[American Institute of Aeronautics and Astronautics 43rd AIAA Aerospace Sciences Meeting and Exhibit - Reno, Nevada ()] 43rd AIAA Aerospace Sciences Meeting and Exhibit - Transition

American Institute of Aeronautics and Astronautics

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Transition Onset and Turbulent Heating Measurements for

the Mars Science Laboratory Entry Vehicle

Brian R. Hollis* and Derek S. Liechty

†

NASA Langley Research Center, Hampton, VA, 23681

Michael J. Wright‡

NASA Ames Research Center, Moffett Field 94035

Michael S. Holden§, Timothy P. Wadhams

**, Matthew MacLean

††

Calspan University of Buffalo Research Center, Buffalo, NY 14225

Artem Dyakonov‡‡

National Institute of Aerospace, Hampton, VA 23666

An investigation of transitional/turbulent heating on the Mars Science Laboratory entry

vehicle has been conducted. Laminar, transitional, and turbulent heating data were

obtained in a perfect-gas, Mach 6 air wind tunnel and in a high-enthalpy shock tunnel in

CO2. Flow field solutions were computed using a Navier-Stokes solver at the test conditions

and comparisons were made between measured and predicted heating levels. Close

agreement was obtained for all laminar perfect-gas cases. For the high-enthalpy CO2 cases,

close agreement with the data was achieved when a fully-catalytic wall boundary condition

was employed, whereas the predictions exceeded the data by more than 25% if a non-

catalytic boundary condition was used. Turbulent heating predictions fell below the perfect-

gas air data by 25% but exceeded the CO2 data by 60%. Transition onset locations were

determined through comparisons with laminar heating predictions, and boundary-layer

parameters from the flow field solutions were used to develop correlations for the transition

onset location and the turbulent heating augmentation on the leeside of the vehicle.

Nomenclature

D = maximum vehicle diameter

H0-Hw = total enthalpy relative to wall conditions

L/D = aerodynamic lift-to-drag ratio

M = free stream Mach number

Me = boundary-layer edge Mach number

m = vehicle mass

mSPECIE = mass fraction of specie

m/(CDA) = ballistic coefficient

p = free stream pressure

q = heat transfer rate

qFR = reference heat transfer rate based on Fay-Riddell theory

* Aerospace Engineer, Aerothermodynamics Branch, AIAA Senior Member

† Aerospace Engineer, Aerothermodynamics Branch

‡ Aerospace Engineer, Reacting Flow Environments Branch

§ Program Director, Aerothermal and Aero-Optics Evaluation Center, AIAA Fellow

** Research Scientist, AIAA Member

†† Senior Research Scientist, AIAA Member

‡‡ Aerospace Engineer, AIAA Member

43rd AIAA Aerospace Sciences Meeting and Exhibit10 - 13 January 2005, Reno, Nevada

AIAA 2005-1437

This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.


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q = dynamic pressure

Rbase = base (maximum) radius

Rcorner = corner radius

Rnose = nose radius

Re = free stream Reynolds Number

Re = boundary-layer momentum thickness Reynolds Number

T = free stream temperature

U = free stream velocity

Vr = relative entry velocity

= angle of attack

= boundary-layer momentum thickness

F = forebody cone half-angle

aft1 = 1st aftbody cone half-angle

aft2 = 2nd aftbody cone half-angle

= free stream density

I. Background

The Mars Science Laboratory (MSL) mission, which is scheduled to be launched in 2009, will perform the first

precision landing (target uncertainty of ±10 km) of a scientific payload on the surface of Mars in 2010. While the

basic vehicle configuration, a 70-deg sphere-cone forebody with a conic or biconic aftbody, will be similar to those

of previous Mars missions such as Viking, Pathfinder or Mars Exploration Rover (MER), the MSL design is both

larger and heavier than

previous designs (Figure

1). To accomplish a

precision landing, the

vehicle will be required to

fly a controlled lifting

trajectory; current designs

call for a lift-to-drag ratio

(L/D) of 0.18, which can

be generated by flight at

an angle-of-attack of 11-

deg. As a result of its

larger mass and size, MSL

will experience heating

levels higher than any of

the previous missions, and

furthermore, because of

the high angle-of-attack

(for a blunt body) flight

requirement, the flow over

the leeside of the forebody

is expected to become turbulent early in the trajectory, which will substantially augment both the heating rates and

loads above the laminar levels.

