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NASA / CR-1998-206947
Euler Technology Assessment for
Preliminary Aircraft Design-Unstructured/Structured Grid NASTD
Application for Aerodynamic Analysis of
an Advanced Fighter/Tailless
Configuration
Todd R. Michal
Boeing Company, St. Louis, Missouri
National Aeronautics and
Space Administration
Langley Research Center
Hampton, Virginia 23681-2199
Prepared for Langley Research Centerunder Contract NAS1-20342
March 1998
Available from the following:
NASA Center for AeroSpace Information (CASI)
800 Elkridge Landing Road
Linthicum Heights, MD 21090-2934
(301) 621-0390
National Technical Information Service (NTIS)
5285 Port Royal Road
Springfield, VA 22161-2171
(703) 487-4650
TABLE OF CONTENTS
SUMMARY ..................................................................................................... 2
INTRODUCTION ........................................................................................... 2
APPROACH .................................................................................................... 4
GRID GENERATION .............................................................................................................. 4
FLOW SOLUTION METHODOLOGY AND PERFORMANCE CHARACTERISTICS ......................... 6
RESULTS ......................................................................................................... 8
BASELINE CONFIGURATION ................................................................................................. 8
CAMBERED WING CONFIGURATION .................................................................................. 10
LATERAL-DIRECTIONAL CHARACTERISTICS ...................................................................... 10
CONTROL DEVICES ............................................................................................................ 10
CONCLUSIONS ........................................................................................... 11
ACKNOWLEDGEMENTS ......................................................................... 13
REFERENCES .............................................................................................. 13
FIGURES ....................................................................................................... 15
SUMMARY
This study supports a NASA project aimed at determining the viability of using
Euler technology for preliminary design use. The primary objective of this study was to
assess the accuracy and efficiency of the Boeing, St. Louis unstructured grid flow field
analysis system, consisting of the MACGS grid generation and NASTD flow solver codes.
Euler solutions about the Aero Configuration/Weapons Fighter Technology (ACWFT)
1204 aircraft configuration were generated. Several variations of the geometry were
investigated including a standard wing, cambered wing, deflected elevons, and deflected
body flap. A wide range of flow conditions, most of which were in the non-linear regimes
of the flight envelope, including variations in speed (high subsonic/transonic, supersonic),
angles of attack, and sideslip were investigated. Several flowfield non-linearities were
present in these solutions including shock waves, vortical flows and the resulting
interactions. The accuracy of this method was evaluated by comparing solutions with test
data and Navier-Stokes solutions. The ability to accurately predict lateral-directional
characteristics and control effectiveness was investigated by computing solutions with
sideslip, and with deflected control surfaces. Problem set up times and computational
resource requirements were documented and used to evaluate the efficiency of this
approach for use in the fast paced preliminary design environment.
The use of unstructured grids was found to significantly decrease the cycle time of
NASTD applications primarily through a reduction in grid generation time. The efficiency
and robustness of this method, while still too slow for generating an entire aerodynamic
database, are sufficient to provide data at a large number of points across the flight
envelope. The accuracy was generally sufficient for preliminary design use up to moderate
angles of attack (~15 degrees). The prediction of aerodynamic effects due to control
surface deflections were of mixed accuracy. Aerodynamic predictions of the elevon
control effectiveness and the lateral-directional characteristics due to asymmetric control
deflections were accurately predicted while the control effectiveness in pitch was
consistently over-predicted. Less accurate aerodynamic predictions were obtained for
control devices that generate a large amount of wake like separation such as the body flap.
Euler technology has strong relevance to preliminary-design applications. This
technology provides a means of predicting non-linear aerodynamic effects that previously
could only be obtained in the wind tunnel. This study has indicated that an un-exploited
potential exists for development of a lateral-directional design tool. However, further work
is needed to determine the parts of the flight envelope where this technology should or
should not be used. Additional work may also be required to develop empirical calibration
for some applications. The greatest benefit of this technology will be realized when it is
tied to advances in multi-disciplinary design tool development.
INTRODUCTION
Over the past decade, great strides have been made in the development of
preliminary design tools in the airplane radar signature and structural analysis disciplines.
In contrast, aerodynamic analysis in preliminary design continues to rely primarily on
linear tools developed several decades ago. The limitations of these methods are becoming
increasingly apparent with the advent of low observable and unmanned aircraft technology.
2
Thesetechnologieshaveled to non-traditionalvehicleshapesand control surfacedevicesthat exhibit highly non-linearaerodynamicbehavior. Aerodynamictools basedon linearaerodynamicmethodsare inadequatefor thesetypesof aircraft,particularly at theedgesofthe flight envelopesuch as high anglesof attack. For such applications,wind tunneltestingand/ornon-linearanalysisis required.
Until recently, Computational Fluid Dynamics (CFD) Euler or Navier-Stokesmethods have been used very little in the preliminary design environment. This is
primarily due to long cycle times and unvalidated accuracy levels. Computational
hardware improvements and CFD technological developments, such as the advent of
unstructured grids, have greatly reduced cycle times making Euler CFD methods a viable
candidate for preliminary design use. Incorporation of these methods into the preliminary
design process enables analytical determination of non-linear aerodynamic properties that
currently can only be assessed in the wind tunnel. These methods could therefore
potentially reduce the amount of costly wind tunnel testing. Another benefit of using CFD
in preliminary design is the potential for the development of a multi-disciplinary design
tool. Such a tool would allow the aerodynamic design to be tightly integrated with the
design of other disciplines.
Despite these potential advantages, CFD methods have not found their way into the
preliminary design environment. One reason for this may be the uncertainty associated
with using a new technology. There are several risks involved in using Euler CFD
methods for preliminary design. Errors are introduced into the analytical predictions by
neglecting the effects of viscosity. The significance of this error varies with the problem
geometry and flow conditions. The ability to account for this error is well documented for
many problems such as attached flows at low to moderate angles of attack, however, for
some cases the consequences of neglecting viscosity are difficult to predict. In addition,
the ability of Euler CFD methods to predict lateral-directional characteristics and the
effects from non-traditional control devices are not proven. It is also unclear whether the
improvements that have been made in CFD cycle time are sufficient to meet the needs of
the fast paced preliminary design environment.
A few years ago, NASA Langley Research Center initiated a project (Ref. 1-6) to
evaluate the viability of a series of Euler CFD methods for use in preliminary design. This
report summarizes the assessment of the Boeing, St. Louis developed CFD tools, MACGS
and NASTD, for use in preliminary design. These tools provide for rapid analysis of
complex configurations using either structured or unstructured grid techniques. This study
was focused on the assessment of the unstructured grid Euler capability of these tools.
NASTD/MACGS applications are performed within an integrated process whereby the
grids are generated directly on the CAD model. A common database file is carried
throughout the process from grid generation through post processing. To avoid the
requirement for a mainframe- or super-computer, which often is not available in the
preliminary design environment, solutions are computed in parallel on a network of
workstations. This provides rapid turnaround and low memory usage. These tools have
been used extensively on Boeing, St. Louis production programs such as the F/A-18 E/F,F-15C and AV-8B.
There were four primary objectives in this study. These were to assess the effects
of viscosity over a range of Mach number and angle of attack, assess aerodynamic
predictions for non-traditional control devices, evaluate lateral-directional analysiscapabilities,anddocumentconvergence-performancecharacteristics.
APPROACH
MACGS and NASTD were applied to the analysis of the AeroConfiguration/Weapons Fighter Technology (ACWFT) 1204 configuration shown in
Figure 1. This configuration was tested extensively (Ref. 7) in the NASA Langley
Research Center 8 ft. Transonic Pressure Tunnel. The ACWFT configuration is
representative of advanced preliminary designs. It is a tailless aircraft with a chined
forebody. Two wing geometries were tested with the ACWFT model. The baseline wing
has +/-30 degree leading and trailing edge sweeps, an aspect ratio of 2.65, and a taper ratio
0.132. The baseline wing cross section consists of a modified NACA 65A004 airfoil with
a sharp leading edge. The alternate wing model consisted of the same planform as the
baseline wing with the addition of camber and twist. The test model contained several
non-traditional control devices including elevons, and body flaps. A flow through duct
connecting the inlet and nozzle was incorporated into the test model. Test results included
force and moment measurements and pressure tap data over the wing surface. In addition,
pressure sensitive paint was used to obtain the global surface pressure distribution over theaircraft.
For this study, CFD solutions were computed on the ACWFT vehicle in six
different configurations as shown in Figure 2. These configurations included the baseline
configuration, baseline fuselage with a cambered/twisted wing, baseline configuration with
aflerbody flap deflected 90 degrees, baseline configuration with symmetric elevon
deflections of-20 degrees, baseline configuration with asymmetric elevon deflections of
+/- 20 degrees, and the baseline configuration with sideslip. The baseline, cambered wing,
and symmetric elevon configurations were modeled assuming symmetry about the fuselage
centerline. While developing the CFD model of the cambered wing configuration, we were
unable to locate the geometric definition of the cambered wing/fuselage interface that was
used on the wind tunnel test model. For this study an interface was made up by blending
the cambered wing geometry into the baseline wing root section. Unfortunately, the results
presented below show that this geometry modification may have influenced the resulting
CFD drag predictions.
A summary of the CFD run matrix is shown in Figure 3. For each configuration,
Euler solutions were computed at flow conditions of Mach 0.6, angles of attack of 1O, 15,
and 20 degrees, Mach 0.9 at angles of attack of 10 and 15 degrees, and Mach 1.2 at 10
degrees angle of attack. In addition to the Euler solutions, Navier-Stokes CFD solutions
were computed about the baseline and cambered/twisted wing configurations at the same
flow conditions to isolate the aerodynamic effects due to viscosity.
