NASA/TM—2011–216423

Basis Function Approximation of Transonic 

Aerodynamic Influence Coefficient Matrix

Wesley W. Li and Chan-gi Pak

Dryden Flight Research Center, Edwards, California

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March 2011
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NASA/TM—2011–216423

Basis Function Approximation of Transonic 

Aerodynamic Influence Coefficient Matrix

Wesley W. Li and Chan-gi Pak

Dryden Flight Research Center, Edwards, California

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National Aeronautics and

Space Administration

Dryden Flight Research Center

Edwards, CA 93523-0273

March 2011
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1

Abstract

A technique for approximating the modal aerodynamic influence coefficients matrices by using basis 

functions has been developed and validated. An application of the resulting approximated modal 

aerodynamic influence coefficients matrix for a flutter analysis in transonic speed regime has been 

demonstrated. This methodology can be applied to the unsteady subsonic, transonic, and supersonic 

aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution 

can be found by the classic methods, such as rational function approximation, k, p-k, p, root-locus 

et cetera. The unsteady aeroelastic analysis for design optimization using unsteady transonic aerodynamic 

approximation is being demonstrated using the ZAEROTM flutter solver (ZONA Technology 

Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter 

speed prediction on an aerostructures test wing 2 configuration with negligible loss in precision in 

transonic speed regime. These results may have practical significance in the analysis of aircraft 

aeroelastic calculation and could lead to a more efficient design optimization cycle.

Nomenclature

AIC aerodynamic influence coefficients

AOA angle of attack

ATW aerostructures test wing

BFA basis function approximation

CEM central executive module

CFD computational fluid dynamics

DFRC Dryden Flight Research Center

FEM finite element model

FTF flight test fixture

g damping

G(x) constraints function

MDAO multidisciplinary design, analysis, and optimization

NASA National Aeronautics and Space Administration

TSD transonic small disturbance

i 1, 2, ..., m, where m is the number of mode shapes

k reduced frequency

m number of mode shapes

n number of basis functions

Q modal AIC matrix

Qij i-th row and j-th column element of the modal AIC matrix Q

Q modal AIC matrix basis functions

Q approximate modal AIC matrix

Qij i-th row and j-th column of the approximate modal AIC matrix Q

v velocity

XCG center of gravity in X direction

YCG center of gravity in Y direction

ZCG center of gravity in Z direction

βk

i modal participation factors of the k-th basis function on the i-th mode shape

βs

i modal participation factors of the s-th basis function on the i-th mode shape
2

βr

j modal participation factors of the r-th basis function on the j-th mode shape

φ structural mode shapes

φi i-th mode shape

Ψ structural mode shapes basis functions

ψ k k-th structural mode shapes basis functions

ω frequency

1. Introduction

Supporting the Aeronautics Research Mission Directorate guidelines, the National Aeronautics and 

Space Administration (NASA) Dryden Flight Research Center (DFRC) is developing an object-oriented 

multidisciplinary design, analysis, and optimization (MDAO) tool (ref. 1). This tool will leverage existing 

tools and practices, and allow the easy integration and adoption of new state-of-the-art software. At the 

heart of the object-oriented MDAO tool is the central executive module (CEM) as shown in figure 1. In 

this module, the user will choose an optimization methodology; provide starting and side constraints for 

continuous as well as discrete design variables and external file names for performance indices, which 

communicate between the CEM and each analysis module; submit script commands to prepare input data 

for each analysis code; execute analyses codes; compute performance indices using post-processor codes;

and compute an objective function and constraint values from performance indices. The structural 

analyses modules such as computations of the structural weight, stress, deflection, buckling, and flutter 

and divergence speeds have been developed and incorporated into the object-oriented MDAO framework.

Figure 1. Object-oriented multidisciplinary design, analysis, and optimization tool.