Because the aeroheating environment of MSL will be more challenging than that of any previous Mars mission,

an extensive investigation of this environment is an important part of the ongoing MSL development program. In

this report, results from experimental heating tests conducted in the NASA Langley Research Center 20-Inch Mach

6 Air Tunnel and in the Calspan University of Buffalo (CUBRC) Large-Energy National Shock (LENS) Tunnel in

CO2 will be presented. These two tests were performed to obtain data on boundary-layer transition onset and

turbulent heating augmentation on the forebody leeside of the MSL. These tests were complemented by a third test

performed in the Graduate Aeronautical Laboratories of the California Institute of Technology (GALCIT) T5

Hypervelocity Shock Tunnel in CO2; that test is discussed in detail in reference 1, and some of the data from T5 will

be employed herein.

Figure 1. Comparison of Mars entry vehicle configurations


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II. Mars Science Laboratory Entry Vehicle Configuration

The current MSL entry vehicle configuration is a 70-deg sphere cone forebody with a biconic (40-deg and 62.18

deg cone angles) aftbody. The max vehicle radius is 2.286 m, the nose radius is 0.5 times the max radius and the

corner radius is 0.05 times the max radius. The current configuration differs from earlier MSL configurations2

through an increase in the max radius and a reduction in the length of the aftbody. Additionally, cavities3,4

in the

forebody heat shield through which the entry vehicle was to be bolted to its carrier vehicle during transit to Mars

were eliminated by moving the attachment point to the aftbody.

III. Experimental Method

A. NASA LaRC 20-Inch Mach 6 Air

Tunnel Test

1. Facility Description

This facility is a blow-down tunnel in

which heated, dried, and filtered air is

used as the test gas. The tunnel has a two

dimensional, contoured nozzle that opens

into a 0.521 m x 0.508 m (20.5-in. x 20.0-

in.) test section. The tunnel is equipped

with a bottom-mounted injection system

that can transfer a model from the

sheltered model box to the tunnel

centerline in less than 0.5 sec. Run times

of up to 15 minutes are possible in this facility, although for the current aeroheating study, run times of only a few

seconds were required. The nominal reservoir conditions of this facility are stagnation pressures of 206.8 to 3447.4

kPa (30 to 500 psia) with stagnation temperatures of 422.2 to 555.5 K (760 ˚R to 1000 ˚R), which produce perfect-

gas free stream flows with Mach numbers between 5.8 and 6.1 and Reynolds numbers of 1.64 106/m to 23.3 10

6/m

(0.5 106/ft to 7.3 10

6/ft). A more detailed description of this facility is presented in reference 5.

2. Wind Tunnel Model and Measurement Techniques

Global surface heating distributions and transition onset locations were measured using the digital optical

measurement method of two color, relative-intensity, phosphor thermography6,7

. In this method, ceramic wind

tunnel models are coated with a phosphor compound that fluoresces in two separate regions (green and red) of the

visible light spectrum. Before and during a wind tunnel run, the phosphor-coated model is illuminated by ultraviolet

(UV) light sources, and the resulting fluorescent intensity of the model is recorded and digitized through a three-

color CCD (charge coupled device) camera. Intensity data are converted to surface temperature values using system

calibrations. Global heat-transfer distributions are then computed from these temperature data using one-

dimensional, constant heat-transfer coefficient conduction theory, while transition onset locations can be determined

from comparison of the measured heating distributions to laminar computational results. As discussed in reference

7, the estimated experimental uncertainty of the heating data is approximately ±13%.

Measurements were made on four different entry vehicle configuration models (Figure 3): two 70-deg sphere

models of 15.24 cm and 17.78 cm diameter, and 50-deg and 60-deg sphere-cone models of 15.24 cm diameter. The

first two configurations were 0.0333 and 0.0389 scale representations of the MSL configuration. These two models

were built to a larger size than that of any large-angle cone models which had previously been tested in the 20-Inch

Mach 6 Air Tunnel because natural boundary layer transition had not been observed in previous testing3 of smaller

(12.70 cm diameter) MSL forebody models. The smaller 15.25 cm diameter model represented a conservative

estimate of the maximum model size that could be tested, while a greater amount of transitional/turbulent data could

be obtained on the larger 17.78 cm diameter model if it did not cause tunnel blockage. Analysis of the data obtained

during this test indicated that neither configuration caused blockage of the wind tunnel flow. The other two models

(the 50-deg and 60-deg sphere-cones) were tested in order to obtain transitional/turbulent data over a wider range of

boundary-layer edge conditions (e.g. Re /Me) than could be generated over the 70-deg sphere-cone models.

