NASA/TM-1999-209526

Unstructured Grid Euler Method

Assessment for Longitudinal and Lateral/

Directional Stability Analysis of the HSR

Reference H Configuration at Transonic

Speeds

Farhad Ghaffari

Langley Research Center, Hampton, Virginia

December 1999

https://ntrs.nasa.gov/search.jsp?R=20000011602 2018-05-14T14:07:26+00:00Z
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NASA / TM-1999-209526

Unstructured Grid Euler Method

Assessment for Longitudinal and Lateral/

Directional Stability Analysis of the HSR

Reference H Configuration at Transonic

Speeds

Farhad Ghaffari

Langley Research Center, Hampton, Virginia

National Aeronautics and

Space Administration

Langley Research Center

Hampton, Virginia 23681-2199

December 1999
Available from:

NASA Center for AeroSpace Information (CASI)

712! Standard Drive

Hanover, MD 21076-1320

(301) 621-0390

National Technical Information Service (NTIS)

5285 Port Royal Road :_

Springfield, VA 22161-2171

(703) 6O5-6OO0
Unstructured Grid Euler Method Assessment for Longitudinal and Lateral/Directional

Stability Analysis of the HSR Reference H Configuration at Transonic Speeds

Farhad Ghaffari

NASA Langley Research Center, Hampton, VA

ABSTRACT

Transonic Euler computations, based on unstructured grid methodology, are performed for a proposed

High Speed Civil Transport (HSCT) configuration, designated as the Reference H configuration within the

High Speed Research (HSR) Program. The predicted results are correlated with appropriate experimental

wind-tunnel data for the baseline configuration with and without control surface deflections for a range of

angle of attack at Moo = 0.95. Good correlations between the predictions and measured data have been

obtained for the longitudinal aerodynamic characteristics on the baseline configuration. The incremental

effects in the longitudinal aerodynamic characteristics due to horizontal tail deflections as well as wing

leading-edge and trailing-edge flap deflections have also been predicted reasonably well. Computational

results and correlations with data are also presented for the lateral and directional stability characteristics

for a range of angle of attack at a constant sideslip angle as well as a range of sideslip angles at a constant

angle of attack. In addition, the results axe presented to assess the computational method performance and

convergence characteristics.

bref

CD

COo

CL

Q

Cm

Cp

Moo

P

Po_

qoo

S_eS

c_

(Sht

_Ze/

8te f

SYMBOLS

reference span, inches

reference chord, inches

drag coefficient, Drag/(qoo S_e/)

skin friction drag at CL = 0

lift coefficient, Lift/(qoo S_/)

rolling moment coefficient, Rolling moment/(qoo S_¢] b_/)

pitching moment coefficient, Pitching moment/(qoo S_/_)

Yawing moment coefficient, Yawing moment/(qoo S_el b_e/)

pressure coefficient, (p - Poo)/qoo

freestream Mach number

static pressure

freestream static pressure

freestream dynamic pressure

Reynolds number per foot

reference area, sq. inches

angle of attack, degrees

sideslip angle, degrees

horizontal tail flap deflection angle, degrees

inboard/outboard wing leading-edge flap deflection angle, degrees

inboard/outboard wing trailing-edge flap deflection angle, degrees

Abbreviations

HSCT

HSR

IGES

LaRC

High Speed Civil Transport

High Speed Research

Initial Graphics Exchange Specification

Langley Research Center
COMPUTATIONAL OBJECTIVE AND APPROACH

The primary objective of the computations is to provide an accuracy assessment of an unstructured grid

Euler method known as USM3D (Ref. 1) for predicting the longitudinal and lateral/directional aerodynamic

characteristics for the baseline configuration and the corresponding effects due to control surface deflections

at transonic speeds. " ...........

A computational matrix is developed to complement the configurations and the conditions tested in

the NASA LaRC 16-Foot Transonic Tunnel. The wind tunnel test was conducted on a 1.675% model (also

referred to as controls model) at an average Reynolds number of about 4 million per foot. Three configurations are selected for the computational method assessment study. Consistent with the experimental

test, each configuration incorporated an extended circular aftbody to represent the wind-tunnel model sting

apparatus and appropriate boundary conditions at the nacelle's inlet and exhaust planes to simulate the

flow through propulsion effects. The selected configurations are:

1) transonic cruise (i.e., 5tel = 0/10; _tef = 0/3; 5hi ---- 0). This geometry is considered as the baseline

transonic cruise configuration.