With the increased complexity of the configuration and the need to increase the fidelity of the 

nonlinear aerodynamic equations to be solved in transonic regime, the computational cost of the unsteady 

aeroelastic analysis required to generate aerodynamic influence coefficients (AIC) matrices for full 

aeroelastic response and design optimizations is very high (ref. 1). Also, these analyses are usually 
3

performed repeatedly to optimize the final design. Even though the computational cost may be reduced by 

the use of advanced algorithms and improved computer hardware processing speeds, these full 

aeroelasticity analyses cannot be incorporated effectively within a preliminary design and optimization 

environment.

The integration of traditional unsteady computational fluid dynamics (CFD) into MDAO involving 

disciplines such as aeroelasticity and aeroservoelasticity is, at present, a costly and impractical venture 

(ref. 1). At industry, aircraft designs using nonlinear CFD codes are mainly based on the manual trial and 

error approach, and this will considerably slow down the whole design procedure. As a result, there is 

considerable motivation to be able to perform unsteady aeroelastic calculations more quickly and 

inexpensively. The development of a MDAO tool to automate this process is needed to reduce time 

required for the early conceptual and/or preliminary design stages.

The results of the rapid computation of a modal AIC matrix and aeroelastic response calculations

using an approximated modal AIC matrix based on basis function approximation (BFA) has been 

previously validated and demonstrated successfully in subsonic speed regime to reduce the computation 

time. This approach reduces the modal AIC computation time using existing AIC matrices during 

optimization procedure by 80 percent (ref. 2). This report extends the study and validation of the 

application of the BFA from the subsonic speed regime to the transonic speed regime.

The transonic speed range introduces some particular flutter problems that are not commonly treated 

by classic linear flutter theory. Classic coupled torsion and bending wing flutter occurs when, with 

increase in dynamic pressure, the bending frequency increases, and the torsional frequency decreases until 

the separation between these two frequencies becomes sufficiently small to allow exchange of energy 

between the two corresponding modes of the structure. In contrast, for wings with attached flow at 

transonic speeds, the flutter mechanism is somewhat different. It is very important to determine the 

location of the shocks, which is known to have fundamental importance on the correct estimation of the 

unsteady aerodynamic loads, for the transonic flutter prediction. The capture of the shock is impossible 

using any linear code at transonic speed. The general aeroelastic problem involving the interaction 

between the aircraft structure and the unsteady transonic flow with the presence of shock waves can only 

be solved using CFD codes, based on transonic small disturbance (TSD) theory, Euler codes, or the 

complete Navier-Stokes equations. The order of complexity and computational costs increases from the 

level of the TSD theory to the Navier-Stokes theory that can accurately predict interactions between 

boundary layers and local shocks on typical aerodynamic surfaces. In this report, CAP-TSDv (ref. 3) code 

is used to capture the shock at steady state and flutter prediction.

This report presents the flutter results of the aerostructures test wing (ATW) 2 model without tip 

boom, in transonic speeds using BFA and traditional methods also called the direct method throughout 

the rest of the report. The MSC NASTRAN (MSC Software Corporation, Santa Ana, California) (ref. 4), 

CAP-TSDv, ZAERO ZONA6 linear code, and ZTAIC nonlinear code (ref. 5) (ZONA Technology 

Incorporated, Scottsdale, Arizona) are used in this study for computing the flutter boundary.

2. Basis Function Approximation

In linear algebra, a basis is a set of vectors that, in a linear combination, can represent every vector in 

a given vector space or free module, and such that no element of the set can be represented as a linear 

combination of the others (ref. 6). In other words, a basis is a linearly independent spanning set. Since 

mode shapes are orthogonal to each other and each mode has a natural frequency associated with the 

eigenvalue, structural mode shapes obtained from different mass or various stiffness configurations of the 

airplane can be used as the basis functions. These basis functions are comparison functions (ref. 7) since 

all the geometric and natural boundary conditions of the airplane are satisfied. The mode shapes of the 

airplane with any of the configurations in the design spaces can be used to generate the basis functions. 