Although these two models differed from the MSL configuration, the flow fields produced by the three geometries

were similar enough that all the boundary-layer transitional behavior data should correlate with local boundary-layer

conditions.

Figure 2. MSL entry vehicle dimensions


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3. Test Parametrics

Wind tunnel operating

conditions are listed in Table

1. The reference heating qFR

values listed were computed

using Fay-Riddell theory8 for

a nose radius of 3.81 cm at a

cold-wall temperature of 300

K. As the operating

conditions were repeatable

with negligible differences

from run-to-run, only the

nominal operating points for

this test are shown as opposed

to the conditions for every

run. Test angles-of-attack

were 11-deg, 16-deg, 20-deg,

and 25-deg. In general, each

model was tested through the range of free stream Reynolds numbers at each angle of attack, but some of the lower

Reynolds number conditions were omitted when it was obvious that the conditions would only produce laminar flow

for a given model geometry and angle-of-attack. Also, only the 60-deg sphere cone was tested at the 25-deg angle-

of-attack.

Table 1 LaRC 20-Inch Mach 6 Air Tunnel Nominal Operating Conditions

Re

(1/ft)

M

P

(Pa)

T

(K)

(kg/m3)

U

(m/s)

H0-Hw

(MJ/kg)

qFR

(W/cm2)

3.95 106

5.99 1131 63.7 0.0620 955.7 0.22 7.82

5.13 106 6.00 1451 63.3 0.0800 955.3 0.22 8.80

5.89 106 6.01 1667 63.3 0.0919 956.5 0.22 9.51

6.35 106 6.02 1860 63.6 0.1020 959.9 0.22 10.20

7.33 106 6.02 2090 63.8 0.1145 960.7 0.22 10.90

B. CUBRC LENS Tunnel Test

1. Facility Description

Aeroheating measurements were made in Leg I of the CUBRC LENS tunnel. The LENS Legs I and II are

reflected shock tunnels with contoured nozzles in which helium or hydrogen driver gas is used to produce test

conditions for full-scale simulation of interceptors at Mach numbers from 7 to 14 in Leg I and from 3 to 8 for Leg II

at altitudes from sea level to 80 km in air test gas. Facility capabilities are discussed in greater detail in reference 9.

For this MSL test, a new capability of employing CO2 as the test gas in place of air was developed. Computational

analyses and calibration tests were performed to define a set of CO2 test conditions with a free stream Reynolds

number range of approximately 1.5 105/ft to 6.5 10

5/ft at a Mach number of 6.2 to 6.8, with a test core greater than

30-inch radius and with test times of several milliseconds.

2. Wind Tunnel Model and Measurement Technique

A 0.133-scale model (30.48 cm max radius) of the MSL configuration (Figure 2) was fabricated from stainless

steel for testing in LENS. The primary instrumentation was 79 thin-film heat-transfer gages on the forebody and

aftbody of the model. Ten surface pressure gages were installed along the forebody centerline, and 10 coaxial

surface thermocouples were also installed along the centerline in order to obtain additional heating data using a

different type of instrumentation for comparison with the thin-film heat-transfer gages.

Figure 3. Forebody parametric geometries for Mach 6 Air

test


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3. Test Parametrics

The 30.48 cm (12.0 inch) radius MSL model was tested at angles-of-attack of 0-deg, 11-deg, and 16-deg at free

stream Reynolds number of 1.5 105/ft to 6.5 10

5/ft. The total enthalpy was ~5 MJ/kg for all but a single 10 MJ/kg

run. The test matrix with the free stream conditions for each run is shown in Table 2.