2) transonic cruise with deflected horizontal tail (i.e., 5zef = 0/10; 5tel = 0/3; 6ht = 6).

3) supersonic cruise configuration where all the control surface deflection angles are set to zero (i.e.,

5_ s = 0/0; _te_ = 0/0; _t = 0).

The above nomenclatures are defined as: the Greek letter (f represents control surface deflection angle; the

subscripts leS, tel, and ht denote the leading-edge flap, trailing-edge flap, and horizontal tail, respectively.

if applicable, the control surface deflection angles, separated by a slash, correspond to those of the inboard

and outboard wing, respectively. For example, 51el = 0/10 denotes that the inboard- and outboard-wing

segmented leading-edge flap deflection angles are set to 0 ° and 10°, respectively. The positive control surface

deflection angles ((_lef, 5te$, and (_ht) are downward ....

The corresponding flow conditions for the selected configurations are then identified from the experimental data analyses at M_ = 0.95. Zero sideslip analyses are performed for all three configurations at

a = 0°, 4 ° , and 8° while the finite sideslip matrix composed of an a-sweep (0 °, 4 °, 8 °, 11 °) at f_ = 3 ° and a

fLsweep (0 °, 3°, 6 °) at c_ = 4 °. The Selected computational matrix is depicted graphically in a chart form

(figure 1) to illustrate the links between Various configurations and the corresponding flow conditions. It

should be noted that the transonic experimental data with finite sideslip angles were only available for the

supersonic cruise configuration when the present computational analysis was initiated. In addition, the

experiment was conducted to only measure the overall forces and moments.

COMPUTATIONAL GRID

The HSR Reference H initial surface geometry was defined in a format known as Initial Graphics

Exchange Specification (IGES), (Ref. 2), which served as the database for all the subsequent grid generation

processes. The GRIDTOOL (Ref. 3) program was used to read in the initial IGES geometry definition and

the corresponding discretization into surface patches. The resulting surface patc-hes-were then read in the

VGRID (Ref. 4) program to generate the initial surface triangulation by the advancing front method. The

initial surface triangles were then read back into GRIDTOOL for the projection onto the initial database

to preserve the proper surface curvature within the interior of a given patch. As a final step, the resulting

front on the surface and the farfield boundary patches were used by VGRID to generate the volume grid

within the computational domain along with the corresponding grid and face connectivity files. A nearfield

and a closeup view of a typical grid, used to compute the flow for the transonic cruise configuration with

deflected horizontal tail (i.e., (ilei = 0/10; (ft_f = 0/3; (fht = 6), are shown in figures 2 and 3 from two

different perspectives. The figures illustrate the triangular meshes on the surface, plane of symmetry, and

the wind-tunnel sting representation. The farfield boundaries, clipped in the figures for clarity, are extended

to about 6-longitudinal and 3-radial body lengths away from the surface.-The computational grid size for

each of the three configurations are as follows: the transonic cruise configuration had about 1.023 million

cells and 36,000 surface triangles; the transonic cruise with deflected horizontal tail configuration had 0.822

2
million cellsand44,000surfacetriangles;andthesupersoniccruiseco_guration had0.997million cellsand

47,000surfacetriangles.Thegrid for thesupersoniccruiseconfigurationwasmirroredaboutthesymmetry

planeto generatethe computationalgrid for all the finite sideslipruns. The resultinggrid approximately

doubledthe grid sizeusedfor the zerosideslip(only one-halfmodeled)computations.

ALGORITHM PERFORMANCE AND CONVERGENCE

USM3D computations were performed on both the Numerical Aerodynamic Simulation (NAS) CrayC90 located at NASA Ames, and the NASA Langley Cray-YMP. All the zero sideslip computations were

conducted on the Cray-C90 using the implicit time integration scheme and the corresponding finite sideslip

cases were executed on the Cray-YMP by using the explicit algorithm. Also, no flux limiters were used to

obtain the solutions presented here. As a general strategy, solutions based on the second-order accurate flux

difference splitting (FDS2) method were first sought; however, for certain cases only the first-order accurate

flux difference splitting (FDS1) could be achieved. In general, the aerodynamic effects due to the difference

in the order of solution accuracy were found to be small, particularly in terms of total forces and moments

and will be discussed later.