Furthermore, each mode shape of the target configuration of the airplane can be approximated as a linear 

combination of a set of basis functions. This basis function approximation has been described in 

reference 2.
4

In general, a time-domain approach is more convenient for analysis tasks and nonlinear problems; on 

the other hand, a frequency-domain approach is better for design tasks and linear problems. Most modern 

CFD codes were developed using the time-domain approach, and therefore, the results of these CFD runs 

are usually the time-histories of pressure distributions, structural displacements et cetera. To use the 

modern CFD computation for our design optimization studies, we are interested in developing an 

approximate unsteady computational aeroelastic analysis in the frequency-domain to dramatically reduce 

the required computation time for optimization runs in the transonic speed regime. To accomplish these 

objectives, the BFA method in reference 2 can be extended and evaluated for approximation in transonic 

speeds. The BFA method requires the unsteady aerodynamics to be represented in the frequency domain. 

The flutter solution can be found by classic methods, such as rational function approximation, k, p-k, p, g 

method, and root-locus et cetera.

The computation of a modal AIC matrix and aeroelastic response calculations in transonic speeds will 

be discussed in this report, and the approximation process is outlined in the flowchart given in figure 2.

Figure 2. Flowchart of the flutter analysis module in the object-oriented MDAO tool.

In step 1, a set of representative basis functions, Ψ , is chosen with the intent to capture salient 

features of the modal responses the airplane is expected to encounter in the various design spaces. In step 

2, the steady state pressure distributions for the transonic speed analysis can be obtained by using a CFD 

code or wind tunnel test.

In step 3, a set of representative basis modal AIC matrices, Q , are computed corresponding to the 

representative basis functions defined in step 1 and steady pressure distributions obtained in step 2 at any 

Mach number and reduced frequencies. These basis modal AIC matrices are used as input for 

approximate modal AIC matrix calculation in step 5. These basis modal AIC matrices are computed only 

once and stored before starting optimization.

In step 4, during optimization, for a set of given structural mode shapes, φ , each mode shape is 

decomposed in a linear combination of the basis functions, Ψ . The i-th mode shape, φi , (i=1, 2, ..., m, 
5

where m is the number of mode shapes) is approximated through the use of a least squares fitting together 

with the following basis functions as shown in equation (1),

φi ≈ βk

i

ψ k

k=1

n

∑ (1)

where ψ k is the k-th basis function and a coefficient, βk

i

, is the modal participation factor of the k-th 

basis function on the i-th mode shape and n is the number of basis functions.

In step 5, an approximate modal AIC matrix, Q , is computed based on a basis modal AIC matrix, Q , 

and modal participation factors in equation (2),

Qij ≈ Qij = βs

i

βr

j

Q

sr

s=1

n

∑ r=1

n

∑ (2)

where Qij is the i-th row and j-th column element of the modal AIC matrix Q.

In step 6, the flutter analysis is performed using the approximate modal AIC matrix in equation (2). 

The flutter speed and frequency could be used for optimization.

3. Applications

In an effort to validate the BFA in transonic speed, a modified ATW2 model was chosen and the 

flutter results using approximate method were compared with the direct flutter results for both subsonic 

and transonic speed. To simplify the problem, especially for CFD mesh generation in CAP-TSDv, the 

wing tip boom is removed from this application. The original ATW2 (ref. 8) with wing tip boom is the 

test article actually designed, built, and tested at NASA DFRC as shown in figure 3. The ATW2 was 

clamped onto a circular plate, which was bolted to a mounting panel. The mounting panel was attached to 

the flight test fixture (FTF) (ref. 9). The FTF was installed on the centerline F-15B (McDonnell Douglas, 

now The Boeing Company, Chicago, Illinois) pylon in flight configuration.
6

Figure 3. Aerostructures test wing 2.