Grid Generation
Grid generation was performed using MACGS (Refs. 8,9), which is a general
purpose, arbitrary topology grid generation system developed at the McDonnell Douglas
Corporation. It supports the generation of multi-zone structured and/or unstructured grids.
MACGS is comprised of three modules: ZONI3G, GMAN, and GPRO. ZONI3G is used
to generate structured and/or unstructured surface grids. GMAN provides the capability to
generatevolume grids, specifyboundaryconditions and generatecoupling informationbetweenstructuredand/orunstructuredgrid zones. GPROmanipulateszones (suchastransforming, splitting, and combining) and supportsinputting and outputting files invariousformats. Theinteractive,graphicaluserinterfacesof ZONI3GandGMAN supportboth thenoviceandexpertuser.
A dual grid approachwasusedin this studywherealI viscouscomputationswereperformed on a pre-existing structuredgrid and all Euler solutionswere computedonunstructuredgrids. Unstructuredgrid generationwasperformedusingthefour stepprocessshownin Figure4. In the designenvironment,the geometryresidingin the CAD systemoften containsdetailedgeometrycomponents or surface gaps that the CFD user does notwant to include in the analysis. The first step in the grid generation process is to modify
the surface geometry to remove these unwanted geometry components and fill any
remaining gaps or holes. An unstructured surface grid is then interactively generated on
the clean surface representation. Grid resolution is controlled by the user through
specification of boundary edge distributions for each surface patch and through several line
and point source options. A tetrahedral volume grid is then generated within MACGS
using a Delauney point insertion approach. Grid swapping and smoothing are used to
ensure the quality of the final grid. Resolution of the volume grid is set by the surface grid
spacing and by two user specified parameters that control the global cell spacing and the
amount of clustering near the geometric surface. The resulting grid is partitioned into
multiple blocks with the METIS algorithm (Ref. 10). The sizes of each block are selected
to balance the solution load on parallel computational systems.
The structured and unstructured surface grids for the baseline ACWFT
configuration are shown in Figure 5. The multi-zone structured grid was generated in
MACGS during a previous study funded by Wright Laboratory (Ref. 7). The inlet and
nozzle ducts were treated differently in the unstructured and structured grids. To simplify
grid generation, the inlet and nozzle ducts were faired over in the structured grid. For all
but two of the unstructured grids, the inlet duct was modeled up to the compressor face
(where a mass flow boundary condition was specified) and the nozzle duct was faired over.
For the symmetric and asymmetric deflected elevon unstructured grids, a flow through duct
was modeled that connected the inlet and nozzle faces. This was done to capture the
effects of the nozzle flow washing over the deflected elevon surfaces.
The sizes of the resulting unstructured surface grids are shown in Figure 6. The
surface grids ranged in size from 150,000 to 300,000 triangles. Cuts through the structured
and unstructured volume grids are shown in Figure 7. In the first of the cuts shown in the
top of Figure 7, the lower surface of the structured grid differs from the unstructured
surface due to the inlet duct fairing. In Figure 8, the surfaces of the unstructured baseline
grid are shown after grid partitioning. In this example, the unstructured grid has been
partitioned into ten equal size zones.
The sizes of the structured and unstructured grids are summarized in Figure 9. The
sizes of the unstructured grids ranged from just over 1 million cells for the baseline
configuration to 2.7 million cells for the asymmetrically deflected elevon grid. The viscous
structured grids contained about 2.7 million nodes. Labor hours for the unstructured grid
generation are shown in Figure 10. The baseline unstructured grid required 30 person
hours to generate. Grid generation for the other five configurations examined in this study
were made by making minor modifications to the baseline surface grid. Thesemodificationsrequiredonly afew personhoursfor eachconfiguration. Thecomputationaltime requiredto generateeachunstructuredgrid is shownin Figure 11. The CPU timesgivenare for a Silicon GraphicsR10,000processorandrangefrom 1.5hoursto almost3hours.
Flow Solution Methodology and Performance Characteristics
The solution computations in this study were performed using the NASTD flow
solver. This flow solver was developed by the McDonnell Douglas Corporation, and can
run on structured, unstructured or a combination of structured and unstructured grids. It
supports multi-block and overlapping (chimera) grids. It runs in serial or parallel on a wide
variety of machines. A complete description of the NASTD structured grid solution
algorithm is given in Reference 11. The structured grid algorithm solves any subset of the
full Reynolds averaged Navier-Stokes equations. Options include Euler, thin layer,
parabolized Navier-Stokes and full Navier-Stokes calculations. Turbulence can be
modeled by a variety of algebraic, one- and two-equation turbulence models. The solution
algorithm can be selected zonally by the user. The default time integration scheme is a
first-order, approximately factored implicit scheme. For inviscid flows (or, under the thin
layer approximation, for directions without viscous terms) the implicit operator is
diagonalized, providing a significant speed-up. Explicit Runge Kutta options of up to
third-order are also available for time accurate flowfields. For steady-state flows, variable
time steps based on local eigenvalues are used to speed convergence. Grid sequencing is
available to speed convergence on large grids. The default explicit spatial operator is a
second-order flux difference splitting scheme, also known as Roe's scheme. The standard
upwind operator has been replaced by a mixed scheme which retains the upwind scheme
stability properties with reduced numerical dissipation. Optionally, the scheme may be
switched to various first- through fifth-order schemes and total variation diminishing
(TVD) limiters may be activated. Other available schemes include standard second-order
central differencing with added second- and fourth-order dissipation.
A complete description of the NASTD unstructured grid algorithm and the parallel
implementation is given in References 12 and 13. This algorithm is a node-based upwind
finite-volume unstructured grid algorithm. The implementation used for this study solves
the Euler equations on tetrahedral cell grids. Higher-order computations are achieved
using a least squares reconstruction scheme with flux limiting. The numerical flux values
are computed at the mid-point of each edge using Roe's approximate Riemann solver.
Flowfield variables are stored at grid nodes and flux computations are performed at each
grid edge. This results in relatively low storage requirements and run times. Time
integration is performed using an explicit point Jacobi or Runge-Kutta algorithm for each
node.
The following NASTD options were used in the computations for this study. The
fluxes were computed using a second-order accurate Roe's scheme with a Total Variational
Dimensioning (TVD) limiter in the case of the structured solutions and a monotone limiter
for the unstructured cases. In the structured grid Navier-Stokes computations, turbulence
was modeled with the Spalart-Almaras turbulence model. An implicit approximate
factorization algorithm was used for the time integration of the structured grid cases and an
explicit Runge-Kutta scheme was used for the unstructured grid cases. Solu.tionconvergencewasdeterminedby monitoringtheintegratedlift, dragandpitching moments.For the viscouscasesthe friction dragwasalso monitored.No attemptwasmadeto findthe maximum CFL numberfor thesecases. Insteada "safe" CFL numberwas selectedbasedon past experience. CFL numbersranged from 0.3 to 0.7 for the unstructuredcomputationsand 1.0to 3.0for thestructuredgrid computations.
Two representativeexamplesof Euler solution convergencefrom this study areshownin Figures 12 and 13. Theseexamplesrepresentthe best (Figure 12), and worst(Figure 13)Euler solutionconvergencehistoriesfrom this study. In thesefigures, the liftcoefficient is plotted versusthe solutioncyclenumber. In Figure 12the convergenceforthebaselineconfigurationat Mach1.2, I0 degreesangleof attackand5 degreessideslipisshown. This solution was restarted from the 0 degree sideslip solution at 700 cycles. Thelift coefficient reaches a steady value after an additional 1800 cycles (for a total of 2500
cycles). In Figure 13 the lift coefficient versus cycle number is shown for the baseline
configuration with -20 degree symmetric elevon deflections at Mach 0.6 and 20 degrees
angle of attack. This solution was run with the first-order accurate scheme for 1800 cycles
and then switched to the second-order scheme. The lift coefficient did not converge to a
steady value but instead oscillated about an average. This behavior was observed in all of
the 20 degree angle of attack cases and is most likely due to a non-steady behavior in theflow field.
The convergence properties of the NASTD Navier-Stokes solutions about the
baseline configuration at Mach 0.6, 10 and 20 degrees angle of attack are shown in Figures
14 and 15. These solutions were run for 680 cycles on a sequenced grid (every other grid
point removed in all three directions). The solution was then switched to the full grid and
reached convergence after another 500 cycles. The Navier-Stokes solutions did not
experience the oscillatory behavior at the high angles of attack seen in the Euler solutions.This could be due to the viscous terms which add additional diffusion to the flow.
Run times for the Euler solutions are summarized in Figure 16. The minimum,
average, and maximum run time are shown for each configuration. The numbers shown
represent the total CPU time and were obtained by multiplying the CPU time/iteration/cell
by the number of cells and the number of iterations. The times do not include the savings
obtained by running on a parallel computational system. For instance the average baseline
configuration solution required 90 hours of CPU time. When this solution was run on a
cluster of ten workstations, the actual clock time was a little under 10 hours. The memory
requirements for each of the Euler solutions are summarized in Figure 17. These memory
requirements are presented as though the solution were run as a single zone on one
processor. The actual memory requirements per machine varied depending on the number
of grid partitions. For ten equal size partitions, the memory requirements are 1/10 the total.
Solution run times and memory requirements for the Navier-Stokes solutions are
summarized in Figures 18a and b. Once again these numbers are given for a single zone
solution on a single processor. The actual requirements were much lower when running in
parallel.
RESULTS
Baseline Configuration
Euler and Navier-Stokes solutions were generated on the baseline ACWFT
configuration. Comparisons of the Euler and Navier-Stokes solutions were made to
identify the error introduced in the Euler solutions by neglecting viscosity. Further
comparisons with test data were made to identify the accuracy of the CFD methods. In
addition, differences between the baseline and other configuration results were used to
measure the incremental effects of each configuration.