The selected computational matrix, along with the corresponding numerical method employed to converge a given solution for each configuration, are summarized in Table 1. In this table, symbols '_/' and

'X' indicate that a solution convergence 'was' and 'was not' achieved, respectively. For zero sideslip cases,

converged solutions with FDS2 were obtained for all configurations at a = 0 ° and 4 °, except for the supersonic cruise configuration at a = 4° where only the FDS1 solution could be achieved. At higher angles of

attack (i.e., a > 4°), no solution convergence based on FDS2 could be achieved for any of the configurations

investigated. It is speculated that the lack of solution convergence could be due to the flow separation

occurring within the boundary layer over the round leading-edges of the inboard wing which is known to

require Navier-Stokes simulation to resolve. As a result, the numerical solutions for a -- 8 ° are obtained

based on FDS1 for all the configurations except for the supersonic cruise configuration for which no solution could be achieved. The supersonic cruise configuration was the most challenging geometry to obtain

solution convergence at the present transonic flow conditions. The solution convergence difficulties were

identified to have been caused primarily by geometrical features particularly with respect to the transition

region between the inboard and the outboard wing. This transition region represents a step discontinuity

from a round to a sharp leading edge of the undeflected flap on the outboard wing. Progress has been made

recently to implement various flux limiters in USM3D as numerical dumping mechanisms to alleviate such

convergence difficulties. As mentioned earlier, the supersonic cruise configuration was the only geometry for

which experimental data were available at transonic speeds with finite sideslip angle. As a result, with the

lessons learned from the zero sideslip computations, only FDS1 solutions were sought for the finite sideslip

analysis. A typical zero sideslip case required about 500 cycles to converge and, for finite sideslip cases,

required about 1000 cycles. The implicit time integration scheme required ,_ 180 words of memory per cell

and _ 30 p-secondcellcycle on the Cray-Cg0. Similarly, the explicit scheme required _ 46 words of

memory per cell and _ 6 p-second/cell/cycle on the Cray-C90. Very little overhead CPU time, generally

less than 5%, is found to be associated with the second-order accurate solution, compared to the first-order

formulation.

Typical convergence characteristics (i.e., Residuals and Lift) for solutions based on FDS2 and FDS1

at a = 0 °,/3 = 0°, for the supersonic cruise configuration are shown in Figures 4(a) and 4(b), respectively.

A typical second-order solution is initiated automatically within USM3D after a certain level of residual

convergence has been achieved by the first-order accurate flux formulation. For example, as shown in figure

4(a), the FDS2 solution is initiated from the FDS1 solution (after 205 iterations) where the residuals have

already dropped 2.5 orders of magnitude. Similarly, figure 4(b) shows the residual and lift characteristics for

the fully converged solution based on the FDS1 formulation. These two particular solutions were obtained

as part of the solution accuracy assessment between first- and second- order flux differencing schemes. A

similar set of convergence history plots are shown in figure 5 for typical solutions obtained for the supersonic

cruise configuration for two different sideslip angles at a = 4 °.

3
RESULTS AND DISCUSSIONS

Surface Pressure Coefficients:

Typical surface pressure coefficient contours computed at a = 4 °, Moo = 0.95, and /3 = 0° for the

transonic cruise, transonic cruise with deflected horizontal tail, and the supersonic cruise configurations,

are shown in figures 6, 7 and 8, respectively. All surface pressure coefficients are contoured over a constant

range from 0.6 to -1.4 with 0.067 increments (i.e., Cp,_,_x = 0:6, Cprn_,_ = -1.4, ACp = 0:067) and shown

from the same vantage point. The surface pressure contours computed for the transonic cruise configuration,

shown in figure 6, serve _as the baseline solution to be correlated with the results obtained for the transonic

cruise configuration with deflected horizontal tail and the supersonic cruise configuration. The latter two

correlations show the effects of horizontal tail deflection and the wing leading-edge and trailing-edge flap

deflections on the computed surface pressure coefficients, respectively. Furthermore, since no experimental

surface pressure data is available, these correlations are only: intended for relative comparisons between

computational results to show, qualitatively, the effects due to a particular control surface deflection.

