NASA/CR{2001{211266

An Inviscid Computational Study of

Three '07 Mars Lander Aeroshell

Congurations Over a Mach Number

Range of 2.3 to 4.5

Ramadas K. Prabhu

Lockheed Martin Engineering & Sciences Company, Hampton, Virginia

December 2001
The NASA STI Program Oce ... in Prole

Since its founding, NASA has been dedicated

to the advancement of aeronautics and space

science. The NASA Scientic and Technical

Information (STI) Program Oce plays a key

part in helping NASA maintain this

important role.

The NASA STI Program Oce is operated by

Langley Research Center, the lead center for

NASA's scientic and technical information.

The NASA STI Program Oce provides

access to the NASA STI Database, the

largest collection of aeronautical and space

science STI in the world. The Program Oce

is also NASA's institutional mechanism for

disseminating the results of its research and

development activities. These results are

published by NASA in the NASA STI Report

Series, which includes the following report

types:

 TECHNICAL PUBLICATION. Reports of

completed research or a ma jor signicant

phase of research that present the results

of NASA programs and include extensive

data or theoretical analysis. Includes

compilations of signicant scientic and

technical data and information deemed

to be of continuing reference value. NASA

counter-part of peer-reviewed formal

professional papers, but having less

stringent limitations on manuscript

length and extent of graphic

presentations.

 TECHNICAL MEMORANDUM.

Scientic and technical ndings that are

preliminary or of specialized interest,

e.g., quick release reports, working

papers, and bibliographies that contain

minimal annotation. Does not contain

extensive analysis.

 CONTRACTOR REPORT. Scientic and

technical ndings by NASA-sponsored

contractors and grantees.

 CONFERENCE PUBLICATION.

Collected papers from scientic and

technical conferences, symposia,

seminars, or other meetings sponsored or

co-sponsored by NASA.

 SPECIAL PUBLICATION. Scientic,

technical, or historical information from

NASA programs, pro jects, and missions,

often concerned with sub jects having

substantial public interest.

 TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientic

and technical material pertinent to

NASA's mission.

Specialized services that help round out the

STI Program Oce's diverse oerings include

creating custom thesauri, building customized

databases, organizing and publishing

research results. . . even providing videos.

For more information about the NASA STI

Program Oce, see the following:

 Access the NASA STI Program Home

Page at http://www.sti.nasa.gov

 E-mail your question via the Internet to

help@sti.nasa.gov

 Fax your question to the NASA Access

Help Desk at (301) 621{0134

 Phone the NASA Access Help Desk at

(301) 621{0390

 Write to:

NASA Access Help Desk

NASA Center for AeroSpace Information

7121 Standard Drive

Hanover, MD 21076{1320
NASA/CR{2001{211266

An Inviscid Computational Study of

Three '07 Mars Lander Aeroshell

Congurations Over a Mach Number

Range of 2.3 to 4.5

Ramadas K. Prabhu

Lockheed Martin Engineering & Sciences Company, Hampton, Virginia

National Aeronautics and

Space Administration

Langley Research Center Prepared for Langley Reseach Center

Hampton, Virginia 23681{2199 under contract NAS1-00135

December 2001
Available from the following:

NASA Center for AeroSpace Information (CASI) National Technical Information Service (NTIS)

7121 Standard Drive 5285 Port Royal Road

Hanover, MD 21076{1320 Springeld, VA 22161{2171

(301) 621{0390 (703) 487{4650
Summary

This report documents the results of a study conducted to compute the inviscid longitudinal aerodynamic

characteristics of three aeroshell congurations with trim tabs of the proposed '07 Mars lander. This was

done in support of the activity to design a smart lander for the proposed '07 Mars mission. In addition to

the three congurations designated as the `shelf ', the `canted', and the `Ames', the baseline conguration

(without tab) was also studied. The unstructured grid inviscid CFD software FELISA was used, and the

longitudinal aerodynamic characteristics of the four congurations were computed for Mach number of 2.3,

2.7, 3.5, and 4.5, and for an angle of attack range of -4 to 20 degrees. Wind tunnel tests had been conducted

on scale models of these four congurations in the Unitary Plan Wind Tunnel, NASA Langley Research

Center. Present computational results are compared with the data from these tests. Some dierences are

noticed between the two results, particularly at the lower Mach numbers. These dierences are attributed

to the pressures acting on the aft body. Most of the present computations were done on the forebody only. Additional computations were done on the full body (forbody and aftbody) for the baseline and the `Shelf '

congurations. Results of some computations done (to simulate 
ight conditions) with the Mars gas option

and with an "eective" 
 are also included.

Nomenclature

CA FA/(q1 Sref ), Axial force coecient

CD CA cos() + CN sin(), Drag coecient

CN FN/(q1 Sref ), Normal force coecient

CL CN cos() - CA sin(), Lift coecient

Cm My/(q1 Sref lref ), Pitching moment coecient

Cp (p - p1)/q1, Pressure coecient

FA Axial force

FN Normal force

lref Reference length

L=D CL/CD

My Pitching moment

M1 Freestream Mach number

p Static pressure

p1 Freestream static pressure

q1 Freestream dynamic pressure

Sref Reference area

x, y, z Cartesian co-ordinates of a given point

 Angle of attack, deg.


 Isentropic index

1
Introduction

The second generation or "smart" landers planned for '07 Mars mission are capable of performing precision

landing, hazard avoidance, and hazard tolerance. (see Ref. [1]) The primary aerodynamic requirement for

such a smart lander is that it should produce an L=D of -0.22 to 0.25 at the trim angle of attack. In an

extensive study aimed at arriving a suitable shape for the aeroshell several shapes had been screened [2].

As a result of this study, two congurations were down-selected, and detailed study was done to compute

their aerodynamics characteristics over a Mach number range. (see Ref. [3]) Ballistic tests had been done

on a related congurations considered to be a possible candidate. Low supersonic wind tunnel test were

done in the NASA Unitary Plan Wind Tunnel on the three candidate aeroshell conguration and also on the

baseline conguration. Longitudinal aerodynamic characteristics were measured for Mach 2.3, 2.7, 3.5, and

4.5 over an angle of attack of -5 to +20 degrees. A limited number of asymmetric tests were also done to

measure the lateral and directional aerodynamic characteristics. In order to supplement this data, a detailed

set of inviscid 
ow computations were done using the software FELISA [4], and longitudinal aerodynamic

characteristics of the four shapes were computed for 2.3, 2.7, 3.5, and 4.5 over an angle of attack of -5{+20

degrees. The present report presents these results and also shows a comparison with the wind tunnel data.

Flow conditions in the Unitary Plan Wind Tunnel are quite dierent from the 
ight conditions in the Mars

atmosphere for the same Mach number. Therefore, a few additional computations were done with 
ow

conditions simulating 
ight in Mars atmospheric gas.

Geometry

The present computational study was done for three aeroshell congurations for the proposed 007 Mars

Lander. These are designated as the `shelf ', the `canted', and the `Ames' congurations. Each of these

three congurations has a tab that makes them trim at non-zero angle of attack. In addition, the baseline

conguration (no tab) was also studied. A sketch of the the four congurations is shown in Fig. 1. Axis

system used for this study is also shown in the gure. The origin is at the nose of the body with the z-axis

along the body axis pointing upstream, the x-axis in the symmetry plane perpendicular to the z-axis, and

the y-axis perpendicular to the symmetry plane. The baseline geometry is a 4.05 m. diameter blunt conical

forebody with a 70-deg. half-cone angle. The nose radius is 0.985 m. and the shoulder radius is 0.0987 m. In the `shelf ' and the `Ames' congurations, the tab is an extension of the conical surface, whereas

in the `canted' conguration the tab is canted forward so that the tab make a 10 deg. angle with the

conical surface. Since all these shapes are symmetric about the x-z plane, only one half of these bodies were

simulated in the computational model. The reference quantities used for reducing the aerodynamic loads to

the non-dimensional form are as follows:

Reference area 12.8825 sq.m.