Contours of the predicted surface pressure coefficient and traces of the streamlines
for the Euler and Navier-Stokes results at Mach 0.6, and angles of attack of 10, 15 and 20
degrees are shown in Figure 19. In this Figure, the Euler solution is shown on the left half
of the aircraft and the corresponding Navier-Stokes solution is shown on the right half of
the aircraft. At 10 degrees angle of attack, the Euler and Navier-Stokes results are very
similar. Both solutions predict a vortex separating off of the chined forebody and another
off of the wing leading edge. At 15 degrees angle of attack the surface pressures are once
again very similar. Both solutions indicate vortices similar to the 10 degree angle of attack
case. At 20 degrees angle of attack the differences between the Euler and Navier-Stokes
solutions are more noticeable both in terms of surface pressure distribution and particle
traces. In the Euler solution the wing leading edge vortex appears to have burst. This
results in significant differences in the surface pressures between the Euler and Navier-
Stokes solutions on the upper surface of the wing.
Total pressure contours at fuselage stations of 260 in., 420 in. and 510 in. are
shown in Figures 20 and 21. The locations and strengths of the predicted vortices for the
Euler and Navier-Stokes solutions are very similar at I0 degrees angle of attack. At 20
degrees angle of attack, however, the Navier-Stokes solution indicates a larger total
pressure loss in the vortex cores and the structure of the vortices are considerably different.
The effect of viscosity on the predicted solutions at different Mach numbers is
presented in Figure 22. In this figure, surface pressure coefficient contours and streamline
traces from the Euler and Navier-Stokes solutions at 10 degrees angle of attack and Mach
numbers of 0.6, 0.9, and ! .2 are compared. The streamline patterns predicted by the Euler
and Navier-Stokes solutions are similar at all three Mach numbers. However, there are
several differences in the predicted surface pressures at Mach 0.9. As expected, the Euler
solution indicates a strong shock over the wing at about 75% chord, while the Navier-
Stokes solution predicts a more diffused footprint of the shockwave that is slightly further
fornvard. This is probably due to shock boundary-layer interaction effects that the Euler
solution is missing. At Mach 1.2, the shock has moved aft of the wing trailing edge and
the Euler and Navier-Stokes solutions agree fairly well.Pressure coefficient contours on the lower surface for the Euler and Navier-Stokes
solutions are compared in Figure 23. The results agree well except near the inlet where the
two solutions have a different treatment of the inlet geometry (faired over for Navier-
Stokes and flow through for the Euler). While having little influence on the upper surface
solution, the different inlet models significantly change the lower surface results.
A comparison of surface pressures from the CFD results and pressure sensitive
paint (PSP) test data is shown in Figures 24 and 25 for flow conditions of Mach 0.6, and
8
Math 0.9 at 15 degreesangleof attack. The PSPdatawas not calibratedto provide aquantitativevaluefor eachcolor. Instead,the color mapusedfor plotting theCFD resultswasselectedto attemptto matchthe colors of the PSPdatathus providing a qualitativecomparisonof the flowfield structuressuchas shockwaves. At Mach 0.6, the Euler,Navier-Stokes,and PSPcontoursarevery similar. At Mach 0.9 the test data comparesfavorably with the Navier-Stokes results while, as expected, the Euler results clearly miss
the shock boundary layer interaction on the wing.
In addition to the PSP pressure data, surface pressure taps were placed at four
spanwise locations on the test model. Comparisons of the CFD surface pressures with
measurements taken at the pressure taps are made in Figures 26 through 31. In Figures 26
through 28, results from the solutions at Mach 0.6, angles of attack of 10, 15 and 20
degrees are shown. At angles of attack of 10 and 15 degrees, the Euler and Navier-Stokes
results are similar with the exception of the suction peak at the leading edge. As expected,
the Euler solution over predicts the acceleration around the sharp wing leading edge. At 20
degrees angle of attack there are significant differences in the Euler and Navier-Stokes
surface pressures. This is consistent with the differences that were observed in the surface
pressure contour plots above. Comparisons of the CFD results with the pressure tap data at
Mach 0.9 are shown in Figures 29 and 30. As expected, the Euler solutions predict a shock
location that is slightly aft of that predicted by the Navier-Stokes solutions. In addition,
there is a considerable amount of smearing of the shock footprint evident in the Navier-
Stokes results and test data that, is not present in the Euler solution. In Figure 31 the CFD
surface pressures are compared with pressure tap data at Mach 1.2 and 10 degrees angle of
attack. At this flow condition, the shock has left the wing surface and the Euler and
Navier-Stokes solutions compare favorably with the test data.
Force predictions were obtained from the CFD solutions by integrating the surface
pressures (and skin friction for the Navier-Stokes solutions) over the aircraft surface.
Corrections were added to the Euler drag estimates to account for the skin friction drag.
The corrections were obtained using the following procedure. First, Euler solutions were
computed over the baseline configuration at angles of attack that resulted in zero lift, and
Mach numbers of 0.6, 0.9, and 1.2. Next, the zero lift drag predicted by each Euler
solution was subtracted from the zero lift drag measured in the test at the same Mach
number to obtain the skin friction contribution to the total drag. The resulting skin friction
estimates were then added to all the Euler estimates. For the Euler solutions that failed to
converge to a steady state, force and moment values were obtained by averaging the
integrated results over the last few hundred cycles of the solution. Error bars are drawn to
indicate the maximum and minimum oscillation about the average. The CFD force and
moment predictions are compared with test data in Figures 32 through 34. The lift and
drag from the Euler solutions match the test data very well with the exception of the Mach
0.9, 15 degrees angle of attack. This is probably due to the missing shock boundary layer
interaction effects in the Euler solution. Surprisingly, the Navier-Stokes force and moment
results are slightly worse than the Euler predictions. The most likely reason for this
discrepancy is the presence of the faired over inlet model used in the viscous computations.
Cambered Wing Configuration
An alternate wing was tested on the ACWFT geometry. This wing was similar to
the ACWFT baseline wing with the addition of camber and twist. Contours of the
predicted surface pressure coefficient and streamline traces are compared for the Euler and
Navier-Stokes cambered wing configuration solutions at Math 0.9, 10 degrees angle of
attack in Figure 35. The comparison is similar to the baseline wing comparisons showing
that the Euler solution is missing the shock boundary layer interaction effects. The force
and moment predictions are compared with the test data in Figures 36-38. Again these
comparisons are very similar to the baseline wing results. The most notable deviations
from the test data occur at flow conditions of Mach 0.6, 20 degrees angle of attack and
Math 0.9, 15 degrees angle of attack. Increments in the force and moment predictions
between the cambered wing and baseline wing configurations are shown in Figures 39-41.
The test data shows a slight increase in lift and decrease in drag with little change in
pitching moment. The Euler and Navier-Stokes results also predict a slight increase in lift,
however, both methods predict an increase in drag. This discrepancy from the test data
may be partially attributed to an increase in interference drag caused by the method
employed to attach the alternate wing in the CFD grid.
Lateral-Directional Characteristics
The ability to compute lateral-directional characteristics is essential for a
preliminary design tool. The baseline configuration was run at 5 degrees sideslip to
evaluate the lateral-directional characteristic prediction capability of the present Euler
method. Streamline traces and surface pressure predictions from the Euler solutions at
Mach 0.6, 15 degrees angle of attack with 0 and 5 degrees sideslip are compared in Figure
42. The sideslip has little effect on the surface pressures. The most notable effect of the
sideslip is the change in track of the vortices downstream of the aircraft with little effect
over the aircraft itself. Force and moment predictions are compared with test data in
Figures 43-45. Increments in the force and moment predictions between the sideslip and
zero sideslip cases are shown in Figures 46-48. The test data indicates large increments in
the lift, drag and moments while the CFD results indicate small changes in the forces and
moments. The test data trends are contrary to the expected behavior of a tailless aircraft.
We suspect that the location of the support strut on the model may have influenced the testdata.
Control Devices
One of the objectives of this program was to evaluate the ability of the CFD method
to compute the effects of control surfaces. Solutions were computed about three control
devices including a symmetrically deflected elevon, asymmetrically deflected elevon, and a
body flap.
The ACWFT solid surface model and a closeup of the surface grid about the
syrmnetrically deflected elevon is shown in Figure 49. The elevon was deflected up 20
degrees on the left and fight sides of the aircraft. Euler solutions were run at all six flow
conditions. A comparison of the surface pressure and streamline traces for the baseline
(undeflected elevon) and symmetrically deflected elevon cases at Mach 0.6, 15 degrees
angle of attack is shown in Figure 50. The elevon deflection primarily affects the solution
10
nearthetail andhaslittle effecton the solution over the wing. The predicted lift, drag andpitching moment from the CFD solutions is compared with test data in Figures 51-53.
Once again the CFD lift and drag predictions agree with the test data, however, the CFD
results over predict the effect of the elevon deflection on pitch up. This can be seen in the
incremental force and moment plots shown in Figures 54-56. The CFD and test data both
indicate a substantial reduction in lift with a slight decrease in drag. The CFD results
overpredict the effect of the elevon deflection on pitching moment.
The ACWFT solid surface model and a closeup of the surface grid about the
asymmetrically deflected elevon and deflected body flap are shown in Figure 57. Surface
pressure and streamline traces from the CFD solutions at Mach 0.6, 15 degrees angle of
attack, with the devices are compared with the baseline solution in Figure 58. The
asymmetrically deflected elevon has little effect on the solution other than in the tail
region, whereas the body flap has a larger influence on the surface pressure of the
surrounding geometry.