Table 2 LENS Test Matrix

Run

(deg)

Re

(1/ft)

M

P

(Pa)

T

(K)

(kg/m3)

U

(m/s)

H0-Hw

(MJ/kg) mCO2 mCO mO2 mO

1 16 3.60 106 6.79 2021 760 1.36 10

-2 2880 5.39 10

6 0.9221 0.0496 0.0283 0.0000

2 16 5.76 106 6.30 4200 937 2.49 10

-2 2833 5.26 10

6 0.9119 0.0561 0.0320 0.0000

3 16 1.39 106 6.79 692 686 5.09 10

-3 2762 4.84 10

6 0.9007 0.0632 0.0361 0.0000

4 16 2.79 106 6.74 1435 710 1.04 10

-2 2760 4.70 10

6 0.9263 0.0469 0.0268 0.0000

5 16 2.22 106 6.64 2152 1157 8.59 10

-3 3658 9.51 10

6 0.7144 0.1818 0.1015 0.0023

6 11 6.43 106 6.21 4510 854 2.70 10

-2 2777 4.88 10

6 0.9315 0.0436 0.0249 0.0000

7 11 1.45 106 6.77 721 687 5.30 10

-3 2753 4.80 10

6 0.9028 0.0619 0.0353 0.0000

8 11 4.39 106 6.25 2869 805 1.82 10

-2 2722 4.72 10

6 0.9239 0.0484 0.0277 0.0000

9 11 2.40 106 6.26 1521 785 9.78 10

-3 2711 4.82 10

6 0.8993 0.0641 0.0366 0.0000

10 11 2.59 106 6.75 1396 732 9.70 10

-3 2817 4.95 10

6 0.9158 0.0536 0.0306 0.0000

11 0 6.10 106 6.23 4365 867 2.57 10

-2 2809 5.00 10

6 0.9285 0.0455 0.0260 0.0000

12 0 1.44 106 6.77 720 691 5.26 10

-3 2761 4.84 10

6 0.9008 0.0632 0.0361 -0.0001

13 0 2.53 106 6.78 1384 741 9.48 10

-3 2849 5.07 10

6 0.9116 0.0563 0.0321 0.0000

C. GALCIT T5 Tunnel Test

Data from heating tests in the GALCIT T5 facility are also discussed in this paper. A detailed description

of the facility and test techniques is presented in reference 1.

IV. Computational Method

Flow field computations at the wind tunnel test conditions were performed using the LAURA code10,11

. The

LAURA (Langley Aerothermodynamic Upwind Relaxation Algorithm) code is a three-dimensional, finite-volume

solver that includes perfect-gas, equilibrium and non-equilibrium chemistry models. The code can be used to solve

the inviscid, thin-layer Navier-Stokes, or full Navier-Stokes equations. For the current study the thin-layer mode

was employed; it was concluded in reference 3 from computations on a similar blunt body that this mode provided

accurate results for attached forebody flows. Time integration to steady-state in LAURA is accomplished through a

point-relaxation scheme. Roe-averaging12

with Harten’s entropy fix13

and Yee’s Symmetric Total Variation

Diminishing limiter14

is used for inviscid fluxes, and a second-order scheme is employed for viscous fluxes. In this

study, a perfect-gas air model was used for the Mach 6 air wind tunnel computations, and a 4-species non-

equilibrium, non-ionizing, carbon-dioxide gas model (CO2, CO, O2, O) with reactions taken from reference 15 was

used for the LENS and T5 cases.

For the LaRC 20-Inch Mach 6 Air Wind Tunnel cases, a uniform, ambient 300 K wall temperature boundary

condition was imposed. The use of a constant wall temperature is valid because the experimental data are reported

in terms of the non-dimensional ratio, q/qFR, which remains constant with wall temperature. The quantity qFR is the

heat-transfer coefficient computed using the Fay-Riddell method with the same nose radius as the MSL model at a

wall temperature of 300 K (540 ˚R).

For the LENS and T5 tunnel cases, the uniform 300 K surface temperature boundary conditions was also

imposed because the test times were on the order of milliseconds and the temperature rise over this interval was

small compared to the total (flow field stagnation) temperature. Both non-catalytic and fully catalytic (full

recombination of CO2) surface boundary conditions were employed in the computations.

Structured, finite-volume, multiple-block forebody grids with a singularity-free nose were employed for the

computations. Grid adaptation was performed (as per the method detailed in reference 11) to align the grid with the

bow shock and to produce nominal wall cell Reynolds numbers on the order of 1.