Reference length 4.05 m.

Pitching moment reference point (0.0, 0.0, -0.8659) m.

The Felisa Software

Computations of the present study were done using the FELISA unstructured grid software. This soft- ware package consists of a set of computer codes for unstructured grid generation, and the simulation of

2
70°

Nose Radius

0.985 m

Shoulder Radius

0.0987 m

Z

X

R=2.025m

Y

X

(a) baseline

2.286 m

(b) shelf

2.286 m

(c) canted

0.754 m

2.506 m

(d) Ames

Figure 1: Geometrical details of the four shapes

3
three-dimensional steady inviscid 
ows using unstructured tetrahedral grids. Surface triangulation and discretization of the computational domain using tetrahedral elements is done by using two separate codes.

Two 
ow solvers are available|one applicable for transonic 
ows, and the other for hypersonic 
ows. The

hypersonic 
ow solver has options for perfect gas, equilibrium air, CF4, CO2, and equilibrium Mars gases.

This solver also has the capability of solving chemical non-equilibrium 
ow, and real gas (chemical and

thermal non-equilibrium) 
ow. The hypersonic 
ow solver with the perfect gas (constant 
) and with the

equilibrium Mars gas options was used for the present computations. Post-processors like the aerodynamic

analysis routine are part of the software package. More information on FELISA software may be found in [4].

A description of the hypersonic 
ow solver may be found in [5].

Grids Used in the Present Study

Starting with the geometry les for the bodies in the iges format, and using the software GridTool [6],

a set of FELISA data les was generated. These les include (1) the FELISA data le that contains all

the information on the body surface denition and the computational domain in a format suitable for the

FELISA grid generator, (2) the FELISA background le that species the grid spacing, and (3) the FELISA

boundary conditions les. The minimum spacing for all the grids was 0.03 m. The computational domain was chosen to be suciently large so that the in
ow boundaries would not aect the 
ow on the body. The computational domain used for the baseline conguration is shown in Figures 2. Similar computational

domains were used for the other three congurations. The computational domain used for the full body

conguration with a sting (as tested in the wind tunnel) is shown in Figures 3. Note, the size shown is the

full scale vehicle size and not the model size actually tested in the wind tunnel. Also, the computational

domian used for the full body and no sting is shown in Figures 4. Only a few computations were done with

these congurations.

Using the FELISA data les, unstructured surface triangulation and tetrahedral grids were generated

using the surface and volume grid generators, respectively. A single grid was built for each conguration.

The properties of these grids are shown in Table 1. All these grids were generated on an SGI ONYX located

in the Aerothermodynamics Branch (AB), NASA Langley Research Center.

Conguration No. of No. of No. of No. of

Tets Points Triangles Surface Points

caseline 9,625,582 1,613,402 9,661 18,477

shelf 9,752,333 1,63,4514 10,124 19,119

canted 9,168,056 1,536,015 11,754 22,131

Ames 9,775,839 1,638,302 10,046 19,092

Table 1: Properties of grids used in the present computations.

4
Front View

4.5 m

Side View

1.9767 m

8.0 m

α°

Figure 2: Computational domain for the forebody. Flow Conditions

The four congurations analyzed here had been tested in the NASA Langley Unitary Plan Wind Tunnel.

The total temperature in the tunnel test section during these tests was about 350 K. At this low temperature,

air can be treated as perfect gas with 
 = 1.4. Therefore, in order to compare the computational results

with the wind tunnel data, a set of the computations were done with the assumption of perfect gas with


 = 1.4. Subsequent to this, a few more computations were done for Mach 4.5 freestream with equilibrium

Mars gas and also with an eective 
. This was done to simulate the 
ight conditions in Mars atmosphere.

The eective 
 is dened as the value of the isentropic index that produces a normal-shock-density-ratio the

same as that in equilibrium Mars gas. This values depends on the freestream Mach number; for Mach 4.5

the value of eective 
 is 1.24.

Flow Solution

The 
ow solutions were computed on an SGI Origin 2000 class parallel computer. The grids were partitioned

so that the problem would run on 64 processors. The FELISA hypersonic 
ow solver with the perfect gas

option with 
 = 1.4, was used for the initial computations. As noted earlier, some cases with the equilibrium

Mars gas, and an eective 
 = 1.24 were also done. Each solution was started with the low-order option,

and after a few hundred iterations, the higher-order option was turned on, and the solution was run to

convergence. After every 100 iteration, the surface pressures were integrated, and the aerodynamic loads,

namely the normal and the axial forces, and the pitching moment acting on the body were computed. The

5
8.0 m

10.5 m 4.05 m

dia.

Support Sting

1.0125 m. dia

Side View

Figure 3: Computational domain for the wind tunnel model.


ow solution was assumed to be converged when these integrated loads reached steady values. The grids

for the forebody required about 45 CPU-hours, whereas the grids for the wind tunnel model conguration

required about 80 CPU-hours. It may be recalled that the wind tunnel model conguration has the aftbody

and the support sting, and a much larger computational domain. The 
ow over the aftbody tends to be

oscillatory, with more iterations required to reach convergence. The computed 
ow solutions were postprocessed to obtain the aerodynamic loads. These loads were non-dimensionalized in the conventional

manner, and the aerodynamic coecients (CN ; CA; CL; CD; L=D; and Cm) were obtained.

It should be recalled at this point that the present computations are inviscid; hence the viscous eects -

skin friction and boundary layer separation are absent. For blunt bodies like the ones studied in this report,

skin friction is a negligible part of the total force on the forebody. Therefore, the present results should

predict the forebody aerodynamic loads well. Flow over the aftbody is viscous dominated, and because of

abrupt changes in the surface slopes and the rearward facing surfaces, there would be separation on the

the aftbody. This precludes simulation of such 
ows using inviscid 
ow solvers. However, the `numerical

viscosity' inherent in inviscid computations sometimes allow such 
ows to be computed. The resulting 
ow

eld shows features of a `separated' 
ow. However, results of such computations should be treated with

caution.

Results and Discussion

Aerodynamic coecients from the present computations on the baseline, the `shelf ', the `canted', and the

`Ames' congurations are summarized in Tables 2{7. Wind tunnel test data in tabular form for CA, CN ,

6
8.0 m

10.5 m

Side View

Base Dia.

0.7087 m.

2.909 m.

Figure 4: Computational domain for the full model with no sting.

Cm, CL, CD, and L=D are available for Mach 2.3, 2.7, 3.5, and 4.5 and  from -5 to +20 degrees for the

four congurations, with the exception of the `Ames' conguration for which test results are not available

for Mach 4.5. In the following sections present computational results are compared with the wind tunnel

data. Wherever dierences are noticed, possible causes are identied. Results of additional computations

in the wind tunnel model conguration (full model with a support sting) and full model without the sting

are also presented. Further, a few results for the 'shelf ' conguration under Mach 4.5 
ight conditions in

equilibrium Mars gas as well as in a perfect gas with an eective 
 are also presented.

The Baseline Conguration:

The results for the baseline conguration are listed in Table 2, and are shown plotted in Figure 5 for Mach

2.3, 2.7, 3.5, and 4.5. These computations were done for only the forebody; the aftbody was absent in the

computational model. It may be observed from Figures 5 that Cm, L=D, and CN vary linearly with , and

do not depend on the Mach number, except at  = 20 degrees. Both CA and CD are maximum at  = 0

degrees, and decreases with increasing angle of attack. With increase in Mach number, these values increase.