Force and moment predictions for the asymmetrically deflected elevon CFD
solutions are compared with test data in Figures 59-61. The comparisons are similar to the
previous results showing good agreement for lift and drag except at Mach 0.9, 15 degrees
angle of attack. Increments in the force and moment predictions with the baseline results
are shown in Figures 62-64. As expected the pitching moment increment is very small at
all three Mach numbers. At Mach 0.6, the CFD results indicate an increase in lift and drag
as the angle of attack is increased whereas the test data a constant increment in lift and a
decreasing increment in drag. These discrepancies may be due to the lack of convergence
of the Euler method at the high angles of attack. The incremental pitch, yaw and roll are
well predicted by the CFD method.
Force and moment predictions for the deflected body flap CFD solutions are
compared with test data in Figures 65-67. Once again the force comparisons are good with
the exception of the transonic cases at Mach 0.9. Increments in the force and moment
predictions with the baseline results are shown in Figures 68-70. The incremental data
indicates a slight decrease in lift and increase in drag (for angles of attack less than 15
degrees). The CFD results underpredict the lift decrease due to the flap particularly at the
higher angles of attack. The CFD drag increments are also much higher than the test data
increments at the higher angles of attack. The discrepancies may be due to lack of
convergence or poor modeling of the large wake like separated region aft of the flap. This
type of flowfield is largely dominated by viscous effects and is not well modeled with an
Euler method. The incremental pitch, roll and yawing moments due to the flap are well
predicted by the CFD results.
CONCLUSIONS
This study has provided an assessment of the viability of using the NASTD
unstructured grid Euler technology in preliminary design. Euler solutions about the
ACWFT 1204 configuration with several geometry variations including baseline wing,
cambered wing, deflected elevons, and deflected body flap were generated. A wide range
of flow conditions, most of which were in the non-linear regimes of the flight envelope,
were evaluated including transonic and high angle of attack flowfields. Several non-
linearities were present in these solutions including shock waves, vortical and separated
11
flows. Comparisonswith testdataandNavier-Stokessolutionswereusedto evaluatetheaccuracyof this Euler methodand to identify viscouseffects for selectedconfigurationsandconditions. Solutionswith sideslipanddeflectedcontrolsurfaceswerecomparedwithtest datato evaluatethe ability to accuratelypredict lateral-directionalcharacteristicsandcontrol effectiveness.
The unstructuredgrid approach facilitated rapid modeling of the ACWFTconfiguration and its variations. Geometryvariations such as flap deflections weremodeledin only a few hours. The methodologyproved to be very robust generatingsolutions for various surfacecontrol devicesand variousflow conditionswith very fewproblems. Run timesweresufficiently fast for usein thepreliminarydesignenvironmentwith overnightrun timespossibleonparallelcomputationalsystems.Thesecycletimesaresufficient to generatedataat a largenumberof pointsacrossthe flight envelope,however,they are still too slow to generatean entire aerodynamicdatabasetypically developedduring thewind tunneltest.
The Eulerresultscapturedseveralnon-linearaerodynamiccharacteristicsof thetestdataat high anglesof attack. Forceandmomentpredictionsweregenerallysufficient forpreliminarydesignuseup to moderateanglesof attack(~ 15degrees)acrossthe examinedMach numberrange. Lessaccuracywasobtainedat high anglesof attackandfor controldevicesthatgeneratedalargeamountof separationsuchasthebody flap. Onesurprisingresult of this studywasthatNavier-Stokesforceandmomentresultswerenot appreciablybetterthan theEulerpredictions. This indicatesthereis little benefit in steppingup to thelonger run times and complexities of a Navier-Stokesmethod for these types ofpredictions.
OnecasewhereNavier-Stokespredictionsmay haveprovidedbetter resultsthantheEuler solutionsis for thedeflectedafterbodyflap. Thiscontroldevicegeneratesa largeseparatedregionaft of theflap that interactswith thefuselagesurfaceto generatechangesin theforce andmomentdistributions. TheEuler methoddid a goodjob of predictingtheflap control effectson theroll, yaw,andpitchingmomentsbut significantlyoverpredictedtheeffect on lift anddrag. Betterpredictionswereobtainedfor theelevoncontrol deviceeffectiveness.The lateral-directionalcharacteristicsdueto asymmetriccontroldeflectionswere accuratelypredictedwhile the control effectivenessin pitch wasconsistentlyover-predicted. With further work, empirical correlations to compensatefor the pitcheffectivenesscouldbedeveloped.
Euler technologyhas strong relevanceto preliminary-designapplications. Thistechnologyprovidesa meansof generatingnon-linearand lateral-directionalaerodynamicdatathat previouslycouldonly beobtainedin thewind tunnel. The ability to analyticallygeneratelateral-directionaldataprovidesanun-exploitedpotentialfor the developmentofa lateral-directionaldesigntool basedonexistingEulcr technology.
It is likely that linearaerodynamictoolswill continueto beusedto developa largeportion of the aerodynamicdatabase.Furtherstudy is necessaryto determinethepartsofthe flight envelopewhere linearand non-linearmethodsare bestapplied. Comparisonsbetweenlinearandnon-linearresultsalthoughnot apartof this study,couldhelpguidethisdetermination. Furtherstudy is also necessaryto developempiricalcalibrationsof Eulerresults for someapplications. This wasevident m the elevoneffectivenesspredictionsobtainedin this study.
12
Oneof the greatest potential benefits of Euler technology may be the ability to tienon-linear aerodynamic data into a multi-disciplinary design tool. The ability to share data
with other disciplines and use a common geometry database is essential if this technology
is to be used successfully in the preliminary design environment.
ACKNOWLEDGEMENTS
This effort was sponsored by NASA-Langley Research Center under Contract
NAS1-20342. Farhad Ghaffari was the contract Technical Monitor and provided
invaluable technical guidance over the course of this study. His assistance is gratefully
acknowledged.
REFERENCES
1.) Finely, D.B.,"Euler Technology Assessment Program for Preliminary Aircraft Design
Employing SPLITFLOW Code With Cartesian Unstructured Grid Method," NASA
CR-4649, March 1995.
2.) Kinard, T.A., Hams, B.W., and Raj, P.," An Assessment of Viscous Effects in
Computational Simulation of Benign and Burst Vortex Flows on Generic Fighter
Wind-Tunnel Models Using TEAM Code," NASA CR-4650, March 1995.
3.) Treiber, D.A. and Muilenberg, D.A.," Euler Technology Assessment for Preliminary
Aircraft Design Employing OVERFLOW Code With Multiblock Structured-Grid
Method," NASA CR-4651, March 1995.
4.) Finely, D.B. and Karman, Jr., S.L.,"Euler Technology Assessment for Preliminary
Aircraft Design - Compressibility Predictions by Employing the Cartesian Unstructured
Grid SPLITFLOW Code," NASA CR-4710, March 1996.
5.) Kinard, T.A. and Raj, P.," Euler Technology Assessment for Preliminary Aircraft
Design - Compressibility Predictions by Employing the Unstructured Grid USM3D
Code," NASA CR-4711, March 1996.
6.) Kinard, T.A., Finley, D.B., and Karman, Jr., S.L.,"Prediction of Compressibility Effects
Using Unstructured Euler Analysis on Vortex Dominated Flow Fields," AIAA Paper-
No. 96-2499, June 17-29,1996.
7.) O'Neil, P.J., Krekeler, G.C., Billman, G.M.. and Creaseman, F.C., "Aero
Configuration/Weapons Fighter Technology (ACX_,'FT) - Summary Technical Report,"
WL-TR-95-3002, December 1994.
8.) Gatzke, T.D., et. al., "MACGS: A Zonal Grid Generation System for Complex Aero-
Propulsion Configurations," AIAA-91-2156, June 199 I.
13
9.) LaBozzetta,W.F., Gatzke, T.D., Ellison, S., Finfrock, G.P., and Fisher, M.S.,"MACGS - Towards the Complete Grid Generation System," AIAA 94-1923, 12thAIAA Applied Aerodynamics Conference, June 20-22, 1994.
10.) Karypis, G., and Kumar, V.,"METIS, Unstructured Graph Partitioning and Sparse
Matrix Ordering System", Users Manual August, 1995.
11 .) Romer, W.W., and Bush, R.H., "Boundary Condition Procedures for CFD Analysis of
Propulsion Systems - The Multi-Zone Problem," AIAA-93-1971, June 1993.
12.) Michal, T., and Halt, D., "Development and Application of an Unstructured Grid Flow
Solver for Complex Fighter Aircraft Configurations," AIAA 95-1785, June, 1995.
13.) Michal, T., and Johnson, J., "A Hybrid Structured/Unstructured Grid Multi-Block
Flow Solver for Distributed Parallel Processing," AIAA 97-1895, June, 1997.
14
Figure1.
r
ACWFT 1204 Three View Drawing.
CASE 1._ Baseline Configuration CASE 3.0: Symmetric Elevons
CASE 1.b: Cambered Wing CASE 3.b: Asymmetric Elevon
CASE 2: AfteYoody Flop CASE4: Basellnewlth Sideslip
Figure 2. ACWFT 1204 Configurations Modeled in Euler Technology Assessment Study.
15
c_
(z
10°
150
200
100
150
200
Case la: BaselineMach 0.6 Mach 0.9 !Mach 1.2
I,V I,V I,V
I,V I,V
I,V
Case 2: After Body FlapMaeh 0.6 Maeh0.9 Mach 1.2
I I I
I I
I
Case 1b: Cambered/Twisted WingMach 0.6 Mach0.9 Mach 1.2
I,V I,V I,V
I,V I,V
I,V
10°
£_ 150200
c_
Case 3a: Symmetric Elevons
100
150
200
Mach 0.6 Mach 0.9 Mach 1.2
I I I
I I
I
Case 3b: Asymmetric Elevons
10°
15o
20°
Mach0.6 Mach0.9 Mach 1.2
I I I
I 1
I
I: Inviscid V: Viscous
Case 4: Baseline with Sideslip
o_
10°
150
200
Mach0.6 Mach0.9 Mach 1.2
I I I
I I
I
Figure 3. Euler Technology Assessment CFD Run Matrix.