For the direct method, the flutter analysis of the target design configuration is performed using the 

ZAERO linear subsonic method (ZONA6) and nonlinear transonic method (ZTAIC) with the first 10 

structural modes and 16 reduced frequencies at 4 Mach numbers (Mach 0.80, 0.90, 0.95, and 0.98). The 

steady state pressure input data as the initial condition for the ZTAIC method used in the present analysis 

are obtained by using the CAP-TSDv code at all four Mach numbers. In order to prove the setup and the 

process of the ZTAIC method correctly, the flutter results from both the ZONA6 and ZTAIC methods are 

compared and used to verify the process at subsonic Mach numbers. At a subsonic Mach number like 

Mach 0.80, before any shock is formed, both the linear and nonlinear solver should predict similar flutter 

results. Once the process was verified, the ZTAIC method was used for the flutter prediction in the 

transonic speed regime.

The CAP-TSDv calculations were performed on a 150 × 80 × 80 points computational grid with 41 

points along each of the 57 spanwise chords on the wing as shown in figure 4. The steady pressure 

distributions obtained from the CAP-TSDv computations at Mach 0.80, 0.90, 0.95, and 0.98 all at 0.0 deg 

and 1.0 deg of angle of attack (AOA) are shown in figure 5 and 6; and the steady pressure differences 

between the upper and the lower surfaces under 1.0 deg AOA at the same Mach numbers are shown in 

figure 7. With 3-percent structural damping, the primary flutter speed and flutter frequency comparison of 

the ZONA6 and ZTAIC methods are summarized in figures 8(a) and (b). At the subsonic Mach number of 

0.80, the ZTAIC results are in close agreement with those of the ZONA6, as expected, since a transonic 

shock has not formed yet.
7

Figure 4. CAP-TSD wing surface mesh.

Figure 5. Steady pressure distributions on the upper surface using CAP-TSDv code (0-deg angle of 

attack).
8

Figure 6. Steady pressure distributions on the upper surface using CAP-TSDv code (1-deg angle of 

attack).

Figure 7. Steady pressure difference under 1-deg angle of attack using CAP-TSDv code.
9

Figure 8a. Flutter speed.

Figure 8b. Flutter frequency.

Figure 8. Flutter results for the ATW2 without tip boom (ZONA6 compared with ZTAIC).
10

The primary flutter mode is the first bending mode at Mach 0.80. However, the second bending mode 

becomes the primary flutter mode when the Mach number is 0.98. On the other hand, the flutter mode 

interchange does not happen with the ZONA6 computation. At Mach 0.98, the primary flutter mode with 

the ZTAIC computation is the second mode, 81 Hz, as shown in figures 8; while the primary flutter mode 

with the ZONA6 computation is the first mode, 55 Hz. This mode interchange may be caused by the local 

high-pressure zone near the center of the wing as shown in figures 5 and 6.

Based on the pressure differences in figure 7, the aerodynamic load distributions at Mach 0.80 and 

0.90 are fairly similar in shape, which are basically the subsonic distributions (that is high near the 

leading-edge area and low near the trailing-edge area). When the Mach number became 0.95, the area of 

the high aerodynamic loads was extended to the aft section of the wing-root chord, and an even higher 

load area was formed near the leading edge of the wing mid chord. At Mach 0.98, aerodynamic load 

distributions are changed drastically, and compared to the Mach 0.95 case, highly concentrated 

aerodynamic load distributions are observed from the leading edge of the wing tip chord to the trailing

edge of the wing root chord. From these observations, we may conclude that the transonic effects start to 

happen between Mach numbers of 0.90 and 0.95, and strong shocks are formed at Mach 0.98 as shown in 

figures 5 and 7. Probably the second bending mode in figure 9 can be easily excited by the concentrated 

aerodynamic load distributions at Mach 0.98.

Figure 9. V-g and V-ω plots for the ATW2 at Mach 0.98 using ZONA6 and ZTAIC methods.