As the Mach number increases, the forebody surface pressures increase which result in increased CA. The

drag coecient, CD, which is essentially a component of CA, varies similar to CA. The magnitudes of CN

are small compared to the magnitudes of CA. Figures 6 to 9 show comparisons of the present results with the tunnel data for Mach numbers 2.3, 2.7,

3.5, and 4.5, respectively. As noted before, the wind tunnel model was a full model, and had a sting at the

base to support the model in the wind tunnel test section. Most of the present computations, however, were

7
done on the forebody only, and the aftbody was absent (see Figure 2). This would lead to dierences in the

test data and CFD results in those cases where the pressures acting on the aftbody contributed signicantly

to the loads. An examination of the gures 6{9 shows that the computed Cm agrees with the test data at

Mach 2.3. At higher Mach numbers there are small dierences between the two; the computed points lie

above the test data. Computed L=D values agree with the test data for all Mach numbers and angles of

attack except at  = 20 degrees. The computed values of CA are smaller than the measured values by about

0.16 at Mach 2.3. This dierence decreases with increase in Mach number, and at Mach 4.5 the dierence is

0.03 which is about 2%. The wind tunnel CN vs.  curves exhibit a peculiar trend at low angles of attack.

At  = 0, the curve has a negative slope, and the CN values are negative between 0 and 2 degrees, This

trend is very prominent at Mach 2.3, and disappears gradually as the freestream Mach numbers increases

to 4.5. A similar trend had been observed in the past during tests on a similar model (see [7]). The exact

cause of this is not been fully understood. Also, at angles of attack greater than 12 degrees, CN increases

rapidly. The present computations do not show any of this trend.

Figure 10 shows the Cp contours on the symmetry plane for Mach 4.5 and  = 8, 12, 16, and 20 degrees.

Figure 11 presents Cp contours on the symmetry plane for  = 15 degrees and Mach 2.3, 3.5, and 4.5.

Wind Tunnel Model Conguration:

In an attempt to reproduce the conditions in the wind tunnel, computations were done for the baseline

body in the wind tunnel model conguration (full model with a support sting) as shown in Figure 3. These

computations were done for Mach 2.3 at  = -4 to 20, and for Mach 4.5 at  = 0 to 5 degrees. The results

are also shown in Figures 6 & 9. It may be noticed that Cm and L=D do not change from the forebody

values at Mach 2.3 and 4.5. The agreement between the computed CA as well as CD and the wind tunnel

data is much better at Mach 2.3. The CN exhibits a non-linear trend with angle of attack. However, the

trend noticed in the wind tunnel data at low angles of attack is not seen in the results for Mach 2.3. At

 = 12 degrees and higher, the computed results agree very well with the wind tunnel data.

Examination of the 
ow over the body and the symmetry plane shows what appears to be a separated


ow on the aftbody (see Figures 12 to 14). It should be recalled here that the present computations are

inviscid, and the boundary layer is absent. Nevertheless, the numerical viscosity in the solution algorithm

causes the 
ow to "separate" similar to what happens in viscous 
ow. Figure 12{14 shows the streamlines

on the body as well as on the symmetry plane for Mach 2.3, and  = 0, 5, 8, 12, 16, and 20 degrees. The

separated 
ow region for low angles of attack is nearly symmetric about the longitudinal axis. The pressures

on the aftbody contribute not only to CA but also to CN . This is evident in the CN vs.  curve for Mach

2.3. For  greater than 8 degrees, there is a marked asymmetry in the separated 
ow region. The resulting

pressure distribution on the aftbody gives rise to a normal force. The computed values of CN for  greater

than 10 degrees agree very well with the tunnel data. The dierences between the computational results for

forebody and full model are small (0.02 at  = 20 degrees), and do not seem to contribute to the pitching

moment.

The Shelf conguration:

The results for the `shelf ' conguration are shown plotted in Figure 15 for Mach 2.3, 2.7, 3.5, and 4.5. The variations of the aerodynamic coecients with angle of attack and Mach number in general are very similar

to what was observed for the baseline conguration. The pitching moment varies linearly with angle of attack

for the four Mach numbers. The curves shift upwards as the Mach number is increased from 2.3 to 4.5. The

8
trim angles of attack are 13.1 degrees for Mach 2.3 and 14.3 degrees for Mach 4.5. The corresponding L=D

values are -0.22 for Mach 2.3 and -0.24 for Mach 4.5

These results are also shown compared with the wind tunnel data. See Figures 16 to 19 The computed

Cm values are consistently below the wind tunnel data, and the two curves are nearly parallel to each

other at Mach 2.3 and 2.7. At higher Mach numbers, the computations come closer to the test data. The

dierences between the computed and the tunnel data are larger (0.003) at Mach 2.3, and nearly zero at

Mach 4.5. The computed L=D values agree with tunnel data well for all angles of attack at the four Mach

numbers. The trend of CA, CD, CN , and CL for this conguration are similar to that observed for the

baseline conguration.

Figure 20 shows the sonic lines for  = -4 to 20 degrees, and Mach 2.3, 3.5, and 4.5.

Wind Tunnel Model Conguration; 
 = 1.4:

Four additional computations were done for the `shelf ' conguration in the wind tunnel model conguration

and 
 = 1.4 at Mach 4.5 and  = 8, 12, 16, and 20 degrees. The purpose of these computations was to

check how well the full model computations predict the aerodynamic data for the full model at Mach 4.5.

The results of these computations are listed in Table 5, and also shown in Figure 21. The agreement

between these results and the wind tunnel test data is much better than the forebody alone. The axial force

coecients dier by less than 0.02 at  = 8 degrees, and this dierence decreases to zero at  = 20 degrees.

The steep rise in the normal force coecient beyond  > 12 degrees is also predicted by the computations.

It may be recalled that a similar comparison was noted for the baseline congurations for Mach 2.3.

Computations with Mars Gas and an eective 
: Two additional computations were done for the `shelf ' forebody with equilibrium Mars gas with 
ight

conditions for Mach 4.5. The freestream conditions are as follows:

Velocity 967.6 m=s

Density 2.183E-03 kg=m3

Temperature 184.4 K

Further, two more computations were done with 
 = 1.24 to see how well the results of computations

with an eective 
 compare with the Mars gas. The eective 
 is dened as the value of isentropic index that

produces a normal-shock-density-ratio equal to that for the equilibrium Mars gas. This value depends on

the freestream conditions; for Mach 4.5 
ight conditions the value is 1.24. The results of these computations

are listed in Table 5 and shown graphically in Figure 22. An examination of this gure shows that the Mars

gas and 
 = 1.24 produce results that are almost indistinguishable from each other. Further a comparison

of these results with the results for 
 = 1.4 show dierences in Cm and CA. There is an increase in Cm and

also in CA. The increased Cm would increase the trim angle by about 2 degrees. The increase in CA (and

consequently CD) is 0.06 at  = 10 degrees. The CN does not seem to be aected noticeably. Computations were done on the full model with and without the support sting to study the eect of

the aftbody on the results with 
 = 1.24. Results of these computations are also listed in Table 5 and are

shown in Figure 22. The eect of the aftbody is seen primarily in CA, CD, and in CN . The pressures on the

aftbody do not seem to aect the Cm. Both axial and normal force coecients increases due to the aftbody;

changes in CD are similar to those in CA. The normal force coecient seems to rise steeply with increase in

. These are consistent with the observation made on the baseline conguration. Also, the presence of the

sting on the aerodynamic coecients under these conditions (M = 4.5, 
 = 1.24) is negligible.