Surface Grid Generation
Volume Grid Generation
Figure 4. MACGS Unstructured Grid Generation Process.
16
35O
_' 300
250
200
¢_ 150
"6_oo
_ 5o
0
Figure 6. Unstructured Surface Grid Sizes, Six ACWFT 1204 Configurations.
18
IF l 111111i_lfl/I III UlliliIfJIIlklI III lllitlllllillllll Illlr)[_i IIilllllUllmlllllll
_llIIIYl,!lll r I f\_ I It Ill 11IIIIINnlllllll II_[llJfJt_/ll.l(l I\l I I IllUIIItmmHIIIII II
iiiiiiiiiiiiiii,_1 lIII] llll II '_l_llit_Nllttlllll IIIILllllllllnlllllII__/HIIIllllllllllllllll II_ii_jiiiiVlllllllh'ltmillllllldiilllllllll II ""_I]ITTlll f_ILII ] flfdlllllllllllll]llllll111II II
_ I _iI IIIIIIl[[llll_llllllIII I illlll
_q_l_'_llJ]]JiBI]llllllllllli[llll [JillIII I........ I.III
]I I
l_,i i
i
Figure 7. Cuts Through Volume Grids, ACWFT 1204 Baseline Configuration.
19
\ \
Figure 8. ACWFT 1204 Unstructured Surface Grid After Partitioning.
2O
3000 I
2500
2000
1500
1000
50O
• Cells •Nodes I
30
Figure 9. Volume Grid Sizes.
25
20
"J 10
5
Figure 10. Unstructured Grid Labor Hours.
21
3
2.5O
® 2
¢_ 1.5
m o.5
Figure 11. Unstructured Grid Computation Time.
"E
o
o68
0 67
0_16
Oe6
O64
063
5OO
Figure 12.
I ! | i n * * | i
I tOO 1500 2000 2500 3000 _7 4_3: 45[]0 SOOt] SSO0
Cycte
Lift Convergence, Unstructured Grid Euler Solution, Mach 1.2,
10 degrees angle of attack, 5 degrees sideslip.
22
3D
25
20
ou,
15
1
Figure 13.
148
i i i
1000 2000 3000 4000 5000 8000 7000
Cycle
Lift Convergence, Unstructured Grid Euler Solution,
Mach 0.6, 20 degrees angle of attack.
148
138 i i i i i i i840 720 800 080 gso _040 1120 1200
Cycle
Figure 14. Lift Convergence, Navier-Stokes Solution,
Mach 0.6, 10 degrees angle of attack.
3_
3oo
®5
2_
292 I i i I i i i ,_0 720 000 O_ gSD _._ 1120 1200 12_C 1&60
Cycle
Figure 15. Lift Convergence, Navier-Stokes Solution,
Mach 0.6, 20 degrees angle of attack.
23
1000
900
800
700
600
500
400'I,.,.
I_ 300
200
100
0
Figure 16.
5O
45
4O0" 35x
300E• 25
o 20
_ 150
_ 10
5
Figure 17.
um• average •maximum
NASTD Euler Solution Computation Times.
NASTD Euler Solution Memory Requirements.
24
I minimum• maximum200
1000ILl
800
600
400
II:--- 200t_¢n
0
• average140
120
100
:E 80
0E 6O
411
2O
0
a.) Single Processor Run Time b) Single Processor Memory
Requirements
Figure 18. NASTD Navier Stokes Solution Time and Memory Requirements.
25
M = 0.6, _ = 100 M = 0.6, _ = 15 ° M = 0.6, c_= 20 °
Cp
I .5• 0.0
-0.5-1.0
15
Navier-Stokes Euler
Navier-Stokes Euler
Navier-Stokes
Figure 19. Viscous Effects Versus Angle of Attack, ACWFT Baseline Configuration,
Mach 0.6, NASTD Euler and Navier-Stokes Solutions.
Euler Navler-Stokes Pl/Pto
I 1.0
0.9
0.8 260 .............
0.7
420 ...........................
Euler, Navier-Stokes
Comparison
Figure 20. Euler and Navier-Stokes Comparison, Total Pressure Contours, ACWFT
Baseline Configuration, Mach 0.6, 10 Degrees Angle of Attack.
26
Euler Navier-Stokes Pt/Pto
I 1.0
0.9
0.8 260 ............
0.7
420 .............................
Euler, Navier-Stokes
COMPARISON
Figure 21. Euler and Navier-Stokes Comparison, Total Pressure Contours, ACWFT
Baseline Configuration, Mach 0.6, 20 Degrees Angle of Attack.
Cp
0,0
-0.5_'_-1 0
M = 0.6, _ = 10 0
Euler_| Navier-
okes
M = 0.9, _ = 10 0
Euler
M = 1.2, a = 10 0
EulerNavier-Stokes
Figure 22. Viscous Effects Versus Mach Number, ACWFT Baseline Configuration, 10
Degrees Angle of Attack, NASTD Euler and Navier-Stokes Solutions.
27
Lower Surface
M "0.9, a = 10° M= 1.2, ='= 10°
Invlscld Viscous Invlscld Vbcous
Cp
0.30
-0.25
-0.80
Figure 23. Pressure Coefficient Contours on Lower Surface of ACWFT Baseline
Configuration, Euler and Navier-Stokes Solutions.
Experimental - Computational Cp Comparison
Euler Solution
(M 0.6, c_15 °)
Pressure
(M 0.6, a 16 °)
Nsvler-Stokes
(M 0.6, c_15 °)
Figure 24. CFD Surface Pressure Comparison With Pressure Sensitive Paint Test Data,
ACWFT Baseline Configuration, Mach 0.6, 15 Degrees Angle of Attack.
28
Experimental - Computational Cp Comparison
Euler Solution Pressure Sensitiv( Navier-Stokes Sol
(M 0.9, _ 15°) (M 0.9, a 16°) (M 0.9, _ 15°)
Figure 25. CFD Surface Pressure Comparison With Pressure Sensitive Paint Test Data,
ACWFT Baseline Configuration, Mach 0.9, 15 Degrees Angle of Attack.
i0
1o_5
g: 0
_'_? _ TestData
'_',,_, - _ NAS'T'DEulerSolution
--- NASTD Newer-Stokes Solution
_0% span
, , , , ,
¢_ cut at 70% spa_ o
_e t_ 20 21 zz
x/c
jo
ioiT
t 55% span
| o¢
2:b e_so
te tg _o Zl 22 z_
_c
i
cut at 88% span
Ie4 _ee !e2 _le l_o _ _$ _1_ _ e_c
Figure 26. CFD Surface Pressure Comparison With Test Data, ACWFT Baseline
Configuration, Mach 0.6, 10 Degrees Angle of Attack.
29
qt,,1-
!lb
-1 o
J . . .
i
Testo=a. -- NASTD 6.Jler Solutbon_, ....... N/L_TD N_vier.Slokes Solution
,,, !
,_e e- cutat40% _oan
m
|
d .
I
>o •
cutat88% s_n
Figure 27. CFD Surface Pressure Comparison With Test Data, ACWFT Baseline
Configuration, Mach 0.6, 15 Degrees Angle of Attack.
_a
er
",. _ list U_ta: " I"¢_TD Euler Solutioni ",., ...... NASTD Nav_r-Stokes Solution
! \,,
p
c_
(_ °,7 /
__""-cut"" _t 88% span
Figure 28. CFD Surface Pressure Comparison With Test Data, ACWFT Baseline
Configuration, Mach 0.6, 20 Degrees Angle of Attack.
30
'7't,
(_ lest Data_._..: N/kS'TO Euler Solution
...... NAS'T'D Navier-Stokes Solubon
•. i; t_ w m: _ ..... , ,,I
s_
• cutmtTO%spsn
• = 'D _: =t =
_ _ ,
!
'_ t_o c_ 55% _n
lo [
II.I; 11; l_.s la,4 1_+: z;'t, _t;= z'Ll ZZ4 =_Z Z_;
|=
cut _t 88% span
Figure 29. CFD Surface Pressure Comparison With Test Data, ACWFT Baseline
Configuration, Mach 0.9, 10 Degrees Angle of Attack.
it •:._.,',+
i
.+f:
F
._ i+1:
(_ T_ E)alaf'_ I_ASTC Euler Solution
_-_-__..... ....... I_A_ IIS NaVl_r-_StoltesSolut=on
° ":,"'m'-- m _
"% _I
11 tI0,+ ,_+ ,i_ ,! ,I :) 21 LI l,; II =; _/,. _; III I II 14 •+i_ /,;_ L; ,I_.| ;11.| gg" +:+.Z Z_L+'
O 1_ 10
(]I
(-
Figure 30. CFD Surface Pressure Comparison With Test Data, ACWFT Baseline
Configuration, Mach 0.9, 15 Degree Angle of Attack•
31
¢._
Zl
_ " cut at 40% span
¢_ T,,_._ _ iNASTD Euler Solution
...... NASTD Na,ttet-Stokes Solution|
_4
)c
J_
)4
, t .....
"*--Z" "--_ "°_'_" "--''_" _ _ ",
! cut at 56% span
_--o_ T .... _..Q__.._..._.._=_._
cut at 70% span
• i
r _ _ a
-_ c4
_"- t" cut at 88% spmn
i
Figure 31. CFD Surface Pressure Comparison With Test Data, ACWFT Baseline
Configuration, Mach 1.2, 10 Degrees Angle of Attack.