In this ATW2 example, five different design configuration points in the parameter design space are 

selected for generating basis functions for the approximate flutter solution. The design variations used in 

this study are variations in wing thickness and stiffness that are effected by variations in the number and 

orientation of the composite laminate plies. These design variations are intended to capture salient 

features of the modal responses resulting from mass and stiffness variations in the design space. The 
11

structural mode shapes obtained from those five design configurations are used as the basis functions. The 

finite element model and the first three mode shapes of the target configuration are shown in figure 10. 

The total weight and center of gravity of the various fictitious and target design configurations are listed 

in table 1. The target configuration in this report is simply an arbitrary design point within the design 

space and is used as a check case.

Figure 10. The aerostructures test wing 2 without tip boom mode shapes.

Table 1. Summary of total weight and center of gravity location for the ATW2 without the tip boom 

model with different fictitious design configurations.

Fictitious configuration Config. 1 Config. 2 Target Config. 3 Config. 4 Config. 5

Total weight, lb 2.38 2.53 2.81 2.88 3.66 4.09

XCG, in 12.56 12.70 12.93 12.98 13.41 13.58

YCG, in -7.17 -7.24 -7.35 -7.38 -7.59 -7.67

ZCG, in -3.12 -2.95 -2.66 -2.59 -2.05 -1.85

First, 10 structural mode shapes for each design configuration (a total of 50 basis functions) are 

generated using MSC/NASTRAN. The natural frequencies of the various fictitious and target design 

configurations are listed in table 2. Next, the size of 50 by 50 basis modal AIC matrix corresponding to 

those 50 basis functions is computed for each of the 16 reduced frequencies at Mach 0.80, 0.90, 0.95, and 

0.98 using ZONA6 and ZTAIC. To solve the modal participation factors, the mode shapes of the target 

design configuration are fitted using the basis functions with a least squares method. Then the 

approximate modal AIC matrices are computed based on the modal participation factors. Finally, the 

approximate flutter boundaries can be solved by using any linear frequency domain flutter solution code 

and the approximated modal AIC matrices. It is important to point out that in order to provide meaningful 

comparisons of the direct method with the BFA method, the ZTAIC method is used to predict the flutter 

boundaries in this report. Given that the BFA approach can be used with any kind of CFD code, ZTAIC is 
12

selected for CFD simulation in this report due to limited resources. The focus in this report is to 

demonstrate that the BFA method can be used to predict adequate flutter speed and flutter frequency at 

transonic speed regime regardless the level of CFD code.

Table 2. Summary of natural frequencies (Hz) of the ATW2 without the tip boom model with different 

design configurations.

Mode Config. 1 Config. 2 Target Config. 3 Config. 4 Config. 5

1 23.08 22.77 22.06 21.88 20.01 19.13

2 85.61 85.76 84.85 84.49 79.34 76.50

3 120.67 126.71 131.30 131.62 126.99 122.32

4 147.99 168.97 194.28 194.17 183.32 177.44

5 193.02 195.30 194.80 198.63 223.60 228.76

6 261.81 277.74 291.27 292.60 286.04 277.39

7 315.54 322.42 323.05 322.07 304.26 295.76

8 399.82 423.26 442.11 443.74 434.38 426.50

9 459.62 472.40 472.64 470.78 442.33 431.65

10 560.99 582.69 595.67 595.64 572.74 561.21

As shown in table 3, using ZTAIC, both methods predicted similar flutter speed and flutter frequency

at Mach 0.80, 0.90, 0.95, and 0.98. The speed versus damping, V-g, and speed versus frequency, V-ω, 

curves from the direct and BFA methods at Mach 0.90, 0.95, and 0.98 are given in figures 11, 12, and 13. 