9
The `canted' and the `Ames' congurations:

The results for the `canted' and `Ames' congurations are shown plotted in Figure 23 & 28. These results

are also shown compared with the wind tunnel test data Figures 24 to 27 and 29 to 31. It may be seen in

these gures that the trends in all the aerodynamic coecients at the four Mach numbers is similar to what was seen for the 'shelf ' conguration. The `canted' body behaves very similar to the `shelf '. The `Ames'

body has a larger tab, and hence trims at a 17 angle of attack, and produces a trim L=D of -0.28 at M=4.5.

Conclusion

Inviscid longitudinal aerodynamic characteristics of the three aeroshell conguration for the proposed '07

Mars lander are presented. These shapes are designated the `shelf ', the `canted', and the `Ames'. Aerodynamic data for the baseline conguration are also computed, and are included in this report. The results

are for Mach 2.3, 2.7, 3.5, and 4.5, and for  from -4 to +20 degrees. These results are compared with the

data from model tests in the NASA Unitary Plan Wind Tunnel. The comparison showed that computed

aerodynamic coecients compare well with wind tunnel data at Mach 4.5. At lower Mach numbers the

comparison is not as good, particularly in CA and CN . The dierences between the computed and wind

tunnel data are attributed to the pressures acting on the aft part of these bodies. Most of the present computations were done on the forebody only; the aftbody was absent in these computational models. A few

computations done for computational models with the aftbody and sting show that, at lower Mach numbers,

the aftbody pressure contribute signicantly to CA and CN . The aftbody pressures account for most of

the dierences between the wind tunnel data and the (forebody only) axial force coecients. The pitching

moment coecient is not aected noticeably by the aftbody pressures. This study also indicated that 
ow

"separation" on the aftbody at angles of attack greater than 8 degrees leads to nonlinear trends in CN with

. A few computations were done at Mach 4.5 in Mars gas environment as well as with an eective 
. There was a very good agreement between these two sets of results, indicating that, at these low Mach numbers,

the computations with an eective 
 yield results that are nearly the same as would be found in Mars gas.

Acknowledgments

The work described herein was performed at Lockheed Martin Engineering & Sciences Company in Hampton,

Virginia, and was supported by Aerothermodynamics Branch, NASA Langley Research Center under the

contract NAS1-00135. The technical monitor was Dr. Kenneth Sutton. The tunnel data in the form of curve

ts were provided by Ms. Kelly Murphy of the Aerothermodynamics Branch.

References

[1] Lockwood, M.K., Powell, R.W., Graves, S.G., and Carman, G.L.: \Entry System Design Considerations

for Mars Landers," AAS 01-023, 24th Annual AAS Guidance and Control Conference, January{February, 2001.

[2] Prabhu, R.K.: \Inviscid Flow Computations of Several Aeroshell Congurations for a '07 Mars Lander,"

NASA/CR-2001-210851, April 2001.

10
[3] Prabhu, R.K.: \Inviscid Flow Computations of Two '07 Mars Lander Aeroshell Congurations Over a

Mach Number Range of 2 to 24," NASA/CR-2001-210852, April 2001.

[4] Peiro, J., Peraire, J., and Morgan, K.: \FELISA System Reference Manual and User's Guide," Technical

Report, University College, Swansea, Wales, 1993, (Also NASA CP 3291, May 1995.)

[5] Bibb, K.L., Peraire, J., and Riley, C.J.: \Hypersonic Flow Computation on Unstructured Meshes," AIAA

Paper 97-0625, January 1997.

[6] Samereh, J.: \GridTool: A Surface Modeling and Grid Generation Tool," NASA CP 3291, May 1995.

[7] Blake, W.W.: \Experimental Aerodynamic Characteristics of the Viking Entry Vehicle Over the Mach

Number Range 1.5{10," Martin Marietta Contractor Report TR-372106, April 1971.

11
Mach No. , deg. CN CA Cm CL CD L=D

2.3 -4.0 -9.9624E-03 1.3743 7.3799E-03 8.5928E-02 1.3716 6.2646E-02

2.3 0.0 -2.1217E-04 1.3782 5.1236E-05 -2.1217E-04 1.3782 -1.5395E-04

2.3 4.0 9.5892E-03 1.3743 -7.2949E-03 -8.6300E-02 1.3716 -6.2919E-02

2.3 8.0 1.9304E-02 1.3626 -1.4550E-02 -1.7052E-01 1.3520 -1.2612E-01

2.3 12.0 2.8956E-02 1.3428 -2.1710E-02 -2.5086E-01 1.3195 -1.9012E-01

2.3 16.0 3.8496E-02 1.3145 -2.8710E-02 -3.2532E-01 1.2742 -2.5532E-01

2.3 20.0 4.7998E-02 1.2768 -3.5574E-02 -3.9159E-01 1.2162 -3.2197E-01

2.7 -4.0 -1.0074E-02 1.4150 7.4349E-03 8.8656E-02 1.4123 6.2776E-02

2.7 0.0 -1.8801E-04 1.4191 4.4844E-05 -1.8801E-04 1.4191 -1.3249E-04

2.7 4.0 9.6443E-03 1.4150 -7.3095E-03 -8.9085E-02 1.4122 -6.3081E-02

2.7 8.0 1.9471E-02 1.4026 -1.4629E-02 -1.7592E-01 1.3917 -1.2641E-01

2.7 12.0 2.9187E-02 1.3813 -2.1809E-02 -2.5864E-01 1.3572 -1.9057E-01

2.7 16.0 3.8827E-02 1.3506 -2.8842E-02 -3.3495E-01 1.3090 -2.5589E-01

2.7 20.0 4.8565E-02 1.3091 -3.5818E-02 -4.0210E-01 1.2468 -3.2252E-01

3.5 -4.0 -1.0094E-02 1.4612 7.4142E-03 9.1859E-02 1.4583 6.2988E-02

3.5 0.0 -1.7304E-04 1.4656 3.1700E-05 -1.7304E-04 1.4656 -1.1807E-04

3.5 4.0 9.7103E-03 1.4612 -7.3231E-03 -9.2242E-02 1.4583 -6.3252E-02

3.5 8.0 1.9542E-02 1.4476 -1.4605E-02 -1.8212E-01 1.4362 -1.2680E-01

3.5 12.0 2.9351E-02 1.4242 -2.1791E-02 -2.6740E-01 1.3992 -1.9111E-01

3.5 16.0 3.9205E-02 1.3897 -2.8894E-02 -3.4537E-01 1.3467 -2.5646E-01

3.5 20.0 4.9530E-02 1.3415 -3.6186E-02 -4.1228E-01 1.2775 -3.2271E-01

4.5 -4.0 -1.0096E-02 1.4894 7.3686E-03 9.3824E-02 1.4865 6.3118E-02

4.5 0.0 -1.8435E-04 1.4941 3.2561E-05 -1.8435E-04 1.4941 -1.2339E-04

4.5 4.0 9.7301E-03 1.4895 -7.3010E-03 -9.4196E-02 1.4866 -6.3365E-02

4.5 8.0 1.9527E-02 1.4751 -1.4510E-02 -1.8596E-01 1.4635 -1.2707E-01

4.5 12.0 2.9340E-02 1.4500 -2.1620E-02 -2.7277E-01 1.4244 -1.9150E-01

4.5 16.0 3.9389E-02 1.4121 -2.8792E-02 -3.5136E-01 1.3683 -2.5680E-01

4.5 20.0 5.0618E-02 1.3574 -3.6641E-02 -4.1669E-01 1.2929 -3.2231E-01

Table 2: Summary of aerodynamic coecients for `baseline' forebody.

