 as expa.a502 and, optionally, expa.aic.op4.Additional A502 input sections were created to define the mapping betweenA502 networks and the planar components needed to satisfy the spliningrestrictions of MSC/NASTRAN. The file expa.a502 contains the componentgeometry, the base force and force increment data (as well as the set ofaerodynamic extra points for which force increments are defined) and theAIC matrix. (An option in the expa.a502 file format allows the AIC matrixto be stored in a binary file expa.aic.op4 as an alternative to save diskspace.)

A DMAP alter was developed that utilized the ISHELL process spawningDMAP module to invoke the program that creates the necessary datablocks.The MSC/NASTRAN execution prior to the ISHELL fork creates thenecessary structural geometry datablocks on output2 files and the ISHELLprocess reads those data and the associated A502 files to create a set pair ofoutput2/output4 files that contain the necessary matrix and tabular data tocomplete the execution of the aeroelastic analysis. The ISHELL process is

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executed once for each Mach number in the TRIM subcases of theMSC/NASTRAN run.

The user of MSC/NASTRAN must supply some information to drive theprocess. Notably, in the implementation discussed here, the user must supplya dummy CAERO1 entry for each of the planar components. While thegeometry of these entries is not important, the number of degrees of freedomassociated with each component must be specified by the user. Thisrequirement comes from the fact that the SPLINE1/2 entries that the usermust also supply will refer to the external grid point identification numbersof the input aerodynamic model during input data checking operations.Consequently, the input model must be correct in supplying the externalaerodynamic grid point idís. The actual geometry of these data isunimportant, since the ISHELL process brings in the correct geometry priorto the computation of the spline transformation matrices.

The user of this procedure does NOT need to supply dummy aerodynamicextra points. The ISHELL procedure creates the MSC/NASTRAN XLISTdatablock from the data supplied by A502. This simplifies the definition ofthe aerodynamic model to just the dummy CAEROi entries, the SPLINEidata and the TRIM entries needed to specify the desired maneuver trimanalyses. As many TRIM entries can be input as are desired with theappropriate unique Mach numbers causing an ISHELL invocation to retrievethe appropriate data. It is assumed in the procedure that the modelgeometry is invariant with Mach number over a given MSC/NASTRANexecution (just as is the case in regular MSC/NASTRAN aeroelasticanalyses).

Example

As a test case for the bypass procedure, a simplified light aircraft (thestructural model shown in Figure 5 and the A502 aerodynamic model inFigure 6) was chosen. This example exercises several important features ofthe developed bypass procedure. The wing of this aircraft, for example, iscomprised of spanwise-varying airfoil sections providing for aerodynamictwist. In addition, there is structural twist. The result, from the standpoint ofDoublet-Lattice modelling, is that the W2GJ Direct Matrix Input is requiredto match the pressure distribution at zero angle of attack. Also, the fuselageis heavily cambered and not particularly "slender," so that slender bodymodeling is very difficult.

Due to the heavy body camber, other modelling methods (for example,treating the fuselage as a set of cruciform lifting surfaces) might be appliedusing correction factors (WKK input) to adjust for the resultingoverprediction of the total vehicle lift. These additional data items require

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extra work to produce and typically require other aerodynamic data toprovide the "target" answers against which the corrections are confirmed.Furthermore, the resulting model is only appropriate for certain datacomponents. For example, a simplistic (uncorrected) Doublet Lattice model isshown in Figure 7 that is appropriate for symmetric maneuver analysis. Itcannot predict any side forces and cannot model the vertical tail forces.Another model would be needed to perform antisymmetric or asymmetricanalyses.

The A502 model of Figure 6, however, is a reasonably faithful representationof the geometry and includes the camber effects on the wing and body andcan model (to the limit of potential flow theory) the wing/body interferenceeffects. Further, this model is appropriate for both symmetric andantisymmetric maneuver and asymmetric maneuver requires only mirroringthe geometry.

Some representative results for a Mach 0.6, 2G symmetric pullup are shownin Table 1 for the Doublet Lattice model and the A502 model. Similar wingsplines are used, but the Doublet Lattice model uses beam splines to load thefuselage keel, while the A502 model uses surface splines to an appropriateset of points on the skin.

CL0 CMo CLα /rad CMα/rad CMδ/rad αtrim(deg) δ trim(deg)

Rigid Flex Rigid Flex Rigid Flex Rigid Flex Rigid Flex Rigid Flex Rigid Flex

A502 0.398 0.402 -0.942 -0.412 5.063 2.383 -9.30 -11.46 -2.630 -3.058 -3.250 -2.895 -8.790 -8.03

DLM 0.0 0.0 0.0 0.0 5.313 5.984 6.952 5.883 -2.382 -2.121 0.268 0.252 1.158 1.106

As can be seen in the table of results, the answers obtained are quitedifferent for the trim settings and, therefore, the applied, trimmed load.Obviously, the DLM results are not the best that can be obtained, since noattempt was made to match the intercept values by modelling the wingcamber effects. Consequently, the center of pressure for the configuration isnot correctly predicted by the flat plate model and the resulting trim settingsare different. The magnitudes and signs of the rigid slopes, however, arereasonably comparable while the elastic effects are profoundly different forthe two modeling approaches. The elastic differences in the lift slope arisefrom the elastic corrections of the overly effective lifting surfacerepresentation of the body attached to a relatively small number of fuselagekeel points. The moment differences arise from a combination of the spliningeffects and the lack of agreement in the center of pressure.