1.2 1.2
1.0
0.8
"6
I
o,4
0.2 !
0.0 ,
0 5 10 15 20
Angle of Attack (deg)
1.2
1.0
o.B
_ 0.6
_ 0.4]
00
0.0 0.1 0.2 0.3 0.4 0.5
Drag Coefficient
1.0
0.8eO
0.6L)
=t.I 0.40.2
I
0.0
0.20 0.15 0.10 0.05 0.00
Pitching Moment Coefficient
Figure 32. CFD Force and Moment Comparisons With Test Data, ACWFT Baseline
Configuration, Mach 0.6.
32
"6
1.2
1.0
0.8
0.6
0.4
0.2
0.0
1.2 1.2 ................................................ _-_
1.0
"6 "6
""i -t 0.4
lestda_ 0.2
H I o.o0 5 10 15 20 0.0 0.1 0.2 0.3 0.4 0.5 0.20 0.15 0.10 0.05 0.00
Angle of Attack (deg) Drag Coefficient Pitching Moment Coefficient
Figure 33. CFD Force and Moment Comparisons With Test Data, ACWFT Baseline
Configuration, Mach 0.9.
1.0
0.9
0.8
0.7
0.6
0.5
= 0.4-J
0,3
0.2
0.1
O0
5 10 15
Angle of Attack (deg)
1.0
0.9
0.8
0.7
0.6
0.5
I 0 0.4! =I :I O,3
2O
0.2
0.1
0.0
0.0
m
E
0,1 0.2 0.3 0.4 0.5
Drag Coefficient
1.0
0.9
0.8
0.7
0.6
0.5
0.4
0.3
0.2
0.1
0.0
0.10 0.05 o.oo -o.o5 -oJo
Pitching Moment Coefficient
Figure 34. CFD Force and Moment Comparisons With Test Data, ACWFT Baseline
Configuration, Mach 1.2.
33
Cambered Wing
M = 0.9, _ = 10°
Cp
Inviscid Viscous
0.5
I0.0
-0.5
-1.0
-1.5
Figure 35. Cambered Wing Surface Pressure Coefficient and Streamline Traces, Mach
0.9, 10 Degrees Angle of Attack.
1.2 1.2 1.2 ....... =.........
1.0 1.0 1.0
_ 0.8 _ 0.8 _ 0.5
o, _ o., -_ o,
0.2 0 2 ! 0.2
0.0 0.0 t _ , 0.0
0 5 10 15 20 0.0 0.1 0.2 0.3 0.4 0.5 0.20 0.15 0.10 0.05
Angle of Attack (deg) Drag Coefficient
0.00
Pitching Moment Coefficient
Figure 36. CFD Force and Moment Comparisons With Test Data, ACWFT Cambered
Wing Configuration, Mach 0.6.
34
t.2
1.0
0.8.2
0.6
0.4
0.2
0.0
5 10 15 20
Angleof Attack(deg)
1.2
1.0
0.8"6
_ 0,6
_ 0.4
0.2
0,0
0,0 0.1 0.2 0.3 0.4 0.s
Drag Coefficient
1.2
1.0
- 0.8
mo
0.6
0.4
0.2-
0.0
0.20
J
0,15 0.1o oo5 oo0
Pitching Moment Coefficient
Figure 37. CFD Force and Moment Comparisons With Test Data, ACWFT
CamberedWing Configuration, Mach 0.9.
1.2
1.0
i 0.83Z
0.6
0.4
0.2 =
0.0
0 5 l0 t5 20
Angle of Attack (¢leg)
,2]i
1.0 _-
/
ool II • _ro.,_= I0.0 0.1 0.2 0.3 0.4 0.5
Drag Coefficient
1.2 _ ...............................................
1.o
._ 0.8
==8 0.8¢J
=.J
0.4
0.2
0.0 ' P
0.10 o,05 0._ -o.os -o.I0
Pitching Moment Coefficient
Figure 38. CFD Force and Moment Comparisons With Test Data, ACWFT Cambered
Wing Configuration, Mach 1.2.
0.30
. 0.20e=JE
0.10
=_-_ 0.00
P
-0.10- -0.20
-0.30
IN.I clara
• f,_STD Euler
• N_SI"D Namer.Stokes
0.10
0.08
i |0.08_ 0.0,1
0.02
_ -0.02-0.04
-0,1_
-0.08
-0.10
0
0.4
0.3
i 0.2
i 0.1
0.0
-0.1
E_ -0.2
! -0.3
-0.4
5 10 15 20 S 10 15 2O 5 10 lS 20
AngleOfAttack ((:leg) Angleof Attack(deg) Angleof Attack(dog)
Figure 39. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
Cambered Wing Configuration, Mach 0.6.
35
0.30
. 0.20c
_o
0.10
0.00m
c
_ -0.100
-0.20
-0.20
i 0.10]
F 0.08
_o 0.06
( ) 0.04
0.02a_ 0.00
E =) -0.02E
,i°°' tT
i _-o.08-0.08i
-0.I0
0
0.4 _r"..................................
0.3
0,2
)(oo_" o-0.1
6
G_=
-0.2
-0.3 t
-0.4I r5 10 15 20 5 10 15 20 0 5 10 15 20
Angle of Attack (deg) Angle of Attack (deg) Angle of Attack (deg)
Figure 40. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
Cambered Wing Configuration, Mach 0.9.
0.30
- 0,20co
0.10¢,.)
_ 0.00
-0.I0
= -0.20
-0.30
0.10
0.08
0.06u
_= 0.04
0.02
_=_0.00'D'_ -0.02=
-0.04®
-0.08
-0 08
-0.100 5 10 15 20
Angle ol Attack (deg)
0'l- 0.30.2o i
!)o,0.0°-0.1
I_ -0.2
-_ -o.3
-0.4
T..... m
5 10 15 20 0 5 10 15 20
Angle of Attack ((leg) Angle of Attack (dell)
Figure 41. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
Cambered Wing Configuration, Mach 1.2.
36
Beta = 5° Beta = 0°
(M = 0.6, ct = 150)
Cp
0.5
0.0
-0.5-1.0
-1.5
Figure 42. Surface Pressure Coefficient and Streamline Traces, ACWFT Baseline
Configuration, Mach 0.6, 0 and 5 Degrees Side Slip Angles, 15 Degrees Angle of Attack.
12 t2 12
10
--O8c
¢:
U
.-J04
00
10
_08"
/m _ o_/ (J
£_ (]4.
02,
00'
0 5 10 15 2O O0
Angle of Attack (deg)
mI
J
10
" O8c
04
02
O0
!ov 02 03 04 05 020 0_o oco -olo
Drag Coefficient Pitching Mommnt Coefficient
Figure 43. CFD Force and Moment Comparisons With Test Data, ACWFT Baseline
Configuration, 5 Degrees Sideslip, Mach 0.6.
37
,2F..............................................1.o i.o | I.O _I_
I!°'I'°"0.6 0.6 _ 0.6
0.4 _ 0.4 _ 0.4
02 021 02oo_/ I 00I i .=.,_, il 000 s 10 15 2o 0.0 0.1 0.2 0.3 0.4 o.s 0.10 o.os 0.00 -o.os -0.10
Angle of Attack ((leg) Drag Coefficient Pitching Moment Coefficient
Figure 44. CFD Force and Moment Comparisons With Test Data, ACWFT Baseline
Configuration, 5 Degrees Sideslip, Mach 0.9.
1.2 " 1.2
1.0
0,8-
0,6
¢.)E
0,4-
1.0
i q i
5 10 15
Angle of Attack (deg)
1.2
1.0
0.8
0.6
I0,4 t
0.2
0.0
0.0 0.1 0.2 0.3 0.4 0.5
DrxgCoeff_lem
0.8JR
0.6
0
E04
0.2 " 0.2
0.0 0.0
0 2O 0.10
i
0.05 O.OO -O.(Y3 -0.10
Pitching Moment Coefficient
Figure 45. CFD Force and Moment Comparisons With Test Data, ACWFT Baseline
Configuration, 5 Degrees Sideslip, Mach 1.2.
38
0.30 ....................................
= 0.20
0.10¢.)
0.00 ___m
|_ -0.I0
.._ -0,20
-0.30
5 10 15 20
Angle of Attack (deg)
0.10
0.08
0.06 I0.04 +
0.02
=0.04 t
=0,06 1
0 5 10 15 20
Angle Of AtWck (aeg)
0,4
0.3
_o.1
o.o-0.1 _•0.2
--_ -0.3
.0.4
0 5 10 15 20
Angle of AIlaCk ((leg)
I 0.05 i ! 0.05I
0.04 II 0,04
0.03 J 0.03
"°" foo,I 0.01 ]•0.0, _ .0.0,_ _,
z -o.o2 z _o.o21 i
i-0.05 .............. J .0.05 '- J
5 10 15 20 0 5 10 15 20
AIlgle of Atlack (deg) Angle of Attack (de.g)
Figure 46. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
Baseline Configuration, 5 Degrees Sideslip, Mach 0.6.
0.30
i 0.200,10
;5 0.00 _--_- -
]
i .0,10
--= -0.20
-0.30
0 5 10 15 20
Angle of Attack (deg}
0.05 ...........................
_ 0,03U
0.02
oo,0
.0,01 • •
.002
-0.03.0_0.4
-0.05 ..............
5 10 15 20
Angle of Attack (deg)
0_10 1
_ o.o8io.o6
_ o.ooL--.-_'-I-_"-0.02_
-0.06 _tmdm
.0.10
0 5 10 15 20
Angle of _tack (Oeg)
0.05 ,
0.02
o.0_ Jr
•0.02 _
.0.03 -_ . ,
•0.05L ........ --_0 5 10 15 20
Angle of Attack (deg)
0,4 ........................................