12
Mach No. , deg. CN CA Cm CL CD L=D

2.3 0.0 -2.1035E-04 1.5584 -2.1060E-05 -2.1035E-04 1.5584 -1.3498E-04

2.3 1.0 3.1944E-03 1.5582 -1.9248E-03 -2.4000E-02 1.5580 -1.5404E-02

2.3 2.0 5.6125E-03 1.5580 -3.9234E-03 -4.8764E-02 1.5572 -3.1314E-02

2.3 3.0 8.7543E-03 1.5571 -5.8740E-03 -7.2750E-02 1.5554 -4.6772E-02

2.3 4.0 1.2411E-02 1.5562 -7.8081E-03 -9.6174E-02 1.5533 -6.1917E-02

2.3 5.0 1.6191E-02 1.5544 -9.7277E-03 -1.1935E-01 1.5499 -7.7002E-02

2.3 8.0 2.5503E-02 1.5459 -1.5446E-02 -1.8989E-01 1.5344 -1.2376E-01

2.3 12.0 2.8358E-02 1.5374 -2.2170E-02 -2.9191E-01 1.5097 -1.9335E-01

2.3 16.0 4.6564E-02 1.5116 -2.9373E-02 -3.7189E-01 1.4659 -2.5370E-01

2.3 20.0 6.9515E-02 1.4755 -3.6283E-02 -4.3933E-01 1.4103 -3.1152E-01

4.5 0.0 -2.5092E-04 1.5276 6.1677E-06 -2.5092E-04 1.5276 -1.6426E-04

4.5 1.0 2.7317E-03 1.5273 -1.9026E-03 -2.3924E-02 1.5271 -1.5666E-02

4.5 2.0 5.6031E-03 1.5266 -3.8278E-03 -4.7678E-02 1.5259 -3.1246E-02

4.5 3.0 8.5654E-03 1.5254 -5.7405E-03 -7.1280E-02 1.5238 -4.6779E-02

4.5 4.0 1.1813E-02 1.5236 -7.6431E-03 -9.4497E-02 1.5207 -6.2140E-02

4.5 5.0 1.4996E-02 1.5210 -9.5442E-03 -1.1762E-01 1.5165 -7.7562E-02

Table 3: Summary of aerodynamic coecients for `baseline' full model conguration.

13
Mach No. , deg. CN CA Cm CL CD L=D

2.3 -4.0 -3.0764E-02 1.4358 3.6454E-02 6.9467E-02 1.4344 4.8428E-02

2.3 0.0 -2.0239E-02 1.4387 2.8035E-02 -2.0239E-02 1.4387 -1.4068E-02

2.3 4.0 -9.5881E-03 1.4332 1.9514E-02 -1.0954E-01 1.4290 -7.6653E-02

2.3 8.0 1.1545E-03 1.4192 1.0945E-02 -1.9637E-01 1.4055 -1.3971E-01

2.3 12.0 1.1964E-02 1.3964 2.3462E-03 -2.7863E-01 1.3684 -2.0362E-01

2.3 16.0 2.2839E-02 1.3645 -6.2407E-03 -3.5415E-01 1.3179 -2.6872E-01

2.3 20.0 3.3705E-02 1.3228 -1.4728E-02 -4.2075E-01 1.2546 -3.3538E-01

2.7 -4.0 -3.1640E-02 1.4782 3.7328E-02 7.1551E-02 1.4768 4.8450E-02

2.7 0.0 -2.1078E-02 1.4813 2.8907E-02 -2.1078E-02 1.4813 -1.4229E-02

2.7 4.0 -1.0359E-02 1.4755 2.0363E-02 -1.1326E-01 1.4712 -7.6985E-02

2.7 8.0 5.1076E-04 1.4606 1.1721E-02 -2.0277E-01 1.4465 -1.4018E-01

2.7 12.0 1.1418E-02 1.4363 3.0888E-03 -2.8746E-01 1.4073 -2.0426E-01

2.7 16.0 2.2375E-02 1.4021 -5.4999E-03 -3.6496E-01 1.3540 -2.6955E-01

2.7 20.0 3.3510E-02 1.3563 -1.4122E-02 -4.3239E-01 1.2860 -3.3624E-01

3.5 -4.0 -3.2626E-02 1.5262 3.8322E-02 7.3916E-02 1.5248 4.8477E-02

3.5 0.0 -2.2010E-02 1.5296 2.9907E-02 -2.2010E-02 1.5296 -1.4389E-02

3.5 4.0 -1.1188E-02 1.5233 2.1322E-02 -1.1742E-01 1.5188 -7.7311E-02

3.5 8.0 -2.5775E-04 1.5072 1.2703E-02 -2.1002E-01 1.4925 -1.4072E-01

3.5 12.0 1.0771E-02 1.4805 4.0289E-03 -2.9728E-01 1.4504 -2.0496E-01

3.5 16.0 2.1975E-02 1.4421 -4.6743E-03 -3.7637E-01 1.3923 -2.7033E-01

3.5 20.0 3.3764E-02 1.3895 -1.3667E-02 -4.4351E-01 1.3173 -3.3669E-01

4.5 -4.0 -3.3176E-02 1.5556 3.8892E-02 7.5418E-02 1.5541 4.8528E-02

4.5 0.0 -2.2539E-02 1.5593 3.0518E-02 -2.2539E-02 1.5593 -1.4455E-02

4.5 4.0 -1.1726E-02 1.5527 2.1982E-02 -1.2001E-01 1.5481 -7.7520E-02

4.5 8.0 -8.3281E-04 1.5356 1.3426E-02 -2.1454E-01 1.5205 -1.4109E-01

4.5 12.0 1.0232E-02 1.5071 4.8315E-03 -3.0334E-01 1.4763 -2.0547E-01

4.5 16.0 2.1627E-02 1.4652 -3.9038E-03 -3.8307E-01 1.4144 -2.7084E-01

4.5 20.0 3.4397E-02 1.4059 -1.3585E-02 -4.4852E-01 1.3329 -3.3651E-01

Table 4: Summary of aerodynamic coecients for `shelf ' forebody.

14
Gas Mach  CN CA Cm CL CD L=D Cong.

Model No. deg.


=1.4 4.5 8.0 -3.304E-03 1.5789 1.363E-02 -2.230E-01 1.5631 -1.4268-01 Body+Sting


=1.4 4.5 12.0 1.321E-02 1.5515 5.190E-03 -3.096E-01 1.5203 -2.037E-01 Body+Sting


=1.4 4.5 16.0 3.158E-02 1.5092 -3.519E-03 -3.856E-01 1.4594 -2.642E-01 Body+Sting


=1.4 4.5 20.0 5.286E-02 1.4481 -1.312E-02 -4.456E-01 1.3788 -3.232E-01 Body+Sting


=1.24 4.5 10.0 3.213E-03 1.5895 1.202E-02 -2.729E-01 1.5659 -1.742E-01 Forebody


=1.24 4.5 14.0 1.497E-02 1.5459 3.449E-03 -3.595E-01 1.5036 -2.391E-01 Forebody

Mars Gas 4.5 10.0 2.966E-03 1.5888 1.214E-02 -2.730E-01 1.5652 -1.744E-01 Forebody

Mars Gas 4.5 14.0 1.421E-02 1.5477 4.006E-03 -3.606E-01 1.5052 -2.396E-01 Forebody


=1.24 4.5 10.0 7.313E-03 1.6362 1.244E-02 -2.769E-01 1.6126 -1.717E-01 Body+Sting


=1.24 4.5 14.0 2.713E-02 1.5930 4.342E-03 -3.591E-01 1.5522 -2.313E-01 Body+Sting


=1.24 4.5 10.0 8.283E-03 1.6418 1.282E-02 -2.769E-01 1.6183 -1.711E-01 Body, No Sting


=1.24 4.5 14.0 2.717E-02 1.5975 4.507E-03 -3.601E-01 1.5566 -2.313E-01 Body, No Sting

Table 5: Summary of aerodynamic coecients for `shelf ', 
ight conditions.