The surface pressure coefficients Computed for the baseline transonic cruise configuration (figure 6)

can be characterized by two dominant flow features, one that can be attributed to typical flow expansions

caused by the physical geometry and the other to shock waves due to the transonic flow effects. The typical

fl0w expansions on the upper Surface (figure 6(a)) occur at the leading edges, particularly on the inboard

wing, and the hinge lines associated with the deflected leading-edge and trailing-edge flaps on the outboard

wing. The transonic flow shock waves (i.e., highlighted by a concentration of several pressure contour lines)

on the upper surface appear in two locations. The first location is very close, nearly spanning in parallel, !

to the entire wing trailing edge, and the other location is Situated approximately along the mid-chord,

spanning the horizontal tail and extending to the fuselage and onto the vertical tail. The lower surface

pressure contours (figure 6(b)) indicate fairly benign flow characteristics with the exception of a shock wave

that appears in between the two nacelles and the compression rings on the wing just ahead of each nacelle

caused by the pylons. :........ _ __

The comparative assessments of the results shown in figures 6 and 7 indicate only very small effects

due to_t-he'horizontal _tail clei]ectlon'0n the computed Surface pressure coeffi_nt: cont6_S for the transonic

cruiseconfigUration. The most not-iceab]e change appears to be the flip of the shock wave from the tipper

surface of the undeflected horizontal tail to the lower surface of the deflected horizontal tail. Similarly, a i

Comparison of the"resultsshoWn in figures 6 and 8 indicate the effects of wing leading-edge and trailing'edge i

flap deflections on the computed surface pressure coefficient contours for the transonic cruise configuration.

As expected, the comparison indicates that the hinge line flow expansions disappear on the upper surface of

the undeflected wing leading-edge and trailing-edge flaps. In addition, the shock waves on the wing upper

surface near the wing trailing edge and the horizontal tail, for the supersonic cruise configuration, are not

as crisply predicted as those of the transonic cruise configuration. This smearing of the shock waves is

speculated to have been caused primarily by the excess dissipation generally associated with the first-order :

accurate solutions. The comparison shows very little difference on the computed lower surface pressure

coefficient contours due to wing leading-edge and trailing-edge flap deflections.

Forces and Moments and Correlations with Data: .:

:: All predicted inviscid drag Coefficients have been corrected with an appropriate CDo to account for

the skin friction using the experimental data obtained in the NASA LaRC i6-Foot Transonic Tunnel. This

Coo was determined by a linear extrapolation from the plotted C_ vs. CD curve for the supersonic cruise

configuration as shown in figure 9. This Coo provides a minimum skin friction drag Which, by definition, is

independent of the configuration camber and twist. Through this analysis, a CDo of 0.0126 was obtained,

which was subsequently added to all the inviscid drag predictions.

Solution method assess .mcnt - As discussed earlier, the supersonic cruise configuration was the most

challenging geometry for ol_tai_ning solution convergence at thepresent transonic Mach number. As a result,
the FDS2solutioncouldonly beachievedfor a = 0 °, and the FDSi Solution could only be achieved for

= 4°. This prompted a solution method assessment study to be performed at a = 0 ° for which both firstand second- order solutions could be obtained. The results from this study are presented in figure 10 for

the longitudinal aerodynamic characteristics. The correlations at a = 0° indicate that the solution order of

accuracy has only a small effect on the predicted lift and drag, with relatively more pronounced effect on

the pitching moment estimate.

Horizontal tail deflection effects - The predicted and measured longitudinal aerodynamic effects due to

the horizontal tail deflection at Moo = 0.95 and/3 = 0° are shown in figure 11. The measured aerodynamic

effects due to the horizontal tail deflection on both the lift and drag coefficients appear to be small. The

pitching moment characteristics remain nearly the same over the measured range of flow conditions, and

as expected, the horizontal tail deflection causes roughly a constant positive shift in the pitching moment

coefficients. The overall lift and drag coefficients have been predicted reasonably well with the Euler method

both in terms of incremental effects due to the horizontal tail deflection and magnitudes. Although the

overall magnitudes of the pitching moment coefficients are not predicted as well, the incremental effects due

to the horizontal tail deflection are predicted reasonably well.