These figures indicate that an interchange between the first bending and the second bending flutter modes 

results when Mach number increased from 0.90 to 0.98. At Mach 0.90, the first bending mode, 56 Hz, is 

predicted as the primary flutter mode. This flutter mode became the hump mode at Mach 0.95 and stable 

mode at Mach 0.98. Meanwhile, the stable second mode at Mach 0.90 became unstable at Mach 0.95 and 

later became primary flutter mode at Mach 0.98. With 3-percent structural damping, the primary flutter 

speed and flutter frequency are summarized in table 3, and figures 14(a) and (b). The largest percentage 

error of flutter speed and frequency is 1.60 percent at Mach 0.98 and 1.27 percent at Mach 0.80 

respectively. Although a different CFD code may generate different unsteady AIC and predict different 

flutter results, the important point is that the results of the BFA method is following the one predicted by 

the direct method. Overall, the resulting flutter boundaries prediction using the BFA method is well 

compared with the direct method in transonic speed. The flutter boundaries prediction based on the BFA 

method is essentially identical to those using direct solution on the target design configuration.

Table 3. Summary of the ATW2 without the tip boom model flutter results comparison for BFA and the 

direct method using ZTAIC. 

Flutter speed, KEAS Flutter frequency, Hz Mach

BFA Direct Error BFA Direct Error

0.80 621.77 622.94 -0.19% 59.85 60.62 -1.27%

0.90 612.39 612.38 0.00% 56.56 56.56 0.00%

0.95 641.94 640.68 0.20% 53.63 54.01 -0.70%

0.98 557.56 566.60 -1.60% 81.23 81.10 0.16%
13

Figure 11. V-g and V-ω plots for the ATW2 at Mach 0.90 using direct ZTAIC and BFA ZTAIC methods.

Figure 12. V-g and V-ω plots for the ATW2 at Mach 0.95 using direct ZTAIC and BFA ZTAIC methods.
14

Figure 13. V-g and V-ω plots for the ATW2 at Mach 0.98 using direct ZTAIC and BFA ZTAIC methods.
15

Figure 14a. Flutter speed.

Figure 14b. Flutter frequency.

Figure 14. Flutter results for the ATW2 without tip boom.
16

Table 4 lists the computational cost of computing the modal AIC matrices with given structural mode 

shapes at one Mach number and 16 reduced frequencies using the direct method and the BFA method. 

The size of the ATW2 model is about 260 degrees of freedom. The computational cost of generating the 

modal AIC using the direct method and the BFA method is about 43 s and 8 s using ZONA6, and 49 s 

and 8 s using ZTAIC respectively in this ATW2 example. As a result, the BFA approach reduces the 

modal AIC computation time using existing AIC matrices by 80 percent. The comparison was done on an 

Intel® CoreTM2 2.80GHz CPU computer (Intel Corporation, Santa Clara, California). The computational 

cost of generating the basis functions before the parametric study is not included in the comparison.

Table 4. Computational cost comparison of BFA and the direct method.

ZONA6 ZTAIC

Direct method elapsed time 43 s 49 s

BFA method elapsed time 8 s 8 s

4. Concluding Remarks

Many of the most interesting engineering applications are very complex systems, such as aeroelastic 

analyses and optimization in which the analysis is achieved through the use of expensive and time 

consuming numerical simulations. The processing times on even the most powerful workstations may be 

too expensive. A technique for approximating the modal AIC matrix by using basis function 

approximation has been applied in transonic speed regime, and a process for using the resulting AIC 

matrix in aeroelastic analyses has been developed and demonstrated. The technique presented has been 

shown to offer consistent flutter speed prediction on an ATW2 configuration with insignificant small loss 

in accuracy, less than 1.6-percent error in the ATW2 test case, and the results are essentially identical to 

those using direct solution. The BFA approach has been demonstrated successfully in the transonic speed 

regime to reduce the computation time of generating modal AIC using basis functions. The BFA 

approach has shown an 80-percent reduction in the modal AIC computation time using existing AIC 

matrices. These results may have practical significance in the aeroelastic analysis of an aircraft and could 

lead to a more efficient design optimization cycle in a transonic speed regime. For the design optimization 

and parameter study, hundreds or even thousands of analyses have to be performed. Creation of an 

approximated modal AIC matrix using basis functions allows for analysis, and a parameter optimization 

cycle of the system in minutes, rather than hours of computer time on common personal computers.
17

References

1. Pak, Chan-gi, and Wesley Li, “Multidisciplinary Design, Analysis and Optimization Tool 

Development Using a Genetic Algorithm,” Proceedings of the 26th Congress of International Council 

of the Aeronautical Science, Anchorage, Alaska, September 14–18, 2008.