15
Mach No. , deg. CN CA Cm CL CD L=D

2.3 -4.0 -1.9757E-02 1.4228 3.0188E-02 7.9541E-02 1.4207 5.5986E-02

2.3 0.0 -1.0144E-02 1.4273 2.2479E-02 -1.0144E-02 1.4273 -7.1071E-03

2.3 4.0 -4.5003E-04 1.4236 1.4745E-02 -9.9754E-02 1.4201 -7.0244E-02

2.3 8.0 9.2349E-03 1.4117 7.0185E-03 -1.8733E-01 1.3992 -1.3388E-01

2.3 12.0 1.8938E-02 1.3915 -6.4701E-04 -2.7078E-01 1.3650 -1.9837E-01

2.3 16.0 2.8555E-02 1.3623 -8.1480E-03 -3.4805E-01 1.3174 -2.6420E-01

2.3 20.0 3.8130E-02 1.3237 -1.5485E-02 -4.1690E-01 1.2569 -3.3169E-01

2.7 -4.0 -2.0062E-02 1.4645 3.0924E-02 8.2145E-02 1.4623 5.6174E-02

2.7 0.0 -1.0406E-02 1.4693 2.3223E-02 -1.0406E-02 1.4693 -7.0823E-03

2.7 4.0 -6.6418E-04 1.4655 1.5467E-02 -1.0289E-01 1.4619 -7.0382E-02

2.7 8.0 9.0764E-03 1.4530 7.7716E-03 -1.9323E-01 1.4401 -1.3418E-01

2.7 12.0 1.8806E-02 1.4313 1.5063E-04 -2.7919E-01 1.4039 -1.9886E-01

2.7 16.0 2.8501E-02 1.3998 -7.3465E-03 -3.5844E-01 1.3534 -2.6484E-01

2.7 20.0 3.8326E-02 1.3572 -1.4756E-02 -4.2818E-01 1.2885 -3.3232E-01

3.5 -4.0 -2.0385E-02 1.5121 3.1754E-02 8.5143E-02 1.5098 5.6392E-02

3.5 0.0 -1.0694E-02 1.5173 2.4069E-02 -1.0694E-02 1.5173 -7.0480E-03

3.5 4.0 -9.3172E-04 1.5132 1.6373E-02 -1.0648E-01 1.5094 -7.0546E-02

3.5 8.0 8.8554E-03 1.4997 8.7310E-03 -1.9995E-01 1.4863 -1.3452E-01

3.5 12.0 1.8620E-02 1.4760 1.1970E-03 -2.8866E-01 1.4476 -1.9941E-01

3.5 16.0 2.8494E-02 1.4407 -6.2665E-03 -3.6972E-01 1.3927 -2.6546E-01

3.5 20.0 3.8968E-02 1.3908 -1.3961E-02 -4.3906E-01 1.3203 -3.3256E-01

4.5 -4.0 -2.0527E-02 1.5412 3.2182E-02 8.7032E-02 1.5389 5.6555E-02

4.5 0.0 -1.0875E-02 1.5468 2.4612E-02 -1.0875E-02 1.5468 -7.0306E-03

4.5 4.0 -1.1375E-03 1.5426 1.6999E-02 -1.0874E-01 1.5388 -7.0668E-02

4.5 8.0 8.5640E-03 1.5284 9.5159E-03 -2.0423E-01 1.5147 -1.3483E-01

4.5 12.0 1.8293E-02 1.5032 2.1760E-03 -2.9464E-01 1.4742 -1.9987E-01

4.5 16.0 2.8305E-02 1.4646 -5.1801E-03 -3.7649E-01 1.4157 -2.6595E-01

4.5 20.0 4.0012E-02 1.4072 -1.3724E-02 -4.4369E-01 1.3360 -3.3210E-01

Table 6: Summary of aerodynamic coecients for `canted' forebody.

16
Mach No. , deg. CN CA Cm CL CD L=D

2.3 -4.0 -3.7264E-02 1.4494 4.6583E-02 6.3932E-02 1.4485 4.4138E-02

2.3 0.0 -2.6234E-02 1.4510 3.7411E-02 -2.6234E-02 1.4510 -1.8080E-02

2.3 4.0 -1.4999E-02 1.4439 2.8026E-02 -1.1568E-01 1.4393 -8.0373E-02

2.3 8.0 -3.6459E-03 1.4282 1.8536E-02 -2.0238E-01 1.4138 -1.4314E-01

2.3 12.0 7.7690E-03 1.4036 9.0156E-03 -2.8423E-01 1.3745 -2.0678E-01

2.3 16.0 1.9200E-02 1.3698 -5.1140E-04 -3.5911E-01 1.3220 -2.7164E-01

2.3 20.0 3.0558E-02 1.3265 -9.9064E-03 -4.2497E-01 1.2570 -3.3810E-01

2.7 -4.0 -3.8172E-02 1.4928 4.7903E-02 6.6053E-02 1.4918 4.4277E-02

2.7 0.0 -2.7072E-02 1.4946 3.8698E-02 -2.7072E-02 1.4946 -1.8113E-02

2.7 4.0 -1.5735E-02 1.4871 2.9257E-02 -1.1943E-01 1.4824 -8.0567E-02

2.7 8.0 -4.2594E-03 1.4703 1.9693E-02 -2.0884E-01 1.4554 -1.4350E-01

2.7 12.0 7.2872E-03 1.4441 1.0073E-02 -2.9312E-01 1.4141 -2.0729E-01

2.7 16.0 1.8818E-02 1.4078 5.1821E-04 -3.6995E-01 1.3585 -2.7233E-01

2.7 20.0 3.0366E-02 1.3607 -8.9679E-03 -4.3685E-01 1.2890 -3.3890E-01

3.5 -4.0 -3.9154E-02 1.5418 4.9361E-02 6.8492E-02 1.5408 4.4453E-02

3.5 0.0 -2.7935E-02 1.5438 4.0107E-02 -2.7935E-02 1.5438 -1.8095E-02

3.5 4.0 -1.6496E-02 1.5357 3.0616E-02 -1.2358E-01 1.5308 -8.0729E-02

3.5 8.0 -4.9222E-03 1.5176 2.1006E-02 -2.1608E-01 1.5021 -1.4385E-01

3.5 12.0 6.7047E-03 1.4889 1.1369E-02 -3.0300E-01 1.4578 -2.0785E-01

3.5 16.0 1.8421E-02 1.4489 1.7477E-03 -3.8166E-01 1.3978 -2.7304E-01

3.5 20.0 3.0466E-02 1.3956 -7.9472E-03 -4.4869E-01 1.3219 -3.3944E-01

4.5 -4.0 -3.9719E-02 1.5719 5.0221E-02 7.0028E-02 1.5708 4.4580E-02

4.5 0.0 -2.8516E-02 1.5739 4.0988E-02 -2.8516E-02 1.5739 -1.8118E-02

4.5 4.0 -1.7038E-02 1.5655 3.1521E-02 -1.2620E-01 1.5605 -8.0872E-02

4.5 8.0 -5.4699E-03 1.5464 2.1961E-02 -2.2063E-01 1.5306 -1.4415E-01

4.5 12.0 6.1666E-03 1.5161 1.2411E-02 -3.0918E-01 1.4843 -2.0831E-01

4.5 16.0 1.7998E-02 1.4730 2.8283E-03 -3.8871E-01 1.4209 -2.7357E-01

4.5 20.0 3.0794E-02 1.4136 -7.2737E-03 -4.5454E-01 1.3389 -3.3949E-01

Table 7: Summary of aerodynamic coecients for `Ames' forebody.