Wing leading-edge and trailing-edge flap deflection effects - The predicted and measured longitudinal

aerodynamic effects due to the leading-edge and trailing-edge flap deflection at Moo = 0.95 and/3 = 0°

are shown in figure 12. The measured aerodynamic effects due to the leading-edge and trailing-edge flap

deflections appear to be small for both the lift and drag coefficients and more pronounced for the pitching

moment characteristics particularly at a < 10°. These aerodynamic effects due to the leading-edge and

trailing-edge flap are reasonably predicted both in terms of overall magnitudes and trends.

Finite sideslip stability analysis - The longitudinal aerodynamic characteristics computed for an asweep (i.e., 0 °, 4°, 8°, and 11°) at/3 = 3 ° and Mc_ = 0.95, along with the experimental measurements for

the supersonic cruise configuration, are shown in figure 13. The correlations indicate excellent agreement

between the computed and measured data throughout the examined range of flow conditions for both

•magnitudes and trends.

The lateral/directional stability characteristics computed for a/_-sweep (i.e., 0 °, 3°, and 6°) at a = 4 °

and Moo = 0.95, along with the experimental measurements for the supersonic cruise configuration, are

shown in figure I4. The correlations indicate a good agreement for the pitching and rolling moments but

not as good for the yawing moment, particularly with respect to the slope (i.e., C,_Z).

CONCLUDING REMARKS

The application viability of an unstructured grid Euler method (USM3D) for predicting the longitudinal

and lateral/directional stability characteristics of the HSR Reference H configuration at a transonic Mach

number of 0.95 have been addressed. Particular emphasis was placed on assessing the ability of the method

to provide consistent and accurate solutions of forces and moments for the baseline configuration with

different control surface deflection settings for the horizontal tail as well as for the leading-edge and trailingedge flaps. The analysis revealed that a consistent set of solutions based on the first-order accuracy was

achievable for a wider range of applicability (in terms of configurations and flow conditions) where the

second-order accurate solutions could not be obtained. The solution method assessment analysis indicated

that the first-order accurate solutions provide comparable results to those of the second-order, particularly

for the aggregate forces and moment characteristics. The computed longitudinal aerodynamic effects due to

control surface deflections and sideslip angles agree reasonably well with the measured data both in terms

of magnitude and trends. Finally, the lateral/directional stability analysis revealed fairly good agreement

between the predictions and measured data for the pitching and rolling moment characteristics but not as

good for the yawing moment. It is believed that the present unstructured grid Euler method is certainly a

viable engineering tool that can be utilized with confidence in the early aerodynamic design and analysis of

the proposed high speed civil transport configuration.
REFERENCES

1 Frink, N. T., "Upwind Schemefor Solvingthe EulerEquationson UnstructuredTetrahedralMeshes,"

AIAA Journal,Vol. 30,No. 1, 1992,pp. 70-77.

2 Smith, B. M.; Brauner,K. M.; Kennicott, P. R.; Liewald,M.; and W'ellington,J.: Initial Graphics

ExchangeSpecification(IGES),Version2.0. NBSIR-82-2631-AF, Feb. 1983. (Available from NTIS as PB

83-137448.)

3 Abolhassani, J. "GRIDTOOL: A Surface Modeling and Grid Generation Tool," NASA-CP-3291, May

1995, pp. 821-832.

4 Parikh, P., Pirzadeh, S., and Lohner, R., "A Package for 3-D Unstructured Grid Generation, FiniteElement Flow Solution and Flow Field Visualization," NASA CR-182090, Sept. 1990.

O

FDS2

4 ° 8 °

FDS 1

0 ° 4 ° o

Transonic cruise

Transonic cruise with 5ht = 6

Supersonic cruise

Table 1.-

4

-4

,4

",1 X

4 x

X X 4

-- 4

4 x

Computational matrix - method of convergence.

Sweep,:oo I Io _ ,xeePe 21iSoi&,o: 4o30 [ "

Figure 1.-....Selected computational matrix - geometries and conditions.
(a) Nearfietd.