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Laminar computations were performed for all Mach 6 Air, LENS CO2, and T5 CO2 cases. Turbulent

computations were performed for all LENS cases and for the Mach 6 cases in which fully-turbulent flow appeared to

be produced. Boundary-layer quantities (Re , Me, , etc) were extracted from the laminar solutions in order to

develop transition onset correlations. These parameters were evaluated at the boundary-layer edge location, which

was defined as the normal distance from the surface at which the local total enthalpy reaches 99.5% of the free

stream total enthalpy. It should be noted that correlations presented herein were based on boundary-layer edge

properties obtained from a Navier-Stokes code, not a boundary-layer solver, as was done in most of the historical

research. It has been shown16

that parameters obtained from boundary-layer codes do not necessarily match those

obtained from Navier-Stokes codes. As the MSL vehicle flight environments will be based entirely on Navier-

Stokes computations, the current approach of obtaining edge parameters from LAURA solutions is consistent with

vehicle development activities.

For turbulent computations, the algebraic Baldwin-Lomax model17

with modifications18

for compressible flow

and the Dhawan-Narashima transition model19

were employed. While it is recognized that more sophisticated

turbulent models than Baldwin-Lomax exist, different models can produce very different results (e.g. reference 20)

and the validation status of any and all turbulence models for hypersonic flows is debatable. The Baldwin-Lomax

model is the standard being used for MSL development activities because it is computationally fast and stable, and

therefore its use for comparisons with the data presented herein will help to define the uncertainty margins which

will be applied to prediction of turbulent environments in flight.

V. Data Presentation and Analysis

A. 20-Inch Mach 6 Air Tunnel Data

Measured centerline heating distributions from each of the four configurations tested in the 20-Inch Mach 6 Air

Tunnel are shown in Figure 4 through Figure 6. The data are presented in the non-dimensional form q/qFR so that

the distributions are Reynolds-number independent as the long as the flow remains laminar. The non-dimensional

laminar heating distributions predicted using the LAURA code are also shown, as are the Re /Me distributions

extracted from the flow field solutions. These Re /Me values will be employed subsequently in the formulation of

transition correlations.

Agreement between the laminar LAURA predictions and the laminar data was generally within ±5% (as shown

for a sample case in Figure 7) and was within ±10% for all cases except in the shoulder region where very strong

flow field gradients were present, and around the leeside sphere-cone junction for some of the cases.

Boundary-layer transition was defined by locating the point where the measured heating level began to increase

above the predicted laminar level. Transition was observed along the leeside centerline of all configurations at most

angles-of-attack and free stream Reynolds numbers. Transition occurred closer to the nose of the vehicle on the 60-

deg sphere-cone and farther from the nose on the 50-deg sphere-cone. Fully-developed turbulent flow (identified by

the transitional heating augmentation reaching a maximum and then decreasing) was produced only on the 60-deg

sphere at = 16-deg (Re = 6.6 106/ft and Re = 7.3 10

6/ft) and at = 20-deg (Re = 7.3 10

6/ft). As the free

stream Reynolds number was increased the transition onset location moved upstream and the transitional/turbulent

heating augmentation above laminar levels increased.

An unexpected phenomena noted in this test was a localized increase in measured heating levels above predicted

laminar levels on the wind side of the model at higher free stream Reynolds numbers. This effect was generally

observed in the region between the wind-side flow field stagnation point and the location of the maximum windside

value of the transition parameter Re /Me. The magnitude of this increase was low and confined to a small region at

lower Reynolds numbers (e.g. the 60-deg sphere-cone at = 20-deg or 25-deg in Figure 5) and became greater and

broader at higher Reynolds numbers (e.g. the 70-deg sphere-cone at = 16-deg in Figure 6). Because the

magnitude of this phenomena increased with Reynolds numbers, and because it occurred near a maxima of the

transition parameter, it was concluded that it represented some form of boundary-layer transition.