17
α, deg.

CD

-4 0 4 8 12 16 20

1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

C

M

-4 0 4 8 12 16 20

-0.04

-0.03

-0.02

-0.01

0

0.01

0.02

Mach 2.3

Mach 2.7

Mach 3.5

Mach 4.5

α, deg.

CA

-4 0 4 8 12 16 20

1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CN

-4 0 4 8 12 16 20

-0.04

-0.02

0

0.02

0.04

0.06

α, deg.

CL

-4 0 4 8 12 16 20

-0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

L/D

-4 0 4 8 12 16 20

-0.4

-0.3

-0.2

-0.1

0

0.1

Figure 5: Inviscid aerodynamic data for the baseline forebody.

18
α, deg.

C

M

-5 0 5 10 15 20 -0.04

-0.03

-0.02

-0.01

0

0.01

0.02

W.T. Data

FELISA (Forebody)

FELISA (Full Body)

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CN

-5 0 5 10 15 20 -0.04

-0.02

0

0.02

0.04

0.06

0.08

α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

α, deg.

CA

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

Figure 6: Comparison of CFD and wind tunnel data for the baseline at Mach 2.3.

19
α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

α, deg.

CN

-5 0 5 10 15 20 -0.04

-0.02

0

0.02

0.04

0.06

0.08

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

C

M

-5 0 5 10 15 20 -0.04

-0.03

-0.02

-0.01

0

0.01

0.02

W.T. Data

FELISA (Forebody)

α, deg.

CA

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

Figure 7: Comparison of CFD and wind tunnel data for the baseline forebody at Mach 2.7.

20
α, deg.

CN

-5 0 5 10 15 20 -0.04

-0.02

0

0.02

0.04

0.06

0.08

α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

α, deg.

C

M

-5 0 5 10 15 20 -0.04

-0.03

-0.02

-0.01

0

0.01

0.02

W.T. Data

FELISA (Forebody)

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CA

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

Figure 8: Comparison of CFD and wind tunnel data for the baseline forebody at Mach 3.5.

21
α, deg.

C

M

-5 0 5 10 15 20 -0.04

-0.03

-0.02

-0.01

0

0.01

0.02

W.T. Data

FELISA (Forebody)

FELISA (Full Body)

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

α, deg.

CN

-5 0 5 10 15 20 -0.04

-0.02

0

0.02

0.04

0.06

0.08

α, deg.

CA

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

Figure 9: Comparison of CFD and wind tunnel data for the baseline at Mach 4.5.

22
α = 8° α = 12° α = 16° α = 20° Figure 10: Symmetry plane Cp contours for the baseline at Mach 4.5.

M = 2.3 M = 2.7 M = 3.5 M = 4.5 Figure 11: Symmetry plane Cp contours for the baseline at Mach  = 16 degrees.

23
M=1

NOTE: Sonic lines are shown by blue dashed lines.

M = 2.3, α = 0°

(a)  = 0

M=1

NOTE: Sonic lines are shown by blue lines.

M = 2.3, α = 5°

(b)  = 5

Figure 12: Streamlines on the body and the symmetry plane for the baseline conguration, M=2.3.

M=1

NOTE: Sonic lines are shown by blue lines.

M = 2.3, α = 8°

(a)  = 8

M=1

NOTE: Sonic lines are shown by blue lines.

M = 2.3, α = 12°

(b)  = 12

Figure 13: Streamlines on the body and the symmetry plane for the baseline conguration, M=2.3.

24
M=1

NOTE: Sonic lines are shown by blue lines.

M = 2.3, α = 16°

(a)  = 16

M=1

NOTE: Sonic lines are shown by blue lines.

M = 2.3, α = 20°

(b)  = 20

Figure 14: Streamlines on the body and the symmetry plane for the baseline conguration, M=2.3.

25
α, deg.

C

M

-4 0 4 8 12 16 20

-0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

Mach 2.3

Mach 2.7

Mach 3.5

Mach 4.5

α, deg.

L/D

-4 0 4 8 12 16 20

-0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CA

-4 0 4 8 12 16 20

1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CD

-4 0 4 8 12 16 20

1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CN

-4 0 4 8 12 16 20

-0.06

-0.04

-0.02

0

0.02

0.04

0.06

α, deg.

CL

-4 0 4 8 12 16 20

-0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

Figure 15: Inviscid aerodynamic data for the `shelf ' forebody.

26
α, deg.

C

M

-5 0 5 10 15 20 -0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

0.06

W.T. Data

FELISA (Forebody)

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CA

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CN

-5 0 5 10 15 20 -0.06

-0.04

-0.02

0

0.02

0.04

0.06

α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

Figure 16: Comparison of aerodynamic data for the `shelf ' conguration at Mach 2.3.

27
α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

α, deg.

CA

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CN

-5 0 5 10 15 20 -0.06

-0.04

-0.02

0

0.02

0.04

0.06

α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

C

M

-5 0 5 10 15 20 -0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

0.06

W.T. Data

FELISA (Forebody)

Figure 17: Comparison of aerodynamic data for the `shelf ' conguration at Mach 2.7.

28
α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

α, deg.

CN

-5 0 5 10 15 20 -0.06

-0.04

-0.02

0

0.02

0.04

0.06

α, deg.

CA

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

C

M

-5 0 5 10 15 20 -0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

0.06

W.T. Data

FELISA (Forebody)

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

Figure 18: Comparison of aerodynamic data for the `shelf ' conguration at Mach 3.5.

29
α, deg.

C

M

-5 0 5 10 15 20 -0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

0.06

W.T. Data

FELISA (Forebody)

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CA

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CN

-5 0 5 10 15 20 -0.06

-0.04

-0.02

0

0.02

0.04

0.06

α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

Figure 19: Comparison of aerodynamic data for the `shelf ' conguration at Mach 4.5.

30
M = 2.3 M = 3.5 M = 4.5

α = 4°

α = -4°

α = 0°

α = 20°

α = 8°

α = 12°

α = 16°

Figure 20: Sonic lines for `shelf ' conguration for Mach 2.3, 3.5, and 4.5.

31
α, deg.

CL

-4 0 4 8 12 16 20

-0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

α, deg.

CN

-4 0 4 8 12 16 20

-0.04

-0.02

0

0.02

0.04

0.06

α, deg.

CA

-4 0 4 8 12 16 20

1.3

1.35

1.4

1.45

1.5

1.55

1.6

1.65

1.7

α, deg.

CD

-4 0 4 8 12 16 20

1.3

1.35

1.4

1.45

1.5

1.55

1.6

1.65

1.7

α, deg.

C

M

-4 0 4 8 12 16 20

-0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

W.T Data

Forebody, γ=1.4

Body+Sting, γ=1.4

α, deg.

L/D

-4 0 4 8 12 16 20

-0.4

-0.3

-0.2

-0.1

0

0.1

Figure 21: Comparison of aerodynamic data for the `shelf ' conguration at Mach 4.5.

32
α, deg.

CA

-4 0 4 8 12 16 20

1.3

1.35

1.4

1.45

1.5

1.55

1.6

1.65

1.7

α, deg.