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To evaluate the accuracy of the A502 splining technique, the overallequilibrium of the two mapped loads was checked. A perfect spline shouldresult in the exact transfer of each of the six components of rigid loadbetween the two models. In this application, however, the two dimensionalnature of the MSC/NASTRAN splines make it extremely difficult to captureall six components. In this model, the splining attempted to capture theimportant axes of lift and pitching moment. The side force was intended to bereasonable and the other axes were ignored. The following table shows thedifference in force for the A502 base solution (thickness and camber effects inthe onset flow) and for the angle of attack perturbation solution. The liftforces are captured to within 2.6% to 5.6%. Roughly 60% of the total sideforce is captured. The drag, on the other hand, is off by as much as 3,200%.It is obvious that a more refined faceting is needed to apply the actual airloads from A502 to this configuration:

∆ Fx ∆ Fy ∆ Fz %∆ Fx %∆ Fy %∆ FzThickness camber -97.09 91.86 -69.19 3246.13 4.96 -2.68α -64.55 2215.11 -1815.52 155.88 40.00 -5.53

Qualitative evaluations were performed using graphical tools developedunder this project as an MSC/PATRAN PCL library. This library allows thedisplay of the flexible pressure distribution and the rigid increments on theaerodynamic geometry. Further, the deformed shape of the aerodynamicmodel can be presented for the A502 models (see Figure 8). These graphicaltools were invaluable in locating a number of problems in the FE model inwhich the attachment of the horizontal stabilizer was artificially weak andaccentuated the elastic effects on the tail to an unreasonable degree. Whilethe MSC/NASTRAN results with the preliminary FE model were clearlyincorrect, it was the graphical tool that clearly identified the tail as theproblem area.

Conclusions

A system for aeroelastic analysis that provides for higher fidelity aeroelasticloads has been developed by coupling the aerodynamics and aerodynamicgeometry from the A502 panel code to MSC/PATRAN V6 andMSC/NASTRAN V69.1 for Solutions 144 and 200. The system is automatedto reduce the level of expertise required to obtain good aeroelastic loads andcan be extended to include unsteady aerodynamic data.

Several important issues have been addressed in this development. First isthe partial removal of the geometric restrictions/assumptions from theaerodynamic geometry. This leads to the most critical area that has beeninvestigated: the means to apply the existing splining (aero/structural

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coupling) methods to the enhanced aerodynamic geometry. The existingsplining methods are inherently "two dimensional" in that they all assumethat the aerodynamic forces are normal to one or two orthogonal planes andthat the aerodynamic geometry consists of collections of points on a plane oralong an axis. Finally, the problem of data visualization has been addressedto allow the analyst to better understand the behavior of the aeroelasticsystem and, hopefully, find and correct modeling problems more quickly.

Using a simple, light aircraft as an example, the system was exercised forstatic aeroelastic analysis. The graphical tools in MSC/PATRAN helped inthe refinement and error correction in the structural FE model and in theevaluation of the splining accuracy. A very simplistic DLM model was alsodeveloped to compare levels of effort and relative qualities of answers.Reasonable results were obtained using the A502 bypass procedure withoutthe need to resort to the MSC/NASTRAN techniques of "correction factors"that are clearly required by the sample (very simplistic) DLM model.

The splining problem was examined and it has been found that, by carefulsubdivision of the A502 model, reasonably accurate force transfer can beobtained. It is by no means, however, a simple task. Subdividing the modelinto 17 regions for the purpose of doing symmetric trim analysis resulted inloads losses on the order of 5%. Clearly, more subdivisions are needed toaccurately capture the total load. In the long term, a different spliningcapability is needed in MSC/NASTRAN that does not require faceting theA502 model. It would need to transfer the full three dimensional airloadfrom the surface mesh to the structural model.

Acknowledgements

The work presented in this paper was performed by The MacNeal-Schwendler Corporation under contract to the Boeing Company. In additionto the authors, a number of people at MSC and at Boeing were involved inthe development and testing of this procedure. In particular at Boeing, Mr.Michael Epton was responsible for the changes to the A502 software. AtMSC, Dr. Erwin Johnson was the project lead with Mr. Dale WongK theprincipal developer of the MNASTA502 bypass procedure workhorse. Finally,Mr. Keane Barthenheier was the author of the MSC/PATRAN PCL libraryfor aerodynamic model visualization.

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References

1. The MacNeal-Schwendler Corporation, Los Angeles, CA, MSC/NASTRANAeroelastic Analysis User’s Guide, Version 68, October, 1994.

2. Epton, Michael A. and Magnus, Alfred E., PANAIR, A Computer Programfor Predicting Subsonic and Supersonic Potential Flows About ArbitraryConfigurations Using a High Order Panel Method, Volume I, TheoreticalManual, Version 3.0, NASA-CR-3251, Revision 1, January, 1992.

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Figure 1: Representative DLM Full Vehicle Model

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Figure 2 A502 Model Mesh Points

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Figure 3 Sample Facets on A502 Mesh

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Figure 5: Structural Model

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Figure 6: A502 Model

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MSC/PATRAN Version 6.2 08-Sep-97 15:32:32

DEFORMATION: SYMMETRIC FLIGHT CONDITIONS, A50, Static Subcase: Displacements, Translational -MSC/NASTRA

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Figure 7: Doublet Lattice Model with Trimmed Structural Deformations

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MSC/PATRAN Version 6.2 08-Sep-97 14:44:03

FRINGE: A502 F06 Results, File: freea502.f06, A502_F06_resdat: Aero Load, Point Forces (VEC-MAG) -PATRAN 2.5

DEFORMATION: A502 F06 Results, File: freea502.f06, A502_F06_resdat: Aero Deformation, At Aero Points -PATRAN 2.5

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Figure 8: Trimmed Deformations on the A502 Model with Force Fringes