E 0.3
I 0.2Z
i Q.10.0 _- 1"T-
-0.1
-0,2
_-- "0.3
-O.l
5 10 15 20
Angle of Attack (deg)
Figure 47. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
Baseline Configuration, 5 Degrees Sideslip, Mach 0.9.
39
0.30
i 0,200.10
0.00
]
-0.10- -0.20
-0.30
0.10
. 0.08
i 0.06
0.04 J
0.00
-0,02
-0.04
-0.06
-0.08
-0.105 10 15 20
Angle of Attack (deg)
0.05 |
j°o:t
i -0,OlI --0.02
-0.03
0 5 10 15 20
Angle of/Umck (deg)
0.4
" 0,3
0.2
_| o1: |oo
_o -0.1Z
_am,am i -02
"_'_" I -_ -0._
-0.45 10 15 20
Angle of Attack (dell)
0.05 t ..................................0.04
0.03 ÷
0.02
0.01 I0
-0.01-0.02
I-o.o4 , _'_ e_-0.05 L
0 5 10 15 20
Angle of Attack (deg)
II
i5 10 15 20
Angle of AUack (dell)
Figure 48. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
Baseline Configuration, 5 Degrees Sideslip, Mach 1.2.
40
-20 ° (up) Elevon Deflection
ACWFT With Symetric Elevon Deflection
Figure 49. Unstructured Surface Grid About Deflected Elevon.
Symmetric Elevon: 5 E = -20 ° Baseline: 6 E = 0 °
(M = 0.6, (_ = 15 0)
Cp
Figure 50. Surface Pressure Coefficient and Streamhne Traces for ACWFT
Configuration with Symmetric Deflected Elevon l'k.flect|ons of 0 and -20 Degrees, Mach
0.6, 15 Degrees Angle of Attack.
41
1.2 1.2
1.0
0.8
J_
0.6O
-J 0,4-
0,2
1.0
5 10 15 20
Angleof Attack(deg)
_E 0.8!o
0.6
E_ 0.4
0.2
0.0 0.0
0 0.0 0.1 0.2 0.3 0.4 o.s
Drag Coefficient
1.2
1.0
._ 0.8
i 0,6
_ 0.4
0.2
0.0 p
0.40 0.35 0.30 0.25 0.20
Pitching Moment Coefficient
Figure 51. CFD Force and Moment Comparisons With Test Data, ACWFT With -20
Degree Symmetric Elevon Deflection, Mach 0.6.
1.2
1.0
0.8
o
0.60
=0.4
0.2
0.0
5 10 15
Angle of Attack(aeg)
1.2 ]i
1,0 i
-- 0.8
Jfa
0.60E
0.4
0.2
0.0
20 0.0 0.1 0.2 0,3 0.4 0.5
Drag Coefficient
1.2 _-.................................... ]
1.0 _-
oeJR
°61S 0,4
0.2-
0,0 , _
0.40 0.35 0,30 0.25 0.20
Pitching Moment Coefficient
Figure 52. CFD Force and Moment Comparisons With Test Data, ACWFT With -20
Degree Symmetric Elevon Deflection, Mach 0.9.
1.2 1.2 1.2
1.0
0,8e.,
_o
=_ 0.6
=:0.4
0.2
0.0
1.0 1.0
I
__ 0.6i
I
0.6
0.4
0.2 0.2
m t
0.0 T 7 - 0.0 _ i
0.0 0.1 0.2 0,3 0.4 0.5 0.30 0,25 0.20 0,15 0.10
Drag Coefficient Pitching Moment Coefficient
O.B
0.6
0
0.4
5 10 15 20
Angleof Attack(deg)
Figure 53. CFD Force and Moment Comparisons With Test Data, ACWFT With -20
Degree Symmetric Elevon Deflection, Mach 1.2.
42
0.30 0.10 ] 0.4 ]
-- 0.08 / 0.3 t0.04 / i 0.2
0.10 0.02 _ / "_ ,.,t_ 0.1
ooo o.oo[ .._, _ " "=._=_o.o
°°21 \ I _o-o.1_-o.1o .0.041 \ T t-O2= -o.o6t i"1-0.20 -008_ ! ,,'_m_ I j _ -o3-0.30 -0.10 I -0.4
0 5 10 15 20 0 5 10 15 20 0 5 10 15 20
Angle of Attack ((:leg) Angle of Attack (deg) Angle of Attack (deg)
Figure 54. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
With -20 Degree Symmetric Elevon Deflection, Mach 0.6.
0.30 0.10 i 0.4
0.08
= I"n ._ o.os_ 0.10 _ 0,040.02 :_ 0.2
.._ 0.00 ' _ 0,00 EmI' -0.02
-0,10 ,_ .0.1
! ioo, ...--= -0.20 • • _ .0.06 [ _uam o
-- • NASTOEu_ c•0.08 1 -- -0.3
•0.30 -0.10 | -0.4
0 5 10 15 20 0 5 10 15 20 0 5 10 15 20
Angle of Attack (deg) Angle of Attack (deg) Angle of Attack (dog)
Figure 55. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
With -20 Degree Symmetric Elevon Deflection, Mach 0.9.
0.30
0.20o
0.10
,1=,,_ 0.00
|-0.10
¢: -0.20
-0.30
m111
B
0.10 |
0.081
0.06 l0.04
0.02
ooo- -.0.02
-0.04-
.0.10
0
0.4
_ 0.1
°iE. oo-0.1
i .0.2
-_ .0.3
-0.4
5 10 15 20 5 10 15 20 5 10 18 20
Angle of Attack (deg) Angle ol Attack (deg) Angle of Attack (deg)
Figure 56. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
With -20 Degree Symmetric Elevon Deflection, Mach 1.2.
43
(down)leftelevon
-20 o (up) right elevon deflection
ACWFT With
Figure 57. Unstructured Surface Grid About Asymmetrically Deflected Elevon andDeflected Body Flap.
Asymmetric Elevon:
8E= +20°/-20 °
(M = 0.6, = 15 °)
Baseline Afterbody Flap: 81== 90°%
0.5O.C
Figure 58. Surface Pressure Coefficient and Streamline Traces, for ACWFT Baseline,
Asymmetrically Deflected Elevon, and Deflected Afterbody Flap Configurations, Mach
0.6, 15 Degrees Angle of Attack.
44
1.2
1.0
0.8
_ 0.6
•,J 0.4
0.2
0.0
.................................... "1
5 10 15 20
Angle 04' Attack (deg)
1.2
T.O
._ 0.8
_ 0.6
_ 0.4
0.2
0,0
0.0 0.1 0.2 0.3 0.4 0.5
Drag Coefficient
1.2
1.0
._ 0.8o
_ 0.6
_ 0.4
0.2
0.0
0.20
i
o.15 o._o 0.05 o.oo
Pltchklg Moment Coefficient
Figure 59. CFD Force and Moment Comparisons With Test Data, ACWFT With
Asymmetric Elevon Deflection, Mach 0.6.
¢J
E.,..i
1.2 t'
1.0
0.8
0.6
0.4
0.2
0.0
1.2
1,0
0.8
_ 0.6
g0.4
1.2 ,I
1.0 i
0.8
_ 0.6
_ 0.4
o.2I ' _"_--0.0 [ '"
0.0 0.1 0.2 0.3 04 0.5
0.2
0.0
5 t0 15 2O 0.20 0.15 0.t0 0,05 0.00
Angle of Attack (deg) Drag Coefficient Pitching Moment Coefficient
Figure 60. CFD Force and Moment Comparisons With Test Data, ACWFT With
Asymmetric Elevon Deflection, Mach 0.9.
1.2
1.0
0.8
i 0.6
_ 0.4
0.2
o.o 10 5 10 15
Angle of Attack (dog)
1.2
1.0 ¸
i _ 0.8
i _ o.8
_J 0.4
o.2t
J0.0 I
20 0.0
Eu_
01 0.2 0.3 0.4
Drag Coefficient
0.5
1.2
1.0
0.8
06
0.4
0.2
o.o I0.10
I
0.05 0.00 -0,05 -010
Pitching Moment Coefficient
Figure 61. CFD Force and Moment Comparisons With Test Data, ACWFT With
Asymmetric Elevon Deflection, Mach 1.2.
45
0.30]-........................................i! o.oa°"°0.20 4 i E
0.10 0.04
__ 0.02
0.00 0.00
-0.10 -0.02
-0.10
0 5 10 15 20
Angle of Auack (deg)
0.05
i 0.040.03
- 0.02
0.01
.0.0t
-0.02
j .0.03.
_ -0.04
.0.05
0 5 10 15 20
Anglo of AtBck [deg)
0.05
i o.o410.03
0.02
I 0.01
g o
_ .0.01E
_ -0.o2
i .0,1_
4?.05
0.4
0.3
0.2
• i _=E o.1
i -0.1
!• w_sraeB - -0.3
-0.4
5 10 15 20 0 5 10 15 20
Ar_le of A_ek (¢_) _ of _et=¢=(¢_)
5 10 15 20
Angle of ,4Lttack (deg)
Figure 62. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
With Asymmetric Elevon Deflection, Mach 0.6.
o.3o ...........................................!
0.20
¢J
0.10
._ 0.00
-0 10
c .020
.030
010
i 008
i ¢, ._ 0.06
i _ o,o4I _ 0.02
ooo! i -002! -004
-0,08
-010
5 10 15 20
Atigle of Attlck ((leg]
005 7¢
004+
003 •
-- 0 32 •
001
I
-001>.