L/D

-4 0 4 8 12 16 20

-0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CD

-4 0 4 8 12 16 20

1.3

1.35

1.4

1.45

1.5

1.55

1.6

1.65

1.7

α, deg.

CN

-4 0 4 8 12 16 20

-0.04

-0.02

0

0.02

0.04

0.06

α, deg.

CL

-4 0 4 8 12 16 20

-0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

α, deg.

C

M

-4 0 4 8 12 16 20

-0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05 W.T Data

Forebody, γ=1.4

Forebody, γ=1.24

Forebody, Mars Gas

Body+Sting, γ=1.24

Body Only, γ=1.24

Figure 22: Comparison of aerodynamic data for the `shelf ' conguration at Mach 4.5.

33
α, deg.

CL

-4 0 4 8 12 16 20

-0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CN

-4 0 4 8 12 16 20

-0.04

-0.02

0

0.02

0.04

0.06

α, deg.

CA

-4 0 4 8 12 16 20

1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CD

-4 0 4 8 12 16 20

1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

L/D

-4 0 4 8 12 16 20

-0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

C

M

-4 0 4 8 12 16 20

-0.02

-0.01

0

0.01

0.02

0.03

0.04

Mach 2.3

Mach 2.7

Mach 3.5

Mach 4.5

Figure 23: Inviscid aerodynamic data for the `canted' forebody.

34
α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

C

M

-5 0 5 10 15 20 -0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

0.06

W.T. Data

FELISA (Forebody)

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CN

-5 0 5 10 15 20 -0.06

-0.04

-0.02

0

0.02

0.04

0.06

α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

α, deg.

CA

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

Figure 24: Comparison of aerodynamic data for the `canted' conguration at Mach 2.3.

35
α, deg.

C

M

-5 0 5 10 15 20 -0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

0.06

W.T. Data

FELISA (Forebody)

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

α, deg.

CN

-5 0 5 10 15 20 -0.04

-0.02

0

0.02

0.04

0.06

0.08

α, deg.

CA

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

Figure 25: Comparison of aerodynamic data for the `canted' conguration at Mach 2.7.

36
α, deg.

C

M

-5 0 5 10 15 20 -0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

0.06

W.T. Data

FELISA (Forebody)

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

α, deg.

CN

-5 0 5 10 15 20 -0.04

-0.02

0

0.02

0.04

0.06

0.08

α, deg.

CA

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

Figure 26: Comparison of aerodynamic data for the `canted' conguration at Mach 3.5.

37
α, deg.

C

M

-5 0 5 10 15 20 -0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

0.06

W.T. Data

FELISA (Forebody)

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CN

-5 0 5 10 15 20 -0.04

-0.02

0

0.02

0.04

0.06

0.08

α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

α, deg.

CA

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

Figure 27: Comparison of aerodynamic data for the `canted' conguration at Mach 4.5.

38
α, deg.

C

M

-4 0 4 8 12 16 20

-0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

0.06

Mach 2.3

Mach 2.7

Mach 3.5

Mach 4.5

α, deg.

L/D

-4 0 4 8 12 16 20

-0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CA

-4 0 4 8 12 16 20

1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CD

-4 0 4 8 12 16 20

1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CN

-4 0 4 8 12 16 20

-0.06

-0.04

-0.02

0

0.02

0.04

0.06

α, deg.

CL

-4 0 4 8 12 16 20

-0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

Figure 28: Inviscid aerodynamic data for the `Ames' forebody.

39
α, deg.

C

M

-5 0 5 10 15 20 -0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

0.06

W.T. Data

FELISA (Forebody)

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CN

-5 0 5 10 15 20 -0.06

-0.04

-0.02

0

0.02

0.04

0.06

α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

α, deg.

CA

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

Figure 29: Comparison of aerodynamic data for the `Ames' conguration at Mach 2.3.

40
α, deg.

C

M

-5 0 5 10 15 20 -0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

0.06

W.T. Data

FELISA (Forebody)

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CN

-5 0 5 10 15 20 -0.06

-0.04

-0.02

0

0.02

0.04

0.06

α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

α, deg.

CA

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

Figure 30: Comparison of aerodynamic data for the `Ames' conguration at Mach 2.7.

41
α, deg.

C

M

-5 0 5 10 15 20 -0.02

-0.01

0

0.01

0.02

0.03

0.04

0.05

0.06

W.T. Data

FELISA (Forebody)

α, deg.

L/D

-5 0 5 10 15 20 -0.4

-0.3

-0.2

-0.1

0

0.1

α, deg.

CA

-5 0 5 10 15 20 1.3

1.4

1.5

1.6

1.7

α, deg.

CD

-5 0 5 10 15 20 1.2

1.3

1.4

1.5

1.6

1.7

α, deg.

CN

-5 0 5 10 15 20 -0.06

-0.04

-0.02

0

0.02

0.04

0.06

α, deg.

CL

-5 0 5 10 15 20 -0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

Figure 31: Comparison of aerodynamic data for the `Ames' conguration at Mach 3.5.

42
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704–0188

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources,

gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this

collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson Davis

Highway, Suite 1204, Arlington, VA 22202–4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704–0188), Washington, DC 20503.

NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE

December 2001

3. REPORT TYPE AND DATES COVERED

Contractor Report

4. TITLE AND SUBTITLE

An Inviscid Computational Study of Three ’07 Mars Lander Aeroshell

Configurations Over a Mach Number Range of 2.3 to 4.5

5. FUNDING NUMBERS

C NAS1-00135

WU 706-85-41-01

6. AUTHOR(S)

Ramadas K. Prabhu

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

Lockheed Martin Engineering & Sciences Company

C/O NASA Langley Research Center

Hampton, VA 23681–2199

8. PERFORMING ORGANIZATION

REPORT NUMBER

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

NASA Langley Research Center

Hampton, VA 23681–2199

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA/CR–2001–211266

11. SUPPLEMENTARY NOTES

Langley Technical Monitor: Kenneth Sutton

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited

Subject Category 02

Distribution: Nonstandard

Availability: NASA CASI (301) 621–0390

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

This report documents the results of a study conducted to compute the inviscid longitudinal aerodynamic

characteristics of three aeroshell configurations of the proposed ’07 Mars lander. This was done in support of the

activity to design a smart lander for the proposed ’07 Mars mission. In addition to the three configurations with tabs

designated as the ‘shelf’, the ‘canted’, and the ‘Ames’, the baseline configuration (without tab) was also studied.

The unstructured grid inviscid CFD software FELISA was used, and the longitudinal aerodynamic characteristics of

the four configurations were computed for Mach number of 2.3, 2.7, 3.5, and 4.5, and for an angle of attack range

of -4 to 20 degrees. Wind tunnel tests had been conducted on scale models of these four configurations in the

Unitary Plan Wind Tunnel, NASA Langley Research Center. Present computational results are compared with the

data from these tests. Some differences are noticed between the two results, particularly at the lower Mach

numbers. These differences are attributed to the pressures acting on the aft body. Most of the present

computations were done on the forebody only. Additional computations were done on the full body (forbody and

aftbody) for the baseline and the ‘Shelf’ configurations. Results of some computations done (to simulate flight

conditions) with the Mars gas option and with an ”effective” 
 are also included.

14. SUBJECT TERMS

Unstructured Grid CFD, Hypersonic Speeds, Mars Gas, Aerodynamic Loads

15. NUMBER OF PAGES

47

16. PRICE CODE

A03

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

18. SECURITY CLASSIFICATION

OF THIS PAGE

Unclassified

19. SECURITY CLASSIFICATION

OF ABSTRACT

Unclassified

20. LIMITATION OF ABSTRACT

UL