The 20-Inch Mach 6 Air test provided a limited set of fully-turbulent data for comparison with computational

predictions. Comparisons with predicted turbulent heating levels on the 60-deg sphere-cone for = 16-deg and 20-

deg are shown in Figure 8. Turbulent predictions are shown for both fully-turbulent flow over the entire vehicle and

for transition to turbulence at the location identified from the test data. The Baldwin-Lomax turbulent heating levels

did not compare as well with the turbulent data as the laminar predictions and data did with each other. On the

windside of the vehicle, the Baldwin-Lomax model predicted a large increase in heating due to transition, when in

fact, little or no augmentation was measured. On the leeside of the vehicle the predicted heating levels were

approximately 15% lower than the measured levels for the fully-turbulent cases. For the cases where transition was


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set at the measured location, the predictions were approximately only 10% lower than the data; however, this

comparison is somewhat unrealistic because the Baldwin-Lomax method cannot predict transition but instead

requires specification of the transition location based on measured data. The under-prediction of leeside turbulent

heating levels was consistent with previous results3 from testing of an earlier MSL configuration where fully-

turbulent heating data were generated due to circular penetrations (attachment points) in the forebody of the model.

It had been theorized that the differences observed in those earlier comparisons might have been due to large-scale

flow field disturbances (separation shocks and vortices) produced by the penetrations. However, the differences in

the current turbulent data sets would seem to seem to invalidate that theory, and given the good agreement between

laminar measurements and predictions, it is reasonable to conclude that the differences can be attributed to a lack of

fidelity in the Baldwin-Lomax model for this type of flow field.

Figure 4. Centerline q/qFR and Re /Me distributions on 3.0-inch radius, 50-deg sphere-cone in 20-Inch Mach 6

Air Tunnel


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Figure 5. Centerline q/qFR and Re /Me distributions on 3.0-inch radius, 60-deg sphere-cone in 20-Inch Mach 6

Air Tunnel


9

Figure 6. Centerline q/qFR and Re /Me distributions on 3.0-inch and 3.5-inch radii, 70-deg sphere-cones in 20-

Inch Mach 6 Air Tunnel


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Figure 7. Sample comparison with uncertainty estimates for 70-deg sphere-cone in 20-Inch Mach 6 Air

Tunnel at = 16-deg

Figure 8. 60-deg sphere-cone turbulent heating comparisons at = 16-deg and 20-deg in 20-Inch Mach 6 Air

Tunnel

B. LENS CO2 Heating Data and CFD Comparisons

Measured heating data from testing of the MSL configuration in the LENS facility are shown in Figure 9 through

Figure 11. Predicted laminar and turbulent heating distributions are also shown for each case, as are the Re /Me

distributions.

As there was no reliable pre-test means of predicting when transition would be produced on this configuration in

CO2, and obtaining transitional/turbulent data was the primary goal of the test program, the MSL model for LENS

was designed to the maximum possible size (24-in diameter) that could be tested in the facility. As can be seen by

comparing the data to the laminar predictions, it appears that transitional or turbulent flow was produced. In fact,

the only case for which the flow remained laminar over the entire model was at the lowest free stream Reynolds


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number at = 0-deg (Run 12). The growth of transitional/turbulent flow on the leeside of the model followed a

similar progression with increasing free stream Reynolds as was observed in the Mach 6 air test. In contrast to the

perfect-gas cases where the turbulent predictions were lower than the data, the turbulent predictions exceeded the

fully-turbulent data (Runs 1, 2, 6 and 8) by 30% to 60% for the LENS CO2 cases.

The characteristics of the windside flow field were somewhat different than those observed in the Mach 6 air

test. While only small areas of windside transitional flow were observed in the Mach 6 air test, there were large

areas of the windside in the LENS CO2 test where the measured heating levels were well above the predicted

laminar levels; furthermore, the windside heating augmentation did not appear to correlate with increasing Reynolds

number. While facility noise in a high-enthalpy shock tunnel such as LENS can be expected to influence transition,

the behavior of the windside flow field is not completely understood at this point. Because the large size model size

precluded obtaining any fully-laminar baseline data at angle-of-attack, additional testing with a smaller model will

be necessary in order to explore the full range of laminar-transitional-turbulent flow on the MSL configuration in

high-enthalpy CO2.