-0.02
i -0.03
E
_ -0.04
-0.05
I
0 5 10 15 20
Angle of Attack ((:leg)
0 05
003 ;
O02JOOti .0o'i
I•005 '
5 10 15 20
Angle of Attack (deg)
0 5 10 15 20
Angle of AXlack (deg)
0.4
,_ 0.3 t
0.2
:E
_'_ oo_-01 I *E-0.2 i
¢ -0.3 j-04
0 5 10 15 20
Angle of Attack (cleg)
Figure 63. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
With Asymmetric Elevon Deflection, Mach 0.9.
46
0.30
._ 0.20u
0.10o¢:
0.00
i -0.10
-- -0,20
-0.30
ig
!!
5 10 15 20
AngM of Attack (deg)
0.05 ................
| o,. iI 0.03 i
o 0.02 t
o.01
0 .... __--_t
-0,01
-0.02
-0.03
-0.04
-0,05
0 5 10 15 20
Angle of AtzJck (deg)
0.10
0.08
0.06
0.04
0.02
0.00
-0.02
-O,04
-0.06
-0.08
-0.10
I---I5 10 15 20
Angle of AttIck (dog)
0.05 !
0.04 t0.03
0.02 I •0.01
E
| -0.°21
I '03 t--'' IS -0.04 • NAS'rDEder•0.05 ' " I
0 5 10 15 20
Angle of Atlack (deg)
0.4 .................
°_02
io,
_ -0.2
--_ -0.3
-0.4
0 5 10 15 20
Angle ol Anlck ((leg)
Figure 64. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
With Asymmetric Elevon Deflection, Mach 1.2.
1.2
1.0
. 0.8c.,t
0.6
O
0.4
0.2
0.0
o 5 10 15 20
Angle of Attack (deg)
1.2 ...................... _ 1.2
1.0 I i i 1.0o.5 i "E 0.8
!°°I °'-J 0.4 I 0.4
/
°21 I 02_dm
0.0 0,0
0.0 0.1 0.2 0.3 0.4 0.5 0.20
Drag Coefflckmt
o.15 o.1o o.o5 o.oo
Pitching Moment Coefficient
Figure 65. CFD Force and Moment Comparisons With Test Data, ACWFT With
Deflected Body Flap, Mach 0.6.
47
1,2 .......................................
1'0 B_0.8
0,6
"_ 0"4 t
0.2
0,0 ,
1.2
1.0
-- 0.8
_ 0.4-
0,2-
1.2 ............................
1.0
_ 0.8
i 0,6
0.4
0.2
0.00.0 L b I _ i
0 5 10 15 20 0.0 0.1 0.2 0.3 0.4 0.5 0.20 0.15 0.10 0.g5 0.00
Angle of Attack (deg) Drag Coefficient Pitching Moment Coefficient
Figure 66. CFD Force and Moment Comparisons With Test Data, ACWFT With
Deflected Body Flap, Mach 0.9.
1.2 ¸
1.0
_ 0.8c
0.6
S 0.4
0.2
0.0 ,
0 5 10 15
_gkD of Attack (dag)
1.2 ........................................ ,
20 0.5
1.0 _"
0.8 ÷
0.6
0,4
0.2 "t
0.0
0.0 0.1 0.2 0.3 0.4
Drag Coefficient
1211.0 I
0.8
0.6
0.4
0.2
0.0
0.10 0.05 0.00 -0.05 -0.10
Pitching Moment Coefficient
Figure 67. CFD Force and Moment Comparisons With Test Data, ACWFT With
Deflected Body Flap, Mach 1.2.
48
0,30 ......................................
E 0.20
i 0.10o
-0.I0
--_ -02D
-0.30
5 10 15 20
Angle of Attack (deg)
0.05 ............................
0.04
0,03
0.02
0.01 ;0
>_ .0.01
"i -0.02
-0.03
-0.04
.0.05
0 5 10 15 20
Ang4e of _tack (oog)
0.10
0.08
0.06
0.04
0.020.00
.0.02
-0.04
o: I .,-°,,.I-0.10
5 10 15 20
Angle of Attack (de9)
0.05 |
::o°;I0.02
0.01
-0.01
.0.02 1
.o.o3I-- .0.05
0 5 10 15 20
Ang_ of Attack (deg)
0.4 ...............................
i 0.30.2
=_ o.1
_ o.o_o -oij.o:
--_ -0.3
-0.4
t
5 10 15 20
Angle of Attack (deg)
Figure 68. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
With Deflected Body Flap, Mach 0.6.
0,30 ................
0.20
0.10t.J
0.00 ------4-------+---------
-0.30 .......................
5 10 15 20
Angle of Attack (deg)
0.05=
0.04
:_ 0.03o- 0.02
o.01
_ -0.01
_ -O.02
-0,03|.0.04
!;
-0.055 10 15 20
Angle of Attack (dog}
0.10 ,
0.08 _'
oo,t:.°0:L_o:oo
..0.06 1 _*"'_'-- .0.08 ; ! I _' I
0 5 10 15 20
Angle of Atlack (deg)
0.04 ÷
0.03 t
°°2I Io.o,_ /
-0.01 _ I
•0.02 4
•0.03 Jl
_lWtdatl I.0.04 i II _110 F.uler
-0.05(0 5 10 15 20
Angkl of AnaCk (dog/
0,4 ,
j:::I
I! 0.1
0,0
-0,1
-0.2
-- -0.3 ¢
-0,4 [0 5 10 15 20
Angle Of Atlack (deg)
Figure 69. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
With Deflected Body Flap, Mach 0.9.
49
0.30
0.20
0.10cJ¢=
..= 0.00
i -0.10
--_-0.20
-0.30 ...................................
5 10 15 20
Angle of Aeack (e_g)
0,0S -
-_ 0.04.
0.03
0.020.01
J __ 0 .... _
_ 1.01
_ 41.02
-0.03-0.04
_).055 10 15 20
Angle of Allz_k (deg)
0.10
0.08
.._ 0.06
0.02
0.00 _
i -0.02
.0.04
•0.06 _u_== I
•0.08 , I _,,¢Io =''-' I-0.10
5 10 15 20
Angle of Anack (deg)
: 0.05 ]-i oo4t l--u== I
io.o3t I " N_Eu=I;oo t
o.ol _
_ o_
-0.0t [
-0.02 1
-0.05 I
0 5 10 15 20
Angle of Attack (deg)
0.4
=_ 0.1
E_ o.o
O .0.1-0.2
--_ .0.3
-0.4
0 5 10 15 20
Angle of Attack (deg)
Figure 70. CFD Incremental Force and Moment Comparisons With Test Data, ACWFT
With Deflected Body Flap, Mach 1.2.
50
REPORT DOCUMENTATION PAGE For,.Ap_o,.dOMB No. 0704-0188
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! 1. AGENCY USE ONLY(Leave blank) i 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED
I March 1998 Contractor Reporti4. TITLE AND SUBTITLE 15. FUNDING NUMBERS
Euler Technology Assessment for Preliminary Aircraft Design -
Unstructured/Structured Grid NASTD Application for Aerodynamic
Anal_'sis of an Advanced Fighter{Tailless Configuration
6. AUTHOR(S)Todd R. Michal
7. PERFORMING ORGANIZATION NAME(S) AND ADORESS(ES)
Boeing Company
St. Louis, Missouri
9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)
National Aeronautics and Space Administration
Langley Research Center
Hampton, VA 23681-0001
NAS1-20342, Task 13
WU 522-22-11-01
8. PERFORMING ORGANIZATIONREPORT NUMBER
10. SPONSORING/MONITORINGAGENCY REPORT NUMBER
NASA/CR-1998-206947
11. SUPPLEMENTARY NOTES
Langley Technical Monitor: Farhad Ghaffari
12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE
Unclassified-Unlimited
Subject Category 02
Distribution: Standard
Availability: NASA CASI (301) 621-0390
13. ABSTRACT (Maximum 200 words)
This study supports the NASA Langley sponsored project aimed at determining the viability of using Euler
technology for preliminary design use. The primary objective of this study was to assess the accuracy and efficiency
of the Boeing, St. Louis unstructured grid flow field analysis system, consisting of the MACGS grid generation andNASTD flow solver codes. Euler solutions about the Aero Configuration/Weapons Fighter Technology (ACWFT)
1204 aircraft configuration were generated. Several variations of the geometry were investigated including a standard
wing, cambered wing, deflected elevon, and deflected body flap. A wide range of flow conditions, most of which
were in the non-linear regimes of the flight envelope, including variations in speed (subsonic, transonic, supersonic),
angles of attack, and sideslip were investigated. Several flowfield non-linearities were present in these solutions
including shock waves, vortical flows and the resulting interactions. The accuracy of this method was evaluated by
comparing solutions with test data and Navier-Stokes solutions. The ability to accurately predict lateral-directional
characteristics and control effectiveness was investigated by computing solutions with sideslip, and with deflected
control surfaces. Problem set up times and computational resource requirements were documented and used to
evaluate the efficiency of this approach for use in the fast paced preliminary design environment.
14. SUBJECT TERMS
Computational Fluid Dynamics, Euler Formulation, Preliminary Aircraft Design,
Advanced Fighter Design, Vortical Flows, Unstructured/Structured Grid NASTD
17. SECURITY CLASSIFICATIONOF REPORTUnclassified
¼SN 1540-_)1-280-5500
18. SECURITY CLASSIFICATIOhOF THIS PAGEUnclassified
19. SECURITY CLASSIFICATIOI_OF ABSTRACTUnclassified
IS. NUMBER OF PAGES
55
16. PRICE CODE
A_420. LIMITATION
OF ABSTRACT
_tandard Form 298(Re¥. 2-89)PrescribedbyANSIStd Z39-18298-102