(b) Closeup.

Figure2.- Typical computationalgrid - Uppersurfaceview. _lef _-- 0/10, _tef = 0/3, Sht = 6.

7
\

(b) Closeup.

Figure 3.- Typical computational grid - Lower surface view. _lef = 0/10, *t,f = 0/3, *ht = 6.
FDSI <-

"-_ 0

-2

v

_-I --3

0

-> FDS2

i ' i

I i I , I i

200 400 600 800

0

-1

-2

-3

-4

I ' I ' I '

, I , l i I i

0 I00 200 300 400

C L

0.00

-0.02

-0.04

-0.06

-0.08

' I ' I ' I ' 1

0.00

-0.02

-0.04

-0.06

-0.08

' I '

, I , I , I ,

, I , I ,,, I

0 200 400 600 800 0 I00 200 300

Iterations ]t.erations

(a)FDS2. (b)FDSI.

400

Figure 4.- Typical convergence characteristics - _lef = 0/0, _fef = 0/0, _hf = 0; Moo = 0.95, a = 0°,/9 = 0°.

0

2

O_

o

--3

0.0

' I ' I ' I ' I '

f , I , I , I , I ,

1 ' I ' I ' I ' I '

0

-1

-2

-3

0.0

xl03 , , , , , , , , , xl03

0.2 0.4 0.6 0.8 1.0 0.2 0.4 0.6 0.8 1.0

C L

1.0

0.8

0.6

0.4

0.2

0.0

I I I ' i ' 1.0 I I ' I ' I

0.8

0.6

0.4

0.2

0.0

, i , I , i , , , xl03 , I , i , i , i , xl03

0.2 0.4 0.6 0.8 1.0 0.2 0.4 0.6 0.8 1.0

Iterations Iterations

(a) _ = 3° . (b) _ = 6°.

Figure 5.- Typical FDS1 convergence characteristics - 6_f = 0/0, 6tel = 0/0, _ht = 0; Moo = 0.95, a = 0°.
(a) Upper surface.

=

Figure 6.-

(b) Lower surface.

Computed surface Cp contours - Transonic cruise geometry. Moo 0.95, (_ = 4 °, _ = 0°.

10 !
(a) Upper surface.

(b) Lower surface.

Figure 7.- Computed surface Cp contours - Transonic cruise geometry with 5ht = 6. M_ = 0.95, a = 4 °, _ = 0%

11
(a) Uppersurface.

Figure8.-

(b) Lower surface.

Computed surface Cp contours - Supersonic cruise geometry. M_¢ = 0.95, _ = 4 °, _ = 0 °.

12
0.020

0.018

0.016

Co

0.014

0.011

0.01C

0.00

unnel

Cf / _lation to exclude configuration

-""" camber and twist contribution to zero-lift drag

."< Coo value of 0.0126 is used to correct the Euler

drag predictions to account for the skin friction

2 0.02 0.03 0.04

0.01 CL

Figure 9.- CDo determination strategy- Supersonic cruise geometry, Moo = 0.95

0.051

0.02:

Cm

0.0_

-0,025

-0.o_

%

5o :o::5 b._o' 0._5 o._o 0._5 t._o

CL

1.00 I20

Experiment USM3D

(LaRC16'T'r) VDS2 R)S]

[] • •

t Experimental to the inviscid CDo predictions has been added

0,50

CL

-0.00 0.50 -0.00

DOO DE]

-°_.?_oo o.h.... o.og.... o.h " o.'_6" o._o

(z, deg CD

Figure 10.- Solution method accuracy assessment - Supersonic cruise geometry, M_ = 0.95, 3 = 0 °-

13
0.05_

O.O'_d

Cm

0.000

-0.02

-o_o_ _o.25 o.oo o.2s 0.50 o.V5

CL

'l._O

1.00 1.00

Experiment (t_c _6.Tn USM3D

8_ = 60 [] •

8.t = 0 ° O •

Experimental C_ has been added

to the inviscid predictions

0.50

CL

-0.00

.o.5.%o 0.20

0.50 -0.00

o._ " _5._o" ,oloo ....,51oo -°'_._o

a,deg

• - "0.{M 0.08 '0.12_ ' '0.16'

CD

Figure 11.- Horizontal tail deflection effects and predictions - Transonic cruise geometry, Moo = 0.95,/3 = 0°.