The appropriate wall recombination boundary condition for high-enthalpy testing is another area that requires

further investigation. The laminar data and predictions shown in Figure 9 through Figure 11 agreed to within ±10%

with full surface recombination of CO2 enforced as a boundary condition. It can be expected that the combination of

the cold model surface temperature (the temperature rise is minimal with respect to the total temperature over the

several milliseconds of test time) and the stainless-steel surface would tend to drive dissociated species toward

recombination; however the close agreement with 100% recombination predictions was unexpected. In order to

investigate the influence of recombination on heating, computations were performed with a non-catalytic boundary

condition. The predicted non-catalytic heating levels were consistently much lower (25% in the example shown in

Figure 12) than the fully-catalytic predictions and the measured data. Note that Figure 12 also includes a sample of

the pressure data, which was matched very well by the predictions, and a sample of the coaxial thermocouple data,

which showed greater scatter than, but generally agreed with, the thin-film data Cleary, the best agreement was

obtained with the fully-catalytic boundary condition, however, this close agreement is not conclusive proof that the

fully-catalytic boundary-condition is correct, because other systemic errors, such as in the non-equilibrium reaction

rate models used in the flow field computations, cannot be ruled out. An investigation of material catalytic

properties independent of the aerothermodynamics of a wind tunnel test will be necessary to resolve this issue.

As a check on the computational methods employed in this study, independent computations for selected LENS

conditions were performed using the DPLR code21

. Comparisons of DPLR and LAURA results for laminar non-

catalytic and fully-catalytic, and turbulent fully-catalytic cases for Runs 7 and 11 are shown in Figure 13. The two

codes were found to be in close agreement for all cases. This result provided increased confidence in the

fundamental algorithms employed in each code, however much of the thermochemical data in these codes derives

from the same sources, so the agreement does not completely validate the computational methods.

As previously noted, a test program similar to that conducted in LENS was conducted in the T5 facility. As a

additional check on the computational methods, LAURA was used to compute several T5 cases (shots 2272 and

2259) with the same assumptions as in the LENS cases. The results are shown in Figure 14 and are consistent with

those from the LENS test: the laminar, fully-catalytic predictions were in close agreement with the laminar data,

while the turbulent, fully-catalytic predictions exceeded the turbulent data. Examination of the T5 data also revealed

some evidence of windside turbulence (T5 shots 2263, 2264, 2274 and 2275), but only of magnitude similar to that

seen in the Mach 6 air tests.


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Figure 9. Centerline q/qFR and Re /Me distributions on MSL model at = 0-deg in LENS


13



14



15

Figure 12. Comparison of fully-catalytic and non-catalytic predictions for LENS Run 7

Figure 13 . Comparison of DPLR and LENS predictions for LENS Runs 7 and 11

Figure 14. LAURA predictions for T5 Shots 2272 and 2259


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C. Transition Onset and Turbulent Heating Augmentation Analysis

Data from the LENS CO2 and 20-Inch Mach 6 Air tests, as well as from the T5 CO2 tests discussed in reference

1, were compared to the flow field properties extracted from the CFD solutions to determine transition onset

locations and turbulent heating augmentation above laminar levels. These three facilities provided data in two test

gases (air and CO2) at perfect-gas and non-equilibrium conditions across a wide range of total enthalpy levels (0.2

MJ/kg in Mach 6 air, 5 MJ/kg in LENS, and 5 MJ/kg to 15 MJ/kg in T5). The Mach 6 air data set was of the highest

fidelity, due to the repeatability of operating conditions, well-defined thermodynamic environment (perfect-gas) and

the accuracy of the transition onset locations obtained from the global imaging system. Also, transition

measurements in this tunnel have been shown22

to correlate well with those from purposely-designed “quiet”

tunnels, even though this facility was not designed as a quiet tunnel. However, the test conditions (Mach 6, perfect-

gas air) are not directly applicable to hypervelocity flight in the CO2 atmosphere of Mars. The high-enthalpy, CO2

test gas data from LENS and T5 are more relevant to Martian entry conditions, but the model instrumentation

density was much lower, which made it difficult to precisely define the location of transition onset. Additionally,

the effects of the highly dynamic operating characteristics on free stream turbulence levels and transition onset are

hard to characterize, but presumably tend to produce transition earlier than would be observed at identical conditions

in flight.