2. Pak, Chan-gi, and Wesley W. Li, “Application of Approximate Unsteady Aerodynamics for Flutter 

Analysis,” AIAA–2010–3085, Proceedings of the 51st AIAA Structures, Structural Dynamics, and 

Materials Conference, Orlando, Florida, April 12–15, 2010.

3. Batina, John T., David A. Seidel, Robert M. Bennett, and Samuel R. Bland, Preliminary User’s 

Manual for CAP–TSD (version 1.0), 1988.

4. MSC/NASTRAN Quick Reference Guide Version 69, The MacNeal Schwendler Corporation, 1996.

5. ZAERO Theoretical Manual Version 8.3, ZONA Technology, Inc., 2007.

6. Kahaner, David, Cleve Moler, and Stephen Nash, Numerical Methods and Software, Prentice–Hall, 

Englewood Cliffs, New Jersey, 1989.

7. Meirovitch, Leonard, Analytical Methods in Vibrations, The Macmillan Company, New York, 1967.

8. Lung, Shun-fat, and Chan-gi Pak, “Updating the Finite Element Model of the Aerostructures Test 

Wing Using Ground Vibration Test Data,” AIAA–2009–2528, Proceedings of the 50th

AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, Palm 

Springs, California, May 4-7, 2009.

9. Richwine, David M., “F–15B/Flight Test Fixture II: A Test Bed for Flight Research,” NASA–TM–

4782, 1996.
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

1. REPORT DATE (DD-MM-YYYY)

01-03-2011

2. REPORT TYPE 

Technical Memorandum

 4. TITLE AND SUBTITLE

Basis Function Approxmiation of Transonic Aerodynamic Influence 

Coefficient Matrix

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Li, Wesley W. and Pak, Chan-gi

 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

 8. PERFORMING ORGANIZATION

 REPORT NUMBER

H-3082

10. SPONSORING/MONITOR'S ACRONYM(S)

NASA

13. SUPPLEMENTARY NOTES

Li and Pak, NASA Dryden Flight Research Center. Also presented as meeting paper 072 at the 27th Congress of the International 

Council of the Aeronautical Sciences, Nice, France, September 19-24, 2010.

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified -- Unlimited

Subject Category 01 Availability: NASA CASI 443-757-5802 Distribution: Standard

19b. NAME OF RESPONSIBLE PERSON

STI Help Desk (email: help@sti.nasa.gov)

14. ABSTRACT

A technique for approximating the modal aerodynamic influence coefficients matrices by using basis functions has been developed 

and validated. An application of the resulting approximated modal aerodynamic influence coefficients matrix for a flutter analysis in 

transonic speed regime has been demonstrated. This methodology can be applied to the unsteady subsonic, transonic, and supersonic 

aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution can be found by the 

classic methods, such as rational function approximation, k, p-k, p, root-locus et cetera. The unsteady aeroelastic analysis for design 

optimization using unsteady transonic aerodynamic approxmination is being demonstrated using the ZAERO flutter solver (ZONA 

Technology Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter speed prediction 

on an aerostructures test wing 2 configuration with negligible loss in precision in transonic speed regime. These results may have 

practical significance in the analysis of aircraft aeroelastic calculation and could lead to a more efficient design optimization cycle.

15. SUBJECT TERMS

Aerodynamic influence coefficient matrix, Basis function approximation, Flutter analysis, Least squares method, Multidisciplinary 

design optimization

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(443) 757-5802

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NASA/TM-2011-216423

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