0.05_

0.02:

Cm

O.(XX

-0.02.:

-0.05_ _o :o.'25 b_ " 0).5 0.k'" b.7i " ',._o

CL

I.(30

Experiment USM3D

(LaRC 16' "IT)

51el=0/0 , _tef=0]0 O •

5]¢f---'0/10; _Stef--O/3 O •

Experimental Coo has been added l

to the inviscid predictions J

1.00

E L

0.50 0.00

-0"5t_:00 .... 0.(30 ..... 5,00 .... 10100 .... 15100

oc,deg

0.50 ( (_/3:32(_ (I13 0.00

-058._o" o.o_ o.o_ 'o.;_ o._ 0.20

CD

Figure 12.- Leading/trailing-edge flap deflection effects and predictions - Transonic cruise geometry, M_ = 0.95, fl =

14
E L

0.05C

0.025 (__(_

Cm

0.017

-0.02

°°-_5o -0.'25 o._ " 0.'.,5 0.;o 0.7i '_._o

CL

1.00

-0.00

-0"-_:oo ....o._o'''5._o ...._oloo' ' );oo

a,deg

Experiment USM3D

(L_RC16'rr_ (mS1)

O •

!

Experimental Ct_ has been added [

to the inviscid predictions J

1.00

o.2o

Figure 13.- Longitudinal aerodynamic characteristics and predictions - Supersonic cruise geometry, Moo = 0.95,/3 = 3 °.

igure 14.-

°°21 Pitching Moment

o.ol Experiment USM3D

(LaRC 16' Tr) (FDSI)

Cm°°°[ 0 0 qbO 0 C?O 0 _0 0 % 0 •

"0"0126!0

,o.o "":4'.0 ...S.o_"o'.o....21o....;.o....6:0

C I

0.03

0.02

0.0_

020

-0.01

-0.02

-0,036. 0

Rolling Moment ]

00 _00 _

"0 00IDo O OD

'''-4'.o ...-_.o....olo' f.o....,/.o"L_6:0

13,deg

Cn

0.03

0,02

0.01

0,00

-0.01

-0.07

-0.0_¥ 0

Yawing Moment 1

O O

O

O

O

O

oOqD

0

" " -/.o.... -_.o " o'.o.... i.o.... ,io.... 6'.0

13,deg

Lateral/directional stability characteristics and predictions - Supersonic cruise geometry, Moo = 0.95, a = 4 °.

15
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December 1999 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Unstructured Grid Euler Method Assessment for Longitudinal and Lateral/

Directional Stability Analysis of the HSR Reference H Configuration at WU 537-07-20-21

Transonic Speeds WBS 4.3.1

6. AUTHOR(S)

Farhad Ghaffari

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

NASA Langley Research Center

Hampton, VA 23681-2199

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National Aeronautics and Space Administration

Washington, DC 20546-0001

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L-17927

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NASA/TM-1999-209526

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13. ABSTRACT (Maximum 200 words)

Transonic Euler computations, based on unstructured grid methodology, are performed for a proposed High Speed

Civil Transport (HSCT) configuration, designated as the Reference H configuration within the High Speed

Research (HSR) Program. The predicted results are correlated with appropriate experimental wind-tunnel data for

the baseline configuration with and without control surface deflections for a range of angle of attack at Mo_ = 0.95.

Good correlations between the predictions and measured data have been obtained for the longitudinal aerodynamic

characteristics of the baseline configuration. The incremental effects in the longitudinal aerodynamic characteristics due to horizontal rail deflections as well as wing leading-edge and trailing-edge flap deflections have also been

predicted reasonably well. Compuational results and correlations with data are also presented for the lateral and

directional stability characteristics for a range of angle of attack at a constant sideslip angle as well as a range of

sideslip angles at a constant angle of attack. In addition, the results are presented to assess the computational

method performance and convergence characteristics.

14. SUBJECT TERMS

Computational fluid dynamics; Euledinviscid method; High speed research/High

Speed Civil Transport Reference H; Unstructured grid; Transonic speeds; Longitudinal

and lateral/directional

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