Values of the transition parameter Re /Me at the location of transition onset along the leeside centerline along

with uncertainty estimates for the data are shown for the Mach 6 air data set in Figure 15 and for the LENS and T5

CO2 data sets in Figure 16. The transition onset location was identified as the point where the data began to diverge

from the predicted laminar distribution. This location was fairly easy to determine in the Mach 6 data set since the

phosphor thermography technique employed to acquire the data provided an essentially continuous distribution. In

the LENS and T5 data sets, precise identification of the transition onset location was not possible because the data

were obtained from discrete heat-transfer gages. The values reported were based on averages between the two gages

which were judged to be immediately before and immediately after the transition location, and as a result the

uncertainty estimates were much higher for the LENS and T5 data. In the Mach 6 air test, transition onset occurred

at Re /Me values of 350 to 425. In the LENS CO2 test, transition onset occurred at Re /Me values of 175 to 375,

while in the T5 CO2 test the values of Re /Me varied from 375 to 525 at transition onset. These ranges of data are

consistent with Re /Me values for smooth-body transition of 200 to 400 which are often cited in the literature (e.g.

reference 23).

The transition data from all three facilities are plotted together against total enthalpy, free stream Reynolds

number, and dynamic pressure in Figure 17 through Figure 19. With respect to total enthalpy, there did not appear to

be a consistent trend in any of the data sets. There did appear to be a trend of increasing Re /Me with increasing free

stream Reynolds number or dynamic pressure in each of the individual data sets, but these trends did not correlate

between the three data sets. The three data sets did correlate reasonably well when plotted against the quantity

q Re ,R, as shown in Figure 20; however a wider range of data (including gases other than CO2 and air) will be

required to determine whether this correlation is valid.

Heating augmentation values (the ratio of measured transitional/turbulent heating to predicted laminar heating

along the leeside centerline) from the three data sets are shown in Figure 21. Separate hyperbolic functions were

fitted to the air and CO2 data sets. These fits match the character of the available data, but the maxima defined for

each fit are purely conjectural.


17

Figure 15. Transition data from 20-Inch Mach 6 Air Tunnel

Figure 16. Transition data from LENS and T5 CO2 tests


18

Figure 17. Transition data plotted vs. total enthalpy

Figure 18. Transition data plotted vs. dynamic pressure


19

Figure 19. Transition data plotted vs. free stream Reynolds number

Figure 20. Transition data plotted against free stream Reynolds number times dynamic pressure


20

Figure 21. Transitional/turbulent heating augmentation factor

VI. Conclusions

Experimental data were obtained from a perfect-gas wind tunnel and a high-enthalpy, reflected shock tunnel in

CO2 in support of the Mars Science Laboratory entry vehicle development program. Comparisons between the data

and computational predictions showed good agreement for laminar, perfect-gas air cases. For laminar, reacting CO2

cases, the comparisons also showed close agreement if the assumption that the test model surface was fully-catalytic

was made. Turbulent comparisons under-predicted by significant amounts the perfect-gas data and over-predicted

the reacting-gas CO2 data. While the turbulence model employed in these computations was relatively simple, it is

the model being used in development of the MSL vehicle, and therefore the comparison errors shown in the research

will have to be taken into account in the formulation of uncertainty margins for MSL.

Transition onset locations along the leeside centerline of the vehicle were also obtained and reported in terms of

the transition parameter Re /Me. The transition data from each of the three facilities could be correlated separately

in terms of the free stream dynamic pressure or length-Reynolds number. A tentative form of a global correlation of

all data sets in terms of the product of dynamic pressure times free-stream length-Reynolds number was also

developed, but more data will be required to validate this correlation. Additionally, turbulent heating augmentation

correlations for the leeside centerline of the MSL were developed for air and CO2.

Because of its large ballistic coefficient and high-angle attack during entry into the Mars atmosphere, leeside

turbulent heating will be an important design consideration in the development of the vehicle’s heat shield. Based

on the results of comparisons between computational methods and experimental data shown in this study, it is

apparent that large uncertainty factors will have to be applied to turbulent heating predictions for the vehicle in order

to ensure mission success. It is clear that further development of computational methods for turbulent blunt-body

flows will be required to reduce design uncertainties for future Mars missions.


21

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[American Institute of Aeronautics and Astronautics 43rd AIAA Aerospace Sciences Meeting and Exhibit - Reno, Nevada ()] 43rd AIAA Aerospace Sciences Meeting and Exhibit - Transition